Free Essay: Fluid Mechanics - 97973 Words

Fluid Mechanics

Egon Krause Fluid Mechanics

Egon Krause

Fluid Mechanics
With Problems and Solutions, and an Aerodynamic Laboratory

With 607 Figures

Prof. Dr. Egon Krause RWTH Aachen Aerodynamisches Institut W¨ llnerstr.5-7 u 52062 Aachen Germany

ISBN 3-540-22981-7 Springer Berlin Heidelberg New York
Library of Congress Control Number: 2004117071 This work is subject to copyright. All rights are reserved, whether the whole or part of the material is concerned, speciﬁcally the rights of translation, reprinting, reuse of illustrations, recitation, broadcasting, reproduction on microﬁlm or in other ways, and storage in data banks. Duplication of this publication or parts thereof is permitted only under the provisions of the German Copyright Law of September 9, 1965, in its current version, and permission for use must always be obtained from Springer-Verlag. Violations are liable to prosecution under German Copyright Law. Springer is a part of Springer Science+Business Media springeronline.com c Springer-Verlag Berlin Heidelberg 2005 Printed in Germany

The use of general descriptive names, registered names, trademarks, etc. in this publication does not imply, even in the absence of a speciﬁc statement, that such names are exempt from the relevant protective laws and regulations and therefore free for general use. Typesetting: Data conversion by the authors. Final processing by PTP-Berlin Protago-TEX-Production GmbH, Germany Cover-Design: design & production GmbH, Heidelberg Printed on acid-free paper 62/3020Yu - 5 4 3 2 1 0

Preface

During the past 40 years numerical and experimental methods of ﬂuid mechanics were substantially improved. Nowadays time-dependent three-dimensional ﬂows can be simulated on high-performance computers, and velocity and pressure distributions and aerodynamic forces and moments can be measured in modern wind tunnels for ﬂight regimes, until recently not accessible for research investigations. Despite of this impressive development during the recent past and even 100 years after Prandtl introduced the boundary-layer theory, the fundamentals are still the starting point for the solution of ﬂow problems. In the present book the important branches of ﬂuid mechanics of incompressible and compressible media and the basic laws describing their characteristic ﬂow behavior will be introduced. Applications of these laws will be discussed in a way suitable for engineering requirements. The book is divided into the six chapters: Fluid mechanics I and II, exercises in ﬂuid mechanics, gas dynamics, exercises in gasdynamics, and aerodynamics laboratory. This arrangement follows the structure of the teaching material in the ﬁeld, generally accepted and approved for a long time at German and foreign universities. In ﬂuid mechanics I, after some introductory statements, incompressible ﬂuid ﬂow is described essentially with the aid of the momentum and the moment of momentum theorem. In ﬂuid mechanics II the equations of motion of ﬂuid mechanics, the Navier-Stokes equations, with some of their important asymptotic solutions are introduced. It is demonstrated, how ﬂows can be classiﬁed with the aid of similarity parameters, and how speciﬁc problems can be identiﬁed, formulated and solved. In the chapter on gasdynamics the inﬂuence of variable density on the behavior of subsonic and supersonic ﬂows is described. In the exercises on ﬂuid mechanics I and II and on gasdynamics the material described in the previous chapters is elaborated in over 200 problems, with the solutions presented separately. It is demonstrated how the fundamental equations of ﬂuid mechanics and gasdynamics can be simpliﬁed for the various problem formulations and how solutions can be constructed. Numerical methods are not employed. It is intended here, to describe the fundamental relationships in closed form as far as possible, in order to elucidate the intimate connection between the engineering formulation of ﬂuid-mechanical problems and their solution with the methods of applied mathematics. In the selection of the problems it was also intended, to exhibit the many diﬀerent forms of ﬂows, observed in nature and technical applications. Because of the special importance of experiments in ﬂuid mechanics, in the last chapter, aerodynamics laboratory, experimental techniques are introduced. It is not intended to give a comprehensive and complete description of experimental methods, but rather to explain with the description of experiments, how in wind tunnels and other test facilities experimental data can be obtained. A course under the same title has been taught for a long time at the Aerodymisches Institut of the RWTH Aachen. In the various lectures and exercises the functioning of low-speed and supersonic wind tunnels and the measuring techniques are explained in experiments, carried out in the facilities of the laboratory. The experiments comprise measurements of pressure distributions on a half body and a wing section, of the drag of a sphere in incompressible and compressible ﬂow, of the aerodynmic forces and their moments acting on a wing section, of velocity proﬁles in a ﬂat-plate boundary layer, and of losses in compressible pipe ﬂow. Another important aspect of the laboratory course is to explain ﬂow analogies, as for example the

VI

Preface

so-called water analogy, according to which a pressure disturbance in a pipe, ﬁlled with a compressible gas, propagates analogously to the pressure disturbance in supercritical shallow water ﬂow. This book was stimulated by the friendly encouragement of Dr. M. Feuchte of B.G. TeubnerVerlag. My thanks go also to Dr. D. Merkle of the Springer-Verlag, who agreed to publish the English translation of the German text. Grateful acknowledgement is due to my successor Professor Dr.-Ing. W. Schr¨der, who provided personal and material support by the Aerodyo namisches Institut in the preparation of the manuscript. I am indebted to Dr.-Ing. O. Thomer who was responsible for the preparatory work during the initial phase of the project until he left the institute. The ﬁnal manuscript was prepared by cand.-Ing. O. Yilmaz, whom I gratefully acknowledge. Dr.-Ing. M. Meinke oﬀered valuable advice in the preparation of some of the diagrams. Aachen, July 2004 E. Krause

Table of Contents

1. Fluid Mechanics I 1.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2 Hydrostatics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2.1 Surface and Volume Forces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2.2 Applications of the Hydrostatic Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.2.3 Hydrostatic Lift . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3 Hydrodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3.1 Kinematics of Fluid Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3.2 Stream Tube and Filament . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3.3 Applications of Bernoulli’s Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.4 One-Dimensional Unsteady Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5 Momentum and Moment of Momentum Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5.1 Momentum Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5.2 Applications of the Momentum Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5.3 Flows in Open Channels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5.4 Moment of Momentum Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.5.5 Applications of the Moment of Momentum Theorem . . . . . . . . . . . . . . . . . . 1.6 Parallel Flow of Viscous Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.6.1 Viscosity Laws . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.6.2 Plane Shear Flow with Pressure Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.6.3 Laminar Pipe Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7 Turbulent Pipe Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.1 Momentum Transport in Turbulent Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.2 Velocity Distribution and Resistance Law . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.3 Pipes with Non-circular Cross Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2. Fluid Mechanics II 2.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2 Fundamental Equations of Fluid Mechanics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.1 The Continuity Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.2 The Navier-Stokes Equations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.3 The Energy Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.4 Diﬀerent Forms of the Energy Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3 Similar Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3.1 Derivation of the Similarity Parameters with the Method of Dimensional Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3.2 The Method of Diﬀerential Equations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3.3 Physical Meaning of the Similarity Parameters . . . . . . . . . . . . . . . . . . . . . . . 2.4 Creeping Motion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 2 2 3 5 5 5 7 8 10 12 12 13 17 18 19 20 21 22 24 25 26 27 29 31 31 31 31 32 35 36 38 38 40 41 42

VIII Table of Contents Vortex Theorems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.5.1 Rotation and Circulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.5.2 Vorticity Transport Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6 Potential Flows of Incompressible Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6.1 Potential and Stream Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6.2 Determination of the Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6.3 The Complex Stream Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6.4 Examples for Plane Incompressible Potential Flows . . . . . . . . . . . . . . . . . . . 2.6.5 Kutta-Joukowski Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6.6 Plane Gravitational Waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.7 Laminar Boundary Layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.7.1 Boundary-Layer Thickness and Friction Coeﬃcient . . . . . . . . . . . . . . . . . . . 2.7.2 Boundary-Layer Equations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.7.3 The von K´rm´n Integral Relation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . a a 2.7.4 Similar Solution for the Flat Plate at Zero Incidence . . . . . . . . . . . . . . . . . 2.8 Turbulent Boundary Layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.8.1 Boundary-Layer Equations for Turbulent Flow . . . . . . . . . . . . . . . . . . . . . . . 2.8.2 Turbulent Boundary Layer on the Flat Plate at Zero Incidence . . . . . . . . . 2.9 Separation of the Boundary Layer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.10 Selected References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.11 Appendix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3. Exercises in Fluid Mechanics 3.1 Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.1 Hydrostatics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.2 Hydrodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.3 Momentum and Moment of Momentum Theorem . . . . . . . . . . . . . . . . . . . . 3.1.4 Laminar Flow of Viscous Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.5 Pipe Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.6 Similar Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.7 Potential Flows of Incompressible Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.8 Boundary Layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.1.9 Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2 Solutions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.1 Hydrostatics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.2 Hydrodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.3 Momentum and Moment of Momentum Theorem . . . . . . . . . . . . . . . . . . . . 3.2.4 Laminar Flow of Viscous Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.5 Pipe Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.6 Similar Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.7 Potential Flows of Incompressible Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.8 Boundary Layers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.2.9 Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.5 44 44 45 46 46 48 48 49 53 54 55 56 56 58 59 61 61 62 64 66 66 69 69 69 71 76 80 83 86 88 91 92 96 96 99 106 113 119 123 126 132 134

4. Gasdynamics 139 4.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139 4.2 Thermodynamic Relations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139

Table of Contents One-Dimensional Steady Gas Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.1 Conservation Equations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.2 The Speed of Sound . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.3 Integral of the Energy Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.4 Sonic Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.5 The Limiting Velocity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.3.6 Stream Tube with Variable Cross-Section . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.4 Normal Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.4.1 The Jump Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.4.2 Increase of Entropy Across the Normal Compression Shock . . . . . . . . . . . . 4.4.3 Normal Shock in Transonic Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5 Oblique Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5.1 Jump Conditions and Turning of the Flow . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5.2 Weak and Strong Solution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.5.3 Heart-Curve Diagram and Hodograph Plane . . . . . . . . . . . . . . . . . . . . . . . . . 4.5.4 Weak Compression Shocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.6 The Prandtl-Meyer Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.6.1 Isentropic Change of Velocity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.6.2 Corner Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.6.3 Interaction Between Shock Waves and Expansions . . . . . . . . . . . . . . . . . . . . 4.7 Lift and Wave Drag in Supersonic Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.7.1 The Wave Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.7.2 Lift of a Flat Plate at Angle of Attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.7.3 Thin Proﬁles at Angle of Attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.8 Theory of Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.8.1 The Crocco Vorticity Theorem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.8.2 The Fundamental Equation of Gasdynamics . . . . . . . . . . . . . . . . . . . . . . . . . 4.8.3 Compatibility Conditions for Two-Dimensional Flows . . . . . . . . . . . . . . . . . 4.8.4 Computation of Supersonic Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.9 Compressible Potential Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.9.1 Simpliﬁcation of the Potential Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.9.2 Determination of the Pressure Coeﬃcient . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.9.3 Plane Supersonic Flows About Slender Bodies . . . . . . . . . . . . . . . . . . . . . . . 4.9.4 Plane Subsonic Flow About Slender Bodies . . . . . . . . . . . . . . . . . . . . . . . . . 4.9.5 Flows about Slender Bodies of Revolution . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.10 Similarity Rules . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.10.1 Similarity Rules for Plane Flows After the Linearized Theory . . . . . . . . . . 4.10.2 Application of the Similarity Rules to Plane Flows . . . . . . . . . . . . . . . . . . . 4.10.3 Similarity Rules for Axially Symmetric Flows . . . . . . . . . . . . . . . . . . . . . . . . 4.10.4 Similarity Rules for Plane Transonic Flows . . . . . . . . . . . . . . . . . . . . . . . . . . 4.11 Selected References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5. Exercises in Gasdynamics 5.1 Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.1 One-Dimensional Steady Flows of Gases . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.2 Normal Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.3 Oblique Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.4 Expansions and Compression Shocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.5 Lift and Wave Drag – Small-Perturbation Theory . . . . . . . . . . . . . . . . . . . . 4.3

IX 141 141 142 143 144 145 145 147 147 149 150 151 151 153 154 155 156 157 158 159 160 161 161 161 163 163 164 166 167 170 170 171 172 174 175 178 178 180 182 183 184 185 185 185 188 191 193 196

X

Table of Contents 5.1.6 Theory of Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1.7 Compressible Potential Flows and Similarity Rules . . . . . . . . . . . . . . . . . . . Solutions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.1 One-Dimensional Steady Flows of Gases . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.2 Normal Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.3 Oblique Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.4 Expansions and Compression Shocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.5 Lift and Wave Drag – Small-Perturbation Theory . . . . . . . . . . . . . . . . . . . . 5.2.6 Theory of Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2.7 Compressible Potential Flows and Similarity Rules . . . . . . . . . . . . . . . . . . Appendix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198 199 203 203 208 211 214 217 219 221 225 233 233 233 234 235 237 243 247 247 248 249 251 253 253 257 258 259 262 262 263 265 267 268 271 271 273 278 280 283 283 283 287 295 299 299

5.2

5.3

6. Aerodynamics Laboratory 6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) . . . . . . . . . . . . . . . o 6.1.1 Preliminary Remarks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.1.2 Wind Tunnels for Low Speeds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.1.3 Charakteristic Data of a Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.1.4 Method of Test and Measuring Technique . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.1.5 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2 Pressure Distribution on a Half Body . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2.1 Determination of the Contour and the Pressure Distribution . . . . . . . . . . . 6.2.2 Measurement of the Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2.3 The Hele-Shaw Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3 Sphere in Incompressible Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3.1 Fundamentals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3.2 Shift of the Critical Reynolds Number by Various Factors of Inﬂuence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3.3 Method of Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4 Flat-Plate Boundary Layer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4.1 Introductory Remarks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4.2 Method of Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4.3 Prediction Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.4.5 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5 Pressure Distribution on a Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5.1 Wing of Inﬁnite Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5.2 Wing of Finite Span . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5.3 Method of Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.5.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.6 Aerodynamic Forces Acting on a Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.6.1 Nomenclature of Proﬁles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.6.2 Measurement of Aerodynamic Forces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.6.3 Application of Measured Data to Full-Scale Conﬁgurations . . . . . . . . . . . . 6.6.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.7 Water Analogy – Propagation of Surface Waves in Shallow Water and of Pressure Waves in Gases . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.7.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Table of Contents 6.7.2 The Water Analogy of Compressible Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.7.3 The Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.7.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.8 Resistance and Losses in Compressible Pipe Flow . . . . . . . . . . . . . . . . . . . . . . . . . . 6.8.1 Flow Resistance of a Pipe with Inserted Throttle (Oriﬁcee, Nozzle, Valve etc.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.8.2 Friction Resistance of a Pipe Without a Throttle . . . . . . . . . . . . . . . . . . . . . 6.8.3 Resistance of an Oriﬁce . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.8.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.8.5 Problems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.9 Measuring Methods for Compressible Flows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.9.1 Tabular Summary of Measuring Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.9.2 Optical Methods for Density Measurements . . . . . . . . . . . . . . . . . . . . . . . . . 6.9.3 Optical Setup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.9.4 Measurements of Velocities and Turbulent Fluctuation Velocities . . . . . . . 6.9.5 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge . . . . . . . . . . . . . . 6.10.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10.2 Classiﬁcation of Wind Tunnels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10.3 Elements of a Supersonic Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10.4 The Oblique Compression Shock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10.5 Description of the Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.10.6 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.11 Sphere in Compressible Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.11.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.11.2 The Experiment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.11.3 Fundamentals of the Compressible Flow About a Sphere . . . . . . . . . . . . . . 6.11.4 Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.11.5 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Index

XI 299 304 305 307 307 307 311 313 316 320 320 320 325 326 327 329 329 329 332 333 337 339 342 342 342 344 348 348 351

1. Fluid Mechanics I

1.1 Introduction
Fluid mechanics, a special branch of general mechanics, describes the laws of liquid and gas motion. Flows of liquids and gases play an important role in nature and in technical applications, as, for example, ﬂows in living organisms, atmospheric circulation, oceanic currents, tidal ﬂows in rivers, wind- and water loads on buildings and structures, gas motion in ﬂames and explosions, aero- and hydrodynamic forces acting on airplanes and ships, ﬂows in water and gas turbines, pumps, engines, pipes, valves, bearings, hydraulic systems, and others. Liquids and gases, often termed ﬂuids, in contrast to rigid bodies cause only little resistance when they are slowly deformed, as long as their volume does not change. The resistance is so much less the slower the deformation is. It can therefore be concluded, that the arising tangential stresses are small when the deformations are slow and vanish in the state of rest. Hence, liquids and gases can be deﬁned as bodies, which do not build up tangential stresses in the state of rest. If the deformations are fast, there results a resistance proportional to the friction forces in the ﬂuid. The ratio of the inertia to the friction forces is therefore of great importance for characterizing ﬂuid ﬂows. This ratio is called Reynolds number. In contrast to gases, liquids can only little be compressed. For example, the relative change in volume of water is 5 · 10−5 when the pressure is increased by 1 bar, while air changes by a factor of 5 · 10−1 under normal conditions in an isothermal compression. If liquid and gas motion is to be described, in general, not the motion of single atoms or molecules is described, neither is their microscopic behavior taken into account; the ﬂowing medium is considered as a continuum. It is assumed to consist out of very small volume elements, the overall dimensions of which, however, being much larger than the intermolecular distances. In a continuum the mean free path between the collisions of two molecules is small compared to the characteristic length of the changes of the ﬂow quantities. Velocity, pressure and temperature, density, viscosity, thermal conductivity, and speciﬁc heats are described as mean values, only depending on position and time. In order to be able to deﬁne the mean values, it is necessary, that the volume element is small compared to the total volume of the continuum. This is illustrated in the following example for the density of air ﬂowing in a channel with a cross-sectional area of 1 cm2 . At room temperature and atmospheric pressure one cubic centimeter of air contains 2.7 · 1019 molecules with a mean free path of about 10−4 mm. A cube with length of side of 10−3 mm, – the 1012 th part of a cubic centimeter – still contains 2.7 · 107 molecules. This number is suﬃciently large such that a mean value of the density can be deﬁned for every point in the ﬂow ﬁeld. The basis for the description of ﬂow processes is given by the conservation laws of mass, momentum, and energy. After the presentation of simpliﬁed integral relations in the ﬁrst chapter in Fluid Mechanics I these laws will be derived in Fluid Mechanics II for three-dimensional ﬂows with the aid of balance equations in integral and diﬀerential form. Closed-form solutions of these equations are not available for the majority of ﬂow problems. However, in many instances, approximate solutions can be constructed with the aid of simpliﬁcations and idealizations. It will be shown, how the magnitude of the various forces per unit volume and of the energy contributions, appearing in the conservation equations, can be compared with each other by introducing similarity parameters. The small terms can then be identiﬁed and dropped, and only the important, the largest terms are retained, often leading to simpliﬁed, solvable conservation equations, as is true for very slow ﬂuid motion, at times referred to as creeping motion. In this

2

1. Fluid Mechanics I

case the inertia forces can be neglected in comparison to the friction forces. As an example the solution of the simpliﬁed conservation equations for very slow motion will be derived for the ﬂow in a friction bearing. If the inertia and pressure forces are dominant, the terms describing the friction forces per unit volume can be dropped. Inviscid ﬂows can be shown to possess a potential, and with the aid of the potential theory the surface pressure distribution of external ﬂows about rigid bodies can be determined. The potential ﬂow theory forms the basis for determining the lift of aerodynamically shaped bodies. Several applications will be discussed with the aid of analytic functions. In the vicinity of rigid walls, in general, the friction forces cannot be neglected. It will be shown how their inﬂuence on the ﬂow can be determined with Prandtl’s boundary-layer theory, as long as the region, in which the viscous forces act, is thin compared to length of the body. The similar solution of the boundary-layer equations will be derived for the case of the ﬂat plate. The importance of a non-vanishing pressure gradient will be elucidated for the case of separating ﬂows. In ﬂows of gases at high speeds marked changes of the density occur. They have to be taken into account in the description of ﬂow ﬁelds. The laws governing compressible ﬂows will be described in the chapter Gasdynamics.

1.2 Hydrostatics
1.2.1 Surface and Volume Forces A continuum is said to be in equilibrium, if the resultant of the forces, acting on every arbitrary part of the volume vanishes. The forces are called surface, volume, and inertia forces, as their magnitude is proportional to the surface, volume, or mass of the part of the continuum considered. The surface forces act normal to the surface, as long as the ﬂuid is at rest. The corresponding stresses (normal force per surface element) are – after Euler (1755) – called ﬂuid pressure or abbreviated pressure. The equilibrium condition is derived for an arbitrarily chosen triangular prismatic volume element. For the surface forces to be in equilibium, the sum of the vertical and horizontal components must be equal to zero. If the forces per unit area, the pressures on the surface, are denoted by p1 , p2 , and p3 , then the forces can be written as products of the pressures and the areas, on which they act.The following sketch shows the prismatic element with the surface forces indicated. If another geometric shape of the volume element would have been chosen, the equilibrium condition would always require the vanishing of the sum of the vertical and horizontal components of the surface forces. p1 a d − p3 c d cos(a,c) − ρ g p2 b d a b p1 − = = = p3 c d cos(b,c) = 0 c cos(a,c) c cos(b,c) p2 = p3 = p . abd =0 2

(1.1)

For c → 0 the volume forces vanish. It follows for every point in a ﬂuid which is in equilibrium, that the pressure p does not depend on the direction of the surface element on which it acts. The equilibrium condition for a cylinder with inﬁnitesimally small cross-sectional area A, and with its axis normal to the positive direction of the gravitational force, yields the following relation p1 A = p2 A ⇒ p1 = p2 = p . (1.2)

1.2 Hydrostatics The axis of the cylinder represents a line of constant pressure. If the cylinder is turned by 90 degrees, such that its axis is parallel to the direction of the gravitational force, then the equilibrium of forces gives for the z-direction −ρ g A dz − (p + dp) A + p A = 0 . (1.3)

3

It follows from this relation, that in a ﬂuid in the state of rest, the pressure changes in the direction of the acting volume force according to the diﬀerential equation dp = −ρ g dz . (1.4)

After integration the fundamental hydrostatic equation for an incompressible ﬂuid (ρ = const.) and with g = const. is obtained to p + ρ g z = const. . (1.5)

If ρ is determined from the thermal equation of state ρ= p RT , (1.6)

for a constant temperature T = T0 (isothermal atmosphere) the so-called barometric height formula is obtained p = p0 e gz −R T 0

.

(1.7)

The diﬀerential form of the fundamental hydrostatic equation is valid for arbitrary force ﬁelds. With b designating the acceleration vector, it reads grad(p) = ρ b . 1.2.2 Applications of the Hydrostatic Equation Assume, as shown in the sketch, that in a ﬂuid-ﬁlled vessel the hatched parts are solidiﬁed without any change of density. The equilibrium of the ﬂuid remains unchanged (Principle of solidiﬁcation, Stevin 1586). By the process of solidiﬁcation communicating vessels are generated. This principle is, for example, applied in liquid manometers and hydraulic presses. (1.8)

U-tube manometer

Single-stem manometer

4

1. Fluid Mechanics I

The fundamental hydrostatic equation yields p1 − p2 = ρ g (z2 − z1 ) . (1.9) The pressure diﬀerence to be measured is proportional to the diﬀerence of the heights of the liquid levels. Barometer One stem of the U-tube manometer is closed and evacuated (p2 = 0). The atmospheric pressure is pa = p1 = ρ g (z2 − z1 ) . Hydraulic Press For equal pressure on the lower sides of the pistons the force F2 is F2 = F1 A2 A1 . (1.11) (1.10)

Communicating vessels can be used for generating large forces, if A2 >> A1 .

Hydrostatic Paradox (Pascal 1647) The force acting on the bottom of all vessels is independent of the shape of the vessels and of the weight of the ﬂuid, as long as the surface area of the bottom A and the height h of the vessels are constant. FB = (pB − pA ) A = ρ g h A (1.12)

Force on a Plane Side Wall The ﬂuid pressure results in a force acting on the side wall of the vessel: Fs = Fs =
0

A h

(p − pa ) dA B ρ g z dz = ρ g h 2 B h2 2 (1.13)

= ρgA

1.3 Hydrodynamics The location of the point of application of force follows from the balance of moments zs = 1.2.3 Hydrostatic Lift 2 h . 3

5

(1.14)

A body immerged in a ﬂuid of density ρF experiences a lift or an apparent loss of weight, being equal to the weight of the ﬂuid displaced by the body (Archimedes’ principle 250 b. C.) Fs = ρF g
A

(z1 − z2 ) dA = ρF g τ

.

(1.15)

τ is the volume of the ﬂuid displaced by the body. Hence the weight of a body – either immersed in or ﬂoating on a ﬂuid – is equal to the weight of the ﬂuid displaced by the body. Balloons and ships are examples for the application of Archimendes’ principle.

1.3 Hydrodynamics
1.3.1 Kinematics of Fluid Flows Two methods are commonly used for the description of ﬂuid motion, the Lagrangian method and the Eulerian method. Lagrangian Method (Fluid Coordinates) The motion of the ﬂuid particles is described by specifying their coordinates as a function of time. The line connecting all points a particle is passing through in the course of time is called path line or Lagrangian particle path. The path line begins at time t = t0 at a point deﬁned by the position vector r 0 = i x0 + j y0 + k z0 . (1.16)

The motion of the ﬂuid is completely described, if the position vector r is known as a function of time: r = F (r 0 ,t) or in components x = f1 (x0 ,y0 ,z0 ,t) ; y = f2 (x0 ,y0 ,z0 ,t) ; z = f3 (x0 ,y0 ,z0 ,t) . (1.17)

(1.18)

The velocity is obtained by diﬀerentiating the position vector with respect to time v= lim ∆t → 0 r2 − r1 ∆t = dr dt (1.19)

6 where

1. Fluid Mechanics I

dx dy dz dr =i +j +k =iu+jv+kw . (1.20) dt dt dt dt In the last equation u, v, and w are the components of the velocity vector. The components of the acceleration vector are d2 x d2 y d2 z bx = 2 ; b y = 2 ; b z = 2 . (1.21) dt dt dt For most ﬂow problems the Langrangian method proves to be too laborious. Aside from a few exceptions, not the path-time dependence of a single particle is of interest, but rather the ﬂow condition at a certain point at diﬀerent times. Eulerian Method (Space Coordinates) The motion of the ﬂuid is completely determined, if the velocity v is known as a function of time everywhere in the ﬂow ﬁeld v = g(r,t) , or written in components u = g1 (x,y,z,t) ; v = g2 (x,y,z,t) ; w = g3 (x,y,z,t) . (1.22)

(1.23)

If the velocity v is independent of time, the ﬂow is called steady. The Eulerian method oﬀers a better perspicuity than the Lagrangian method and allows a simpler mathematical treatment. Particle Path and Streamline Particle paths designate the ways, the single ﬂuid particles follow in the course of time. They can be determined by integrating the diﬀerential equations: dx dy dz =u ; =v ; =w (1.24) dt dt dt The integrals are identical with the functions f1 , f2 , and f3 , given previously in (1.18). x = y = z = u dt = f1 (x0 ,y0 ,z0 ,t) , v dt = f2 (x0 ,y0 ,z0 ,t) , w dt = f3 (x0 ,y0 ,z0 ,t) (1.25)

y v v streamline α

Streamlines give an instantaneous picture of the ﬂow at a certain time. Streamlines are deﬁned by the requirement that in every point of the ﬂow ﬁeld their slope is given by the direction of the velocity vector. For plane ﬂows there results tan(α) = dy v = dx u . (1.26)

u

In a steady ﬂow particle paths are identical with streamlines.

x

1.3 Hydrodynamics Reference Frame and Form of Motion

7

Certain unsteady motions can be viewed as steady motions with the aid of a coordinate transformation. For example, the ﬂow about the bow of a ship appears to be unsteady to an observer standing on land, while the ship is passing by. The observer sees the single particle paths of the ﬂow. The picture of the streamlines is diﬀerent for every instant of time. However, for an observer on board of the ship the ﬂow about the bow appears to be steady. Streamlines and particle paths are now identical, and the picture of the streamlines does not change in the course of time.

particle path

streamline
Observer at rest (Unsteady motion) Observer aboard of the ship (Steady motion)

1.3.2 Stream Tube and Filament Continuity Equation Streamlines passing through a closed curve form a stream tube, in which the ﬂuid ﬂows. Since the velocity vectors are tangent to the superﬁcies, the ﬂuid cannot leave the tube through the superﬁcies; the same mass is ﬂowing through every cross section. m=ρvA ; ˙ ρ v1 A1 = ρ v2 A2 (1.27)

The product v A is the volume rate of ﬂow. A stream tube with an inﬁnitesimal cross section is called stream ﬁlament. Bernoulli’s Equation (Daniel Bernoulli 1738) It follows from Newton’s law m dv = dt Fa (1.28)

that, if pressure, volume, and friction forces act on a element of a stream ﬁlament, the equilibrium of forces is

ρ dA ds

∂p dv = − ds dA + ρ g cos α ds dA − R dt ∂s

(1.29)

8 with

1. Fluid Mechanics I

ds cos α = −dz and R = there results ∂p dz dv =− −ρg −R . dt ∂s ds If the velocity depends on the path s and on the time t, then the total diﬀerential is ρ ∂v ∂v dt + ds . ∂t ∂s With v = ds/dt the substantial acceleration is dv = R dA ds

(1.30)

(1.31)

(1.32)

(1.33)

dv ∂v ∂v ∂v 1 ∂(v 2 ) = +v = + , (1.34) dt ∂t ∂s ∂t 2 ∂s Therein ∂v/∂t is the local and v(∂v/∂s) the convective acceleration and the diﬀerential equation becomes ∂v ρ ∂v 2 ∂p dz ρ (1.35) + + +ρg = −R . ∂t 2 ∂s ∂s ds With the assumption of inviscid (R = 0), steady (v = v(s)), and incompressible ﬂuid ﬂow, the last equation can be integrated to yield the energy equation for the stream ﬁlament (Bernoulli’s equation 1738). ρ p + v 2 + ρ g z = const. . (1.36) 2 According to this equation the sum of the mechanical energies remains constant along a stream ﬁlament. The equilibrium of forces can be formulated for arbitrary force ﬁelds, if the acceleration vector b is known: ∂v ρ ∂v 2 ∂p ρ + − ρ b cos(b,s) = −R (1.37) + ∂t 2 ∂s ∂s 1.3.3 Applications of Bernoulli’s Equation Measurement of the Total Pressure (Pitot Tube) If in an inviscid ﬂow the velocity vanishes in a point, then this point is called stagnation point.

stagnation point

Pitot tube

It follows from Bernoulli’s equation that the pressure in the stagnation point (total pressure p0 ) is equal to the sum of the static pressure p∞ and of the dynamic pressure ρ u2 /2 of the ∞ oncoming ﬂow. The total pressure can be determined with the Pitot tube, the opening of which is positioned in the opposite direction of the ﬂow.

1.3 Hydrodynamics Measurement of the Static Pressure

9

The so-called Ser’s disc and a static pressure probe are used to measure the static pressure in a ﬂow ﬁeld. In contrast to the Pitot tube relatively large measuring errors result from small angles of attack.

Measurement of the Dynamic Pressure (Prandtl’s Static Pressure Tube) Because of the friction in the ﬂuid the measured dynamic pressure deviates from that of the inviscid ﬂow. The deviation depends on the Reynolds number and the ratio of the diameters d/D. It can be corrected with the factor given in the diagram above.

Prandtl’s static pressure tube combines the static pressure probe and the Pitot probe for the measurement of the dynamic pressure, which can be determined with Bernoulli’s equation from the diﬀerence of the total and the static pressure. p0 − p∞ = β Outﬂow from a Vessel The outﬂow velocity is v2 =
2 v1 + 2 g h +

ρ 2 u 2 ∞

with β = β

Re,

d D

.

(1.38)

2 (p1 − p2 ) ρ

.

(1.39)

For A1 >> A√and with p2 = p1 the outﬂow velocity 2 becomes v2 = 2 g h (Torricelli’s theorem 1644).

The actual outﬂow velocity is smaller, caused by the friction forces. The cross section of the stream, in general, is not equal to the geometric cross section of the opening. The stream experiences a contraction Ψ = Ae /A, which is called stream contraction; it depends on the shape of the outﬂow opening and on the Reynolds number.

10

1. Fluid Mechanics I

The volume rate of ﬂow is then ˙ Q=ΨA 2gh . (1.40)

Measurement of the Volume Rate of Flow in Pipes The volume rate of ﬂow of a steady pipe ﬂow can be obtained by measuring a pressure diﬀerence. A suﬃciently large pressure diﬀerence must be enforced by narrowing the cross section. If m = A2 /A1 designates the area ratio, one obtains for the velocity in the cross section 2 v2 = 2 (p1 − p2 ) 1 (1 − m2 ) ρ (1.41)

and for the volume rate of ﬂow ˙ Qth = v2 A2
Venturi nozzle

.

(1.42)

In technical applications mainly the Ventury nozzle, the standard nozzle, and the standard oriﬁce are used for measuring the volume rate of ﬂow. The inﬂuence of the friction in the ﬂuid, of the area ratio, and of the shape of the contraction is taken into account in the discharge coeﬃcient α. ˙ Q = α A2 2 ∆pw ρ (1.43)

standard nozzle

standard orifice

The pressure diﬀerence ∆pw is called diﬀerential pressure. The dimensions, the positions of the pressure measurements, the directions for installing, the tolerances, and the discharge coeﬃcients are laid down in the ﬂow-measurement regulations.

1.4 One-Dimensional Unsteady Flow
If the velocity in a stream tube not only changes with the path length s but also with the time t, then the volume rate of ﬂow also changes with time. Since ˙ Q(t) = v(s,t) A(s) , it follows that ˙ ∂v ∂Q = A . ∂t ∂t (1.45) (1.44)

1.4 One-Dimensional Unsteady Flow

11

Consequently, the interdependence between pressure and velocity is changed. If all other assumptions remain the same as in the derivation of Bernoulli’s equation, one obtains the energy equation for unsteady ﬂow for a stream ﬁlament to ρ ∂v ρ ds + p + v 2 + ρ g z = k(t) . (1.46) ∂t 2 If the integral over the local acceleration is small in comparison to the other terms in the above equation, the ﬂow is called quasi-steady. Oscillation of a Fluid Column A ﬂuid oscillates in a U-tube after displacement from its equilibrium position by the amount ξ0 , under the inﬂuence of gravity. The energy equation for unsteady inviscid ﬂow gives ρ g ξ = −ρ g ξ + ρ dv l dt . (1.47)

With v = −dx/dt the diﬀerential equation describing the oscillation is d2 ξ 2 g + ξ = 0, dt2 l and with the initial conditions v(t = 0) = 0, ξ(t = 0) = ξ0 the solution is 2g . l Therein α is the eigenfrequency of the oscillating ﬂuid column. ξ = ξ0 cos(α t) with α = Suction process in a plunger pump The pumping process of a periodically working plunger pump can be described with the energy equation for unsteady ﬂows with some simplifying assumptions. In order to avoid cavitation the pressure in the intake pipe should not fall below the vapor pressure pV . During the suction stroke the pressure attains its lowest value at the piston-head. If it is assumed, that the velocity only depends on the time t, the pressure at the piston-head pP H is obtained to pa − ρ g h ξ0 pP H = + ρ ω 2 l2 ρ ω 2 l2 l cos(ω t) − 1 ξ0 2 l 1 − 3 cos2 (ω t) . (1.51) (1.50) (1.49) (1.48)

In general the piston stroke ξ0 is much smaller than the length of the intake pipe l; with this assumption the angular velocity ωV , at which the pressure at the piston-head reaches the value of the vapor pressure pP H = pV , is ωV = pa − ρ g h − pV ρ ξ0 (l − ξ0 ) ⇒ . (1.52)

The mean volume rate of ﬂow is 2 π/ω ω ˙ Q= v A dt 2 π π/ω

ω ˙ Q = ξ0 A π

.

(1.53)

12

1. Fluid Mechanics I

1.5 Momentum and Moment of Momentum Theorem
The momentum theorem describes the equilibrium between the time rate of change of momentum and external forces and the moment of momentum theorem the equilibium between their moments. For steady as well as for time-averaged ﬂows these theorems involve only the ﬂow conditions on the boundaries of a bounded ﬂuid domain. 1.5.1 Momentum Theorem According to the momentum theorem of mechanics the time rate of change of the momentum is equal to the sum of the acting external forces dI = dt F . (1.54)

For a system with n particles with masses mi and velocities v i it follows with n I= i=1 mi v i

(1.55)

d dt

n

mi v i = i=1 F

.

(1.56)

If the particles are assumed to form a continuum with density ρ(x,y,z,t) the sum changes into a volume integral. The rate of change of momentum is then d dI = dt dt τ (t)

ρ v dτ

.

(1.57)

The volume τ , which always contains the same particles, changes in a time interval from τ (t) to τ (t + ∆t). d dt ρ v dτ = lim 1 ∆t ρ v (t + ∆t) dτ − ρ v(t) dτ (1.58) (1.59) (1.60)

τ (t)

∆t→0

τ (t+∆t)

τ (t)

ρ v (t + ∆t) = ρ v(t) + dI = dt τ (t)

∂ (ρ v) dτ + lim ∆t→0 ∂t

∂(ρ v) ∆t + ... ∂t 1 ρ v dτ ∆t ∆τ (t)

The last integral can be changed into a surface integral over the surface A(t). Since dτ = (v · n) dA ∆t , it follows that dI = dt τ (t)

(1.61)

∂ (ρ v) dτ + ∂t

A(t)

ρ v(v · n) dA . (1.62)

For steady ﬂows the time rate of change of momentum is given by the surface integral of the last equation. The surface A of the volume τ considered is called control surface. The external forces, which are in equilibrium with the time rate of change of momentum, are volume and surface forces, for example, volume forces due to the gravitational force:

1.5 Momentum and Moment of Momentum Theorem Fg = τ 13 (1.63)

ρ g dτ

.

The forces which act on the surface are given by the pressure and friction forces. The pressure force is described by the integral Fp = − p n dA .
A

(1.64)

The friction forces are given by the surface integral extended over the components of the stress ¯ tensor σ Ff = −
A

¯ (σ · n) dA .

(1.65)

If a part of the control surface is given by a rigid wall, then a supporting force F s is exerted by the wall on the ﬂuid. The supporting force is equal to the force the ﬂuid exerts on the wall, but acts in the opposite direction. The momentum theorem for steady ﬂows then reads
A

ρ v (v · n) dA = F g + F p + F s + F f

.

(1.66)

The diﬃcult part in the construction of the solution of the momentum theorem mainly consists in the solution of the integrals. If possible, the control surface A has to be chosen in such a way, that the integrals given in (1.63) to (1.66) can be solved. In order to obtain a unique solution the control surface A must be a simply connected surface. 1.5.2 Applications of the Momentum Theorem Force on a Bent Pipe The ﬂow through a horizontal bent pipe is assumed to be inviscid and incompressible. Inlet and outlet cross section, the pressure in the inlet cross section, the external pressure pa , which also prevails in the outlet cross section, and the ﬂow deﬂection angle β have to be known for the solution of the problem. It is advantageous, to choose the control surface as indicated in the sketch by the dashed line. The velocities in the inlet and outlet cross section are determined with Bernoulli’s equation and the continuity equation:

v1 =

2 (p1 − pa ) ρ

1
A1 A2 2

−1

and v2 = v1

A1 A2

.

(1.67)

From the momentum theorem it follows for the x-direction −ρ v1 2 A1 + ρ v2 2 A2 cos β = (p1 − pa ) A1 + Fsx and for the y-direction −ρ v2 2 A2 sin β = Fsy . (1.69) , (1.68)

The supporting force Fs can be determined from the last two equations.

14

1. Fluid Mechanics I

Jet Impinging on a Wall A horizontal plane ﬂuid jet impinges on a plate under the angle β and is deﬂected to both sides without losses. It follows from Bernoulli’s equation that the velocity at both ends of the plate is equal to the jet velocity v1 , if it is assumed, that the streams leave the plate parallel to it. Then the two components of the supporting force are
2 Fsx = −ρ v1 b1 cos β

and Fsy = 0 .

(1.70)

The widths of the streams are b2 = b1 1 + sin β 2 and b3 = b1 1 − sin β 2 . (1.71)

Discontinuous Widening of a Pipe If a pipe is discontinuously widened from the cross-sectional area A1 to A2 , the ﬂuid cannot ﬁll the entire cross-sectional area A2 , when entering the widened part of the pipe. Dead water regions are formed in the corner, which extract momentum with the aid of internal friction from the ﬂuid passing by. This loss of momentum results in a pressure loss. The pressure in the entrance cross section also acts on the frontal area A2 − A1 of the widened part of the pipe. The losses in the ﬂow can be determined with the continuity equation, Bernoulli’s equations, and the momentum theorem, if it is assumed, that downstream from the dead water region the ﬂow properties are constant in every cross section and that the friction caused by the walls can be neglected. The pressure loss is obtained with the control surface shown in the sketch above by the dashed line to ρ 2 ρ 2 (1.72) ∆pl = p01 − p02 = (p1 + v1 ) − (p2 + v2 ) , 2 2 with v1 A1 = v2 A2 and ρ v2 2 A2 − ρ v1 2 A1 = (p1 − p2 ) A2 and ﬁnally to ∆pl = ρ 2 v 2 1 1− v2 v1
2

,

(1.73)

,

(1.74)

.

(1.75)

The pressure loss, nondimensionalized with the dynamic pressure, is called pressure loss coefﬁcient. ζ= A1 ∆pl ρ 2 = 1− v A2 2 1
2

Carnot’s equation

(1.76)

1.5 Momentum and Moment of Momentum Theorem Pressure Loss in an Oriﬁce

15

After what has been said about the ﬂow through the pipe with a discontinuous widening of the cross section, it can be concluded that the ﬂow through an oriﬁce also must generate a pressure loss. It again can be determined with Carnot’s equation. During the passage of the ﬂuid through the oriﬁce the ﬂow contracts and forms a jet, which depends on the shape of the A oriﬁce. (geometric opening ratio m = A1 ). The ratio of the cross section of the bottle neck to the opening cross section of the oriﬁce is called contraction ratio Ψ = A2 . The pressure coeﬃcient A of the oriﬁce, referenced to the conditions of the oncoming ﬂow in the pipe is ζo = ∆pl ρ 2 = v 2 1 1−ψm ψm
2

.

(1.77)

In the following table the pressure loss coeﬃcient of the oriﬁce is given for some characteristic values of the product Ψ m. Ψ m ζo 1 2/3 1/2 1/3 0 1/4 1 4

If the widening of the cross-sectional area is smooth, the losses just discussed can be very much reduced. Then the pressure rises in the ﬂow direction, since the velocity decreases with increasing cross-sectional area. If the opening angle, however, becomes too large, then the ﬂow cannot follow the contour of the widened pipe any longer and a dead-water region similar to the one mentioned previously is generated. Resistance of an Installation in a Pipe For a pipe with constant cross-section area the continuity equation yields v1 = v2 . The resistance of the body installed results in a pressure drop. The control surface is indicated by the dashed line in the sketch. With the friction forces neglected the resistance is Fw = (p2 − p1 ) A. Rankine’s Slip-Stream Theory Performance, thrust, and eﬃciency of wind-driven rotors and propellers (of ships and airplanes) can be determined with the momentum theorem for one-dimensional ﬂow under the following simplifying assumption: The rotation of the ﬂow in the slip stream does not inﬂuence the axial ﬂow velocity; the driving force is uniformly distributed over the cross section of the slip stream, independent of the number of vanes (inﬁnite number of vanes); the ﬂow is decelerated and accelerated without losses. The following ﬁgure shows the ﬂow through a wind-driven rotor. The following relations are valid for the stream tube with p1 = p2 = pa (atmospheric pressure far away from the rotor). (1.78)

16

1. Fluid Mechanics I ρ v1 A1 = ρ v A = ρ v2 A2 ρ 2 ρ 2 v1 + pa = v + p1 2 1 2 ρ 2 ρ 2 v2 + pa = v + p2 2 2 2

(1.79)

streamtube

Continuous velocity variation in the slip stream. Discontinuous pressure change in the cross-sectional plane of the rotor. The force F exerted by the rotor on the ﬂow in the slip stream is determined from the momentum theorem for the small control volume between the cross section 1 and 2 , immediately upstream and downstream from the cross-sectional plane of the rotor (dark area). F = (p2 − p1 ) A (1.80)

The force F can also be obtained for the large control surface,, indicated by the dashed line in the last ﬁgure F = ρ v A (v2 − v1 ) . (1.81)

The velocity in the cross-sectional plane of the rotor is obtained with the aid of Bernoulli’s equation: v = (v1 + v2 ) 2 (Froude’s Theorem 1883) (1.82)

The energy, which can be extracted from the slip stream per unit time is P = ρ A v1 3 4 1+ v2 v1 1−
2 v2 2 v1

. v2 v1

(1.83) = 1 . The maximum 3

The extracted power P attains a maximum value for the velocity ratio extracted power, divided by the cross-sectional area of the rotor, is Pmax 8 3 ρ v1 = A 27 and the corresponding thrust per unit area is F 4 2 = − ρ v1 A 9 . ,

(1.84)

(1.85)

For air with ρ = 1.25kg/m3 the following values are obtained for the maximum extracted power and the corresponding thrust per unit cross-sectional area and time, computed for the wind intensities listed. v1 [m/s] Wind intensity [BF ] Pmax /A [kW/m2 ] F/A [N/m2 ] 1 5 10 15 20 25 30 1 3 6 7 9 10 12 0.00037 0.0463 0.370 1.25 2.963 5.79 10.0 0.555 13.88 55.55 125 222 347 500

1.5 Momentum and Moment of Momentum Theorem 1.5.3 Flows in Open Channels

17

Flows in rivers and channels are called open channel ﬂows. They diﬀer from pipe ﬂows by their free surface, which is exposed to the atmospheric pressure. For a given volume rate of ﬂow and a given width b of the channel, the depth of the water h can change with the velocity. If inviscid steady ﬂow is assumed, Bernoulli´s equation yields the speciﬁc relation for each streamline v2 + h + z = const. 2g (1.86)

The sum of the velocity height v 2 /(2 g) and the depth of water h is called energy height H. The velocity of the water is assumed to be independent of z. With v= there results bottom ˙ Q bh

(1.87)

H =h+

˙ Q2 2 g h2 b2

.

(1.88)

If the volume rate of ﬂow and the energy height H are given, (1.88) gives two physically meaningful solutions for the depth of water h and thereby also for the velocity. These two diﬀerent ﬂow conditions can be found with the relation H = H(h) on both sides of the minimum Hmin. = The corresponding critical depth of water is hcrit. = and the critical velocity is vcrit. = g hcrit. . (1.91)
3

3 2

3

˙ Q2 g b2

(1.89)

˙ Q2 g b2

.

(1.90)

The dimensionless ratio H/Hmin. is
⎡

H h 2 1 = ⎣ + Hmin. 3 hcrit. 2

hcrit. h

2

⎤ ⎦

.

(1.92)

√ The quantity c = g h is the velocity of propagation of gravitational waves in shallow water, and the ratio v/c is called Froude number v . (1.93) c The magnitude of the Froude number determines, which of the two ﬂow conditions prevails. Fr = F r < 1 h > hcrit. F r > 1 h < hcrit. v < vcrit. v > vcrit. subcritical condition supercritical condition

For F r > 1 small disturbances cannot travel upstream. According to Bernoulli’s equation the sum of the geodetical elevation of the bed and the energy height is constant. If the geodetical

18

1. Fluid Mechanics I

bed elevation is suﬃciently increased (undulation of ground), a subcritical open channel ﬂow changes into supercritical motion. If H is the energy height of the channel, the necessary increase of the geodetical bed elevation is Zcrit. = H − Hmin. . (1.94)

Also the process in the opposite direction is observed. A supercritical motion (h < hscrit. ) changes into a subcritical motion with an almost discontinuous increase of the water level (hydraulic jump). This jump is associated with ﬂuid mechanical losses. The water level h2 after the jump can be determined with the momentum theorem and the continuity equation. For an open channel with constant width there results
2 2 ρ (v2 h2 − v1 h1 ) = ρ g

h2 h2 1 − 2 2 2

(1.95)

and h1 v1 = h2 v2 The ratio of the water levels is h2 = h1
2 1 2 v1 1 + − 4 g h1 2

.

(1.96)

.

(1.97)

A decrease of the water level during the jump (h2 /h1 < 1) is not possible, since then the energy height would have to be increased. The diﬀerence of the energy heights H1 − H2 H1 − H2 = h1 4 h1 h2 h2 −1 h1
3

(1.98)

is positive (in the limiting case zero), if h2 /h1 ≥ 1. The hydraulic jump can only occur in supercritical motion. 1.5.4 Moment of Momentum Theorem According to the fundamental theorem of mechanics the time rate of change of the moments of momentum is equal to the sum of the acting external moments. For a system of n particles with masses mi , velocities v i , and the distances r i from a space-bound axis there is obtained d n r i × (mi v i ) = dt i=1 M . (1.99)

Similar to the derivation of the momentum theorem the transition from the particle system to the continuum is achieved by substituting the sum on the left-hand side of (1.99) by a volume integral dL d = dt dt ρ r × v dτ . (1.100)

τ (t)

The moment of momentum is designated as L. Its time rate of change is equal to the sum of the moments of the external forces acting on the ﬂuid considered, referenced to a space bound axis.

1.5 Momentum and Moment of Momentum Theorem For steady ﬂows the volume integral can again be replaced by a surface integral dL = dt ρ (r × v) (v · n) dA .

19

(1.101)

A

The moment of the external forces results from the moments of the volume forces, the pressure forces, the friction forces, and the supporting force: Mg = (r × ρ g) dτ
A

,

τ

Mp = − Mf = −

p (r × n) dA , ¯ r × (σ · n) dA , (1.102)

A

Ms = r s × F s 1.5.5 Applications of the Moment of Momentum Theorem Euler’s Turbine Equation (1754)

If a ﬂuid ﬂows through a duct rotating with constant angular velocity from the outside to the inside in the radial direction (radial turbine),, the moment generated by the ﬂow can be computed with the moment of momentum theorem. The ﬂow is steady with respect to the duct, the walls of which form the control volume. The mass ﬂowing through the duct per unit time is m = ρ v1 A1 sin δ1 = ρ v2 A2 sin δ2 ˙ . (1.103)

With the notation given in the sketch the time rate of change of the angular momentum is ρ (r × v) (v · n) dA = k [−m v1 r1 cos δ1 + m v2 r2 cos δ2 ] ˙ ˙ . (1.104)

A

The moment delivered to the turbine shaft is (Moment of reaction) Md = m [v1 r1 cos δ1 − v2 r2 cos δ2 ] . (1.105) ˙ This relation is called Euler’s turbine equation. The power delivered to the turbine is with the relations u1 = r1 ω and u2 = r2 ω P = Md ω = m (v1 u1 cos δ1 − v2 u2 cos δ2 ) (1.106) ˙ . The largest power output is obtained, when the absolute velocity v2 is normal to the circumferential velocity component u2 , i. e. if cos δ2 = 0. Segner’s Water Wheel (1750) If a ﬂuid ﬂows from a reservoir into a doubly bent pipe, pivoted about its axis, as sketched in the following drawing, the pipe will start to rotate. The ﬂow generates a moment of rotation, which can be picked up at the pipe. A certain part of this moment is used to overcome the bearing friction. If the ﬂuid motion through the pipe is steady, the moment can be determined with the moment of momentum theorem.

20

1. Fluid Mechanics I

With the notation given in the sketch below the out-ﬂowing mass is m = ρ vr A ,and ˙ M = m (vr − ω R) R . ˙

(1.107)

In order to be able to determine the out-ﬂow velocity vr , the ﬂow between the ﬂuid surface in the reservoir and the exit cross section is assumed to be loss-free. The energy equation for the system considered is
4

pa +

1

ρ b ds = pa +

ρ 2 v 2 r

.

(1.108)

The integral can be split into two parts
4 1

ρ b · ds =

2 1

3

ρ g dz +

2

ρ ω 2 r dr

;(1.109)

therein ω 2 r is the centrifugal acceleration. After integration the out-ﬂow velocity is obtained to vr = 2 g h + ω 2 R2 . (1.110)

With the following abbreviations one obtains for M ωR ξ=√ 2gh and M0 = 2 ρ g h A R (1.111)

M = M0

1 + ξ2

1 + ξ2 − ξ

.

(1.112)

The quantity M0 is the starting moment. If the dependence of the friction moment on the rotational speed is known, the moment Md delivered to the rotating pipe can be determined.

1.6 Parallel Flow of Viscous Fluids
When a ﬂuid is deformed a part of the kinetic energy of the ﬂow is converted into heat (internal friction). For example in pipe ﬂow, the internal friction results in a pressure drop in the ﬂow direction. From a macroscopic point of view, the ﬂuid ﬂows in layers (lat. lamina), and the ﬂow is called laminar ﬂow. The velocity changes from layer to layer, and in the limiting case of inﬁnitesimally thin layers, a continuous velocity proﬁle results. The ﬂuid layers ﬂow past each other and the molecular momentum exchange between them causes tangential stresses,, which are closely related to the velocity gradients. They can be described with a phenomenological ansatz. The form of the ansatz depends on kind of ﬂuid (viscosity law). In the close vicinity to a rigid wall the molecules of the ﬂuid loose the tangential component of the momentum to the bounding surface, and as a consequence the ﬂuid adheres to rigid walls (Stokes’ no-slip condition, 1845):

1.6 Parallel Flow of Viscous Fluids 1.6.1 Viscosity Laws Newtonian Fluid

21

Newtonian ﬂuids are ﬂuids in which the tangential stresses are linearly proportional to the velocity gradients. This dependence can be illustrated with the following experiment: The space between two parallel plates is ﬁlled with a Newtonian ﬂuid. If the upper plate is moved with the velocity uw parallel to the lower plate, as shown in the following sketch, then the velocity increases linearly in the y-direction, and the particles in the superjacent layers move with diﬀerent velocities. u(y) = uw y h (1.113)

The angle of shear can be determined from their displacement: ∆γ ≈ − uw ∆t h (1.114)

The rate of strain is obtained by forming the diﬀerential quotient γ = lim ˙ In general, for velocity distributions the rate of strain is γ=− ˙ du dy . (1.116) uw ∆γ =− ∆t h . (1.115)

∆t→0

The relation between the rate of strain and the tangential or shear stress is obtained by a comparison with a shearing test with a rigid body. Shear Experiment The shear stress is proportional to the rate of strain. The constant of proportionality is the dynamic shear viscosity µ; it depends on the medium, the pressure, and on the temperature. The ratio ν = µ is called kinematic viscosity. ρ

Rigid body τ = f (γ); γ = shear action Hooke’s law: τ = G γ

Fluid τ = f (γ); ˙ γ = rate of strain ˙ Newton’s viscosity law: τ = µ γ ˙

The following two diagrams show the temperature dependence of the dynamic and the kinematic viscosity of water and air at atmospheric pressure.

22

1. Fluid Mechanics I

dynamic shear viscosity ν = f (T )

kinematic viscosity ν = f (T )

Non-Newtonian Fluids Many ﬂuids (for example high-polymeric ﬂuids and suspensions) do not follow Newton’s viscosity law. In order to be able to describe the diﬀerent ﬂow processes with simple relations, numerous empirical model laws were proposed. The Bingham model describes the ﬂow process of ﬂuids, which below a certain shear stress behave as a rigid body (tooth paste) τ = µ γ ± τ0 ˙ . (1.117)

For |τ | < τ0 ist γ = 0. The Ostwald-de Waele ˙ model can describe nonlinear ﬂow processes τ = η | γ |n−1 γ ˙ ˙ . (1.118)

For n = 1 this model is identical with Newton’s viscosity law. The three models are shown in the diagram. The deviation of n from unity indicates the deviation of the ﬂuid from Newtonian behavior. For n < 1, the behavior is called pseudoplastic, for n > 1 the ﬂuid is called dilatant. 1.6.2 Plane Shear Flow with Pressure Gradient In the shear experiment described only shear stresses act in the ﬂuid. In the following example it is shown, how normal stresses together with shear stresses inﬂuence the ﬂow. In order to simplify the derivation it is assumed, that the normal stresses are caused by a pressure change in the x-direction, and that the shear stresses change only in the y-direction. In this parallel shear ﬂow between two plates τ = τ (y) p = p(x) (1.119)

the shear stress is assumed to be positive in the x-direction, if the normal of the bounding surface points in the negative y-direction. Assume that L is the length in the x-direction, over which the pressure changes from p1 to p2 . The equilibrium of forces can then be written down for Newtonian and non-Newtonian ﬂuids. The velocity distribution is obtained by inserting the viscosity law and integration in the y-direction. For a Newtonian ﬂuid with τ = −µ du dy (1.120)

1.6 Parallel Flow of Viscous Fluids there results from the equilibrium of forces dp dτ =0 . + dx dy Integration yields τ (y) = (p1 − p2 ) y + c1 L .

23

(1.121)

(1.122)

The velocity distribution is u(y) = p2 − p 1 2 y + c1 y + c2 2µL . (1.123)

The constants of integration c1 and c2 are obtained from the boundary conditions. y=0: u=0 y = h : u = uw Stokes’ no-slip condition (1.124)

u(y) = h2

p2 − p1 2µL

y h

2

−

y y (1.125) . + uw h h

The velocity distribution is determined by the wall velocity uw and the pressure diﬀerence in the x-direction p1 − p2 .

u >0 1 w

p1 − p2 = 0 p1 − p2 > 0 p1 − p2 < 0

u =0 2 w u >0 4 w

u > 0 p1 − p2 = 0 3 w

The wall shear stresses are obtained by diﬀerentiation h p 1 − p 2 uw + L 2µ h h p 1 − p 2 uw y = h : τw = µ − L 2µ h y = 0 : τw = −µ The volume rate of ﬂow is ˙ Q = b h 0

(1.126) . (1.127)

u(y) dy = h3

p 1 − p 2 uw h + 12 µ L 2

.

(1.128)

24

1. Fluid Mechanics I

1.6.3 Laminar Pipe Flow A pipe with circular cross section and radius R is inclined at the angle α. Laminar ﬂow ﬂows though it. It is assumed that the shear stress depends only on the radial coordinate. The equilibrium of forces yields − dp 1 d + ρ g sin α − (τ r) = 0 . dx r dr (1.129)

The shear stress is obtained by integration τ =+ r 2 p1 − p2 + ρ g sin α L . (1.130)

For a Newtonian ﬂuid with τ = −µ du the velocity distribution for the Stokes’ no-slip condition is dr R2 4µ p1 − p2 + ρ g sin(α) L 1− r R
2

u(r) =

.

(1.131)

The velocity attains its maximum value at the axis of the pipe (r = 0) umax = R2 4µ p1 − p2 + ρ g sin α L . (1.132)

The volume rate of ﬂow through the pipe is ˙ Q=
R 0

u(r) 2 π r dr =

π R4 8µ

p1 − p2 + ρ g sin α L

(1.133)

(Derived by Hagen and Poiseuille about 1840 for α = 0). With the volume rate of ﬂow a mean velocity um can be deﬁned um = ˙ Q R2 = A 8µ p1 − p2 umax + ρ g sin α = L 2 . (1.134)

The pressure diﬀerence p1 − p2 is a measure for the wall shear stress, when the gravitational force can be neglected. τw = R p1 − p2 2 L . (1.135)

The pressure diﬀerence referenced to the dynamic pressure of the mean velocity is p1 − p2 64 µ l = ρ 2 um ρ um D 2 2 In the last equation the dimensionless expression 64 8 τw = ρ u2 Re m ρ um D µ

. is the Reynolds number Re,

(1.136)

.

(1.137)

The quotient 8 τw /(ρ u2 ) is called the pipe friction coeﬃcient λ. It is proportional to the wall m shear stress, nondimensionalized with the dynamic pressure of the mean velocity. λ= 8 τw ρ u2 m . (1.138)

1.7 Turbulent Pipe Flows Flow in the Intake Region of a Pipe

25

If the velocity depends only on the radial coordinate r, as for example in the Hagen-Poiseuille law, then the pipe ﬂow is called fully developed. This ﬂow condition is reached only at the end of the intake length Li , which is approximately Li = 0.029 Re D . (1.139)

In the intake region the velocity proﬁles change as indicated in the following drawing.

In the intake region an additional pressure loss arises, which can be described by a loss coeﬃcient (ζi ≈ 1.16) p1 − p2 L = λ + ζi ρ 2 um D 2 . (1.140)

In long pipes the intake losses can be neglected, but they have to be taken into account in viscosity measurements in capillary viscometers. The viscosity can be determined with the aid of the equation derived by Hagen and Poiseeuille, (1.133) for α = 0: The validity of this equation was conﬁrmed with extreme accuracy, so that it can be used for viscosity measurements. The experimental set up consists of a pipe with a small diameter and a large diameter-to-length ratio, with the axis positioned horizontally. Then the gravitational acceleration acts normal to the ﬂow direction, and the pressure drop is solely proportional to the dynamic shear viscosity µ and iversely proportional to the volume rate of ﬂow. The viscosity can be determined by measuring the pressure drop and the volume rate of ﬂow. In order to avoid errors, caused by the variation of the velocity proﬁle in the intake region, the pressure must be measured further downstream. This can be done by computing the entrance length with (1.139) and comparing it with the distance between the position of the boreholes for the pressure measurements and the entrance of the pipe.

1.7 Turbulent Pipe Flows
The Hagen-Poiseuille law looses its validity when the Reynols number exceeds a certain value. Experiments show, that irregular velocity ﬂuctuations set in, which cause an intensive intermixing of the various layers in the ﬂow. The momentum exchange normal to the axis of the pipe increases markedly. The proﬁles of the timeaveraged velocity become fuller, and the wall shear stress increases. The pressure drop no longer is proportional to ˙ ˙ Q, but approximately to Q2 . The ﬂow is then called turbulent. In technical pipe ﬂows the transition from laminar to turbulent ﬂow in general occurs at a Reynolds number Re = 2300. Under special experimental conditions pipe ﬂows can be kept laminar up to Reynolds number Re = 20000 and higher, with the diameter of the pipe again used as the reference length. The ﬂow is then very susceptible to small perturbations and is diﬃcult to maintain.

26

1. Fluid Mechanics I

1.7.1 Momentum Transport in Turbulent Flows Following Reynolds (1882) the velocity in turbulent ﬂows can be described with an ansatz, which contains a time-averaged value and a ﬂuctuating part. For the fully developed turbulent pipe ﬂow the splitting of the component in the axial direction yields u(r,Φ,x,t) = u(r) + u (r,Φ,x,t) , ¯ and normal to it v(r,Φ,x,t) = v (r,Φ,x,t) . (1.142) (1.141)

The time-averaged value of the velocity is determined in such a way, that the time-averaged value of the ﬂuctuations vanishes. In general the time-averaged mean value of the squares and the product of the two components (correlation) of the ﬂuctuation velocity do not vanish. When they are multiplied by the density, they have the dimension of stress. The time-averaged correlation can be interpreted as the transport of momentum per unit area in the radial direction, while the squares represent normal stresses. The turbulent momentum transport is mainly determined by the correlation of the velocity ﬂuctuations uv = 1 T
T 0

(u v ) dt.

(1.143)

This result is obtained with the momentum theorem, written for the control surface indicated by the solid line shown in the drawing above. τ ∂ ρ v dτ + ∂t

A

ρ v (v · n) dA = F p + F f

(1.144)

With the velocity v = i (¯ + u ) + j v u The time-averaged momentum equation is obtained to ρ
A

(1.145)

u v dA = 2 π r L ρ u v = π r2 (p1 − p2 ) − τ 2 π r L ,

(1.146)

which yields with Newton’s viscosity law (p2 − p1 ) r d¯ u = −ρ u v + µ 2L dr (1.147)

The quantity ρ u v is called the apparent or turbulent shear stress. If the velocity proﬁle u(r) ¯ is to be determined, the correlation u v has to be known. Since the apparent stresses cannot be obtained from (1.147), an additional hypothesis has to be introduced, which constitutes a relation connecting the velocity correlation with the time-averaged velocity. Prandtl’s Mixing-Length Hypothesis In his hypothesis about the turbulent momentum transport Prandtl assumes, that small agglomerations of ﬂuid particles move relative to the surrounding ﬂuid in the main ﬂow direction and normal to it and exchange momentum with their surroundings.

1.7 Turbulent Pipe Flows

27

The distance along which the agglomerations loose their excess momentum is called mixing length l. Normal to the direction of the main ﬂow the change of velocity between two layers, by the distance l apart from each other, is ∆u = u(y + l) − u(y) ≈ l ¯ ¯ d¯ u dy . (1.148)

This velocity diﬀerence can be assumed to be equal to the velocity ﬂuctuation u u =l d¯ u dy . (1.149)

The ﬂuctuations in the main ﬂow direction cause ﬂuctuations of the normal velocity component v , which are of equal order of magnitude. A positive ﬂuctuation u most of the time originates a negative ﬂuctuation v , as can be reasoned from continuity considerations. Hence v = −cu , (1.150)

where c is a positive constant, which can be included in the mixing length. The turbulent shear stress τt then follows as the mean value of the product of both ﬂuctuations u and v τt = ρ u v = ρ l 2 d¯ d¯ u u dy dy . (1.151)

The absolute value of d¯/dy is introduced, in order to ensure, that the sign of the turbulent u shear stress changes with the sign of the velocity gradient. 1.7.2 Velocity Distribution and Resistance Law With the equation for the turbulent shear stress, (1.151), the momentum equation for the turbulent pipe ﬂow reads with a shift of coordinates to y = R − r − p2 − p1 d¯ u d¯ d¯ u u (R − y) = µ + ρ l2 2L dy dy dy . (1.152)

For y = 0 the right-hand side of (1.152) represents the wall shear stress τw . The turbulent shear stress vanishes at the wall, as the velocity ﬂuctuations have to vanish there because of Stokes’ no-slip condition. The mixing length l therefore has also to vanish in the vicinity of the wall. Prandtl assumed, that in the layer nearest to the wall the mixing length is proportional to the distance from the wall, l=ky . (1.153)

In the above equation k is a constant. Aside from a very thin layer in the immediate vicinity of the wall, the turbulent shear stress is much larger than the laminar contribution, and hence the latter can be neglected in the integration of (1.152). If ﬁnally another of Prandtl’s assumptions is introduced, according to which the turbulent shear stress τt remains constant and is equal to the wall shear stress τw , then, after integration of (1.152), the universal law of the wall for turbulent pipe ﬂow is obtained y u∗ u ¯ 1 +C = ln u∗ k ν . (1.154)

The quantity u∗ = τw /ρ is called friction velocity The constant k = 0.4 is named after von K´rm´n; k and C were determined from experiments. For technically smooth pipes C = 5.5. a a

28

1. Fluid Mechanics I

The logarithmic form of the universal law of the wall shows, that it looses its validity in the immediate vicinity of the wall. Since the mixing length vanishes for y approaching zero, the shear stress in the viscous sublayer is solely given by µ d¯/dy. u With τ = τw = const., the velocity in the viscous sublayer is y u∗ u ¯ = u∗ ν . (1.155)

Although the logarithmic velocity distribution was derived for the region near the wall, the comparison with experimental data shows, that it is approximately valid over the entire cross section. By eliminating the constant of integration C, with u = umax (y = R) there is obtained ¯ ¯ R umax − u ¯ ¯ 1 = ln u∗ k y . (1.156)

The time-averaged velocity averaged over the cross section is um ¯ um ¯ umax ¯ = − 3.75 . u∗ u∗ (1.157)

Universal law of the wall

Time-averaged dimensionless radial velocity disy tribution of the turbulent pipe ﬂow with R = r 1 − R.

Experimental data also conﬁrm (1.157), if the constant 3.75 is replaced by the value 4.07. With the deﬁnition of the pipe friction coeﬃcient λ= 8 u∗ 2 u2 ¯m (1.158)

and the universal law of the wall, (1.158) can be cast into the form √ 1 √ = 2.035 log(Re λ) − 0.91 . λ Equation (1.159) agrees with experimental data, if the constants are changed so that √ 1 √ = 2.0 log(Re λ) − 0.8 . λ (1.160) (1.159)

This law is called the Prandtl universal resistance law. For Reynolds numbers up to Re = 105 also the simple relation of Blasius (1911) can be used λ= 0.316 Re0.25 . (1.161)

The pipes for which these laws are valid, have to be hydraulically smooth.

1.7 Turbulent Pipe Flows Rough Pipes

29

Pipes used in technical applications, in general are not hydraulically smooth. The roughness of the wall inﬂuences the ﬂow, such that the pipe friction coeﬃcient is no longer given by Prandtl’s law. The friction coeﬃcient then depends on the number of roughness elements per unit area, on their shape, and on their distribution. In order to be able to characterize the roughness in a simpliﬁed manner, a parameter k/R (relative roughness) is introduced, which is coordinated with a more accurately deﬁned sand roughness ks /R, such that the pipe friction coeﬃcients agree. The diagram shows the inﬂuence of the relative sand roughness on the pipe friction coeﬃcient. According to this diagram, above a certain Reynolds number the friction coeﬃcient λ depends only on ks /R. For completely rough pipes the quadratic resistance law is valid and the pipe friction coeﬃcient can be determined by the following relation. 1 R √ = 2.0 log + 1.74 ks λ 1.7.3 Pipes with Non-circular Cross Section Experimentally determined resistances of turbulent ﬂows in pipes with non-circular cross sections agree well with the resistance law for circular pipes, if in the equation for the pressure loss and in the deﬁnition of the Reynolds number the hydraulic diameter dh is introduced. 4A U A stands for the cross-sectional area and U for the circumference. L ρ 2 ρ um dh ¯ ∆pl = λ u ¯ . Reh = dh 2 m µ dh = (1.163) (1.162)

(1.164)

As the following diagram shows, the experimental data λ(Reh ) follow the Blasius law for turbulent ﬂows. The pressure loss of turbulent ﬂows in open channels can be obtained in the same manner. The hydraulic diameter is then determined from the cross-sectional area of the ﬂow and the wetted circumference of the channel.

Friction coeﬃcient for laminar and turbulent pipe ﬂow with non-circular cross section

30

1. Fluid Mechanics I

The hydraulic diameter looses its meaning for laminar ﬂows. For laminar ﬂows with non-circular cross section the pipe friction coeﬃcient is determined from the law λ = C/Re, where the shape of the cross section is accounted for in the constant C.

2. Fluid Mechanics II

2.1 Introduction
A ﬂow is completely determined if the velocity vector v and the thermodynamic properties, the pressure p, the density ρ, and the temperature T are known everywhere in the ﬂow ﬁeld. These six quantities, the three velocity components, and the three thermodynamic variables have to be provided for the description of the ﬂow. This can be done by solving the conservation equations for mass, momentum, and energy and the thermal equation of state, which connects the thermodynamic variables with each other. If the latter change markedly, the dynamic shear viscosity, µ, the thermal conductivity λ, and the speciﬁc heat c also have to be prescribed as a function of pressure and temperature. Liquids can be considered as incompressible and their density ρ can be assumed to be constant. This assumption also holds for gases ﬂowing at low speeds. The conservation laws are presented in the form of partial diﬀerential equations ; their solution requires the prescription of initial and boundary conditions.

2.2 Fundamental Equations of Fluid Mechanics
2.2.1 The Continuity Equation The mass m of a ﬂowing medium remains constant in a time-dependent volume τ (t), bounded by the closed surface A(t). dm d = dt dt τ (t)

ρ dτ = 0

(2.1)

The total time derivative of the mass m can be written as d dt With ρ(t + ∆t) = ρ(t) + and dτ = (v · n) dA dt there results d dt τ (t) τ (t)

ρdτ = lim

∆t→0

1 ∆t

τ (t+∆t)

ρ (t + ∆t) dτ −

τ (t)

ρ(t) dτ

.

(2.2)

∂ρ ∆t + . . . ∂t

(2.3)

(2.4)

ρ dτ =

τ (t)

∂ρ dτ + ∂t

A(t)

ρ (v · n) dA .

(2.5)

The surface integral over A(t) can be transformed into a volume integral with the aid of the Gauss divergence theorem, such that the last equation takes on the following form d dt ρ dτ = ∂ρ + ∇ · (ρ v) dτ ∂t . (2.6)

τ (t)

τ (t)

32

2. Fluid Mechanics II

The above volume integral can vanish only, if the integrand vanishes identically everywhere in the ﬂow ﬁeld. Hence ∂ρ + ∇ · (ρ v) = 0 . ∂t (2.7)

The vanishing of the integral by cancellation of negative with positive parts of the integrand is not meant here. Instead it is required, that the integrand vanishes throughout the ﬂow ﬁeld. The above equation is known as continuity equation. It can be written in a diﬀerent form by introducing the total time derivative of the density.

dρ + ρ (∇ · v) = 0 dt

(2.8)

For an incompressible ﬂuid there results ∇ · v = 0. The continuity equation can also be derived with the aid of a mass balance for a space-bound volume element. The change of mass per unit time in the volume element is equal to the diﬀerence between in- and out-ﬂowing mass. − ∂(ρ w) ∂(ρ v) ∂(ρ u) ∂ρ dx dy dz + dx dy dz = dx dy dz + dx dy dz ∂y ∂z ∂x ∂t (2.9)

Again there is obtained ∂ρ + ∇ · (ρ v) = 0 . ∂t (2.10)

The change of mass per unit time in the control volume is expressed by the partial derivative ∂ρ , the mass ﬂow through its surface is given by the expression ∇ · (ρ v). ∂t 2.2.2 The Navier-Stokes Equations The equilibrium of forces acting on a volume element of ﬂuid can be described with the momentum theorem. The time rate of change of the momentum of a closed volume is equal to the sum of the external forces acting on the volume. d dt τ (t)

ρ v dτ =

F

(2.11)

The left-hand side of this equation, which is the total time derivative of the volume integral extended over the momentum ρv, can be rearranged similarly to the derivation of the continuity equation. It follows that d dt ρ v dτ = ∂ρ v + ∇ · (ρ v v) dτ = ∂t ρ ∂v + (v · ∇) v dτ ∂t . (2.12)

τ (t)

τ (t)

τ (t)

The term (v v) is the diadic product of the velocity vector v. The external forces are the volume force, as, for example, the gravitational force Fg = τ (t)

ρ g dτ

,

(2.13)

and the surface forces, which result from the stress tensor σ . ¯ Fσ = −
A(t)

¯ (n · σ )dA

(2.14)

2.2 Fundamental Equations of Fluid Mechanics

33

¯ The term (n · σ ) is the vector product of the normal n and the stress tensor σ . The surface ¯ integral can again be replaced by a volume integral Fσ = − ¯ (n · σ ) dA = − τ (t)

A(t)

¯ (∇ · σ ) dτ

.

(2.15)

The momentum theorem for an arbitrarily closed volume τ then reads ρ ∂v + (v · ∇) v dτ = ∂t ρ g dτ − ¯ (∇ · σ ) dτ . (2.16)

τ (t)

τ (t)

τ (t)

If the three integrals are lumped together into a single integral, again, it can vanish only, if the integrand vanishes identically: ρ ∂v ¯ + (v · ∇) v = ρ g − (∇ · σ ) ∂t (2.17)

The components of this vector equation can be written in arbitrary coordinate systems, suitably chosen for the problem considered. In Cartesian coordinates the following equations result with
⎛

¯ ⎜ σ = ⎝ τyx

σxx τxy τxz σyy τyz ⎟ ⎠ τzx τzy σzz ∂σxx ∂τxy ∂τxz − − ∂z ∂x ∂y ∂τyx ∂σyy ∂τyz − − ∂x ∂y ∂z ∂τzx ∂τzy ∂σzz − − ∂x ∂y ∂z .

⎞

(2.18)

ρ ρ ρ

∂u ∂u ∂u ∂u +v +w +u ∂x ∂y ∂z ∂t ∂v ∂v ∂v ∂v +u +v +w ∂t ∂x ∂y ∂z ∂w ∂w ∂w ∂w +u +v +w ∂t ∂x ∂y ∂z

= ρ gx − = ρ gy − = ρ gz −

(2.19)

These equations describe the relations between the velocity components and the local stresses of the ﬂuid considered. The instantaneous stresses can be related to the velocity ﬁeld with the aid of the Stokes hypothesis, carried over from theoretical mechanics. In Hooke’s law the stresses are assumed to be proportional to the strain, and in ﬂuid mechanics the stresses are assumed to be proportional to the time rate of change of the strain. Before these relations are established, it is noted, that the above equations of motion state that a small volume element moving with the ﬂuid is accelerated or decelerated by the external forces acting on it. It is seen that the momentum balance is completely equivalent to Newton’s second law of motion. It is also noted, that (2.19) is valid for any ﬂuid either Newtonian or non-Newtonian. In order to use (2.19) for determining the velocity and the pressure, the stresses have to be expressed in terms of the derivatives of the velocity components and the viscosity of the ﬂuid. This will be done next for a Newtonian ﬂuid. Stress-Strain Relations For the derivation of the relations describing the dependence of the stresses on the time rate of change of the strain, it is assumed. that the normal stresses σxx , σyy , σzz cause elongations and contractions x , y , z , and the shear stresses τxy , . . . ,τzy cause angular displacements γxy , . . . ,γzy . As indicated in the diagram the components of the time rate of change of the strain are given by the partial derivatives of the velocity components in the direction of the coordinate axes.

34

2. Fluid Mechanics II ˙x = ∂u ∂x ˙y = ∂v ∂y ˙z = ∂w ∂z (2.20)

The sum of the components of the time rate of change of the strain yields the relative change in volume, per time interval dt (volume dilatation) e= ˙ ∂u ∂v ∂w + + = ∇ · v. ∂x ∂y ∂z (2.21)

The angular displacements per time interval dt are γxy = ˙ ∂u ∂v dβ+dα =− + dt ∂y ∂x γyz = − ˙ ∂v ∂w + ∂z ∂y γxy = − ˙ ∂u ∂w + ∂z ∂x . (2.22)

The normal and tangential stresses are related to the time rate of change of the components of the strain and the angular displacements in a linear ansatz. In the state of rest, in which the time rate of change of the strains and the angular displacements vanish, the normal stresses σxx , σyy , σzz are independent of the direction, and solely given by the hydrostatic pressure. This behavior of the ﬂow can be expressed by the following hypothesis, introduced by Stokes ˜ ˙ σxx = p − 2 µ ˙x − λ e , ˜ ˙ σyy = p − 2 µ ˙y − λ e , ˜ ˙ σzz = p − 2 µ ˙z − λ e , and τxy = µ γxy ˙ , ˙ τyz = µ γyz , τxz = µ γxz ˙ . (2.24)

(2.23)

Since for reasons of symmetry τxy = τyx ,τyz = τzy and τxz = τxz , only three normal and three tangential stress components of the Stokes stress tensor are unknown. It can therefore be written as
⎛ ⎜ ¯ σ=p ⎝ ⎛ ⎜ −2 µ ⎜ 1 ⎝ 2
1 2

1 1 1
∂u ∂x ∂u ( ∂y + ( ∂u + ∂z

⎞

⎛

⎟ ˜ ⎜ ⎠−λ ⎝
1 2

∇·v

⎞

∇·v
1 2 1 2

∇·v
∂w ) ∂x ∂w ) ∂y

⎟ ⎠− ⎞ ⎟ ⎟ ⎠

∂v ) ∂x ∂w ) ∂x

( ∂u + ∂y ∂v ∂y 1 ∂v ( + 2 ∂z

∂v ) ∂x ∂w ) ∂y

( ∂u + ∂z ( ∂v + ∂z ∂w ∂z

.

(2.25)

˜ ˜ ˆ The coeﬃcient λ generally is split into the two parts λ = µ − 2 µ. The volume viscosity µ takes ˆ 3 into account the molecular degrees of freedom; it vanishes for monatomig gases. The normal stresses are σxx = p − 2 µ σyy σzz Their mean value is ¯ σ= 1 (σxx + σyy + σzz ) = p − µ (∇ · v) . ˆ 3 (2.27) ∂u 2 − µ − µ (∇ · v) , ˆ ∂x 3 ∂v 2 − µ − µ (∇ · v) , = p−2µ ˆ ∂y 3 ∂w 2 − µ − µ (∇ · v) . = p−2µ ˆ ∂z 3

(2.26)

2.2 Fundamental Equations of Fluid Mechanics

35

For incompressible ﬂows the mean value is simply σ = p. ¯ If the stress-strain relations are substituted in the momentum equations, the Navier-Stokes equations are obtained (1823, 1845). ρ ρ ρ du ∂p ∂ ∂u ˜ ∂ = − + + λ (∇ · v) + 2µ µ dt ∂x ∂x ∂x ∂y ∂p dv ∂ = − µ + ∂y ∂x dt ∂p ∂ dw = − + µ dt ∂z ∂x ∂u ∂v + ∂y ∂x ∂u ∂w + ∂z ∂x + + ∂u ∂v + ∂y ∂x + ∂ µ ∂z ∂u ∂w + ∂z ∂x ∂v ∂w + ∂z ∂y + ρ gx + ρ gy

∂ ∂v ˜ ∂ 2µ µ + λ (∇ · v) + ∂y ∂y ∂z ∂ µ ∂y ∂v ∂w + ∂z ∂y +

∂ ∂w ˜ + λ (∇ · v) + ρ gz 2µ ∂z ∂z

For incompressible ﬂow with constant dynamic shear viscosity µ the above equations reduce to ∂p du = − + µ ∇2 u + ρ gx , dt ∂x dv ∂p + µ ∇2 u + ρ gy , ρ = − dt ∂y ∂p dw = − + µ ∇2 u + ρ gz . ρ dt ∂z ρ 2.2.3 The Energy Equation The Bernoulli equation states that the sum of the mechanical energies is constant in incompressible, loss-free, steady ﬂows. In ﬂows with large density and temperature changes in addition to the change of the mechanical energy also the change of the thermal energy has to be included in the balance equation. The energy E in a closed volume consists out of the internal energy ρ e and the kinetic energy 1 ρ v 2 . 2 E=ρ e+ v2 2 (2.29)

(2.28)

According to the ﬁrst law of thermodynamics, the internal energy of an arbitrary, always the same mass particles containing volume increases, if the amount of heat, added to the gas through the bounding surface, is larger than the work done against the acting volume and surface forces. Internal heat sources are neglected. Since the time rate of change of the various parts of the energy contributing to the balance is considered, often the notion power P is used. The time rate of change of the total energy of a ﬂuid contained in a volume bounded by a closed surface is d dt ρ e+ v2 2 dτ = P . (2.30)

τ (t)

The work to be done by the ﬂuid against the volume forces is given by the work, done during the time interval dt by the forces acting along the various particle paths. If the volume force is given by the gravitational force, the corresponding contribution is described by the inner product of the gravitational acceleration g and the velocity v. Pτ = τ (t)

ρ(g · v)dτ

(2.31)

The work to be done against the surface forces can be determined with the aid of the work necessary for the displacement of the bounding surface of the volume τ . This contribution is ¯ given by the surface integral extended over the inner vector product of the stress tensor σ and the velocity v

36

2. Fluid Mechanics II PA = − ¯ (σ · v) · n dA , (2.32)

A

which can again be transformed into a volume integral. PA = − τ (t)

¯ ∇ · (σ · v) dτ

(2.33)

If, as in the derivation of the Navier-Stokes equations, the pressure is separated from the stress tensor, a second integral results, in which the work of the friction forces per unit volume is ¯ taken into account by the rest of the stress tensor σ and the velocity vector PA = − τ (t)

∇ · (p v) dτ −

τ (t)

¯ ∇ · (σ · v) dτ

.

(2.34)

The ﬁrst integral describes the work of the pressure forces, and the second the work of the friction forces. The contribution to the energy balance by heat conduction is given by the heat ﬂux through the bounding surface q. Pq = − (q · n) dA (2.35)

A(t)

Thermal radiation is not considered here. With Fourier’s law of heat conduction q = −λ ∇ T there is obtained Pq =
A(t)

(2.36)

λ (∇ T · n)dA =

τ (t)

∇ · (λ ∇ T ) dτ

.

(2.37)

The time rate of change of the total energy contained in the volume of ﬂuid can again be written as a volume integral if the assumptions introduced earlier are retained. Then the energy equation in integral form is ρ d dt e+ v2 2 ¯ − ρ (g · v) + ∇ (p · v) + ∇ σ · v − ∇ · (λ ∇ · T ) dτ = 0 .(2.38)

τ (t)

The integral form of the energy equation can be replaced by its diﬀerential form if the integrand vanishes identically ρ d dt e+ v2 2 ¯ = ρ (g · v) − ∇ · (p v) + ∇ · (λ ∇ T ) − ∇ · (σ · v) . (2.39)

2.2.4 Diﬀerent Forms of the Energy Equation The energy equation can be written in diﬀerent forms. The scalar product of the velocity vector and the Navier-Stokes equations yields the following relation ρ d dt v2 2 ¯ = ρ (v · g) − v · (∇ · σ ) . (2.40)

If this equation is subtracted from the energy equation the time rate of change of the internal energy is obtained ρ de p dρ = + ∇ · (λ ∇ T ) + µ Φ . dt ρ dt (2.41)

2.2 Fundamental Equations of Fluid Mechanics

37

p The term ( ρ ) ( dρ ) represents the work of compression or expansion of the pressure. The term dt µ Φ stands as abbreviation for

¯ ¯ µ Φ = v · (∇ · σ ) − ∇ · (σ · v) .

(2.42)

The function Φ is called dissipation function. It represents the amount of the mechanical energy, irreversibly transformed into thermal energy. Another form of the energy equation is obtained by introducing the enthalpy h: h=e+ With h there results ρ d dt h+ v2 2 = ∂p ¯ + ρ (g · v) + ∇ · λ ∇ T − σ · v ∂t (2.44) p ρ (2.43)

In steady ﬂows the stagnation enthalpy remains constant along streamlines, if volume forces can be neglected, and if either the heat conduction and the work done by friction forces per unit volume vanish or if the heat removed by conduction is equal to the work of the friction forces. h0 = h + v2 = const. 2 (2.45)

For perfect gases the enthalpy depends only on the temperature. With dh = cp dT (2.46)

the time rate of change of the temperature in the energy equation can be expressed by the time rate of change of the pressure, the divergence of the heat ﬂux, and the dissipation function. ρ cp dp dT = + ∇ · (λ ∇ T ) + η Φ dt dt (2.47)

In Cartesisan coordinates the equation reads ρ cp ∂T ∂T ∂T ∂T +u +v +w ∂t ∂x ∂y ∂z = + − − ∂p ∂p ∂p ∂p +u +v +w ∂t ∂x ∂y ∂z ∂ ∂T λ ∂x ∂x σxx τxy + ∂ ∂T λ ∂y ∂y + ∂ ∂T λ ∂z ∂z

+ τyz

∂u ∂v + σyy + σzz ∂x ∂y ∂u ∂v + + τxz ∂y ∂x ∂w ∂v + . ∂y ∂z

∂w ∂z ∂u ∂w + ∂z ∂x (2.48)

If the components of the Stokes stress tensor are replaced by the relations introduced earlier, the dissipation function Φ becomes, if µ = 0, ˆ
⎡

∂u Φ = 2⎣ ∂x +

2

+
2

∂v ∂y +

2

+

∂w ∂z
2

2

⎤ ⎦−

2 3

∂u ∂v ∂w + + ∂x ∂y ∂z
2

2

∂u ∂v + ∂y ∂x

∂u ∂w + ∂z ∂x

+

∂w ∂v + ∂y ∂z

.

(2.49)

38

2. Fluid Mechanics II

The dissipation function is always positive, proving, that an irreversible change of mechanical energy always causes a heating of the ﬂuid or gas. The conservation equations for mass, momentum, and energy form a system of ﬁve coupled, nonlinear, partial diﬀerential equations, which require initial and boundary conditions for a solution. For constant density, viscosity, heat conductivity, and heat capacity the energy equation is decoupled from the continuity equation and from the momentum equations, and pressure and velocity can be determined without the use of the energy equation.

2.3 Similar Flows
In order to describe ﬂow processes it is necessary to intergrate the conservation laws just derived. Since the integration of these equations in closed form is, in general, not possible because of the inherent mathematical diﬃculties, ﬂows are often investigated experimentally. Fluid mechanical and thermodynamic data are measured with models geometrically similar to the full-scale conﬁguration, for which the ﬂow is to be determined. However, since in general the models are smaller in size, the measured data have to be applied to the full-scale conﬁguration with the rules of the theory of similitude. This theory makes use of similarity parameters, in which the characteristic quantities with physical dimensions of the ﬂow considered are combined to dimensionless quantities. Two ﬂows about geometrically similar bodies are called similar, if the individual similarity parameters have the same value for both ﬂows. The similarity parameters, that are important for the ﬂow process considered, can either be determined with the method of dimensional analysis applied to the physical properties of the ﬂow or by nondimensionalizing the conservation equations. 2.3.1 Derivation of the Similarity Parameters with the Method of Dimensional Analysis The method of dimensional analysis aims at deriving similarity parameters, which can be used to apply data measured with a model conﬁguration to the geometrically similar full-scale conﬁguration. Thereby the number of necessary experiments can be reduced, which depends on the number the physical quantities inﬂuencing the problem. The dimensional analysis also oﬀers the advantage, that the physical quantities can be combined in such a way, that the results are independent of the measuring units. The physical quantities are combined in a product such that dimensionless combinations result. The number of possible combinations is determined by Buckingham’s Π theorem. According to this theorem m−n dimensionless similarity parameters can be derived from m inﬂuence quantities, if n is the number of the fundamental dimensions, entering the problem considered. In the following the important similarity parameters of the ﬂuid mechanical quantities are derived. If for a ﬂow problem the characteristic values of its length l, time t, velocity v, acceleration b, pressure diﬀerence ∆p, density ρ, and dynamic shear viscosity µ are known, then the dimensions of these quantities can be described by the three fundamental dimensions length, time, and mass. Each of them can be combined with the other quantities (reference quantities) in a product to form a dimensionless similarity parameter. The reference quantities are chosen to be the length l, the velocity v, and the density ρ. With the aid of Buckingham’s Π theorem the following four products for the time t, the acceleration b, the pressure diﬀerence ∆p, and the dynamic shear viscosity µ can be formulated K1 K2 K3 K4 = = = = tα1 ρβ1 v γ1 lδ1 , g α 2 ρβ 2 v γ 2 l δ 2 , ∆pα3 ρβ3 v γ3 lδ3 , µ α 4 ρβ 4 v γ 4 l δ 4 .

(2.50)

2.3 Similar Flows

39

The gravitational acceleration g is used for b. The unknown exponents αi ,...,δi have to be determined in such a way, that the similarity parameters Ki are dimensionless. One of each of the exponents can be chosen arbitrarily, for example α1 = −1, α2 = −0.5, α3 = 1, α4 = −1. Then it follows from the equation for K1 by comparison of the exponents of the fundamental dimensions with [K1 ] = L0 T 0 K 0 Length L : 0 = −3 β1 + γ1 + δ1 Time T : 0 = −1 − γ1 Mass M : 0 = β1 . This system of equations has the solution β1 = 0. γ1 = −1, δ1 = 1 . The similarity parameter K1 is obtained to K1 = l vt . (2.53) (2.52)

(2.51)

The similarity parameters K2 to K4 are determined in similar manner. They are v ∆p ρvl , K4 = K2 = √ , K3 = ρ v2 µ gl The similarity parameters are named after famous scientists. K1 = Sr = K2 K3 K4 l Strouhal number, vt v = Fr = √ Froude number, gl ∆p = Eu = Euler number, ρ v2 ρvl = Re = Reynolds number. µ . (2.54)

Application of the Method of Dimensional Analysis to the Pipe Flow As was shown for the pipe ﬂow, its pressure loss ∆p depends on the density ρ, the mean velocity um , the dynamic shear viscosity µ, the length l, the roughness k, and the diameter D. f (∆p,ρ,um ,µ,l,k,D) = 0 (2.55)

The dimensions of these quantities can be expressed by combinations of the three fundamental dimensions of length, mass, and time. Density, velocity, and length (diameter) are chosen as reference quantities, such that there result four similarity parameters for the remaining four quantities ∆p, µ, l, and k. If the exponents are set αi = 1, then the products can be formulated as K1 K2 K3 K4 = = = = γ1 ∆p ρβ1 um Dδ1 , β2 γ 2 µ ρ u m D δ2 , γ3 l ρβ3 um Dδ3 , β4 γ 4 k ρ u m D δ4 .

(2.56)

The similarity parameters K1 to K4 are obtained as follows K1 = ∆p µ l k , K2 = , K3 = , K4 = ρ u2 ρ um D D D m . (2.57)

40

2. Fluid Mechanics II

The function f can now be written as f1 or ∆p = f2 Re, ρ u2 m l , D k D . (2.59) ∆p , ρ u2 m µ , ρ um D l , D k D =0 (2.58)

The form of the function cannot be determined with the method of dimensional analysis. However, the number of inﬂuence quantities has been reduced from six to three. 2.3.2 The Method of Diﬀerential Equations The similarity parameters can also be determined by nondimensionalizing the variables in the conservation equations with suitable reference quantities. They are denoted by the subscript 1. The similarity parameters then appear as constant coeﬃcients in the diﬀerential equations. To keep the derivation simple, the transport properties µ, λ, and the speciﬁc heat cp are assumed to be constant. With u= ¯ gx = ¯ ρ gx ρ1 g u v1 v= ¯ v v1 x l1 ρ= ¯ ρ ρ1 y l1 p= ¯ p ∆ p1 µ µ= ¯ µ1

¯ t t= t1

x= ¯

y= ¯

(2.60)

the x-component of the momentum equation for two-dimensional incompressible ﬂow, written in physical quantities ρ ∂u ∂u ∂u +u +v ∂t ∂x ∂y =− ∂p + ρ gx + µ ∇2 u ∂x (2.61)

is transformed into dimensionless form Sr ∂u ¯ ∂u ¯ 1 ∂u ¯ ∂p ¯ 1 +u ¯ ∇2 u . gx + ¯ ¯ +v ¯ = −Eu + ¯ ∂t ∂x ¯ ∂y ¯ ∂ x F r2 ¯ Re (2.62)

This equation contains the similarity parameters Sr, Eu, F r, and Re already derived. Additional similarity parameters are obtained from the energy equation for compressible ﬂows. ρ cp ∂T ∂T ∂T +u +v ∂t ∂x ∂y = λ ∇2 T +
⎡

∂p ∂p ∂p +u +v + ∂t ∂x ∂y
2

∂u + 2µ ⎣ ∂x + µ − With the additional reference quantities T ¯ T = T1 2 µ 3

∂v + ∂y
2

2

⎤ ⎦

∂u ∂v + ∂y ∂x

∂u ∂v + ∂x ∂y

2

(2.63)

, ∆ p 1 = ρ1 u 1 2

(2.64)

the dimensionless form of the energy equation is

2.3 Similar Flows ρ Sr ¯ ¯ ¯ ¯ ∂T ∂T ∂T +v ¯ +u ¯ ¯ ∂x ¯ ∂y ¯ ∂t = 1 ¯ ¯ ∇2 T + P r Re ∂p ¯ ∂p ¯ ∂p ¯ ¯ + u +v ¯ ¯ ∂t ∂x ¯ ∂y ¯
⎡
2

41

+ (γ − 1) M a2 Sr
⎧

¯ M a2 ⎨ ⎣ ∂ u + (γ − 1) 2 ∂x ¯ Re ⎩ − 2 3 v ∂ u ∂¯ ¯ + ∂x ∂y ¯ ¯
2

∂¯ v + ∂y ¯
2⎬

2

⎤ ⎦

⎫ ⎭

v ∂ u ∂¯ ¯ + + ∂y ∂x ¯ ¯

.

(2.65)

The above equation contains the additional similarity parameters µ1 cp1 Prandtl number, λ1 v1 v1 Ma = Mach number, =√ a1 γ RT1 cp1 . γ = cv1 Pr =

(2.66)

The Mach number is the ratio of the reference velocity to the speed of sound, which for known isentropic exponent γ and known gas constant R depends only on the reference temperature T1 . The dimensionless variables are the same for all corresponding times and corresponding points of all ﬂow ﬁelds which have similar initial and boundary conditions and identical similarity parameters Sr, Eu, F r, Re, P r, M a, and γ. 2.3.3 Physical Meaning of the Similarity Parameters The similarity parameters derived with the method of dimensional analysis and the method of diﬀerential equations can be interpreted from a physical point of view. The Strouhal number Sr is the ratio of two characteristic times in a time-dependent ﬂow ﬁeld. If the characteristic time t is large compared to l/v, i. e. the time, in which a ﬂuid particle travels the distance l, the ﬂow is called quasi-steady (Sr 0). ∂x If the pressure gradient is positive, the ﬂow is decelerated by friction and pressure forces, and the curvature of the velocity proﬁle at the wall is positive. The wall-shear stress decreases in the x-direction and can become negative. The ﬂuid near the wall then ﬂows in the opposite direction of the main ﬂow, associated with a separation of the ﬂow from the wall (separation and ﬂow reversal) as depicted in the picture below.

2.9 Separation of the Boundary Layer Boundary-layer separation by positive pressure gradient:

65

free streamline

y x

separation point

∆c p

separation points

inviscid flow separation

x

At the x-position, where τw = 0, the ﬂow separates from the body contour. The reversed ﬂow often leads to formation of vortices downstream from the separation point. In the separated region, also called recirculation or dead-water region, the boundary layer equations are no longer valid. The wall-shear stress can only be determined up to the separation point. If ﬂow separation occurs, the actual pressure distribution can markedly deviate from the pressure distribution computed for the inviscid ﬂow. For example, on an airfoil ﬂow separation can give rise to a turbulent wake, which begins at the separation point. At large angles of attack the ﬂow separates on the upper surface of the airfoil near the leading edge, and the lift is greatly reduced.

Flow Around a Circular Cylinder Flows around cylinders and other blunt bodies exhibit substantially diﬀerent separation behavior as a function of the Reynolds number. In very slow motion the ﬂow moves around the cylinder without separating. For Re < 4 the drag coeﬃcient is nearly inversely proportional to the Reynolds number and the drag itself is proportional to the free-stream velocity. This law looses its validity with increasing Reynolds number.

For Reynolds numbers 200 < Re < 105 (subcritical region) the drag coeﬃcient is approximately constant and the drag is proportional to the square of the free-stream velocity. The boundary layer on the cylinder is laminar and separates at an angle of about 83 degrees, measured from the stagnation point. Downstream from the separation points the ﬂow in the wake rolls up into

66

2. Fluid Mechanics II

vortices, which are shed from the cylinder in a alternating pattern (von K´rm´n vortex street). a a df The Strouhal number, based on the separation frequency f of the vortices, Sr = u∞ is 0.2. Laminar-turbulent transition of the ﬂow in the boundary layer occurs for Reynolds numbers 105 < Re < 5 · 105 . Then the kinetic energy of the ﬂow is larger near the wall, and the boundary layer can overcome a larger positive pressure gradient. The separation point moves downstream. The wake, which does no longer exhibit regular vortex shedding, has a smaller cross section. The difference between the actual pressure distribution and that of the potential ﬂow becomes smaller, and the drag coeﬃcient of the cylinder is decreased. Although the drag due to friction is increased, the total drag is decreased, as the pressure or form drag is reduced by the downstream shift of the separation point. Disturbances in the free stream and also the surface roughness inﬂuence the transition between subcritical and supercritical regime. In the latter, beginning at 5 · 105 < Re the drag coeﬃcient increases again with the Reynolds number. Comparable ﬂow situations are observed in the ﬂow around a sphere. However, alternating vortex shedding does not occur.

2.10 Selected References
Becker E.: Technische Str¨mungslehre, Teubner, Stuttgart 1977. o Gersten K., Herwig H.: Str¨mungsmechanik, Vieweg-Verlag Braunschweig, Wiesbaden o 1992. Landau L. D., Lifschitz E. M: Lehrbuch der theoretischen Physik VI - Hydrodynamik, Akademie-Verlag, Berlin 1974. Prandtl L., Oswatitsch K., Wieghardt K.: F¨hrer durch die Str¨mungslehre, Vieweg, u o Braunschweig 1969. Schlichting H., Gersten K.: Grenzschicht-Theorie, Springer, Berlin/Heidelberg 1997. Tietjens O.: Str¨mungslehre, 2 Bde., Springer, Berlin 1960 - 1970. o Truckenbrodt E.: Fluidmechanik, 2 Bde., Springer, Berlin 1980. Wieghardt K.: Theoretische Str¨mungslehre, Teubner, Stuttgart 1974. o Wuest W.: Str¨mungsmeßtechnik, Vieweg, Braunschweig 1969. o

2.11 Appendix
The Nabla operator ∇= ∂ , ∂x ∂ , ∂y ∂ ∂z (2.226)

is formally used as a vector. The gradient of a scalar function p is

2.11 Appendix ∇p= the divergence of a vector a ∇·a= and the curl i ∇xa=
∂ ∂x

67

∂p , ∂x

∂p , ∂y

∂p ∂z

,

(2.227)

∂ , ∂x

∂ , ∂y

a ∂ ⎜ x ⎟ ∂ax ∂ay ∂az + + ⎝ ay ⎠ = ∂z ∂x ∂y ∂z az
⎞ ( ∂az − ∂ay ) ∂z ∂y ⎜ ∂a ⎟ = ⎜ ( ∂zx − ∂az ) ⎟ ⎝ ⎠ ∂x ⎛

⎛

⎞

(2.228)

j
∂ ∂y

k
∂ ∂z

.

(2.229)

ax The following identities are used:

ay

az

( ∂ay − ∂x

∂ax ) ∂y

∇x∇p = 0 ∇ x ∇2 a = ∇2 (∇ x a) a2 (a · ∇) a = ∇ − a x (∇ x a) 2 The total time derivative is given by the operator

.

(2.230)

d ∂ ∂ ∂ ∂ ∂ +u +v +w = + (v · ∇) = ∂t ∂x ∂y ∂z dt ∂t The dyadic product of two vectors is a tensor.
⎛ ⎜ (a b) = ⎝ ay bx ⎞

.

(2.231)

ax bx ax by ax bz ay by ay bz ⎟ ⎠ az bx az by az bz
⎞ ⎛ ⎞

(2.232)

The inner vector product of a vector and a tensor yields a vector.
⎛

¯ (a · γ ) =

ax ay az

⎜ ⎝ γyx

γxx γxy γxz ax γxx + ay γyx + az γzx γyy γyz ⎟ = ⎜ ax γxy + ay γyy + az γzy ⎟ ⎠ ⎝ ⎠ γzx γzy γxz ax γxz + ay γyz + az γzz

(2.233)

In the derivation of the conservation equations the following transformation of the total time derivative of an integral is used (ρ is the density, S a scalar function of space and time). d dt ρ S dτ = = = = ρ S ρ ρ ∂ρ ∂S +S ∂t ∂t dτ +
A(t)

τ (t)

τ (t)

ρ S (v · n) dA

τ (t)

∂S ∂ρ +ρ + S ∇ · (ρ v) + (ρ v · ∇) S dτ ∂t ∂t ∂S + (ρ v · ∇) S dτ ∂t (2.234)

τ (t)

τ (t)

dS dτ dt

3. Exercises in Fluid Mechanics

3.1 Problems
3.1.1 Hydrostatics 1.1 The density of a ﬂuid ρf is to be determined with a U-tube. One stem is ﬁlled with water. water L g h liquid

kg kg ρ2 = 1000 3 m3 m kg ρC = 900 3 a = 0.1 m m Determine the height h! ρ1 = 850 1.4 A cylindrical vessel ﬂoats in another cylindrical vessel, ﬁlled with water. After adding a mass m the water surface is raised by ∆H.

kg m3 1.2 In three communicating vessels pistons are exposed to the forces F1 , F2 and F3 . h = 0.3 m L = 0.2 m ρw = 103

Given: ρ, A, m Determine the diﬀerence in height ∆H! 1.5 A boat with vertical side walls and a weight W0 has a draught in sea water h0 and displaces the volume τ0 . Before entering the mouth of a river the weight of the cargo is reduced by ∆W , in order to avoid the boat running aground. The draught is then h1 and the volume is τ1 . The density of the sea water is ρS , and that of the water in the river ρR .

F1 = 1100 N F2 = 600 N F3 = 1000 N A1 = 0.04 m2 A2 = 0.02 m2 m kg A3 = 0.03 m2 ρ = 103 3 g = 10 2 m s Determine the diﬀerences in height ∆h1 and ∆h2 ! 1.3 A cube ﬂoats in two laminated ﬂuids, one on top of the other. kg kg ρR = 103 3 m3 m W0 = 1.1 · 109 N ∆W = 108 N m h0 = 11 m h1 = 10.5 m g = 10 2 s ρS = 1.025 · 103

70

3. Exercises in Fluid Mechanics Determine (a) the volume τ0 , (b) the area of the deck A, (c) the diﬀerence τ2 −τ1 of the displaced volumes in fresh water and sea water, (d) the draught h2 in fresh water! Determine the force in the screws under the assumption that the weight of the container can be neglected! 1.8 A conical plug with density ρs closes the outlet of a water basin. The base area of the cone levels with the surface of the ﬂuid.

1.6 A diving bell with the weight W is lowered into the sea.

m R = 10−2 m H = 10−2 m g = 10 2 s kg kg ρs = 2 · 103 3 ρ = 103 3 m m How large must the force be to lift the plug? D = 3 m H = 3 m T = 22 m kg kg ρa = 1.25 3 ρW = 103 3 m m N m pa = 105 2 W = 8 · 104 N g = 10 2 m s (a) How high does the water in the bell rise if the temperature remains constant? (b) How large is the force (magnitude and direction) with which the bell must be held? (c) At what depth of immersion is the force zero? 1.7 A container ﬁlled with water is fastened to a plate. It has a small opening in the top. 1.9 A rectangular sluice gate with the width B separates two sluice chambers.

B = 10 m h1 = 5 m h2 = 2 m kg m g = 10 2 m3 s Determine (a) the force acting on the sluice gate, (b) the point of application of force! ρ = 103 1.10 A pivoted wall of a water container with width B is supported with a rod.

R = 1 m ρ = 103

kg m3

g = 10

m s2

h = 3 m B = 1 m α = 30◦ kg m ρ = 103 3 g = 10 2 m s Determine the force in the rod!

3.1 Problems 1.11 The triangular opening of a weir is closed with a plate.

71

1.14 A metereological balloon of mass m and initial volume τ0 rises in an isothermal atmosphere. The envelope is slack until the maximum volume τ1 is attained.

B = 1 m ρ = 103

kg m3

g = 10

m s2

Determine (a) the closing force, (b) the point of application of force! 1.12 A ﬂuid with a free surface rotates in an open circular cylindrical vessel with constant angular velocity, large enough, so that the ﬂuid just reaches the edge of the vessel. When the ﬂuid is at rest, it ﬁlls the vessel up to the height h0 . N m2 m = 2.5 kg p0 = 105 p0 = 1.27 kg m3

Nm τ0 = 2.8m3 τ1 = 10m3 R = 287 kg K m g = 10 2 s (a) How large is the force the balloon must be held with before take oﬀ? (b) At what altitude does the balloon attain the volume τ1 ? (c) How high does the balloon rise? 1.15 A balloon with an inelastic envelope has an opening at the bottom for equalization of the pressure with the surrounding air. The weight of the balloon without the gas ﬁlling is W . Before take oﬀ the balloon is held with the force Fs . W = 1000 N Fs = 1720 N Nm R = 287 kg K m T = 273 K g = 10 2 s Determine the height of rise of the balloon in an isothermal atmosphere! 3.1.2 Hydrodynamics If not mentioned otherwise, the ﬂow is assumed to be loss-free in this chapter. 2.1 Given the velocity ﬁeld u = u0 cos ω t v = v sin ω t with u0 = v0 = 1 m. w w Determine (a) the streamlines for ω t = 0, π , π , 2 4 (b) the path lines, (c) the path line of a particle, which at time t = 0 is in the point x = 0, y = 1 m!

D = 0.5 m h0 = 0.7 m H = 1 m m kg N ρ = 103 3 pa = 105 2 g = 10 2 m m s Determine (a) the height h and the angular velocity ω, (b) the pressure distribution at the wall and on the bottom! Hint: ∂p ∂p = ρ ω2 r = −ρ g ∂r ∂z ∂p ∂p dr + dz dp = ∂r ∂z 1.13 Determine the pressure as a function of the height z (a) for an isothermal atmosphere, (b) for a linear temperature variation T = T0 − α z, (c) for an isentropic atmosphere, (d) for a height of 3000 m, 6000 m und 11000 m! z=0: Nm R = 287 T0 = 287 K γ = 1.4 kg K m N K g = 10 2 p0 = 10 2 α = 6.5 · 103 m m s

72

3. Exercises in Fluid Mechanics

2.2 Under an iceberg a steady downward ﬂow is initiated by the cooling of the water by the ice. Determine the ﬂow velocity v in the depth h under the assumption, that the cold water (density ρc ) does not mix with the warm water (density ρw ).

N D = 6 · 10−3 m ∆p = 125 2 m kg p = 103 3 m
6 2.5 ≤ Re ≤ 250 : β = 1 + Re 250 ≤ Re : β=1

h = 50 m

ρc − ρw m = 0.01 g = 10 2 ρc s

2.3 Hot exhaust air of temperature Ti ﬂows through an open smokestack with a large suction scoop into the atmosphere. The external temperature is Ta .

2.5 In order to determine the velocity in a pipe ﬂow the pressure diﬀerence ∆p is measured. The pressure diﬀerence deviates from the dynamic pressure of the undisturbed ﬂow, if there is a large blockage in the pipe.

Plot v∞ /

2 ∆p ρ

as a function of

d ! D

2.6 Water ﬂows out of a large reservoir under the inﬂuence of gravity into the open air. Ti = 450 K Ta = 300 K m H = 100 m g = 10 2 s Determine the discharge velocity, taking into account the inﬂuence of compressibility! Hint: Use the Bernoulli equation in differential form: 1 dp + v dv + g dz = 0 ρ 2.4 Determine the free-stream velocity v∞ of a Prandtl static pressure tube, taking into account the inﬂuence of the viscosity for Ns (a) µ = 10−3 2 , m Ns (b) µ = 10−2 2 , m Ns (c) µ = 10−1 2 . m h = 0.1 m H = 1.5 m D = 0.1 m What is the diameter d of the water stream at the position H below the opening? 2.7 Water ﬂows out of a large pressure tank into the open air. The pressure diﬀerence ∆p is measured between the cross sections A1 and A2 .

3.1 Problems

73

A2 = 0.1 m2 A3 = 0.2 m2 N kg h = 1 m ρ = 103 3 pa = 105 2 m m N m g = 10 2 ∆p = 0.64 · 105 2 s m The outﬂow pipe was provided with a variable cross-section distribution in order to enable the measurement for the pressure. The pressure diﬀerence ∆p is measured in the cross sections indicated in the sketch. Determine (a) the velocities v1 , v2 , v3 , (b) the pressures p1 , p2 , p3 , and the pressure p above the water surface! A = 0.3 m2 2.8 Water ﬂows out of a large vessel through an opening of width B and height 2a into the open air.

Ad = 0.1 m2 h = 5 m N kg H = 80 m pa = 105 2 ρ = 103 3 m m m g = 10 2 s (a) How large is the volume ﬂow per unit time? (b) Sketch the curve of the static pressure in the pipe! (c) At what size of the cross section of the exit will vapor bubbles be formed, if the vapor pressure is N pv = 0.025 · 105 2 ? m A = 1 m2 2.10 Air ﬂows out of a large pressure tank through a well-rounded nozzle and a diﬀuser into the open air.

For

a → 0 the volume ﬂow per unit h √ ˙ time is Q0 = 2 a B 2 g h. Determine ˙ ˙ Q0 − Q 1 a the relative error = , for ˙ h 4 Q 1 3 , ! 2 4

kg N ∆p = 10 2 m3 m Determine the velocity in the throat of the nozzle as a function of the ratio of A the cross sections AD (a) for loss-free ﬂow, (b) for an eﬃciency of the diﬀuser of ρ = 1.25 ηD = ρ 2

pa − pD = 0.84 ! 2 2 (vD − va )

2.9 Two large reservoirs, one located above the other, are connected with a vertical pipe, with a nozzle attached to its end.

(c) What is the maximum velocity that can be attained for this eﬃciency of the diﬀuser? 2.11 Water ﬂows through a nozzle mounted in a pipe (cross-sectional ratio mD , discharge coeﬃcient αD ) and an oriﬁce

74

3. Exercises in Fluid Mechanics (mD , αB ). The mercury gauges show diﬀerences in height of hD and hB . Determine the time necessary for equalizing the water levels, if the dividing wall is lifted by the amount f h0 ! Neglegt the contraction of the ﬂow! 2.14 Water ﬂows out of a large reservoir into a lower reservoir, the discharge opening of which is suddenly reduced to one third. mD = 0.5 αD = 1.08 mB = 0.6 hD = 1 m hB = 1.44 m D = 0.1 m kg kg ρW = 103 3 ρHg = 13.6 · 103 3 m m m g = 10 2 s Determine 1. the volume ﬂow per unit time, 2. the discharge coeﬃcient of the oriﬁce!

2.12 A sluice gate is suddenly opened.

A = 0.03 m2 AB = 1 m2 m h = 5 m g = 10 2 s Determine the time, in which the water level rises to the quadruple value of its initial height h! 2.15 A pipe ﬁlled with gasoline is held vertically in water and closed at the upper end with a top. A small, well-rounded outlet in the top is opened.

As = 3000 m2 h(t = 0) = h0 = 5 m m g = 10 2 s How large must the cross section of the opening A be so that the water level of the bordering lake is attained within 10 minutes, if quasi-steady ﬂow is assumed? 2.13 Two equally large reservoirs, one of which is ﬁlled with water, are separated from each other by a wall.

B = 20 m h(t = 0) = h0 = 5 m m f = 0.05 m g = 10 2 s

D = 0.1 m d = 0.01 mm L = 0.8 m kg kg m ρB = 800 3 ρw = 103 3 g = 10 2 m m s Since the gasoline is lighter than water, it will begin to ﬂow upward through the small hole. The surface of the water surrounding the pipe is very large compared to the cross-sectional area of the pipe. How long will it take, until the pipe is completely emptied from gasoline?

3.1 Problems 2.16 A ﬂuid ﬂows through a pipe with a wellrounded intake. Its velocity is v0 (t).

75

Show that the integral of the local acceleration can be approximated as follows:
L −∞

∂v ds = ∂t

D √ +L 2

dv0 dt

D Hint: Assume that for s < − √8 the ﬂuid ﬂows radially towards the intake with ˙ the velocity v = Q π s2 , and that for 2 D s ≥ √8 the velocity is equal to v0 ! For D s = − √8 , v = v0 .

L = 20 m D h = 5 m L1 = 5 m m 3 kg p = 10 g = 10 2 m3 s 1. After what time are 99% of the ﬁnal velocity attained? 2. How much does the pressure at the position 1 diﬀer from its ﬁnal value? 2.19 A piston is moving sinusoidally in a pipe s = s0 sin ω t.

2.17 Liquid ﬂows out of a large container through a hose, lying horizontally on the ground, in steady motion into the open air. The end of the hose is suddenly lifted up to the height of the liquid level.

N L = 10 m D h = 2m m2 kg N s0 = 0.1 m ρ = 103 3 pv = 2500 2 m m m g = 10 2 s At what angular velocity ω is the pressure at the piston head equal to the vapor pressure pv ? pa = 105 2.20 In a hydraulic ram the valve I and the valve II are alternatively opened and closed. A part of the water is pumped from the height h1 to the height h2 . The other part ﬂows through the valve I.

L = 10 m h = 5 m D = 0.16 m m g = 10 2 s Determine 1. the velocity v0 immediately after lifting up the hose, 2. the time, in which the velocity decreases to v20 , 3. the ﬂuid volume that ﬂowed through the hose during this time! 2.18 The discharge pipe of a large water container is led to a lake. The throttle valve at the end of the pipe is suddenly opened.

h1 = h2 = 5 m L = 10 m D A = 0.1 m2 AB T1 = 1 s m g = 10 2 s

76

3. Exercises in Fluid Mechanics (a) First the valve I is opened for the time T . Determine the volume QI of the water discharged! (b) After closing the valve I, the valve II is opened until the velocity in the pipe is decreased to zero. How large is the discharge volume QII ? (c) the closure time of the valve so that the excess prssure does not exceed the safe value ∆psaf e ! 3.1.3 Momentum and Moment of Momentum Theorem In this chapter the friction forces are neglected in comparison to the volume, pressure, and inertia forces, but not the pressure losses, resulting from ﬂow separation. 3.1 Water ﬂows out of a bifurcated pipe into the open air. The pressure in the inﬂow stem is higher by the amount ∆p than in the surrounding air.

2.21 The ﬂap at the end of a discharge pipe of a large container is suddenly opened.

N a pa = 105 2 L1 = L2 = 5 m m D1 = 0.1 m D2 = 0.05 m h = 2 m kg m ρ = 103 3 g = 10 2 s m Determine (a) the time T , in which the velocity attains 99% of its ﬁnal value, (b) the volume of the ﬂuid discharged, (c) the pressures pA and pB immediately after opening the ﬂap and at time T ! (d) Sketch the pressure at the positions A and B as a function of time! 2.22 The pressurized pipe system of a storage power station is closed with a valve. During the closure (shut-down time Ts ) the discharge volume decreases linearly ˙ from Q0 to zero.

A1 = 0.2 m2 A2 = 0.03 m2 A3 = 0.07 m2 α2 = 30◦ α3 = 20◦ N kg ∆p = 104 2 ρ = 103 3 m m Determine (a) the velocities v1 , v2 , v3 , (b) the force F in the cross section 1, (c) the angle α3 , for which Fsy vanishes! 3.2 Water ﬂows out of a large container through a pipe under the inﬂuence of gravity in steady motion into the open air. Downstream from the nozzle the water jet is deﬂected by 180◦ . The ﬂow is assumed to be two-dimensional.

h = 200 m L = 300 m A = 0.2 m2 kg m m3 ˙ ρ = 103 3 g = 10 2 Q=3 s m s N ∆psave = (p1 − pa )save = 2 · 107 2 m Determine (a) the excess pressure in front of the open valve for steady ﬂow, (b) the pressure variation p1 (t) during closure of the valve (Sketch the result!),

A = 0.2 m2 AD = 0.1 m2 h = 5 m kg m ρ = 103 3 g = 10 m s Determine the forces retaining the pipe and the guide vane (a) for the sketched conﬁguration,

3.1 Problems (b) for the case that the inlet of the pipe and the nozzle are removed! 3.3 Water ﬂows out of a two-dimensional nozzle in steady motion with the velocity v0 against a guide vane, moving with the velocity vr .

77

Given: v1 , vA , ρ1 , ρA , AR Determine (a) the mass of air displaced, (b) the thrust and the net performance! 3.6 The constant free-stream velocity of a propeller is v1 . A certain distance downstream from the propeller the velocity in the slipstream is v2 , outside of it v1 .

m A = 0.1 m2 v0 = 60 2 β = 45◦ s kg ρ = 103 3 m (a) At what velocity vr does the performance of the vane attain its maximum value? (b) How large is then the force acting on the vane? 3.4 Two two-dimensional cascades (inﬁnitely many blades) with width B and spacing t deﬂect a ﬂow by the angle α.

m m A = 7.06 m v1 = 5 v2 = 8 s s kg ρ = 103 3 m Determine (a) the velocity v in the cross-sectional plane of the propeller, (b) the eﬃciency! 3.7 A ducted propeller is positioned in a free stream with constant velocity. The inlet lip is well rounded.

cascade I

cascade II

Given: ρ, v1 , α, B, t Determine (a) the velocity v2 , (b) the pressure diﬀerence p1 − p2 , (c) the pressure loss p01 − p02 , (d) the force exerted by the ﬂow on a blade! 3.5 A rocket moves with constant velocity. The air ﬂowing past the rocket is displaced in the radial direction. The velocity in the jet is vA , around it v1 .

N m p1 = 105 2 A = 1 m2 v1 = 10 s m kg N ρ = 103 3 p1 = 1.345 · 105 2 m m (a) Sketch the variation of the static pressure along the axis! Determine (b) the mass ﬂow, (c) the thrust, (d) the power transferred by the propeller to the ﬂow! 3.8 Two blowers, drawing air from the surroundings, diﬀer in the shape of their inlets.

78

3. Exercises in Fluid Mechanics kg m3 ˙ ρ = 1.25 3 Q2 = 4 s m Determine (a) the velocity v2 and the pressure p2 , (b) the velocities v1 and v3 , (c) the power of the blower, (d) the retaining force of the blower casing (traction or compressive force?)! 3.11 Water ﬂows out of a large frictionless supported container through a pipe, with a discontinuous increase of the cross section, into the open air.

Given: ρ, A, ∆p Determine (a) the discharge volume, (b) the power of the blowers, (c) the retaining force! 3.9 A pipe with an inserted nozzle is positioned in a free stream with constant velocity. N h = 5 m A = 0.1 m2 pa = 105 2 m kg m ρ = 103 3 g = 10 2 m s (a) For what cross-sectional area A2 does the volume rate of ﬂow attain its maximum value? With A2 determined under a) compute (b) the pressure p1 , (c) the cutting forces Fs1 , Fs2 , Fs3 (Traction or compressive forces?)! 3.12 A pump is feeding water from a lake into a large pressurized container. The volume rate of ﬂow is measured with a standard nozzle (discharge coeﬃcient α).

m A1 = 0.2 m2 A2 = 0.1 m2 v∞ = 40 s kg ρ = 1.25 3 m Determine (a) the velocity in the cross sections A1 and A2 , (b) the retaining force! 3.10 A jet apparatus, which is driven with a blower, sucks the volume rate of ﬂow ˙ Q2 through a ring-shaped inlet.

A1 = 0.1 m2 N pa = 105 2 m

A3 = 0.2 m2

H = 5 m h = 3 m d = 0.07 m N N pK = 2 · 105 2 ∆pw = 3160 2 m m N kg m pa = 105 2 ρ = 103 3 g = 10 2 m m s D = 0.1 m α = 1.08

3.1 Problems Determine (a) the velocity in the pipe, (b) the static pressure upstream and downstream from the pump, (c) the net performance of the pump! 3.13 Water ﬂows out of a large container through a pipe with a Borda mouthpiece into a lake. m s2

79

h1 = 0.1 m h2 = 0.2 m g = 10

Determine (a) the velocities v1 and v2 , (b) the Froude numbers F r1 and F r2 , (c) the energy loss H1 − H2 ! 3.16 The volume of water ﬂowing out of a storage pond is controlled with a wicket. h = 1 m g = 10 m s2

Determine (a) the contraction, (b) the out-ﬂow velocity v1 3.14 The volume rate of ﬂow of a ventilation blower is measured with an oriﬁce (discharge coeﬃcient α, contration coeﬃcient Ψ ).

h = 7.5 m g = 10

m s2

N N ∆pw = 300 2 m2 m kg ρ = 1.25 3 m α = 0.7 Ψ = 0.66 A = 10−2 m2 AB m= = 0.5 A (a) Sketch the variation of the static and total pressure along the axis of the pipe! Determine (b) the volume rate of ﬂow, (c) the pressure upstream of the blower, (d) the performance of the blower! pa = 105 3.15 A hydraulic jump occurs in an open channel.

Determine (a) the out-ﬂow velocity v1 as a function of the height of level h1 (why is v1 constant in the cross section?), (b) the height of level, for which the volume rate of ﬂow attains a maximum, (c) the height of level, for which the hydraulic jump does not occur, (d) the depth of water and the velocity downstream from the jump for h1 = 2.5 m! 3.17 The depth of water h1 of an open channel with constant volume rate of ﬂow is controlled by changing the height Zw of a weir. For Zw = 0 the depth of water is h0 .

80

3. Exercises in Fluid Mechanics m ˙ B = 20 m h0 = 2 m Q = 80 s kg ρ=3 3 m (a) Sketch the variation of the depths of water for Zw < Zcrit. and Zw > Zcrit. ! Determine for Zw = 1 m (b) the limiting height Zcrit. of the dam, (c) the depths of water h1 and h2 , (d) the diﬀerence of the energy heights between upper and lower water, (e) the force acting on the weir!
3

H = 10 m h = 1 m R = 0.15 m A = 0.5 · 104 m2 A1 = 1.5 · 104 m2 N |Mr | = 3.6 Nm pa = 105 2 m kg m ρ = 103 3 g = 10 2 α = 30◦ m s Determine (a) the number of revolutions, (b) the rate of volume ﬂow, (c) the pressure p1 , (d) the maximum angular velocity, if the friction torque is assumed to be zero! 3.1.4 Laminar Flow of Viscous Fluids 4.1 Determine the following quantities for a fully developed laminar pipe ﬂow of a Newtonian ﬂuid (a) the velocity distribution u(r) r = f( ) , umax R (b) the ratio um umax , (3.2) (3.1)

Hint: If a hydraulic jump occurs, it will be at the downstream face of the weir; the depth of the lower water is h0 . 3.18 Assume that in a rotating ﬂow pressure and velocity depend only on the radius.

(a) Choose the segment of a circular ring as control surface and, by using the momentum theorem, derive the relation dp v2 =ρ ! dr r (b) For what velocity distribution v(r) does the Bernoulli constant have the same value for all streamlines? 3.19 A lawn sprinkler is fed from a large reservoir. The water jets are inclined to the circumferential direction by the angle α. The friction torque of the bearing is Mr .

(c) the dependence of the pipe friction coeﬃcient on the Reynolds number! 4.2 A Bingham ﬂuid is driven by gravity between two parallel, inﬁnitely wide plates.

3.1 Problems Given: b, ρ, µ, τ0 , g, dp dz

81

=0

Assume that the ﬂow is fully developed and determine (a) the distance a, (b) the velocity distribution! 4.3 An oil ﬁlm of constant thickness and width ﬂows down on an inclined plate.

Assume fully developed laminar ﬂow and determine (a) the velocity distribution, (b) the ratio of the shear stresses for y = 0 and y = H, (c) the volume rate of ﬂow for a width of the plates B, (d) the maximum velocity for uw = 0, (e) the momentum ﬂux for uw = 0, (f) the wall-shear stress in dimensionless form for uw (g) sketch the velocity and shear stress distribution for uw > 0, uw = 0, and uw < 0 ! 4.6 A Newtonian ﬂuid ﬂows between two coaxial cylinders.

δ = 3 · 103 m B = 1 m α = 30◦ Ns kg µ = 30 · 10−3 2 ρ = 800 3 m m m g = 10 2 s Determine the volume rate of ﬂow! 4.4 An oil ﬁlm is driven by gravity. Given: R, a, µ, dp dx

Assume fully developed laminar ﬂow and determine (a) the velocity distribution (sketch the result!), (b) the ratio of the shear stresses for r = a and r = R, (c) the mean velocity! Given: δ, a, ρ, µ, g. Determine the velocity distribution in the oil ﬁlm (a) on a plane vertical wall, (a) on a wall of a vertically standing circular cylinder! 4.5 A Newtonian ﬂuid ﬂows in the gap between two horizontal plates. The upper plate is moving with the velocity uw , the lower is at rest. The pressure is linearly decreasing in the x-direction. 4.7 A Couette viscosimeter consists out of two concentric cylinders of length L. The gap between them is ﬁlled with a Newtonian ﬂuid. The outer cylinder rotates with the angular velocity ω, the inner is at rest. At the inner cylinder the torque Mz is measured.

Given: H, uw , ρ, µ,

dp dx

82

3. Exercises in Fluid Mechanics Ra = 0.11 m Ri = 0.1 m L = 0.1 m 1 ω = 10 Mz = 7.246 10−3 Nm s Determine (a) the velocity distribution, (b) the dynamic shear viscosity of the ﬂuid! Hint: The diﬀerential equations for the velocity and shear-stress distributions are: d dr 1 d (rv) = 0 , r dr d v τ = −µ r dr r

Given: R, Tw , λ, µ,

dp dx

4.8 A gas of thermal conductivity λ and speciﬁc heats cp and cv ﬂows in the gap between two horizontal plates. The upper plate is moving with velocity uw at temperature Tw , the lower is at rest and is thermally isolated.

(a) Assume fully developed laminar ﬂow and derive the diﬀerential equations for the velocity and temperature distribution for a ring-shaped volume element for vanishing convective heat ﬂux and constant material properties! State the boundary conditions! (b) Determine the temperature distribution T − Tw Tmax − Tw 4.10 Under a gun slide a plane wall is moving with the velocity u∞ .

Given: uw , H, Tw , ρ, µ, λ, cp , cv

dp dx

= 0, L = 5 · 10−2 m Ns m µ = 10−1 2 h1 = 10−4 m u∞ = 1 s m kg ρ = 800 3 m Determine (a) the similarity parameter of the problem, (b) the volume rate of ﬂow per unit width, (c) the pressure distribution in the gap, (d) the pressure force per unit width bearing the gun slide, (e) the performance loss per unit width t due to bearing friction! x Assume fully developed laminar ﬂow and determine for vanishing convective heat ﬂux and constant material properties (a) the velocity and temperature distribution, (b) the heat ﬂux per unit area through the upper plate! (c) Show that the stagnation enthalpy has the same value everywhere for P r = 1! (d) Determine the time-dependent temperature variation, if both plates are thermally isolated, and if at the time t = 0 the temperature in the ﬂow ﬁeld is T0 ! 4.9 A Newtonian ﬂuid with thermal conductivity λ ﬂows through a pipe. The wall temperature is kept constant by cooling the wall.

h(x) = h1 e− 5 L

4.11 From the momentum equation in integral form dI dt = τ =

∂ (ρ v) dτ + ∂t F

ρ v (v n)dA
A

3.1 Problems derive the diﬀerential form of the momentum equation in the x-direction for an inﬁnitesimally small volume element ρ ∂u ∂u ∂u ∂σxx ∂τxy +u +v + =− ∂t ∂x ∂y ∂x ∂y

83

friction losses in the intake of the pipe and in the nozzle can be neglected. Assume that the ﬂow in the pipe is fully developed!

4.12 The x-component of the friction force acting on a volume element is given by Ff x = ∂ µ ∂x ∂ µ + ∂y ∂ µ + ∂z 2 ∂u 2 − ∇·v ∂x 3 ∂u ∂v + + ∂y ∂x ∂w ∂u . + ∂x ∂x L = 100 m D = 10−2 m + d = 0.5 · 10−2 m pa = 105 p1 = 1.075 · 105 µ1 = 10−3 Ns m2 N m2 N m2 kg ρ = 103 3 m

Reduce this equation for an incompressible ﬂuid with constant viscosity! 3.1.5 Pipe Flows 5.1 The viscosity of an oil is to be measured with a capillary viscosimeter. This is done by measuring the time T , in which a small part of the oil (volume τ ) ﬂows through the capillary. Assume that the ﬂow is loss-free upstream of the position 1!

Determine (a) the velocity in the pipe and at the exit of the nozzle, (b) the pressure in the container, (c) the velocity at the exit of the nozzle for L = 0 and the same pressure in the container! 5.3 Two containers are connected with each other by 25 pipes with diameter D1 and 25 pipes with diameter D2 . A pressure diﬀerence of ∆p is measured between the containers. The pressure-loss coefﬁcient of the intake is ζ.

τ = 10 cm3

kg L = 0.1 m m3 h = 0.05 m D = 1 mm T = 254 s m g = 10 2 s ρ = 900

D1 = 0.025 m D2 = 0.064 m kg L = 10 m ρ = 103 3 λ = 0.025 m N ∆p = 105 2 ζ = 1 m Determine the volume rate of ﬂow! 5.4 A ﬂuid ﬂows through a hydraulically smooth pipe. The pressure drops by the amount ∆p over the length L. N L = 100m D = 0.1m ∆p = 5·104 2 m

5.2 Water ﬂows out of a large container through a hydraulically smooth pipe, with a nozzle ﬁxed to its end. The pressure upstream of the nozzle is p1 . The

84

3. Exercises in Fluid Mechanics Oil : µ = 10−1 Water : µ = 10−3 Ns m2 Ns m2 ρ = 800 kg m3 kg ρ = 103 3 m (a) the momentum ﬂux, (b) the wall shear stress! 5.7 Water ﬂows through a hydraulically smooth pipe with an discontinuous widening of the cross section into the open air.

Determine the velocities of the ﬂow! Hint: Use the Prandtl resistance law for super-critical Reynolds numbers! 5.5 Compressed air is pumped through a hydraulically smooth pipe. The pressure p1 , the density ρ1 and the velocity um1 are assumed to be known in the ¯ intake cross section. N D = 10−2 m p1 = 8 · 105 2 m m um1 = 10 ¯ s Ns kg ρ1 = 10 3 µ = 1.875 · 10−5 2 m m (a) Derive the following relation with the aid of the momentum theorem: λ ρ 2 dp d¯m u + ρ1 um1 + u =0 ¯ ¯ dx dx D 2 m (b) Determine the length, over which the pressure drops by one half of its initial value for compressible ﬂow with constant temperature and for incompressible ﬂow! 5.6 The velocity distribution in the intake region of a laminar pipe ﬂow is described by the following approximation:

D = 0.02 m D2 = 0.04 m L = 0.2 m kg m um1 = 0.5 ¯ ρ = 103 3 s m Ns µ = 10−3 2 m (a) At what length L2 does the pressure diﬀerence p1 − pa vanish? (b) How large is the corresponding pressure loss? Hint: Assume, that the wall shear stress in the widened part of the pipe can be determined with the equations for fully developed pipe ﬂow! 5.8 The feed pipes of a fountain consist out of four straight pipes of total length L, two bends (loss coeﬁcient ζK ) and a valve (ζV ).

u um y δ

=
⎧ ⎨

f 1− 2y − δ 1
2 δ 3 R y 2 δ

y δ

+

1 6

δ R

2

f

=⎩

0 ≤ y ≤ δ(x) δ(x) ≤ y ≤ R

Given: um , R, ρ, µ Determine the following quantities for the intake cross section, the end of the δ intake region, and for R = 0.5

h = 10 m D = 0.05 m L = 4 m ζK = 0.25 ζV = 4.5 λ = 0.025 Determine the volume rate of ﬂow and the height H for dissipative and nondissipative ﬂow with (a) d = D , 2 (b) d = D! Hint: Assume that the ﬂow in the intake and in the nozzle is loss-free and fully developed in the straight pipes!

3.1 Problems 5.9 Water ﬂows through a hydraulically smooth pipe. kg D = 0.1 m Re = 105 ρ = 103 3 m Ns µ = 10−3 2 m Determine (a) the wall shear stress, ¯ (b) the ratio of the velocities uum , ¯max y u∗ (c) the velocity for ν = 5 and for y u∗ = 50, ν u (d) the mixing length for y ν ∗ = 100! 5.10 The velocity distribution of a turbulent pipe ﬂow is approximately described by the ansatz um ¯ y = umax ¯ R
1 7

85

(c) the pressure at the exit of the pump, (d) the net performance of the pump! 5.12 In a fully developed pipe ﬂow with vol˙ ume rate of ﬂow Q a pressure drop ∆p is measured over the distance L.

.

Determine ¯ (a) the ratio of the velocities uum , ¯max (b) the ratio of the momentum ﬂuxes ˙ I ! ρ u2 πR2 ¯ m m ˙ L = 100 m D = 0.5 m Q = 0.393 s N kg ∆p = 12820 2 ρ = 900 3 m m Ns µ = 5 · 10−3 2 m Determine (a) the pipe friction coeﬃcient, (b) the equivalent sand roughness of the pipe, (c) the wall shear stress and the retaining force! (d) How large would the pressure drop be in a hydraulically smooth pipe? 5.13 Air is to be conveyed through a rough pipe with a well rounded intake with the aid of a blower.

3

5.11 A pump is feeding water through a rough pipe (equivalent sand roughness ks ) from the level h1 to the level h2 .

m3 ˙ h1 = 10 m h2 = 20 m Q = 0.63 s L = 20 km D = 1 m ks = 2 mm kg m2 ρ = 103 3 ν = 10−6 m s m 5 N pa = 10 g = 10 2 m2 s (a) Sketch the variation of the static pressure along the axis of the pipe! Determine (b) the pressure at the intake of the pump,

L = 200 m ks = 1 mm Determine the ratio of the blower performance for the diameters D = 0.1 m and D = 0.2 m for the same volume rate of ﬂow and very large Reynolds numbers! 5.14 Water is fed through a system of 100 pipes into a channel with quadratic cross section.

86

3. Exercises in Fluid Mechanics 6.3 A ﬂuid ﬂows slowly and steadily through a hydraulically smooth pipe. The ﬂow is laminar and fully developed. (a) Derive the Hagen-Poisseuille law with the aid of the dimensional analysis from the ansatz ∆p ˙ Q= L α m3 ˙ L = 0.5 m a = 0.1 m Q = 0.01 s kg Ns D = 0.01m ρ = 103 3 µ = 10−3 2 m m For what length does the pressure loss of the channel become equal to that of the pipe system? 5.15 An open channel with quadratic cross section is inclined by the angle α.

µβ D γ

!

(b) Show that the pipe friction coeﬃcient is inversely proportional to the Reynolds number! 6.4 What is the drag of two spheres of different diameter but the same Reynolds number, if one moves in air and the other in water, and if the drag coeﬃcient depends on the Reynolds number only? ρa = 0.125 · 10−2 ρw µa = 1.875 · 10−2 µw

m ˙ Q = 3 · 10−4 a = 0.05 m s kg ρ = 10 3 m m −3 Ns µ = 10 g = 10 2 m2 s Determine the angle of inclination and the pressure loss per unit length!

3

6.5 In an incompressible ﬂow about a circular cylinder the frequency, with which vortices are shed, depends on the freestream velocity, density, and viscosity. Determine the similarity parameters with the aid of the dimensional analysis! 6.6 The pipes in heat exchangers can oscillate due to excitation by the cross-ﬂow they are exposed to. It is known, that in a ﬂow about a circular cylinder, with its axis normal to the direction of the ﬂow, the Strouhal number is constant for 200 ≤ Re ≤ 105 . m m v =3 D = 0.1 m v = 1 s s m2 m2 ν = 10−6 ν = 1.5 · 10−5 s s Determine (a) the minimum diameter of the model cylinder, (b) the excitation frequency f , if for the 1 smallest model f = 600 ! s 6.7 The power needed to overcome the aerodynamic drag of an automobile with quadratic cross-sectional area A is to be determined in a wind-tunnel

3.1.6 Similar Flows 6.1 Derive the dimensionless similarity parameters with the momentum equation for the x-direction ∂u +u ∂t ∂p +µ − ∂x ρ ∂u ∂u +v = ∂x ∂y ∂2u ∂2u + ! ∂x2 ∂y 2

6.2 In hydraulically smooth pipes of different lengths and diameters the pressure loss is measured for diﬀerent velocities, densities, and viscosities. The ﬂow is fully developed, incompressible, and steady. How can the results of the measurements be presented in a single curve?

3.1 Problems experiment. The cross-sectional area of the model cannot exceed Am in order to avoid blockage eﬀects in the windtunnel. m A = 4 m2 Am = 0.6 m2 v = 30 s (a) What speed has to be chosen for the measurement in the wind tunnel? (b) Determine the power needed, if with a larger model the drag F = 810 N is measured! 6.8 Water ﬂows through a model of a valve ˙ (volume rate of ﬂow Q ). Between intake and exit the pressure diﬀerence ∆p is measured. The valve is supposed to be used in an air pipe. A = 0.18 m2 A = 0.02 m2 kg kg ρ = 1.25 3 ρ = 103 3 m m Ns Ns µ = 1.875 · 10−5 2 µ = 103 2 m m N m3 ˙ ∆p = 1.58 · 105 2 Q = 0.2 s m (a) For what volume rate of ﬂow are the ﬂows through the model and the full-scale conﬁguration similar? (b) What is the pressure diﬀerence between intake and exit? 6.9 An axial blower (diameter D, number of revolutions n) is to be designed for air. In a model experiment with water (reduction scale 1:4) the increase of the total pressure ∆p0 is measured. 1 m3 ˙ D = 1 m n = 12.5 Q = 30 s s kg −5 Ns ρ = 1.25 3 µ = 1.875 · 10 m m2 −3 Ns 3 kg ρ = 10 µ = 10 m3 m2 N ∆p0 = 0.3 · 105 2 m Determine (a) the volume rate of ﬂow and the number of revolutions during the experiment, (b) the change of the total pressure of the blower, (c) the power and the torque needed for driving the model and the main conﬁguration!

87

6.10 The power of a propeller of an airplane is to be determined in a wind-tunnel experiment (Model scale 1:4) for the ﬂight velocity v. In the test section of the wind tunnel the velocity can be varied m between 0 and 300 , the pressure bes N N tween 0.5·105 2 and 5·105 2 , and the m m temperature between 250 K and 300 K. The viscosity of air is described by the relation µ µ300K = T 300K
0.75

.

m 1 v = 200 s s Nm T = 300 K R = 287 γ = 1.4 kg K 5 N p = 10 m2 (a) Determine an operating point (v , p , T ) such that the results of the measurements can be applied to the fullscale conﬁguration! (b) What number of revolutions must be used in the experiment? (c) Determine the power, if in the experiment the power P was measured! D = 1 m n = 100 6.11 A model experiment is to be carried out prior to the construction of a tanker (Model scale 1 : 100) in a towing basin. (a) How large would the ratio of the kinematic viscosities ν have to be? ν (b) How large must the towing velocity in water be, if the aerodynamic drag can be neglected and if only the wave drag or only the frictional drag is taken into account? 6.12 A docking pontoon is fastened at the bank of a river. An experiment is to be carried out with a model scaled down to 1 : 16.

88

3. Exercises in Fluid Mechanics L = 3.6 m B = 1.2 m H = 2.7 m m v=3 FD = 4 N h = 2.5 cm s kg ρ = 103 3 m Determine (a) the ﬂow velocity in the model experiment, (b) the force acting on the pontoon, if the force F is measured in the experiment, (c) the drag coeﬃcient of the pontoon, (d) the height of a wave h to be expected at the side of the pontoon, facing the oncoming ﬂow, if the height h was measured in the experiment! m3 ˙ D = 1 m d = 0.4 m Q=1 s kg Ns ρ = 800 3 µ = 10−1 2 m m Ns kg ρ = 103 3 µ = 10−3 2 m m N N ∆pD = 500 2 ∆pl = 400 2 m m Determine (a) the ﬂow velocity and the volume rate of ﬂow in the model experiment, (b) the discharge coeﬃcient and the loss coeﬃcient of the measuring throttle, (c) the diﬀerential pressure and the pressure loss for the full-scale conﬁguration!

6.13 In a reﬁnery oil ﬂows through a horizontal pipe line into a reservoir with pressure pR , with the pressure at the intake being p0 . A safety valve is attached to the end of the pipe line, which in the case of emergency can close the pipe line within the time T . In a model experiment with water (diminution scale 1 : 10) the maximum pressure in front of the safety valve, measured during the shut-down procedure, is pmax . N N p = 1.5 · 105 2 pR = pR = 105 2 m m kg kg ρ = 880 3 ρ = 103 3 m m Ns Ns µ = 10−1 2 µ = 10−3 2 m m N T = 0.5 s pmax = 1.05 · 105 2 m Determine (a) the pressure p0 and the shut-down time in the model experiment, (b) the maximum pressure in the fullscale conﬁguration! 6.14 In a petroleum pipe line (diameter D) the volume rate of ﬂow is to be determined with a measuring throttle (diameter d). In a model experiment with water (diminution scale 1 : 10) a diﬀerential pressure ∆pD and a pressure loss ∆pl is measured at the measuring throttle.

3.1.7 Potential Flows of Incompressible Fluids 7.1 A cyclone is assumed to have the following velocity distribution: vθ (r) = ωr
2 ω r0 r

r ≤ r0 r > r0

vr = 0

1 s kg H = 100 m ρ = 1.25 3 m (a) Sketch vθ (r)! (b) Determine the circulation for a circle around the axis of the cyclone for r < r0 , r = r0 and r > r0 ! (c) Show that for r > r0 the ﬂow is irrotational! (d) How large is the kinetic energy in a cylinder with radius R = 2 r0 and height H? r0 = 10 m ω = 10 7.2 (a) State the deﬁnitions of potential and stream function for two-dimensional ﬂow! What conditions have to be satisﬁed so that they can exist? (b) How are ∇ · v and ∇2 Φ and ∇ x v and ∇2 Ψ related to each other? 7.3 Examine, whether potential and stream function exist for the following velocity ﬁelds! (a) u = x2 y v = y 2 x (b) u = x v = y

3.1 Problems (c) u = y v = −x (d) u = y v = x Determine potential stream-function!

89

Hint: The following relations are valid for polar coordinates: and vr = 1 ∂Ψ ∂Φ = ∂r r ∂θ ∇·v ∇xv = ∂Ψ 1 ∂Φ =− r ∂θ ∂r 1 ∂ (r vr ) 1 ∂vθ = + r ∂r r ∂θ 1 ∂ (r vθ ) 1 ∂vr − k r ∂θ r ∂r vθ =

7.4 A two-dimensional ﬂow is described by the stream function Ψ = U xy. In the L point xref = 0, yref = 1 m the pressure N is pref = 105 2 . m kg m L = 1 m ρ = 103 3 U =2 s m (a) Examine, whether the ﬂow possesses a potential Determine (b) the stagnation points, the pressure coeﬃcient, and the isotachs, (c) velocity and pressure for x1 = 2 m, y1 = 2 m, (c) the coordinates of a particle, which at time t = 0 passes through the point x1 , y1 , for the time t = 0.5 s, (e) the pressure diﬀerence between these two points! (f) Sketch the streamlines! 7.5 Given the potential Φ = yx2 − y3 3 .

7.7 Given the stream function 1 Ψ (r,0) = n rn sin(n θ). (a) Sketch the streamlines for n = 0.5; n = 1 und n = 2 ! (b) Determine the pressure coeﬃcient for the point x = 0, y = 0, if pressure and velocity are known for the point xref = 1, yref = 1! 7.8 Consider a large basin with an outlet. The ﬂow outside of the outlet (r > R0 ) can be described by superposition of a plane sink and a potential vortex. For r = R0 the in-ﬂow angle is α and the depth of water is h0 . The volume rate ˙ of ﬂow of the discharging water is Q.

(a) Determine the velocity components and examine, whether the stream function exists! (b) Sketch the streamlines! 7.6 Determine the velocity ﬁelds c c vr = r vθ = 0 and vr = 0 vθ = r (a) ∇ x v and ∇ · v, (b) potential and stream funktion, (c) the circulation along a curve around the origin!

R0 = 0.03 m h0 = 0.02 m g = 10
3

m ˙ α = 30◦ Q = 0.5 · 10−3 s Determine (a) the circulation Γ (b) the shape of the water surface h(r) for r ≥ R0 , (c) the depth of water at large distances from the outlet! Hint: The discharge volume of the sink is to be determined for the radius r = R0 !

m s2

90

3. Exercises in Fluid Mechanics (a) the curve along which the pressure equals the free-stream pressure p∞ , (b) the pressure on a circle around the origin of coordinates with radius 2R! 7.12 The pressure diﬀerence ∆p between two boreholes in a circular cylinder, with its axis normal to the direction of the oncoming ﬂow, is a measure for the angle between the free-stream direction and the axis of symmetry. Determine (a) the stagnation point and the velocity in the point x = xs , y = h, (b) the contour of the half-body, (c) the pressure distribution on the contour, (d) the isobars, (e) the curve along which the pressure is larger by the amount ρ u2 than 4 ∞ the pressure p∞ of the free-stream, (f) the isotachs, (g) the part of the ﬂow ﬁeld, in which the velocity component v is larger ∞ than u2 , (h) the curve, which is inclined to the streamlines by 45◦ , (i) the maximum deceleration a particle moving along the line of symmetry is experiencing between x = −∞ and the stagnation point!

7.9 The free-stream velocity of a twodimensional half-body with width 2h is u∞ .

(a) What is the relation between the pressure diﬀerence and the angle of attack? (b) At what angle α does ∆p attain its maximum value for every ? 7.13 Consider a ﬂow around a bridge pile with circular with cross section. The free-stream velocity is u∞ . The depth of water far upstream is h∞ .

7.10 Given the stream function Ψ = u∞ y 1− x2 R2 + y2 .

(a) Sketch the streamlines for x2 + y 2 ≥ R 2 ! Determine (b) the pressure distribution on the contour Ψ = 0, (c) the time it takes for a particle to move from the point x = −3 R, y = 0 to the point x = 2 R, y = 0! 7.11 Consider a parallel free-stream with velocity u∞ ﬂowing around a circular cylinder with radius R, with its axis normal to the direction of the oncoming ﬂow being in the origin of coordinates. Determine

m h∞ = 6 m R = 2 m s kg m ρ = 103 3 g = 10 2 m s Determine (a) the height of the water surface at the wall of the pile as function of θ, (b) the height of the water surface at the stagnation points, (c) the lowest depth of water, measured from the ground! Hint: Assume two-dimensional ﬂow! u∞ = 1

3.1 Problems 7.14 The roof (weight G) of a hangar of length L and semi-circular cross section rests on the walls of the hangar, without being ﬁxed. The hangar is completely closed except for a small opening on the leeward side.

91

L = 100 m H = 10 m G = 107 N kg m ρ = 1.25 3 α = 45◦ u∞ = 50 s m Investigate, whether the roof has to be anchored to the walls! Hint: Assume two-dimensional ﬂow! 7.15 Consider a ﬂow around a rotating circular cylinder of length L with its axis normal to the direction of the freestream with velocity u∞ . The circumferential velocity of the ﬂow on the surface of the cylinder caused by the rotation is vt .

Determine (a) the transition point for Recrit. = 5 · 105 , (b) the velocity at the point x = 0.1 m, y = 2 · 10−4 m with the aid of the Blasius solution! At what coordinate y does the velocity for x = 0.15 m attain the same value? Sketch (c) the variation of the boundary-layer thickness δ(x) and a velocity proﬁle for x < xcrit. and x > xcrit. , (d) the wall-shear stress as a function of x for dp dp dp 0 ! dx dx dx 8.3 The surface of a ﬂat plate is parallel to the dircetion of a free stream of water. Formulate the momentum thickness in terms of an integral over the wall-shear stress for a laminar boundary layer − x 0

τ (x; y = 0) dx dar! ρ u2 ∞

(a) Determine the circulation! (b) Discuss the ﬂow ﬁeld for vt = u∞ ! (c) Determine the force acting on the cylinder!

8.4 Air moves past a ﬂat plate (length L, width B), with its surface parallel to the direction of the free stream. m L = 0.5 m B = 1 m u∞ = 10 s kg m2 ρ = 1.25 3 ν = 1.5 · 10−5 m s (a) Sketch the velocity proﬁles u(y) for several values of x! (b) State the boundary conditions for the boundary-layer equations! (c) Sketch the distribution of the shear stress τ (y) for the position x! (d) Compute the boundary-layer thickness at the trailing edge of the plate and its drag! 8.5 The velocity proﬁle in the laminar boundary layer of a ﬂat plate at zero incidence (length L) can be approximated by a polynomial of fourth degree u y = a0 + a1 + a2 u∞ δ 3 y y + a3 + a4 δ δ y δ
4 2

3.1.8 Boundary Layers 8.1 Show that the drag coeﬃcient of the ﬂat plate at zero incidence is proportional 1 to √ReL for a laminar boundary layer! 8.2 The surface of a ﬂat plate is parallel to the direction of a free stream of air. m2 m ν = 1.5 · 10−5 u∞ = 45 s s

+

.

92

3. Exercises in Fluid Mechanics (a) Determine the coeﬃcients of the polynomial! (b) Prove the validity of the following relations δ1 3 = , δ 10 37 δ2 = , δ 315 δ 5.84 = √ , x Rex 1.371 cD = √ ReL Determine (a) the retaining forces F1 and F2 for L1 = L2 , (b) the lengths L1 and L1 for F1 = F2 ! 9.2 Two quadratic plates are exposed to a ﬂow, one at zero incidence, the other with its surface normal to the direction of the free stream.

8.6 The velocity proﬁle in a laminar boundary layer on a ﬂat plate at zero incidence of length L is approximated by A) a polynom of fourth degree u 3 = u∞ 2 y 1 − δ 2 y δ
3

and B) a sinosoidal ansatz u π y = sin u∞ 2 δ . u∞ = 5 m2 s (a) Explain the diﬀerence between frictional and pressure drag! (b) How large must L2 be, so that both plates generate the same drag? (c) How does the drag depend on the free-stream velocity? ν = 15 · 10−6 9.3 Two ﬂat rectangular plates are exposed to a parallel ﬂow. The plates have the same lateral lengths L1 and L2 . The edge of plate 1, with lateral length L1 is parallel to the direction of the free stream, and of plate 2 the edge with lateral length L2 . The free-stream velocity is u∞ . m2 s (a) Determine the ratio of the friction m m m forces for u∞ = 0.4 ; 0.8 ; 1.6 ! s s s (b) How large would the free-stream velocity for the plate 2 have to be, if the free-stream velocity of plate 1 is u∞ = 0.196 m , and if both drag cos L1 = 1 m L2 = 0.5 m ν = 10−6 m s L1 = 1 m ρ = 1.25 kg m3

(a) Determine δ1 , δ2 , δ, and cw ! (b) Compute the boundary-layer thickness at the trailing edge of the plate and the drag for m u∞ = 1 L = 0.5m B = 1m s kg m2 ! ρ = 103 3 ν = 10−6 m s 3.1.9 Drag 9.1 Two ﬂat plates at zero incidence, one downstream from the other, are exposed to the free-stream velocity u∞ .

u∞ = 1

m s

L = L1 + L2 = 0.36 m kg m3 ν = 10−6 m2 s

B = 1 m ρ = 103

3.1 Problems eﬃcients are supposed to have the same value? Hints: cD = 0.074 ReL
1 5

93

−

1700 ReL

for 5 · 105 < ReL < 107 . 9.4 A kite (surface area A, weight W ) generates the force F in the kite string at an angle of attack α. m L = 3.75 m B = 0.5 m u = 1.5 s kg Ns ρw = 103 3 µw = 10−3 2 m m kg ρa = 1.25 3 m Ns µa = 1.875·10−5 2 h = 4m b = 2m m Neglect the wave drag , the drag of the sail mounting, and the frictional drag of the upper side of the surfboard and determine (a) the wind speed u∞ , (b) the drag of the sail! (c) How large would the frictional drag of the upper side of the board be? Hints: 0.074 1700 Board : cD = 1 − ReL Re 5
L

m A = 0.5 m2 W = 10 N u∞ = 10 s α = 10◦ β = 55◦ F = 42.5 N kg ρ = 1.25 3 m Determine the lift and drag coeﬃcient of the kite! 9.5 Assume that the ﬂow around a circular cylinder separates at α = 120◦ , that the pressure distribution up to the separation point can be determined with the potential ﬂow theory, and that the pressure in the dead water region is constant!

(for : 5 · 105 < ReL < 107 ) Sail : cD = 1.2 (for : Re > 103 ) 9.7 How large must the surface of the equivalent drag of a parachute at least be, in order to avoid the sinking speed in quiescent air to exceed v? m v=4 W = 1000 N cD = 1.33 s kg ρ = 1.25 3 m 9.8 A sphere and a circular cylinder of the same material fall with constant velocity in quiescent air. The axis of the cylinder is normal to the direction of the gravitational acceleration. For 0 < Re ≤ 0.5 the drag coeﬃcient of a 24 sphere is given by cD = Re and that of a circular cylinder with its axis normal to the oncoming ﬂow by 8π cD = . Re(2 − ln Re)

Neglect the frictional drag and determine the drag coeﬃcient of the cylinder! 9.6 A surfboard (width b) moves with the velocity u over the surface of quiescent water. The height of the triangular sail is h and its width b.

94

3. Exercises in Fluid Mechanics ρ = 800 kg m3 ρa = 1.25 m2 s kg m3 Ds = 5 · 10−5 m ρs = 2.5 · 103 m s2 kg m3 Ns µa = 1.875 · 10−5 2 m

νa = 15 · 10−6

g = 10

Determine (a) the maximum diameters, for which these relations are valid, (b) the corresponding sink velocities! 9.9 Determine with the aid of the diagram for the drag coeﬃcient of a sphere

kg ρa = 1.25 3 m m g = 10 2 s (a) At what velocity of the air v1 are the dust particles suspended? (b) How large is the velocity of the dust particles if the velocity of the air is m va = 3 ? s Hint: Assume that the dust particles do not inﬂuence each other! cD = 3 24 (1 + Re) Re 16 for 0 < Re < 1

(a) the steady sink velocity of a spherical rain drop of diameter D in air, (b) the steady ascending velocity of a spherical bubble of air of diameter D in water! kg D = 1 mm ρw = 103 3 m m2 νw = 10−6 s m2 kg ρa = 1.25 3 νa = 15 · 10−6 m s m g = 10 s 9.10 Spherically shaped dust particles (density ρs ) are to be conveyed with a stream of air against the gravitational force.

9.11 A spherically shaped fog droplet (diameter D) is being suspended by an upward motion of air. At time t = 0 the air ﬂow stops and the droplet begins to sink. kg D = 6 · 10−5 m ρw = 103 3 m Ns kg ρa = 1.25 3 µa = 1.875 · 10−5 2 m m m g = 10 2 s (a) How large is the velocity of the air ﬂow prior to the lull? (b) After what time does the droplet attain 99% of its steady sink velocity? 24 Hint: For Re ≤ 0.5 the law cD = Re is valid for steady and unsteady ﬂow. 9.12 A sphere is falling in steady motion in quiescent air with the velocity v1 . A downwash squall increases the velocity to v2 .

m D = 0.35 m G = 4.06 N v1 = 13 s m kg ρa = 1.25 3 v2 = 18 m s m2 νa = 15 · 10−6 s

3.1 Problems (a) How large is the drag coeﬃcient prior to the squall? (b) What steady ﬁnal velocity does the sphere attain after dying out of the squall? 9.13 A spherically shaped deep-sea probe is heaved with constant velocity from the depth H to the surface of the sea in the time T1 .

95

D = 0.5 m H = 4000 m T1 = 3 h kg Ns m ρ = 103 3 µ = 10−3 2 g = 10 2 m m s Assume constant density of the water, neglect the weight of the cable rope, and determine (a) the power needed for heaving the probe, if the cable force is F1 = 2700 N, (b) the weight of the probe and the shortest heaving time, if the cable can take twice the value of the force, (c) the power then needed! Hint: Use the diagram of problem 9.12! 9.14 A sphere of diameter D and density ρs is vertically shot into quiescent air with initial velocity v∞ . Assume a constant drag coeﬃcient and determine (a) the height of rise, (b) the rise time, (c) the velocity at impact on the ground, (d) the falling time! (e) What values do these quantities attain for a wooden sphere of density ρw and for a metal sphere of density ρm , if cD = 0.4 and cw = 0? m m g = 10 2 D = 0.1 m v0 = 30 s s kg kg ρa = 1.25 3 ρw = 750 3 m m kg ρm = 7.5 · 103 3 m

96

3. Exercises in Fluid Mechanics AH − AB h1 = A (H + ∆H)− AB h2 m ∆H = ρA 1.5 (a) τ0 = (b) ∆W = ρM A (h0 − h1 ) g A = 1.95 · 104 m2 (c) W0 −∆W = ρM τ1 g = ρF τ2 g W0 − ∆W ρM τ2 − τ1 = −1 ρM g ρF = 2.44 · 103 m3 (d) τ2 − τ1 = (h2 − h1 ) A h2 = 10.625 m 1.6 W0 = 1.07 · 105 m3 ρm g

3.2 Solutions
3.2.1 Hydrostatics 1.1

p = pa + ρw g h = pa + ρF g (h − L) h kg = 3 · 103 3 ρF = ρw h−L m 1.2 pi = p2 = ∆h1 = ∆h2 = 1.3 FL = W FL = ρ1 ha2 g + ρ2 (a − h) a2 g W = ρK a3 g ρ2 − ρ K h = a = 6.67 · 10−2 m ρ 2 − ρ1 1.4 Fi + pa Ai p1 + ρ g ∆h1 = p3 − ρ g ∆h2 p2 − p1 = 0.25 m ρg p3 − p2 = 0.33 m ρg

(a) pa π 2 π D H = p D2 (H − h) 4 4 p = pa + ρw g (T − h) ρw g (T + h) + pa 2 ρw g ρw g(T +h) + pa 2ρw g
2

h = −

−T H

= 2m W1 = FL1 = ρ AB h1 g W2 = FL2 = ρ AB h2 g W2 = W1 + m g The volume of the water remains constant. (b) F = FL − Wair − W π 2 D g [(H − h) ρw − H ρa ] = 4 − G = −9.58 · 103 N

3.2 Solutions (c) 0 = FL − Wair − W ρa 4W − h0 = H 1 − ρw π D 2 ρw g pa T0 = h0 1 + ρw g (H − h0 ) = 18.3 m 1.7 (b) F zF = zF = 1.10 dF = [pi (z) − pa ] dA sin α dA = R dα 2 π R cos α π 3 F = R ρ g = 1.05 · 104 N 3 1.8 s=L z=h1 z=0 1 h3 1 3 h2 1

97

(a) dF z z h2 [p1 (z) − p2 (z)] B dz h1 : p1 (z) = pa + ρg(h1 −z) h2 : p2 (z) = pa + ρg(h2 −z) z ≤ h1 : p2 = pa 1 ρ g B(h2 − h2 ) F = 1 2 2 6 = 1.05 · 10 N = ≤ ≤ ≤

z dF

− h3 2 = 1.86 m − h2 2

FL = s=0 s dF

dF = [pi (s) − pa ] B ds pi (s) = pa + ρ g [L sin α − z(s)] ρ g B h2 = 3 · 104 N F = 6 sin α dFL = ρ g (h − z) 2π r dr R r = H −h+z H π 2 R H ρg FL = 6 F = W − FL π 2 ρ R H (ρs − ) g = 3 2 = 1.57 · 10−2 N 1.9 1.11

(a) dF = [pi (z) − pa ] b(z) dz pi (z) = pa + ρ g (H − z) 1 ρ g B 3 = 2.5 · 103 N F = 4 (b) z=H F zF = zF

z dF z=0 √ 3 B = 0.217 m = 8

98 1.12

3. Exercises in Fluid Mechanics (b) ρ(z) = p(z) R (T0 − α z) αz p = p0 1 − T0
1 κ

g Rα

(c) ρ(z) = ρ0 Boundary condition: r = 0, z = h : p − pa p = pa 1 = ρ g (h − z) + ρ ω 2 r2 2 (d) p/p0 a) b) c) 3000 m 0.695 0.686 0.681 6000 m 0.483 0.457 0.442 11000 m 0.263 0.215 0.186

p p0 1−

p = p0

γ − 1 gz γ R T0

κ κ−1

(a) Surface: p = pa z0 (r) = h + r=R: z0 = H ω2 = 2 g ω2 r2 2g 1.14 (a)

H −h R2

F = FL (z = 0) − W = (ρ0 τ0 − m) g = 10.6 N r2 R2 (b) for z ≤ z1 : p 1 τ1 = p 0 τ0 gz p1 − 1 = e R T0 p0 τ1 p0 ln z1 = ρ0 g τ0 = 10.0 km

z0 (r) = h + (H − h) Volume of the water: πR2 h0 =
R 0

z0 (r) 2 π r dr

1 = πR2 h + πR2 (H −h) 2 (c) h = 2 h0 − H = 0.4 m 4g 1 ω = (H − h0 ) = 13.9 R2 s (b) r = R : p = pa +ρ g (H − z) ρ z = 0 : p = pa +ρ g h + ω 2 r2 2 1.13 dp = −ρ(z) g dz (a) ρ(z) = p(z) R T0 p = p0 e gz − RT 0

FL (z2 ) = W ρ0 τ 1 p0 ln z2 = ρ0 g m = 12.8 km 1.15 FL (z) FL (0) FL Ggas ρgas (z) ρgas (0) = = = = W + Wgas (z) W + Wgas (0) + Fs ρτ g ρgas τ g gz = e− RT RT Fs ln 1 + g W = 7.84 km

z =

3.2 Solutions 3.2.2 Hydrodynamics 2.1 (a) dy v v0 tan(ω t) = =− dy u u0 v0 y = − tan(ω t) x + c u0 Straight lines with slopes 0,−1,−∞ (b) x(t) = u dt + c1 u0 sin(ω t) + c1 = ω v dt + c2 v0 cos(ω t) + c2 ω
2

99

2.3

y(t) = =

dp 1 + d (v 2 ) + g dz = 0 ρ 2 p = ρ RT inner: RTi ln outer: RTa ln v= 2.4 ∆p = β (a) Assumption: ρ v∞ D m > 250 v∞ = 0.5 µ s ρ v∞ D = 3000 µ ρ 2 v 2 ∞ p1a p0a +gH = 0 Ti −1 Ta = 31.6 m s p1i p0i + v2 +gH =0 2

ω u0 ω v0

(x − c1 )2 + (y − c2 )2 = 1

2

Circles with radius 1 m. (c) Circle around the origin 2.2

2gH

p1 + ρc g h = p2 + ρc

v2 2 p2 = p1 + ρw g h ρ c − ρw v = 2gh ρc m = 3.16 s

(b) Assumption: m ρ v∞ D > 250 v∞ = 0.5 s µ ρ v∞ D = 300 µ (c) Assumption: 2.5 < v∞ = 0.45 m s ρ v∞ D < 250 µ ρ v∞ D = 27 µ

100 2.5

3. Exercises in Fluid Mechanics v2 = 2 ∆p ρ 1−
A2 A1 2

v1

m = 12 s m = 4 v3 s m = 6 s

∆p = v∞ D 2 v∞
2 ∆p ρ

ρ 2 v 2 = v (D2 − d2 ) d2 = 1− 2 D

(b) p2 + ρ 2 ρ 2 v = p3 + v3 2 2 2 5 N p3 = pa = 10 2 m p2 = 0.46 · 105

N m2 N p1 = 1.1 · 105 2 m ρ 2 p + ρ g h = pa + v3 2 N p = 1.08 · 105 2 m 2.8

2.6

ρ 2 ρ g (h + H) = ρ g H + v1 2 ρ 2 v = 2 2 2 v1 D = v2 d2 h = 0.05 m d=D 4 h+H 2.7 (a) p1 + ρ 2 ρ 2 v = p2 + v2 2 1 2 v1 A1 = v2 A2 = v3 A3

pa + ρ g (a + h) = pa + ρ g z + ˙ Q = =
2a 0

ρ 2

v(z)2

v(z) B dz (h + a)3 − (h−a)3 3 (1 + a 3 ) h a h a (1 − h )3

2 2gB 3

˙ ˙ Q0 − Q = ˙ Q a/h ˙ ˙ Q0 −Q ˙ Q

−

−1

0.25 0.264% 0.75 2.728%

0.5 1.108% 1.0 6.066%

a/h
˙ ˙ Q0 −Q ˙ Q

3.2 Solutions 2.9 (a) vD AD = va A

101

p0 = pa + ∆p = pa + vD = (b) 2 ∆p A ρ AD

ρ 2 v 2 a A m = 4 AD s

p0 −pa = (p0 − pD ) − (pa − pD ) ρ 2 ρ 2 ρ 2 v − ηD ( v D − v a ) = 2 D 2 2 2∆p A ρ AD A AD
2

(a) pa + ρ g H = p5 − ρ g s + p5 = pa + ρ g s ˙ Q = Ad (b) 2gh=4 m3 s (c) ρ 2 v 2 5

vD = · vD =

(1 − ηD ) + ηD 4
A AD 2

−1 2

0.16

A AD

+ 0.84

A →∞: AD 2.11 (c) Vapor bubbles are formed, if p2 = p3 = pv .
∗ v5 A∗ = v ∗ A d

vD = 10

m s

pa = pv + ρ g h + A∗ d

ρ ∗2 v 2 h pa − pv − = A ρgH H = 0.244 m2

(a) π D2 ˙ αD Q = mD 4 2 (p1 − p2 )D ρw

2.10

p1 +ρw g h1 = p2 + ρw g (h1 −hD ) + ρHg g hD m3 ˙ Q = 0.07 s (b) π D2 2 (p1 − p2 )B ˙ αB Q = mB 4 ρw mD hD αB = αD = 0.75 mB hB

102 2.12

3. Exercises in Fluid Mechanics T = −
0 dh B √ 4f h0 2 g h h0 B = 100 s = 2f 2 g

2.14

pa + ρ g z0 +

ρ 2 v = p1 + ρ g z1 2 0 ρ 2 + v1 2 p1 + ρ g z1 = pa + ρ g z2 pa + ρ g z0 + ρ 2 ρ 2 v = pa + v2 2 0 2 2 2 v0 v2

The assumption of quasi-steady ﬂow reA 2 2 quires that As 1: v0 v1 . v1 A = v0 As = − T =− As A
0 h0

dh As dt dh As 2 h0 √ = A g 2gh
2 h0 g

Volume rate of ﬂow: A dz0 AB = v1 A − v2 dt 3 v1 = 2gh
4h 3 AB dz √ 0√ T =√ · 2g A h 3 h − z0 √ √ 3 AB =√ · 2 (3 h − z0 )− 2g A √ √ 4h √ −3 h ln(3 h − z0 ) h

A = As 2.13

T

= 5 m2

AB T =6 A 2.15 pa + ρ g z0 + ρ 2 v = p1 + ρ g z1 2 0 ρ 2 + v1 2 p1 + ρ g z1 = pa + ρ g z2

h [3 ln 2 − 1] = 108 s 2g

2 2 v0 v1 dz0 B B = − v1 f = v0 2 dt 2 dz2 dz0 = − dt dt dh d(z0 − z2 ) dz0 = = 2 dt dt dt

p1 + ρB g z1 +

ρB 2 ρB 2 v = pa + v 2 1 2 2 pa = p1 + ρw g z1 2 2 v1 v2 2 v2 d = v1 D 2 dz1 2 D = dt

3.2 Solutions T = D d
0 −L 2

103

ρB 2 g (ρw − ρB )

dz √ 1 −z1
2

s1 s1

= 2.16

D d

ρB ρ w − ρB

2L = 80 s g

∂v ds ∂t dv2 ∂v D 1 ds ≈ L L ∂t dt v0 2 dv2 T = −2 L 2 v2 v0 2L = √ =2s 2gh +ρ s1 T

s1

(c) Q = A
0

v2 dt v0 2

= −2 A L = 2.18

π L D2 ln 2 = 0.279 m3 2

v0

dv2 v2

L −∞

∂v ds = ∂t + =

D −√

8

−∞ L
D −√ 8

∂ ⎝ v0 π D ⎠ 4 ds + ∂t 2 πs2 ∂v0 ds = ∂t (a) pa + ρB g (h + z2 ) = pa + ρ g z2 ρ 2 + v 2 2 s2 ∂v ds + ρ s0 ∂t s2 ∂v dv2 ds ≈ L dt s0 ∂t √ 0.99 2 g h dv2 T = 2L 2 2 g h − v2 0 L = √ 2gh √ √ 0.99 2 g h 2 g h + v2 ln √ 2 g h − v2 0 = 10.6 s (b) pa = p1 + ρ g h1 + + ρ L1 ρ 2 v 2 2

⎛

2

⎞

D dv0 √ +L dt 2

2.17

(a) v0 = m 2 g h = 10 s

(b) pa + ρ 2 ρ 2 v = pa + v2 2 1 2

pa = p1e

dv2 dt ρ 2 = ρ g h1 + v2e 2

104

3. Exercises in Fluid Mechanics for: 1 dv2 = dt L v2 = 0.99 gh− v2 2 (a) pa +ρ g h1 = pa + QI = A
0

ρ 2 dv v +ρL 2 dt vdt vI 0

2gh

TI

p1 − p1e = ρ g h 1 − 0.992 · L1 1− · L N = 746 2 m 2.19

= 2 AL

vdv 2 g h1 − v 2 v2 = −AL ln 1− I 2 g h1

Determination of vI : TI = 2 L dv 2 g h1 − v 2 √ 2 g h1 + vI L = √ ln √ 2 g h1 2 g h1 − vI
0 vI

vI =

2 g h1

e

TI TI

√

√

2 g h1 L 2 g h1 L

−1 +1

e QI = 0.240 m3 (b) pa = pp + ρ g h + ∂v ds + ρ s1 ∂t ∂v dvp s0 ≈ L ( ∂t dt L sp ρ 2 v 2 p

pa + ρ g h1 = pa + ρ g (h1 + h2 ) dv ρ 2 + v +ρL 2 dt
TII

sp s1

1) QII = A

v dt
TI 0

pp = p a − ρ g h + ρ s0 ω 2 × s0 L sin ωt − cos2 ωt 2 ppmin = pv pp = ppmin with cos ωt = 0 dpp = 0) dt 1 pa − pv − ρ g h = 8.8 ρ s0 L s (follows from 2.21

vdv 2 g h2 + v 2 2 vI = AL ln 1 + 2 g h2 = 0.194 m3 = −2 AL vI ω= 2.20

(a) pa + ρ g h = pa + ρ 2 v 2 2 dv1 dv2 + L2 + ρ L1 dt dt

3.2 Solutions
2 2 v1 D1 = v2 D2 2

105 gh

pB = pa +

1 − 0.992 1+
L1 L2 D2 D1

2ρ

D2 T = 2 L1 D1 √ · =
0.99 0 2

+ L2 · dv2 2 2 g h − v2 ·
0.99 0

= 1.003 · 105 (d)

N m2

2gh

L1

√

D2 D1

+ L2

√ 2 g h + v2 · ln √ 2 g h − v2 = 5.231 s (b) Q = · L1
0.99 0

2gh

√

2gh

D2 D1 √

2

+ L2 A2 · v2 dv2 2 2 g h − v2
2

2.22

2gh

= − L1 ·

D2 D1

2 ln 2 g h − v2

0.99: 0

+ L2 A2 · √
2gh

= 0.048 m3 (c) ρ 2 v = pa + 2 1 ρ 2 pB + v2 = pa + 2 pA + dv2 ρ 2 v + ρ L2 2 2 dt dv2 ρ 2 v 2 + ρ L2 2 dt (a) pa + ρ g h = p1s + v1s = ˙ Q0 A ⎛ ρ 2 v 2 1s
⎞

as shown under (a):
2 2 g h − v2 dv2 = 2 dt D2 2 L1 D1 + L2

p1s − pa = ρ ⎝g h −

˙ Q2 ⎠ 0 2 A2 N m2

= 18.875 · 105 (b)

t = 0: pA = pB = pA + 1+ = 1.16 · 105 t = T: pA = pa + ρ g h 0.992 1 − + 1 − 0.992 1+
L1 L2 D2 D1 2 4 D2 4 D1

ρgh
L1 L2 D2 D1 2

pa +ρ g h = p1 +

N m2

ρ 2 v + 2 1 Dv1 + ρL dt ˙ ˙0 1 − t Q(t) = Q Ts ˙ Q0 − A TS t TS
2

p1 (t) = pa +ρ g h+ρ L − ρ 2 ˙ Q0 A
2

1−

= 1.187 · 105

N m2

106

3. Exercises in Fluid Mechanics (c)
∗ A2 sin α2 − A3 sin α3 = 0 ∗ α3 = 12.37◦

3.2 (a)

1. ∆pzul = p1max − pa = ρgh+ρL TS = 0.25 s 3.2.3 Momentum and Moment of Momentum Theorem 3.1 Fs1 ˙ Q0 A TS
2 2 −ρ vD AD − 2 ρ v1 A1 = Fs1

vD vD AD ρ 2 pa + vD 2 = −4 ρ g h AD

= 2gh = 2 v1 A1 ρ 2 = pa + v1 2 = −2 · 104 N

2 ρ vD AD = (pa + ρ g h − pa ) A + + Fs2 Fs2 = ρ g h(2 AD − A) = 0

(a) ρ 2 ρ 2 ρ 2 p1 + v1 = pa + v2 = pa + v3 2 2 2 ∆p = p1 − pa v1 A1 = v2 A2 + v3 A3 v1 = 2 ∆p ρ 1
A1 A2 +A3 2

(b)

−1

m = 2.58 s v2 = v3 = A1 v1 A2 + A3 m = 5.16 s

2 ρ vD AD = (pa + ρ g h − pa ) A Fs1 = −2 ρ g h A = −2 · 104 N Fs2 = 0

(b)
2 2 ρ v3 A3 cos α3 + ρ v2 A2 cos α2 2 −ρ v1 A1 = (p1 − pa ) A1 + Fsx Fsx = −866.4 N 2 2 ρ v2 A2 sin α2 − ρ v3 A3 sin α3 = Fsy Fsy = −238.4 N

3.3

3.2 Solutions (a) P = −FS vr Moving control surface:
2 −ρ ve A ve ρ 2 pa + ve 2 ve A 2 − 2 ρ va A∗ cos β = Fs = v0 − vr ρ 2 = pa + va 2 = 2 v a A∗

107

(c) ρ 2 2 p01 −p02 = (p1 −p2 ) + (v1 −v2 ) 2 ρ 2 v1 tan2 α = 2 (d)
2 Fx = −Fsx = ρ v1 Bt tan2 α 2 Fy = −Fsy = −ρ v1 Bt tan α

P = ρ (v0 −vr )2 vr A (1+cos β) dP = 0 dvr v0 m vr = = 20 3 s (b) 4 2 Fs = − ρ v0 A (1 + cos β) 9 = 3.08 · 105 N 3.4

Cascade II (b) p1 + ρ 2 ρ 2 v = p2 + v2 2 1 2 ρ 2 p1 − p2 = v1 tan2 α 2

(c) p01 − p02 = 0 (d) Momentum equ. as for cascade I Fx = Fy ρ 2 v Bt tan2 α 2 1 2 = −ρ v1 Bt tan α

3.5 (a) v1 B t = v2 B t cos α v1 v2 = cos α Cascade I (b)
2 2 −ρ v1 B t + ρ v1 B t cos2 α = (p1 − p2 ) B t + Fsx 2 ρ v2 B t sin α cos α = Fsy Force normal to the blade: Fsx tan α = − Fsy 2 p1 − p2 = −ρ v1 tan2 α

(a) ρ1 v1 A∞ = ρ1 v1 (A∞ − AR ) + ∆m ˙ ∆m = ρ1 v1 AR ˙ (b)
2 2 −ρ1 v1 A∞ + ρ1 v1 (A∞ − AR ) + 2 +ρA vA AR +

ρ1 vx vr dA = Fs
AM

108

3. Exercises in Fluid Mechanics For
AM

A∞ AR

1

:

vx = v1 ρ1 vr dA
AM

(b) ρ 2 ρ 2 v = p1 + v1 2 1 2 2 2 (p1 − p1 ) + v1 v1 = v2 = ρ kg m = ρ A v1 = 13 · 103 ˙ s p1 + (c)

ρ1 vx vr dA = v1

= v1 ∆ m ˙ 2 Fs = ρA vA AR P = Fs v1 2 = ρA vA v1 AR 3.6

2 2 −ρ v1 A∞ − ∆m v1 + ρ v2 A ˙ 2 +ρ v1 (A∞ − A) = Fs

See problem 4.5. ρ v1 A∞ + ∆m = ρ v1 (A∞ − A) ˙ +ρ v2 A Fs = ρ v2 (v2 − v1 ) A = 0.39 · 105 N (d) ˙ P = Q (p02 − p01 ) ˙ = Q (p1 − p1 ) = 448.5 kW 3.8

(a) ρ 2 ρ v = p1 + v 2 2 2 1 ρ ρ 2 p2 + v 2 = pa + v2 2 2 v1 A1 = v A = v2 A2 pa + 0 = − + − (p1 − p2 ) A + Fs 2 2 ρ v1 A∞ + ρ v2 A2 2 ρ v1 (A∞ − A2 ) ∆ m v1 = Fs ˙

See problem 4.3
2 2 ˙ ρ v1 A∞ + ∆ m = ρ v2 A2 2 + ρ v1 (A∞ − A2 ) v1 − v2 v = 2 m = 6.5 s

(a) ρ 2 v 2 1 ∆p = p2 − p1 = pa − p1 2∆p ˙ Q = vA= A ρ pa = p1 + (b) ˙ P = Q (p02 − p01 ) = 2∆p ∆p A ρ

(b) η= 3.7 (a) v1 Fs v1 = = 0.769 Fs v v

(c) ρ v 2 A = Fs = 2 ∆p A

3.2 Solutions 3.10

109

(d) ρ v 2 A = (pa − p1 ) A ∆p = p2 − p1 = pa − p1 ∆p ˙ A Q=vA = ρ (e) ˙ P = Q (p02 − p01 ) = (f) ρ v 2 A = Fs = ∆p A 3.9 ∆p ∆p A ρ

(a) v2 = pa p2 (b)
2 2 2 −ρ v1 A1 − ρ v2 A2 + ρ v3 A3 = (p2 − pa ) A3 v1 A1 + v2 A2 = v3 A3

˙ Q2 m = 40 A3 − A1 s ρ 2 = p2 + v2 2 N = 0.99 · 105 2 m

⎛

⎞

v1

⎜ = ⎝1 +

1 2
A1 A3

1−

A1 A3

⎟ ⎠ v2

= 96.6 v3 = v1

m s

A1 A1 + v2 1 − A3 A3 m = 68.3 s

(a) ρ 2 ρ 2 v = p2 + v2 2 ∞ 2 2 2 −ρ v2 A2 + ρ v1 A1 = (p2 −p∞ ) A1 v2 A2 = v1 A1 p∞ + v2 = 1−2 m = 56.6 s m = 28.3 s v∞
A2 A1 2

(c) ˙ P = Q1 (p01 − p01 ) ˙ = Q1 (p1 − p1 ) p1 = p2 ρ 2 pa = p1 + v1 2 ρ 2 2 p = v1 − v2 v1 A1 2 = 46.6 kW (d)
2 ρ v1 A1 = (pa − p2 ) A1 + Fs v2 2 Fs = ρ v1 − 2 A1 2 = 1066 N

+2

A2 A1

v1 (b)

2 2 −ρv3 A1 +ρv1 A1 = (p3 −p∞ )A1 +Fs

p∞ + Fs =

ρ 2 2 (v − v∞ ) A1 = −100 N 2 1

ρ 2 ρ 2 v = p3 + v3 2 ∞ 2 v3 = v1

(Traction force)

110 3.11

3. Exercises in Fluid Mechanics 3.12

(a)
2 2 −ρ v1 A1 + ρ va A2 = (p1 − pa ) A2 ρ 2 pa + ρ g h = p1 + v1 2 v1 A1 = va A2

(a) π D2 π D2 ˙ =α Q = v1 4 4 v1 = α d2 D2 2 ∆ pw ρ

√ ˙ Q = 2
A1 A2

2ghA
2

2 ∆ pw m = 1.33 ρ s

−2

A1 A2

+1

(b)
2 2 −ρ v2 A2 + ρ v1 A1 = (p2 −pE ) A1 ρ 2 pa = p2 + ρ g H + v2 2 v2 A2 = v1 A1

˙ dQ = 0 dA2 A2 = 2 A1 = 0.2 m2 (b) N p1 = pa − ρ g h = 5 10 m2
4

pE = pa −

ρ 2 ⎣ D2 −1 v 2 1 d2

⎡

2

⎤

+ 1⎦

(c)

− ρ g H = 0.482 · 105

N m2 N m2

pA = pK + ρ g h = 2.3 · 105 (c) P = v1

π D2 (poA − poE ) 4 3 = 1.9 · 10 kW

2 2 −ρ v1 A1 + ρ va A2 = (p1 − pa ) A1 +Fs1

3.13

Fs1 = −ρ g h A1 = −5 103 N (Traction force) Fs2 = 0 2 ρ va A2 = Fs3 Fs3 = 2 ρ g h A1 = 104 N (Compressive force)

(a)
2 ρ ve Ae = (p0 − pe ) A ρ 2 p0 = pe + ve 2 Ae = 0.5 Ψ = A

3.2 Solutions (b)
2 ρ v1 A = (p0 − p1 ) A p0 = pa + ρ g (h + h1 ) p1 = pa + ρ g h1 m v1 = g h = 3.16 s

111

(a)
2 2 −ρ v1 B h1 +ρ v2 B h2 = ρ g B

h2 1 − 2 2 h − ρgB 2 2

3.14 (a)

v1 B h1 = v2 B h2 g h2 v1 = (h1 + h2 ) 2 h1 m = 1.73 s m v2 = 0.87 s (b) v1 F r1 = √ = 1.73 gh F r2 = 0.61 (c)

(b) ˙ Q = αmA 2 ∆ pw ρ m3 = 7.67 · 10−2 s

H1 − H2 = h1 − h2 + = 3.16

1 2 (v 2 − v2 ) 2g 1

(h2 − h1 )3 = 0.0125 m 4 h1 h2

(c)
2 2 −ρ ve Ae + ρ v1 A = (pe − p1 ) A ρ 2 pa = pe + ve 2

˙ Q = v1 A1 = ve Ae = ve Ψ m A p1 = pa − α 2 m2 · 1 2 +1 · ∆ pw 1 − Ψm N = 0.998 · 105 2 m (d) ˙ P = Q (pa − p1 ) = 14.4 W 3.15

(a) For each streamline it is, pa + ρ g h = p + ρ g z + p + ρ g z = pa + ρ g h1 v = 2 g (h − h1 ) = v1 (independent of z) (b) ˙ Q = v1 B h1 ˙ dQ = 0 dh1 2 = h = 5m 3 ρ 2 v 2

h1max (b)

v1 F r1 = √ =1 gh 2 h1gr = h=5m 3

112

3. Exercises in Fluid Mechanics (c) ρgB h2 h2 1 − 2 2 2 v1 B h1
⎛
2 = −ρ v1 B h1 + 2 +ρ v2 B h2 = v2 B h2

(e)
2 2 −ρ v1 B h1 + ρ v3 B h3 2 2 h1 h3 − =ρgB + Fs 2 2

⎞

h2 =

h h1 ⎝ 1+16 −1 −1⎠ h1 2

F = − Fs =

= 5.93 m m v2 = 4.22 s 3.17 (a) 3.18

ρg B (h1 − h0 ) · 2 ˙ 2 Q2 h1 + h0 − · 2h h gB 1 0 = 2.60 · 105 N

(b) Zcrit. + Hmin = H0 ˙ Q2 3 − 2 g B 2 h2 2 0 = 0.446 m ˙ Q2 g B2

Zcrit. = h0 +

3

(a) − ρ v 2 dr dθ dθ − ρ v 2 dr = 2 2 dp dr dr = p− r− dθ − dr 2 2 dp dr dr r+ dθ + − p+ dr 2 2 dθ + 2 p dr 2 v2 dp =ρ dr r (b) d ρ p + v2 = 0 dr 2 v2 dp dv = ρ = −ρ v dr dr r const. v = r

(c) zW > zcrit. : h2 = hcrit. =
3

˙ Q2 g B2

= 1.170 m H1 = ZW + Hmin H1 ZW = + 1 = 1.570 Hmin Hmin from diagram: h1 = 2.26 hcrit. h1 = 2.64 m (Elevation of the upper water) (d) H1 − H 3 = H 1 − H 0 = ZW − Zcrit. = 0.554 m

3.2 Solutions 3.19 (a) (c) pa + ρ g (h + H) = p1 + v1 = ˙ Q A1 ρ 2 v 2 1

113

p1 = 1.095 · 105 (d) Mr = 3 ρ vr A (r x v a )z = 3 ρ vr A R(−vr cos α + ω R) Determination of vr : ξ0 = ω0 1 = 1.73 tan α 1 = 163 s

N m2

3.2.4 Laminar Flow of Viscous Fluids 4.1

pa = pa +

s2 ρ 2 v − ρ (b · ds) 2 r s0 ρ 2 ω 2 R2 = pa + vr −ρ gH + 2 2

(a) (p1 − p2 ) π r2 − τ 2 π r L = 0 τ = −µ du dr
2

vr = with

2 g H + 2 ω 2 R2

ωR and ξ = √ 2gh M0 = −3 ρ A R 2 g H cos α : ξ Mr = 1+ξ 2 1+ξ 2 − M0 cos α 2
Mr M0 Mr M0

r=R: u = 0 u(r) r = 1− umax R (b) ˙ Q u(m) = umax umax π R2
1

ξ2 =
⎡ ⎢ ⎛

+ tan2 α − 1 · 2 tan2 α ⎤
⎞2 ⎥ ⎠ − 1⎥ ⎦

= 2 × = 1 2

·⎢ 1 + ⎝ ⎣

1− 2
Mr M0

0

1−

r R

2

×

2 tan α

+ tan2 α − 1

r r d R R

ξ = 0.07 n = 1.05 (b) m3 ˙ Q = 3 vr A = 2.13 · 10−3 s 1 s (c) 8 τw ρ u2 m 4 µ um τw = R 64 λ = Re λ =

114

3. Exercises in Fluid Mechanics ˙ Q = B δ 0

4.2 For y ≤ a the ﬂuid behaves as a rigid body. (a)

u(y) dy

dτ = ρ g sin α dy du τ = −µ dy y=0: u = 0 y=δ: τ = 0 ρ g sin α y2 δy − 2 µ ρ g B sin α 3 ˙ δ Q = 3µ m3 = 1.2 · 10−3 s

u(y) = ρ g a ∆z = τ0 ∆z τ0 a= ρg (b) 4.4 (a)

a≤y≤b:

dτ = ρg dy τ = −µ dw + τ0 dy

y=a: y=b:

τ = τ0 w = 0 − dτ +ρg = 0 dy τ = −µ d2 w ρ g = 0 + dy 2 µ y=0: w = 0 y=δ: τ = 0 ρg 2 δ µ y 1 − δ 2 y δ
2

ρg [(b − a)2 − (y − a)2 ] w(y) = 2µ 0≤y≤a: 4.3 w(y) = ρg (b − a)2 2µ

dw dy

w(y) =

3.2 Solutions (b) (b) uw + τ (y = H) = τ (y = 0) uw − (c) ˙ Q = um B H = B
2 H 0 1 dp 2 µ dx 1 dp 2 µ dx

115

H2 H2

u(y) dy BH

= τ 2 π dz − dτ dr − τ+ dr 2 π (r + dr) dz + (e) dIx =B dt (f) (d)

dp H uw − 2 dx 12 µ

umax = −

dp H 2 dx 8 µ

+ρ g π (r + dr)2 − r2 dz = 0 −r dτ −τ +ρgr = 0 dr τ = −µ d dr dw dr

H 0

ρ u(y)2 dy =

6 2 ρ u BH 5 m

dw ρg r=0 + dr µ r=a : w=0 r =a+δ : τ =0 r ρg × 2µ × (a + δ)2 ln r a2 − r2 + a 2

ρ 2

τw 12 = u2 Re m

(g)

w(r) =

4.5 (a)

dp dτ + = 0 dx dy τ = d2 u = dy 2 y=0: u = y=H: u = u(y) = 1 dp 2 H 2 mu dx y + uw H du −µ dy 1 dp µ dx 0 uw y H
2

4.6

−

y + H

116

3. Exercises in Fluid Mechanics (a) dp 1 d(τ r) + =0 dx r dr Compare problem 5.4b τ = −µ du dr (b) µ = τ −r d dr v r r=Ri 2

1 d du 1 dp =0 r − dr µ dx r dr r=a: u=0 r=R: u=0 u(r) = −
⎡

2 Mz = −τ (r = Ri ) 2π Ri L Mz Ri µ = 1− 2 Ra 4 π ωRi L Ns = 10−2 2 m

4.8

1 dp 2 (R − a2 ) × 4 µ dx r R a R

× ⎣

R2 − r2 ln − R2 − a2 ln

⎤ ⎦

ρ

de dx dy = dt ∂q dy ∂y u+ dx + τ u dx − ∂u dy ∂y dx

= (b) a R 2 a2 ln R + R2 − a2 τ (r = a) = a τ (r = R) a 2 R2 ln R + R2 − a2

q− q+ τ+

−

∂τ dy ∂y

(c) um ˙ Q = 2 − a2 ) π (R R 1 = u(r) 2πr dr π (R2 − a2 ) a a 2 1 dp 2 R 1+ = − + 8µ dx R + 1− ln a R a R 2

dτ = 0 dy ∂T ∂y ∂u τ = −µ ∂y q = −λ ρ ∂2T de = λ +µ dt ∂y 2 ∂u ∂y
2

⎤ ⎥ ⎦

(a) It follows from problem 5.5a): u(y) = uw y H ∂T ∂T ∂T +u +v ∂t ∂x ∂y

4.7 (a) d 1 d (r v) = 0 dr r dr r = Ri : r = Ra : v(r) = v=0 v = ω Ra

ρ

2 2 ω Ra r2 − Ri 2 − R2 r Ra i

µ uw 2 λ H H : T = Tw 0: q=0 µ uw T (y) = Tw + 2λ H −

de = dt = d2 T = dy 2 y = y =

ρ cv 0

2

(H 2 −y 2 )

3.2 Solutions (b) qw = −λ (c) h0 = cp T + h0 = ∂h0 = ∂x ∂h0 = ∂y = = Pr = (d) ∂T = 0 ∂y µ uw 2 dT = dt ρ cv H t = 0 : T = T0 µ uw T (t) = T0 + ρ cv H 4.9 (a) v u = cp T + 2 2 const: ∇ h0 = 0 2 ∂( u ) ∂T 2 cp + =0 ∂x ∂x u2 ∂( 2 ) ∂T + cp ∂y ∂y du dT +u cp dy dy µ cp uw 2 1− y λ H ∂h0 µ cp =1: =0 λ ∂y
2 2

117

dT dy

= y=H µ u2 w H

r=R : u=0 du =0 r=0 : dr Temperature distribution: 0 = q 2 π r dx + τ u 2π r dx − dq − q+ dr 2π (r + dr) dx − dr dτ du dr dr · u+ − τ+ dr dr · 2 π (r + dr) dx + dp + p− p+ dx · dx · u π (r + dr)2 − r2 dT dr 2 du dT r + µr =0 dr dr r=R : T = Tw dT =0 r=0 : dr q = −λ

λ

d dr

(b) It follows from problem 5.1a u(r) r 2 = 1− umax R r µ u2 max T − Tw = 1− 4λ R r=0: Tmax − Tw = T − Tw Tmax − Tw 4.10 (a) Re (b) du dr ˙ h(x) Q = u(x,y) dy B 0 = const. dp ∂2u = µ 2 dx ∂y y=0 : u = u∞ y = h(x) : u=0 h1 L
2

2

t

4

µ u2 max 4λ r = 1− R

4

Velocity distribution: τ +r dτ dp +r = 0 dx dr τ = −µ

= 1.6 · 10−3

r

d dp du −µ r dx dr dr

=0

118

3. Exercises in Fluid Mechanics u(x,y) = u∞ − y − h(x) dp y 1− h(x) dx 1− = 3K h1 µ u∞ 1 + 3 1− 2 h(x) h(x) y × 1−2 h(x) ×

h2 (x) y 2 η h(x) ˙ Q u∞ h(x) h3 (x) dp = − B 2 12 µ dx x=0: p = p∞ x p(x) = p∞ + 6 µ u∞ − 12 µ ˙ Q B

0 x 0

dx − h2 (x ) dx h3 (x )

P L = 20 µ u2 × ∞ B h1 1 9K 2 (e 5 − 1) × e5 − 1 − 16 W = 55.9 m 4.11

With p(x = L) = p∞ yields: ˙ u∞ Q = B 2
L dx 0 h2 (x) L dx 0 h3 (x)

2 ˙ 3 e5 − 1 Q u∞ h1 3 = 4 B e5 − 1 2 −5 m = 4.49 · 10 s

(c) With K = e5 − 1 e5 − 1
2x 3 2

dIx = dt =

τ

∂(ρ u) dτ + ∂t Fx

A

ρ u (v · n) dA

: τ p(x) = p∞ + 15 · (d) Fpy = B L h1
L 0

e 5L

µ u∞ L · h1 h1 3x − K e 5L + K − 1

∂ρ ∂(ρ u) ∂u dτ = ρ +u dx dy dz ∂t ∂t ∂t ρ u (v · n) dA = dy dz−

A

−ρ u2 + ρ u2 + (p(x) − p∞ ) dx

∂(ρ u2 ) dx ∂x ∂(ρ uv) dy ∂y

− −ρ uv + ρ uv +
2

dx dz

= 15µ u∞

5 2 (e 5 −1)+K −1 6 = ρu N m

= 3035 (e) P = u∞ B τxy = −µ
L 0

∂u ∂u + ρv + ∂x ∂y ∂(ρ u) ∂(ρ v) + +u ∂x ∂y × dx dy dz

×

τxy (x,y = 0) dx

Continuity equation: ∂ρ ∂(ρ u) ∂(ρ v) + + =0 ∂t ∂x ∂y

du dr

3.2 Solutions Fx = − ∂σxx ∂τxy + ∂x ∂y (τyx = τxy ) ∂u ∂u ∂u +u +v ∂t ∂x ∂y ∂σxx ∂τxy + =− ∂x ∂y u2 ¯m1 dx dy dz (a) p1 + ρ 2 ρ u = p a + u2 ¯ ¯ 2 m1 2 m2

119

ρ

π D2 π d2 = u2 ¯m2 4 4 2(p1 −pa ) u2 ¯m1 = ρ m = 1 s m = 4 s

1
D d 4

−1

4.12 µ = const. : ∂2u ∂2u + + Frx = µ ∂x2 ∂y 2 µ ∂ (∇ · v) + 3 ∂x ρ = const. : ∂2u ∂2u Frx = µ + + ∂x2 ∂y 2 3.2.5 Pipe Flows 5.1 ρ um D ¯ Re 4τ um = ¯ π D2 T ρ L ρ g (h + L) = (1 + λ + ζe ) u2 ¯ D 2 m µ = Assumption: Laminar ﬂow ζe = 1.16 λ = 11.92 64 = 5.37 Re = λ Intake region: Le = 0.029 Re D = 0.16 · 10−3 m Ns µ = 8.40 · 10−3 2 m 5.2 5.4 (c) um2 = ¯ 5.3 um2 ¯ ∂ u + ∂z 2
2

(b) p0 = p1 + (1 + λ L ρ 2 ) u ¯ D 2 m1

∂2u ∂z 2

Re1 =

ρ um1 D ¯ = 104 µ 0.316 = 0.0316 λ = √ 4 Re N p0 = 2.66 · 105 2 m

m 2 (p0 − pa ) = 18.22 ρ s

π 2 2 ˙ u ¯ Q = 25 (¯m1 D1 + um2 D2 ) 4 L ρ ¯ ∆p = (1 + λ + ζ) u2 D1 2 m1 L ρ = (1 + λ + ζ) u2 ¯ D2 2 m2 um1 = ¯ 2 ∆p L ρ (1 + λ D1 + ζ) 2 ∆p L ρ (1 + λ D2 + ζ)

L

um2 = ¯

m3 ˙ Q = 0.518 s

∆p = λ um ¯

L ρ 2 u ¯ D 2 m 2 ∆p D 1 = √ ρ L λ

120

3. Exercises in Fluid Mechanics um ¯ √ ρD √ λ = Re λ µ 2 ∆p D ρ D = ρ L µ √ 64 = 384 (Re λ)crit = 2300 2300 √ (b) Compressible ﬂow: dρ d¯m u + um ¯ =0 dx dx (Continuity equation) ρ ρ1 p (T = const.) p1 ρ1 dp p1 dx ρ1 p1 u2 dp ¯m1 p2 dx ¯m1 λ ρ1 p1 u2 2D p 0

Oil: Re λ = 283 √ 1 Re λ √ = 64 λ m um = 1.56 s

ρ = dρ = dx dp − dx + =

Water: √ Re λ = 3.16 · 104 √ 1 √ = 2.0 log(Re λ) − 0.8 λ m um = 2.59 ¯ s
2

Re = Re1 = 0.533 · 105 0.316 = 0.0208 λ = √ 4 Re p dp − ρ1 p1 u2 ¯m1 +
2 1

5.5 (a)

1

dp + p
2 1

λ ρ1 p 1 u 2 ¯m1 2D = 0
⎡

dx

L = − ρ um ¯ dρ dx × = ρ+ dx d¯m u dx ¯ × um + dx = ρ1 um1 ¯ d¯m u dx dx
2

D p1 ⎣ p2 1− λ ρ1 u 2 ¯m1 p1 D p1 ln λ p2
2

2

⎤ ⎦−

= 287.9 m

Incompressible ﬂow: d¯m u = 0 dx dp λ = − dx D D L = 2 λ 5.6 (a) r = R−y ˙ I =
R 0

ρ+

dρ dx dx

um + ¯

πD2 − 4

ρ1 2 u ¯ 2 m1 p1 − p2 = 384.7 m ρ1 u 2 ¯m1

−ρ u2 ¯m

πD2 = −τw π D dx+ 4 πD2 dp + p− p+ dx dx 4

8 τw λ = ρ u2 ¯m λ ρ 2 dp d¯m u + ρ1 um1 + u =0 ¯ ¯ dx dx D 2 m

ρ u2 2π r dr
1 0

= 2ρ u2 π R m

u um

2

r r d R R

3.2 Solutions u =1 um ˙ I = ρ u2 π R 2 m δ = R: u r 2 = 2 1− um R ˙ I = 1.33 ρ u2 π R2 δ = 0: m 121 ×

L2 =

L1 D2 D2 2 1 − λ1 2 λ2 D2 D1 4 D2 −2 × 4 D1 = 1.0 m

(b) ∆pv = p1 −pa + 1−
4 D1 4 D2

R : δ = 2 ⎧ ⎨ 96 r 1− r u = ⎩ 17 R 24 R um
17

ρ 2 u ¯ 2 m1

R 2

≤r≤R
R 2

= 117.2 5.8 ρgh = + λ

N m2

0≤r≤

˙ I = 2ρ u2 π R2 × m 0.5 24 2 r r d + × 17 R R 0 1 96 r r 2 + 1− × R 0.5 17 R r r × d R R = 1.196 ρ u2 π R2 m (b) τw = µ u m
2 δ 2

L + 2 ζK + ζv D

ρ 2 u 2 d 2 ud d = um D 2 ¯ π d2 ˙ Q = ud 4 2gh 1+ d D 4

ρ 2 u + ¯ 2 m

=

λ

L D

+ 2 ζK + ζV

π d2 4

1−

2 δ 3 R

+

1 6

δ R

H=

u2 d 2g

δ→0 : δ→R : δ→ R : 2

τw → ∞ µ um τw = 4 R µ um τw = 5.65 R

with losses: (a) m3 ˙ Q = 5.79 · 10−3 s (b) H = 6.96 m

5.7 (a) p1 + L2 ρ 2 u ¯ = p a + 1 + λ2 × 2 m1 D2 ρ 2 × u + ¯ 2 m2 L1 ρ 2 u ¯ + ζ + λ1 D1 2 m1 2 D2 ζ = 1− 1 2 D2 2 2 um1 D1 = um2 D2 ¯ ¯

m3 ˙ Q = 9.82 · 10−3 s loss-free: (a) m3 ˙ Q = 6.94 · 10−3 s (b) m3 ˙ Q = 27.77 · 10−3 s

H = 1.25 m

H=h

Re1 = 104 λ1 = 0.0316 Re2 = 5 · 103 λ1 = 0.0376

H=h

122

3. Exercises in Fluid Mechanics (b) λ = um ¯ τw (b) um ¯ umax ¯ = − 4.07 u∗ u∗ u∗ 2 λ = 8 um ¯ um ¯ 1 = 0.84 = umax ¯ 1 + 4.07 λ
8

5.9 (a) 0.316 √ 4 Re µ Re = ρD N λ ρ u2 ¯m = 2.22 2 = 8 m

˙ I = ρ um π R 2 ¯

R 0

ρ u2 2 π r dr ¯ ρ u2 π R 2 ¯m
2

= 2 · = 5.11 (a)
1 0

umax ¯ u ¯ 1−

·
2 7

r R

r d R

r R

50 49

(c) u ¯ y u∗ =5= ν u∗ (Viscous sub-layer) (b) λ m um = 0.236 ¯ u=5 ¯ 8 s (for y = 0.11 mm) y u∗ = 50 ν (Logarithmic velocity distribution) u ¯ y u∗ = 2.5 ln + 5.5 u∗ ν m u = 0.720 ¯ s (for y = 1.1 mm) (d) l = 0.4 y = 0.4 = 0.85 mm 5.10 (a) r = R−y ˙ um ¯ Q = umax ¯ π R2 umax ¯ 1 r = 2 1− R 0 49 = 60 (d) ˙ P = Q(p2 − p1 ) = 160.5 kW
1 7

pa + ρ g h1 = p1 + 1 + λ

L 2D

ρ 2 u ¯ 2 m

π D2 ˙ Q = um ¯ 4 um D ¯ = 8 · 105 Re = ν R = 250 ks (From diagram, page 29) λ = 0.024 p1 = 1.22 · 105 N m2

y u∗ ν

ν λ 8

(c) um ¯ p2 = pa + ρ g h2 + λ = 3.77 · 105 N m2 L ρ 2 u ¯ 2D 2 m

r r d R R

3.2 Solutions 5.12 5.14 λps L ρ 2 Lch ρ 2 = λch u ¯ u ¯ D 2 mps dh 2 mch dh = a

123

Rech = Reps (a) L ρ 2 u ¯ ∆p = λ D 2 m π D2 ˙ Q = um ¯ 4 π 2 ∆p D5 = 0.0356 λ = ˙ 8 ρ L Q2 (b) Re = ρ um D ¯ = 1.8 · 105 η R = 60 ks ks = 4.2 mm (c) ∆p π D2 − τw π DL = 0 4 N D = 16 2 τw = ∆p 4L m π D2 = −2517 N F = −∆p 4 λ = 0.016 (From diagram page 29) ∆p = 5.8 · 103 5.13 P1 = P2 1 + λ1 1+
L D1 L λ2 D2 2 ρ 2 ρ 2

˙ ρa Q = 105 µ a2 ˙ ρD Q = µ 100 π D2 4

λch = 0.018 λch = 0.030

= 1.27 · 104 Lch = 13.57 m 5.15

(From diagram, page 29) ρ g L sin α + dh ρ 2 L u = 1+λ ¯ dh 2 m 4 a2 = 3a ρ 2 u ¯ 2 m

˙ ρ dh Q = 8 · 103 λ = 0.033 µ a2 α = 0.02◦ ˙ 2 1 ρ Q ∆pv N = λ = 3.61 3 L dh 2 a2 m Reh = 3.2.6 Similar Flows 6.1 u= ¯ u v p ρ , v= , p= ¯ ¯ , ρ= , ¯ v1 v1 ∆p1 ρ1 µ x y ¯ t µ= , x= ¯ ¯ , y= ¯ , t= µ1 L1 L1 t1 ∂u ¯ +u ¯ ∂t ∂p ¯ µ ¯ = −Eu + ∂ x Re ¯ ρ Sr ¯ Sr = ∂u ¯ ∂u ¯ +v ¯ ∂x ¯ ∂y ¯ ∂2u ∂2u ¯ ¯ + ∂ x2 ∂ y 2 ¯ ¯

(d)

N m2 u2 ¯m1 u2 ¯m2

πD ˙ Q = um ¯ 4 R 1 √ + 1.74 = 2.0 log ks λ L 1 + λ1 D1 D2 4 P1 = = 39.2 L P2 1 + λ2 D2 D1

∆p1 L1 , Eu = , 2 v1 t1 ρ1 v 1 ρ1 v1 L1 Re = µ1

124 6.2

3. Exercises in Fluid Mechanics 6.6 (a) f1 (Re,Eu) = 0 : Eu = f2 (Re) D D ∆p = 2 Eu = f3 (Re) λ= L L ρ u2 2 m (b) Sr = Sr : Remin = 200 ν = 1 mm v

Dmin = Remin

6.3 (a) L3 T −1 = (M L−2 T −2 )α · (M L−1 T −1 )β Lγ α = 1 β = −1 γ = 4 ∆p D4 ˙ Q ∼ Lµ (b) D ∆p λ = L ρ u2 2 m ˙ ∆p D um µ Qη ∼ ∼ L D3 D 1 λ ∼ Re 6.4
2 cDa π Da ρ2a u2 FDa ∞L 4 = π 2 W FDw cDw 4 Dw ρ2 u2 ∞w w cDa µ2 ρ2 Re2 a a = cDw µ2 ρ2a Re2 w w Rea = Rew : cDa = cDw FDa = 0.281 FDw

Sr = Sr : 6.7 (a)

f=

v Dmin 1 f =2 v D s

Re = Re :

v =

A v A

Small, so that ﬂow in wind-tunnel remains incompressible: A = Am : v = 77.46 (b) FD F v FD D cD ρ v 2 A 2 F v = cD ρ v 2 Am D 2 Re = Re : cD = cD P = FD v = 24.3 kW P = 6.8 (a) Re = Re : ˙ Q = vA ˙ Q vA µρ DA ˙ = Q µ ρ DA 2 A D = A D m3 ˙ Q = 9 s m s

6.5 F (f,µ,ρ,u∞ ,D) = 0 K1 = f α1 ρβ1 uγ1 Dδ1 ∞ Take α1 = 1 : β1 = 0. γ1 = −1, δ1 = 1 fD K1 = = Sr u∞ α 2 β2 γ2 K2 = µ ρ u ∞ D δ 2 Take α2 = −1 : β2 = 1, γ1 = 1, δ2 = 1 ρ u∞ D = Re K2 = µ

(b) Eu = Eu : ∆p = ρ v2 ∆p ρ v2 ˙ ρ Q2 A 2 ∆p = ˙2 ρ Q A2 N = 4.94 · 103 2 m

3.2 Solutions 6.9 (a) ˙ Re = Re : Q = = = Sr = Sr : n = = (b) Eu = Eu : ∆p0 ρ v2 = ∆p0 ρ v2 N = 527 2 m = 0.75 (b) cD = cD FD = ˙ P = Q∆p0 = 15.82 kW ˙ P = Q ∆p0 = 15 kW P = 201 Nm M = 2πn P M = = 179 Nm 2πn 6.10 (a) T =T : Ma = Ma : Re = Re : µ v p m s ρ D = p p= ρ D N = 4 · 105 2 m = v = 200 v D 1 n = 400 vD s = µ300
K ρ 2 v B 2 ρ v2B 2

125

(b) v D ˙ Q v D2 η ρD ˙ Q ηρ D 3 m 0.5 s v D n vD 1 13.3 s
2

Fr = Fr : oder Re = Re : 6.12 (a)

v = v

L = 0.1 L

L v = = 100 v L

F rL = F rL : m s

v =v

L L

(c)

H F H D

= 1.64 · 104 N (c) cD = 1.12 (d) v v √ : h = 16 h = 0.4 m =√ gh gh 6.13 (a) Eu = Eu , Re = Re : p0 − pR ρ v2 ρ D2 µ 2 = = 2 p0 − pR ρv ρ D2 µ2 N p0 − pR = 440 2 m Sr = Sr : vD T = 0.57 s T = v D (b) Analogously to a)

(b) Sr = Sr : (c) P = 6.11 (a) Re = Re : Fr = Fr : ν ν = 10−3 v L ν = ν vL L v = v L (!) ρ 2 ρ 2

n =

v 3 D2 P =4P v3D2 6.14 (a)

Pmax = 6.68 · 105

N m2

Re = Re :

v

m3 ˙ Q = 8 · 10−4 s

ρµ D ρ µD m = 0.1 s =

π 4

˙ Q D2

126

3. Exercises in Fluid Mechanics (c) α=
4 ˙Q π d2 2 ∆ pD ρ

Γ = =√
D d 2

v ds = 0 c ω = 0 (d) E =

2 EuD

EuD = EuD : α = α = 0.6366 Eul = Eul : ζ = ζ = 77.1 7.2 (a) ρ v2 = ∆pD ρ v2 = 0.625 · 105 ∆pl = N m2

ρ 2 v H 2 π r dr 2 θ 2 2 = π ρ H ω r0 (0.25 + ln 2) = 3.7 · 108 Nm
0

2 r0

(d) ∆pD

u=

∂Φ ∂x

v=

∂Φ ∂y

Potential Φ exists, if ∂Ψ u= ∂y ∇xv=0. ∂Ψ v=− ∂x

ρ v2 ∆pl ρ v2 N = 0.5 · 105 2 m

Stream function exists, if, ∇·v =0 .

3.2.7 Potential Flows of Incompressible Fluids 7.1 (a)

(b) ∇ · v = ∇2 Φ ∇ x v = −∇2 Ψ k

7.3
a) b) c) d)

(b) Γ = c 2π

∇·v 4xy 2 0 0

|∇xv| y 2 − x2 0 -2 0

v ds
0

Stream function exists, for c) and d), the potential for b) and d). Determination of the stream function: (a) Ψ = u dy + f (x) = y2 + f (x) 2

= = (c)

vθ (r) r dθ 2 πω r2 r ≤ r0 2 2 πω r0 r > r0

v = − Ψ = (b)

∂Ψ = −f (x) = −x ∂x

1 2 (x + y 2 ) + c 2

1 Ψ = (y 2 − x2 ) + c 2

3.2 Solutions Determination of the potential: (c) Φ = v = u dx + f (y) = x + f (y) 2
2

127

(d) dx L x2 = ln u U x1 x2 = 5.44 m Ψ = const : x1 y1 = x2 y2 y2 = 0.74 m t= x1 x2

∂Φ = f (y) = y ∂y 1 2 (x + y 2 ) + c Φ = 2 (d) Φ = xy + c 7.4 (a) | ∇ x v |= 0 : (b) u= U x L v=− U y L Potential exists.

(e) p1 − p2 = (cp1 − cp2 ) ρ 2 v 2 ref N = 0.442 · 105 2 m

(f)

Stagnation points: u=v=0: x=y=0 7.5 (a)
2

Pressure coeﬃcient: p − pref cp = ρ 2 v 2 ref = 1− = 1− Isotachs: v 2 = u2 + v 2 = U L
2

u +v 2 u2 + vref ref x2 + y 2 2 2 xref + yref

2

u = 2xy v = x2 − y 2 ∇·v = 0 : Stream function exists. (b) Ψ = x y2 − Streamlines: Ψ = const. x3 +c 3

(x2 + y 2 ) vL U
2

x2 + y 2 =

x3 k + 3 x Asymptotes: y=± x → ±∞ : x → ±0 :

(x = 0)

Circles around the origin of coordinates with radius |v| L U (c) m m v1 = −4 s s m | v 1 |= 5.66 s ρ = pref + cp1 v 2 2 ref N = 0.86 · 105 2 m u1 = 4

x y = ±√ 3 y → ±∞

p1

128 7.6

3. Exercises in Fluid Mechanics

vr =

c r

vθ = 0

(a) ∇xv=0 ∇·v =0 (b) Φ = vr dr + f1 (θ) Parallel ﬂow: n = 1 : Ψ = r sin θ = y 1 n = 2 : Ψ = r2 sin(2 θ) = xy 2 See problem 8.4. (b) = 0: f2 (r) = k2 cp = (c) Circle with radius r:
2π

= c ln r + f1 (θ) ∂Φ = 0 : f1 (θ) = k1 ∂θ Ψ = ∂Ψ ∂r r vr dθ = cθ + f2 (r)

p − pref v2 =1− 2 ρ 2 v v ref 2 ref

Γ =

0

vr = 0

vθ r dθ = 0 c vθ = r

vθ = −rn−1 sin(n θ) vr = rn−1 cos(n θ) cp = 1 − n = 1: n > 1: n < 1: x2 − y 2 2 cp (0.0) = 0 cp (0.0) = 1 cp (0.0) = −∞ n−1 (a) ∇xv = 0 ∇·v = 0 7.8 (a) (b) Φ = c θ + k3 Ψ = −c ln x + k4

(c) Γ =2πc 7.7 (a) n = 0.5: Ψ = 2 √ r sin θ 2 θ = 0.2 π c 2 θ r= sin−2 2 2 Ψ = vr vθ Γ r E θ− ln 2π 2π R E = 2πr Γ = 2πr

Ψ = 0: Ψ = c:

3.2 Solutions tan α = − vr vθ Stagnation point: u = v = 0 : r=R0 129

˙ Q E = − h0 ˙ m2 Q Γ = = 4.33 · 10−2 h0 tan α s (b) ρ ρ ρ g h + v 2 = ρ g h0 + v 2 2 2 0 2 2 2 v = vr + vθ v2 = v2 0 r=R θ 1 h(r) = h0 + 8g ˙ Q π R0 h0 sin α R0 · 1− r
2

h xs = − , ys = 0 π π2 u(xs ,h) = u∞ 1 + π2 π2 v(xs ,h) = u∞ 1 + π2 (b) Contour: Streamline through stagnation point rC = with θ h θ h π−θ = π sin θ π sin θ

= π−θ

·
2

(c) cp = 1 − cpC = h u2 + v 2 h 2x + π =− 2 2 u∞ π x + y2 2

(c) lim h(r) = h0 + r→∞ ˙ 1 Q 8g π R0 h0 sin α = 2.35 · 10−2 m
2

sin(2 θ ) sin θ − θ θ

(d) cp = h π cp
2

const:
2

7.9 x+ + y 2 = (1−cp ) h π cp

Circles around − πh p .0 with radius c √ h 1−cp π cp

(e) cp = 1 2h : x+ 2 π
2

+ y2 = 2

h π

2

(a) Ψ = u∞ y + E θ+c 2π h y = u∞ y + arctan π x x h u = u∞ 1 + π x2 + y 2 h y v = u∞ 2 π x + y2

+c

130

3. Exercises in Fluid Mechanics (f) √ 7.10 (a) u2 + u∞
2

v2

= k kh π 2

Ψ =0: y = 0 x2 + y 2 = R 2 (Parallel ﬂow) r= x2 + y 2 → ∞ : Ψ → u∞ y

x−

h π

k2

−1

+ y2 = h .0 π (k2 −1)

k2

−1

Circles around kh dius π (k2 −1) (g) v = u∞

with ra-

Stream function describes the ﬂow around a cylinder.

y h u∞ > π x2 + y 2 2 h π
2

x2 + y −

<

h π

2

(b) cp = 1 − vr = u∞
2 2 vr + vθ 2 u∞ R 1− r

2

cos θ R r
2

vθ = −u∞ 1 + r=R: v =1 u 2 2 2 h h 1 h x+ + y− = 2π 2π 2 π tan α = (c)
2R

sin θ

(h)

cp = 1 − 4 sin2 θ dx u(x.0) R2 1− 2 x
−2R −3R

∆t =

−3R

u(x.0) = u∞

∆t =

x−R 1 R ln x+ u∞ 2 x+R R = (1 + 0.5 ln 1.5) u∞

7.11 Determination of velocity components see problem 8.10. cp = (i) Acceleration along the x-Axis: du =u b = dt h = −u∞ π db =0: dx ∂du ∂dx 1 h 1 + x2 π x3 3 h 2 π 4 π 2 u = − 27 h ∞ (a) cp = 0 r = √
2 R r 2

2 cos(2 θ) −

R r

2

R 2 cos(2 θ) √ −

or
2

xmax = − bmax

2x R

2y R

=1

Hyperbola

3.2 Solutions 7.14

131

(b) cp = 7.12 (a) 7 − sin θ 16

dFy = (pi − p) LH sin θ dθ ρ pi = p ∞ + u2 2 ∞ ρ p = p ∞ + c p u2 2 ∞ ρ = p∞ + (1 − 4 sin2 θ) u2 2 ∞ Fy =
3π 4 π 4

2 ρ u2 LH sin3 θ dθ ∞ 1 sin2 θ cos θ− 3

= 2 ρ u2 LH − ∞ − ρ ∆p = (cp1 −cp1 ) u2 2 ∞ cp = 1 − 4 sin2 α sin2 α2 − sin2 α1 = sin(α1 + α2 ) · sin(α2 − α1 ) ∆p = 2 ρ u2 sin(2 α) sin(2 ) ∞ (b) π α= 2 7.13 (a) ρ g h∞ + ρ 2 ρ u = ρ g h(θ) + v 2 2 ∞ 2 2 r = R : v 2 = vθ = 4u2 sin2 θ ∞ u2 h(θ) − h∞ = ∞ (1−4 sin2 θ) 2g vr 2 cos θ 3
6
3π 4 π 4

= 7.37 · 10 N < G Mooring not necessary. 7.15 (a) Ψ = u∞ r sin θ 1 − r Γ ln 2π R R = u∞ 1 − r − R r
2

−

2

cos θ R r
2

vθ = −u∞ 1 + + Γ 2πr vθ

sin θ ∗

r = R:

vortex

= vt =

Γ 2πR

Γ = 2 π R vt (b) Flow ﬁeld for vt = u∞ : Ψ = u∞ r sin θ 1 − − u∞ R ln vr = u∞ 1 − vθ = u∞ Ψ = 0: r R R r R r
2

(b) Stagnation points: θ = 0 und θ = π u2 h = h∞ + ∞ = 6.05 m 2g (c) θmin hmin = h∞ − π = , 2 3π 2

−

2

cos θ sin θ

3 u2 ∞ = 5.85 m 2g

R R2 − 1+ 2 r r

132

3. Exercises in Fluid Mechanics Contour: Circle around the origin of coordinates with radius R. 2 Stagnation points on the contour (r = R): θs = π , 5π ; nor free Stagna6 6 tion points 3.2.8 Boundary Layers 8.1 cD ∼ τw ¯ FD ∼ ρ u2 BL ρ u2 ∞ ∞ η u∞ µ δ ∼ ∼ ρ u2 ρ u∞ δ ∞

Inertia and frictional forces of equal order of magnitude: ρ u2 τw ¯ ∞ ∼ L δ cD ∼ √ 8.2 (a) xcrit. = (b) dFx = −pLR cos θ dθ dFy = −pLR sin θ dθ ρ p = p∞ + c p u2 2 ∞ r = R: vt − 2 sin θ cp = 1 − u∞ Fx = −LR
2π 0

(c)

1 ReL

ν Recrit = 0.167 m u∞

y Rex = 1.095 η= x u m = 0.36 : u = 16.2 u∞ s from diagram, page 60
2

x = 0.15 m : Rex = 4.5 · 105 y Rex = 1.095 : y = 2.45 · 10−4 m x (c)

ρ 2 u × 2 ∞
2

vt − 2 sin θ × 1− u∞ × cos θ dθ − − LR Fy = −LR × 1−
2π 0 0 2π

×

p∞ cos θ dθ = 0 ρ 2 u × 2 ∞
2

vt − 2 sin θ u∞ × sin θ dθ − − LR
2π

×

(d)

= −2 π ρLR vt u∞ = −ρ u∞ Γ L

0

p∞ sin θ dθ

3.2 Solutions 8.3 (d) From Blasius solution:

133

0.664 5L cf = √ δ(x = L) = √ ReL Rex δ(x = L) = 4.33 mm
L

Fw = 2 u2 ∞ δ 2

0

B τw dx
L 0

−ρ

= ρ u2 B ∞ = 0.144 N 8.5 (a) Boundary conditions: (u∞ − u) dy y = 0: δ y = 1: δ u = 0, u∞ u =1 u∞ x cf dx

δ+ρ

0

u dy + ∆ m u∞ = ˙ =
0

τ (x ,y = 0) dx δ 0

∆m=ρ ˙

v =0 u∞

u u 1− = u∞ u∞ x τ (x ,y = 0) = δ2 = − dx ρ u2 0 ∞
0

δ

From boundary-layer equation: ρ u ∂u ∂u +v ∂x ∂y y =0: δ ∂2 ∂ y =1: δ =η ∂2u : ∂y 2

8.4 (a) ReL = 3.33 105 : Boundary layer laminar

u=v = 0: u u∞ y 2 δ

= 0

∂u ∂u = = 0: ∂x ∂y ∂2 u u∞ y 2 δ

(b) y=0 : y→∞ : (c) From boundary-layer equation: ∂τ = 0 for y = 0 und y = δ ∂y u=v=0 u → u∞

= 0

∂

Inviscid external ﬂow y = 1: δ u = 2 u∞ (b) δ1 = δ = δ2 = δ =
1 0

τ∼

∂ ∂

u u∞ y δ

=0
3

y −2 δ

y δ

+

y δ

4

1−

u u∞

d

y δ

3 10
0

1

u u∞

1−

u u∞

d

y δ

37 315

134

3. Exercises in Fluid Mechanics Von K´rm´n integral relation a a dδ2 τ (y = 0) + =0 dx ρ u∞ τ (y = 0) = − µ u2 d ∞ δ d u u∞

3.2.9 Drag 9.1 (a) F1 = cD1 ρ 2 u 2 L1 B 2 ∞ 1.8 · 105 < Recrit. 1.328 √ = 3.13 · 10−3 ReL1 0.564 N 3.6 · 105 < Recrit. F1 + F2 1.328 ρ 2 √ u 2LB ReL 2 ∞ 0.233 N

y 2 δ

ReL1 = y =0 δ

cD1 = F1 = ReL = Ftot = = F2 = (b)

µ u∞ = −2 δ Integration: 5.84 δ = √ x Rex 2 L τw cD = dx L 0 ρ u2 ∞ 2 L τ (y = 0) = − dx L 0 ρ u2 ∞ 1.371 = √ ReL 8.6 (a) Solution see problem 9.5 A) δ1 3 , = 8 δ δ2 39 = , δ 280 δ 4.641 , = √ x Rex 1.293 cD = √ ReL B) 2 δ1 = 1 − = 0.363 δ π δ2 2 1 = − = 0.137 δ π 2 δ 4.795 = √ 4−π = √ x Rex Rex 2 cD = (b) A) δ(x = L) = 3.288 mm FD = cD ρ u2 2 L B ∞ = 0.91 N B) δ(x = L) = 3.39 mm FD = 0.93 N √ 2− ReL π 2 2 π2

Ftot = 2 F1 L = 0.09 m L1 = 4 L2 = 0.27 m

9.2 (a) The frictional drag results from the shear stresses acting on the body, the pressure drag results from change of the potential pressure distribution caused by the frictional force. (b) cD1 ρ ρ 2 u 2 L2 = cD2 u2 2 L2 1 2 2 ∞ 2 ∞ u ∞ L1 = 3.33 · 105 Re1 = ν 1.328 = 2.30 · 10−3 cD1 = √ Re1

Assumption: Re2 = L2 Re2 (c)
3 1.328 ρ 2 2 u 2 L2 ∼ u ∞ FD1 = √ 1 Re1 2 ∞ ρ FD2 = 1.1 u2 L2 ∼ u2 ∞ 2 ∞ 2

1.310 =√ ReL

u ∞ L2 > 103 : cD2 = 1.1 ν 2 cD1 = L1 = 6.47 · 10−2 m cD2 = 2.16 · 104 > 103

3.2 Solutions 9.3 (a) FD1 cD1 = cD2 FD2 u∞ 0.4 0.8 1.6 m s m s m s

135

9.5

Re1 4 · 105 8 · 105 1.6 · 105 103 cD2
5 5

103 cD1 2.10 2.76 3.19
FD1 FD2

Re1 2 · 10 4 · 10

2.97 2.10 2.76

0.707 1.313 1.156

8 · 105

(b) Re1 = 1.96 · 105 cD1 = cD2 = 3.0 · 10−3 1) Re2 = Re1 : 2) Re2 ≈ 1.3 · 106 : u∞2 = 2.6 3) Re3 ≈ 9 · 106 : u∞2 = 18 m s m s u∞2 = 0.392 m s

dFw (α) u2 DL ∞ dFD (α) = dF (α) cos α D = p(α) L cos α dα 2 ρ 2 p(α) = cp (α) u∞ + p∞ 2 cD = 2 ρ 2

π 0

0≤α≤

2 π: 3 ρ 2 u + p∞ 2 ∞

p(α) = (1 − 4 sin2 α)

2 ≤α≤π: 3 ρ 2 2 1 − 4 sin2 ( π) u + p∞ 3 2 ∞ ρ 2 = −2 u∞ + p∞ √ 2 = 3

(2) and 3) from diagram page 64). 9.4

p(α) =

cD 9.6 (a)

FDS = FDB cDB cL = FL − cL = cD = FD − cD = FL ρ 2 u∞ A 2 W − F sin(β − α) = 0 W + F sin(β − α) = 1.28 ρ 2 u A 2 ∞ FD ρ 2 u∞ A 2 F cos(β − α) = 0 F cos(β − α) = 0.96 ρ 2 u A 2 ∞ ρw 2 ρa u LB = cDS × 2 2 ×(u∞ − u)2 ReL = ρW u L µW = 5.625 · 106 1700 1 − ReL 5 ReL −3 = 3.0 · 10 0.074 hb 2

cDB =

136

3. Exercises in Fluid Mechanics Assumption: Reb = cDS = u∞ = = Reb = (b) FDS = 6.33 N (c)
∗ FDB = c∗ DB

ρa (u∞ − u) b > 103 µa 1.2 cDB ρW 2 LB u 1+ cDS ρa hb m 2.95 s 1.94 · 105 > 103 9.9

Re = 0.5 : DC max = 4.71 · 10−2 mm (b) m s m vC = 0.159 s vS = 0.110

ν D (a) (Lift can be neglected:) v = Re FD = W = ρW FD = = Re √ cD = = π D3 g 6 2 ρa 2 π D v cD 2 4 π 2 cD Re2 ρa va 8 8 FD 1 π ρa νa D 4 ρW Dg 3 ρa νa 217.7 m s

Re∗ L c∗ DB

ρa u∞ − u2 LB 2 ρa (u∞ − u) L = = 3.63 · 105 µa 1.328 = = 2.20 · 10−3 Re∗ L FDB

∗ FDB = 5.45 · 10−3

9.7 ρ cD v 2 A = W 2 W A= = 75.2 m2 cD ρ v 2 2 9.8 (a) FD = W (Lift can be neglected) Sphere: cD
2 3 π DS ρa 2 π D S v =ρ g 2 4 6

=

from diagram: Re = 250 v = 3.75 (b) (Weight can be neglected) FD = FL = ρW FD = Re √ cD = π D3 g 6 2 ρW 2 π D v cD 2 4 D 4 Dg = 115 3 νW 113 m 0.11 s

νa v = Re DS DS =
3

18 Re

2 ρL νa ρ g

Re = v = 9.10

Re = 0.5 : DS max = 6.81 · 10−2 mm Cylinder (Length L): cD ρa 2 π D2 v DC L = ρ Lg 2 4 2 16 Re ρa νa DC = 3 2 − ln Re ρ g

3.2 Solutions W = FD (Lift can be neglected) vrel = va − vs (a) vs = 0 2 3 ρa 2 π D S π DS g v1 cDW = ρS 4 6 2 ρa v1 DS Assumption: Re = 3.6 · 105 Re2 = µ T2 from diagram: cD2 = 0.1 T2 = HD π ρ cD2 8 (F1 + FD1 ) = 241.0 s = 8.3 · 106 H = 1 g vB 0

v dv 1− v vs 2 2

= − vB =

2 vB vs ln 1 − 2g vs vs

1+ (d) TB = − = (e) 1 g

v vs

2

Re2 (d)

vB 0

dv 1− v vs 2

H = 89.64 kW P2 = 2 F1 T2 9.14 (a) ρW π D dv πD g− = −ρS 6 dt 6 ρa 2 π D 2 − cD2 v 2 4 4 ρS D g − 3 ρa c D dz = dt = v
0 v0 3 3

2 vs + vB vs ln g vs − vB

cD = 0.4 wooden metal sphere sphere H [m] TH [s] vB m s TH [s] 37.2 2.64 24.9 2.81 44.0 2.96 29.3 2.98

cw = 0

Introduce steady sinking velocity:
2 vs =

45 3 30 3

−

dv 1 g 1+ v vs H = − =

2

1 g

v dv 1+ v vs 2 2

2 vs v ln 1 + 2g vs

(b) TH = − = 1 g
0 v0

dv 1+ v vs 2

2 v0 vs arctan g vs

(c) ρS π D3 dv ρa π D 2 g = − cD v 2 + 6 dt 2 4 π D3 + ρS g 6 1 dv dz = dt = g 1− v 2 v vs 4. Gasdynamics

4.1 Introduction
In the ﬁrst part of this chapter, which is a continuation of ﬂuid mechanics II, after a short repetition of the important thermodynamic relations, one- dimensional, steady, isentropic ﬂows of compressible gases in a stream tube with variable cross-section are described. Then normal compression shocks in one-dimensional supersonic ﬂows associated with a discontinuous increase of entropy will be explained. It will be shown, that the increase of entropy is due to heat conduction and disspation of mechanical energy in the shock. The description of normal shocks quite naturally leads to oblique shocks, which are discussed as third topic. The properties of strong and weak shocks are explained with the aid of the shock polar diagram, and the jump conditions for weak shocks are subsequently used as starting point for the derivation of the relations for isentropic ﬂows, the expansion around a corner after Prandtl and Meyer, and the reverse, the isentropic compression. The jump conditions for oblique shocks and the Prandtl-Meyer expansion are then employed to determine lift and wave drag of airfoil proﬁles at angle of attack in supersonic ﬂow. In the next chapter an introduction to the computation of supersonic ﬂows is given. Beginning with the fundamental equation of gasdynamics, the theory of characteristics is explained. Thereafter, compressible potential ﬂows are discussed. These considerations include plane and axially symmetric supersonic and subsonic ﬂows about slender bodies. Finally the similarity rules of gasdynamics are discussed. First the similarity rules for twodimensional ﬂows are derived with the aid of the linearized potential equation, together with several examples of their application. The chapter closes with the extension of the similarity rules to axially symmetric and transonic ﬂows. The various topics presented are again supplemented by exercises in gasdynamics in Chap. 5. Several examples are given to illustrate the application of the laws for one-dimensional steady ﬂows. The abrupt changes of the ﬂow quantities across normal and oblique shocks are delt with in the following two subsections. Prandtl-Meyer expansions and their interaction with oblique compression shocks is studied in detail for several ﬂow conﬁgurations, followed by exercises concerned with the determination of lift and wave drag in the frame of the small-perturbation theory. Finally the application of the theory of characteristics and the computation of potential compressible ﬂows together with the similarity rules are demonstrated.

4.2 Thermodynamic Relations
In ﬂows of compressible gases, in addition to the rate of change of the mechanical energy also the rate of change of the thermal energy has to be considered. It is therefore necessary to deﬁne the thermodynamic state of the gas with state variables, e. g. the pressure p, the density ρ, and the temperature T . Their interdependence is described by the thermal equation of state. If the law given by Boyle, Mariotte, and Gay-Lussac is used, i. e. p=ρRT , (4.1) the gas is called thermally perfect. For thermally non-perfect gases other relations must be used, as for example the Van der Waals law. The speciﬁc gas constant R depends on the molecular J weight of the gas. For air it is R = 287 kgK .

140

4. Gasdynamics

Another state variable is the internal energy. It is deﬁned by two thermodynamic quantities, 1 the temperature T and the speciﬁc volumen v = ρ . e = e(v,T ) This relation is known as the caloric equation of state. The total derivative is de = ∂e ∂v dv +
T

(4.2)

∂e ∂T

dT v .

(4.3)

The internal energy of thermally perfect gases depends on the temperature only. It then follows that de = ∂e ∂T dT v , cv =

∂e ∂T

. v (4.4)

∂e The quantity ∂T is called speciﬁc heat cv at constant volume. If cv is constant, the gas is v called calorically perfect, and the internal energy is given by

e = cv T + er

.

(4.5)

The quantity er is a reference value. The enthalpy h is another important state variable: h=e+pv (4.6)

As the internal energy, the enthalpy of thermally perfect gases depends on the temperature only. dh = cp dT The quantity cp is the speciﬁc heat at constant pressure cp = ∂h ∂T . p (4.7)

(4.8)

It follows from the relation for the speciﬁc heats cv and cp cp = cv + R for calorically perfect gases, that cp = is constant. Hence h = cp T + hr , (4.10) (4.9)

where hr is again a reference value. cp The ratio of the speciﬁc heats cv = γ is – according to the gas-kinetic theory – given by the number n of degrees of freedom γ= n+2 n . (4.11)

For monatomic gases (n = 3) γ = 1.667, and for diatomic gases (n = 5) γ = 1.4. At high cp temperatures additional degrees of freedom are excited, and the ratio cv decreases. For air at a temperature of 300 K γ = 1.4, and for 3000 K is γ = 1.292. The entropy s is introduced as another state variable with the second law of thermodynamics T ds = d e + p dv . (4.12)

4.3 One-Dimensional Steady Gas Flow For a thermally perfect gas it follows that ds = cv dρ dp − cp p ρ
⎡

141

.

(4.13)

This relation can be integrated for a calorically perfect gas. s = sr + cv ln ⎣ p pr ⎦ ρ γ ρr

⎤

(4.14)

The quantities pr , ρr are reference values. If the entropy is constant s = sr , there results the isentropic relation p pr = γ ργ ρr . (4.15)

In ﬂows in which the entropy is constant, the pressure varies only with the density ρ, with the relationship depending on the isentropic exponent γ.

4.3 One-Dimensional Steady Gas Flow
Consider the one-dimensional ﬂow of a compressible, inviscid gas without heat exchange. The gas is assumed to be thermally and calorically perfect. 4.3.1 Conservation Equations With the assumptions stated the conservation equations for mass, momentum, and energy are: d [ρ u] = 0 d [ρ u2 + p] = 0 u2 d ρu h+ = 0 . 2

(4.16)

The quantities in the square brackets represent the mass, momentum, and energy ﬂuxes. The system of equations is closed by introducing the thermal equation of state; density, pressure, enthalpy, and velocity can now be determined. p = ρ RT (4.17)

If only continuous changes of the state variables are admitted, it follows from the second law of thermodynamics, that the entropy remains constant. ds = 1 T dh − dp ρ =0 (4.18)

One of the conservation equations can then be replaced by the isentropic relation: p p0 = γ ρ0 ργ (4.19)

The quantities p0 and ρ0 are the values of pressure and density, which the gas attains in the state of rest after isentropic compression.

142

4. Gasdynamics

4.3.2 The Speed of Sound The Mach number M a is an important similarity parameter for the description of compressible ﬂows. It is given by the ratio of the local ﬂow velocity and the speed of sound. Distinguished are subsonic, supersonic, and transonic ﬂows. Since each of the ﬂows exhibits its own particularities, the Mach number has to be known for the description of ﬂows. In a compressible medium pressure disturbances travel with ﬁnite speed. The speed of propagation of small pressure disturbances is the speed of sound. The changes of the state variables caused by the propagation of sound occur at constant entropy. The compression of a gas by an explosion does not take place without losses, and the speed of propagation is larger than the speed of sound. Consider a small disturbance moving with the velocity u through a gas at rest. For an observer moving with the ﬂow, the ﬂow process is steady. The mass and momentum balance ρ u = (ρ + ∆ρ)(u + ∆u) −ρ u2 + (ρ + ∆ ρ)(u + ∆u)2 = p − (p + ∆p) (4.20) yield u2 = 1 ∆p 1 + ∆u ∆ρ u . (4.21)

If the changes of the ﬂow quantities are assumed to be inﬁnitesimally small, then the expression for the speed of sound is a= ∂p ∂ρ , s=const. (4.22)

which for incompressible media is inﬁnitely large as ∆ρ → 0. In the above derivation it was tacidly assumed, that the inﬂuence of friction forces can be neglected; (4.22) can therefore only be applied, if the disturbances are small enough so that the assumption is valid. Disturbances caused by sound are small, and friction forces can be left out. It then follows that the motion in a sound wave can be considered as isentropic, as the entropy production is proportional to the squares of the velocity and temperature gradients. For perfect gases, with the isentropic relation dp/p = γ dρ/ρ and the thermal equation of state p/ρ = R T , one obtains a= γ p = ρ γ RT . (4.23)

The speed of sound depends on the temperature of the gas only. Since it is ﬁnite, the propagation of sound is inﬂuenced by the speed of the sound source. Consider a punctiform sound source: If it is at rest, the sound waves propagate in concentric circles and spread out in the entire space. If the source is set in motion and moves with subsonic speed, the sound waves propagate in excentric circles. If the source moves with the speed of sound, the waves can propagate only

4.3 One-Dimensional Steady Gas Flow

143

in the space downstream from the source, and for supersonic speed, the sound can spread out only in the so-called Mach cone, the half angle of which is given by α = arcsin 1 . Ma (4.24)

An observer can therefore hear a sound source moving with supersonic speed only after it has passed him. 4.3.3 Integral of the Energy Equation It follows from the conservation equations for one-dimensional, steady, compressible ﬂow, that 2 the sum of the kinetic energy u and the static enthalpy h remains constant. The value of this 2 constant is given by the stagnation enthalpy h0 = h + u2 2 . (4.25)

For calorically perfect gases the enthalpy can be replaced by the product of static temperature and the speciﬁc heat at constant pressure cp T . cp T0 = cp T + u2 2 (4.26)

Introducing the thermal equation of state there results γ p u2 γ p0 + = γ − 1 ρ0 γ−1 ρ 2 and with the deﬁnition of the speed of sound a2 a2 u2 0 = + γ−1 γ−1 2 γ−1 γ

(4.27)

.

(4.28)

If the isentropic relation is combined with the expression for the speed of sound, a2 = a2 0 (4.27) takes on the following form γ p0 γ = γ − 1 ρ0 γ−1 p0 ρ0 p p0 γ−1 γ

p p0

,

(4.29)

+

u2 2

.

(4.30)

This equation is often called the Bernoulli equation for compressible ﬂow. Because of the isentropic change of the state the dependence of the static pressure on the kinetic energy is not linear as it is in incompressible ﬂow. If the last equation is divided by the square of the speed of sound, pressure, temperature, density, and speed of sound can be expressed in a simple manner by the Mach number and the stagnation quantities: p = p0 T = T0 ρ = ρ0 a = a0 γ−1 M a2 2 γ−1 M a2 1+ 2 γ−1 M a2 1+ 2 1+ 1+ γ−1 M a2 2 γ − γ−1

−1

1 − γ−1

−1 2

(4.31)

144

4. Gasdynamics

4.3.4 Sonic Conditions If the local ﬂow velocity is equal to the speed of sound, pressure, temperature, and density attain speciﬁc values, which solely depend on the stagnation conditions of the gas. The sonic condition, also refered to as the critical state, is designated by an asterisk: p∗ = p0 T∗ = T0 ρ∗ = ρ0 a∗ = a0 2 γ+1 2 γ+1 2 γ+1 2 γ+1 γ γ−1

1 γ−1

1 2

(4.32)

For air with γ = 1.4 the critical values are as follows: p∗ = 0.528; p0 T∗ = 0.833; T0 ρ∗ = 0.634; ρ0 a∗ = 0.913 a0 (4.33)

Instead of the local speed of sound a also the critical speed of sound can be used to deﬁne a Mach number, which is called the critical Mach number M a∗ = u . a∗ (4.34)

The relation between the local Mach number M a = u and the critical Mach number is given a by the energy equation. If therein a0 is replaced by a∗ , after division by u2 , there is obtained M a∗2 = γ+1 γ − 1 + M2a2 . (4.35)

For M a → ∞ the critical Mach number M a∗ approaches the limiting value lim M a∗ = γ+1 γ−1 . (4.36)

M a→∞

With these relations the ratios of the pressure, temperature, density, and speed of sound, refered to their stagnation values, can be expressed by the critial Mach number : p = p0 T = T0 ρ = ρ0 a = a0 γ−1 M a∗2 1− γ+1 γ−1 M a∗2 1− γ+1 γ−1 M a∗2 1− γ+1 1− γ−1 M a∗2 γ+1 γ γ−1

1 γ−1

1 2

(4.37)

With M a∗ = 1 these relations reduce to those given for the critial state.

4.3 One-Dimensional Steady Gas Flow 4.3.5 The Limiting Velocity

145

A ﬂow attains its maximum velocity, if the gas is expanded into vacuum. The resulting limiting velocity ug depends only on the stagnation conditions. To show this, the Bernoulli equation for compressible ﬂow is solved for the velocity u
⎡

u=

p 2 γ RT0 ⎣ 1− p0 γ−1

γ−1 γ

⎤ ⎦

.

(4.38)

This relation was ﬁrst derived by de Saint-Venant and Wantzel in 1839. If the pressure is lowered to zero, the limiting velocity is given by ug = lim u = p→0 2 γ R T0 γ−1

.

(4.39)

The corresponding Mach number M ag is inﬁnitely large, the critical Mach number has the ﬁnite value, already stated M a∗ = g γ+1 γ−1 . (4.40)

The limiting velocity can also be used to deﬁne the pressure, density, temperature, and speed of sound in terms of the local velocity and the stagnation conditions: p = p0 T = T0 ρ = ρ0 a = a0 1− 1− 1− 1− u2 u2 g u2 u2 g u2 u2 g u2 u2 g
1 γ−1 γ γ−1

1 2

(4.41)

The ratio of the limiting velocity and the speed of sound at stagnation conditions is ug = a0 2 γ−1 . (4.42)

The numerical values of the quantities discussed here are given in the appendix as a function of the Mach number M a. The data can be used to compute one-dimensional isentropic ﬂows, if the stagnation conditions are prescribed. 4.3.6 Stream Tube with Variable Cross-Section The inﬂuence of the compressibility on one-dimensional isentropic ﬂow in a stream tube with slowly varying cross-section A can be investigated with the conservation equations given previously. The continuity equation can be written in the form dρ du dA + + =0 . ρ u A (4.43)

The diﬀerential of the density can be eliminated with the aid of the momentum equation:

146

4. Gasdynamics −u du = dρ dp = a2 ρ ρ (4.44)

There results 1 dA du =− u (1 − M a2 ) A Three cases can be distinguished: 1. Subsonic flow (Ma < 1) The velocity increases with decreasing cross section. 2. Supersonic flow (Ma > 1) The velocity decreases with increasing cross section. 3. Transonic flow (Ma = 1) Sonic velocity can only be attained in the stream tube, if dA vanishes locally. According to this result supersonic ﬂow can only be generated, if the stream tube has a convergent-divergent distribution of the cross section (Laval nozzle). The cross section with minimum area, where the local velocity is equal to the speed of sound, is called critical cross section or throat. The local velocity can be computed with the Bernoulli equation for compressible ﬂow as a function of the pressure ratio:
⎡

.

(4.45)

u=

2 γ p0 ⎣ p 1− γ − 1 ρ0 p0

γ−1 γ

⎤ ⎦

.

(4.46)

The local Mach number is obtained by dividing by the local speed of sound
⎡

Ma =

2 ⎣ p0 γ−1 p

γ−1 γ

⎤

− 1⎦ .

(4.47)

The rate of mass ﬂow through every cross section is constant. For the following derivation it is advantageous to choose the critical rate of mass ﬂow as reference value. ρ u A = ρ∗ a∗ A∗ . (4.48)

With this relation, the interdependence between static pressure, Mach number, and cross∗ sectional area can be formulated. If (4.48) is solved for A and if the stagnation quantities are A introduced as reference values, there results A∗ ρ ρ0 u a0 = A ρ0 ρ∗ a0 a∗ The critical cross-sectional area ratio
A∗ A

. p p0

(4.49)

can be expressed by the pressure ratio
1 γ−1

( pp0 ) γ 1 − ( pp0 ) γ A∗ = γ+1 A γ−1 2 ( γ+1 ) γ−1 2 or by the local Mach number A∗ = A
∗

(4.50)

Ma
2 γ+1

1+

γ−1 2

M a2

γ+1 2 (γ−1)

.

(4.51)

The last equation shows, that A tends to zero for small Mach numbers (M a → 0) and also A for very large Mach numbers (M a → ∞). The critical area ratio is plotted in the following

4.4 Normal Compression Shock

147

diagramm for γ = 1.4 as a function of the Mach number and of the pressure ratio. The maximum ∗ A∗ = 1 is attained for a pressure ratio of p0 = 0.528. The two branches of the curve correspond p A to subsonic and supersonic ﬂow. If the cross section is chosen to ﬁrst converge and then diverge, and if the ﬂow entering the stream tube is subsonic, it is then accelerated and becomes sonic, and ﬁnally in the divergent part supersonic. According to the assumptions introduced, (4.50) and (4.51) are valid only for isentropic ﬂow, since the variation of the density is assumed to be isentropic. Mach number and pressure distribution can be determined for an area distribution prescribed in the streamwise direction. The ﬂow in the nozzle solely depends on the pressure in the exit cross section. If the pressure drop in the exit cross section is small pe1 < pe < p0 , the ﬂow in the nozzle remains subsonic. If the pressure in the exit cross section is reduced to pe1 , the speed in the critical cross section will be equal to the speed of sound. An isentropic supersonic ﬂow in the divergent part of the nozzle can exist only if the pressure in the exit cross section is at least reduced to pe2 . If the pressure in the exit cross section is between pe1 and pe2 , the ﬂow cannot adjust isentropically to the conditions existing in the exit cross section. Then a discontinuous increase of the density, pressure, and temperature occurs in the supersonic part of the nozzle, known as normal compression shock. The ﬂow is decelerated to subsonic speed. This discontinuous process causes losses, which result in an increase in entropy.

4.4 Normal Compression Shock
Compression shocks can be viewed as a discontinuous change of all ﬂow quantities. The corresponding changes can be obtained with the integrals of the conservation equations. 4.4.1 The Jump Conditions Consider a plane supersonic ﬂow with a normal compression shock in a chanel with constant sross-sectional area; the integrals of the conservation equations given earlier can be written in the following form:

148

4. Gasdynamics ρ1 u 1 = ρ 2 u 2 p1 + ρ1 u2 = p2 + ρ2 u2 1 2 u2 u2 2 1 = h2 + h1 + 2 2

(4.52)

ρ In order to determine the density ratio ρ2 the velocity 1 diﬀerence is ﬁrst obtained from the continuity and the momentum equation.

u1 − u2 =

a2 a2 2 − 1 γ u2 γ u1

(4.53)

The local speed of sound is expressed through the critical speed of sound a∗ and the ﬂow velocity u. The intergal of the energy equation yields a2 = γ + 1 ∗2 γ − 1 2 a − u 2 2 . (4.54)

Combining the last equations gives the Prandtl relation u1 u2 = a∗2 or M a∗ = 1 1 M a∗ 2 . (4.55)

The Prandtl relation shows that a normal shock decelerates a supersonic ﬂow always to a subsonic ﬂow. The density ratio follows from the continuity equation ρ2 u1 (γ + 1) M a2 1 = = M a∗ 2 = 1 ρ1 u2 2 + (γ − 1) M a2 1 . (4.56)

The pressure ratio p2 is obtained by ﬁrst solving the momentum equation for the pressure p1 diﬀerence p2 − p1 and expressing the velocity ratio u1 by M a1 . u2 p2 2γ (M a2 − 1) =1+ 1 p1 γ+1 (4.57)

p For high Mach numbers M a1 the pressure ratio p2 tends to inﬁnity, while the density ratio 1 γ+1 approaches the ﬁnite value γ−1 . The increase of entropy across the compression shock follows from (4.14). It is

s2 − s1 p2 = ln ⎣ R p1

⎡

1 γ−1

ρ2 ρ1

−γ γ−1

⎤ ⎦

.

(4.58)

If the Mach number is introduced in the expression for the pressure and density ratio, the entropy diﬀerence across the compression shock becomes
⎨ s2 − s1 2γ (M a2 − 1) = ln 1 + 1 ⎩ γ+1 R ⎧
1 γ−1

(γ + 1) M a2 1 2 + (γ − 1) M a2 1

γ − γ−1 ⎬

⎫ ⎭

.

(4.59)

The condition s2 − s1 ≥ 0 shows, that compression shocks can exist only in supersonic ﬂows. The nozzle ﬂow discussed earlier can now be explained for the case of the non-adjusted pressure in the exit cross section with the aid of the normal-shock relations. If the pressure pe is not equal to the value, required for an isentropic supersonic ﬂow in the divergent part of the nozzle, a normal compression shock is generated at a certain location in the nozzle, with the pressure jump just large enough, such that the pressure in the exit cross section can be matched. The ﬂow velocity between the normal shock and the exit cross section is – according to the shock

4.4 Normal Compression Shock

149

relations – everywhere smaller than the local speed of sound. The following diagram shows the pressure and Mach number distribution in the streamwise direction in the nozzle. The photographic picture taken in a supersonic wind tunnel shows the compression shock as thick dark line in the ﬂow, almost normal to the line of symmetry of the two-dimensional nozzle. Also visible are several Mach lines, originated at the wall of the nozzle. Small disturbances, caused by grooves in the wall propagate along these lines into the ﬂow.

Supersonic ﬂow with compression shock in a Laval nozzle. Mach lines can be noted upstream of the shock, which are originated on the wall of the nozzle.

4.4.2 Increase of Entropy Across the Normal Compression Shock Assuming that the compression shock represents a discontinuity is only an approximation. In reality the shock has a thickness δ of the order of magnitude of several free mean paths. If the gas ﬂowing through the shock can be assumed to be a continuum, the Navier-Stokes equations can be employed for the description of the ﬂow between the upstream and downstream edge of the compression shock. The ﬂow quantities do not change discontinuously in form of a jump, but in a continuous transition from the free-stream conditions to the ﬂow conditions downstream from the shock. The increase of the entropy can now be explained as an action of the frictional forces and the heat conduction within the shock region of ﬁnite thickness. The conservation equations for one-dimensional, steady, viscous, and heat conducting ﬂow are d [ρ u] = 0 4 du d ρ u2 + p − µ = 0 3 dx 4 du u2 dT − µu = 0 . h+ −λ 2 dx 3 dx

d ρu

(4.60)

The terms in the square brackets represent the mass, momentum, and energy ﬂuxes. They remain constant, and the introduction of frictional forces and heat conduction does not aﬀect this result. If the ﬂow variables are constant upstream of and downstream from the shock, the terms describing the shear stresses and the heat ﬂux vanish outside of the shock. The increase of entropy can be explained with the aid of the ﬁrst law of thermodynamics in the form

150

4. Gasdynamics ds = 1 T dh − 1 dp . ρ (4.61)

If the momentum equations is solved for dp and the energy equation for dh and if the resulting expressions are inserted in the second law of thermodynamics, there results
⎡

ρu

ds 1 ⎣d = dx T dx

λ

dT dx
2

+

4 µ 3

du dx

2

⎤ ⎦

(4.62)

As the mass ﬂux is constant, the last equation can be written in the following form: ρ1 u1 (s2 − s1 ) =
2 1

λ T2

dT dx

dx +

4 3

2 1

µ T

du dx

2

dx

(4.63)

The integrals in the above equation represent the contributions of the heat conduction and the dissipation of mechanical energy. They can be determined, if the local gradients du and dT are dx dx known. They, in turn, can be obtained by repeated integration of (4.60), with the boundary cp conditions given by (4.52). If it is assumed, that the Prandtl number is P r = µ λ = 0.75, then the total enthalpy h0 remains constant. The continuity and the momentum equation can then be combined to yield γ+1 ρ1 u1 (u − u1 ) 2γ 1− a∗2 4 du =0 − µ u1 u 3 dx (4.64)

The solution of this equation would provide the velocity distribution u(x) within the shock. The temperature distribution T (x) can be obtained from the energy equation in the form of (4.60) and also the integrands in (4.63). A closed-form solution is, however, not possible. The thickness of the shock can be estimated from (4.64). For transonic speeds it is of the order of magnitude η u1 δ∼ . (4.65) ρ a (u1 − u2 ) 4.4.3 Normal Shock in Transonic Flow The shock relations can considerably be simpliﬁed for low supersonic speeds. If in the relation
⎨ s2 − s1 2γ = ln 1 + M a2 − 1 1 ⎩ R γ+1 ⎧
1 γ−1 γ − γ−1 ⎬

(γ + 1) M a2 1 (γ − 1) M a2 + 2 1

⎫ ⎭

(4.66)

the square of the Mach number M a2 is replaced by 1 + m, with m assumed to be small 1 compared to unity, the expression of the right-hand side of (4.66) can be expanded in a power series. Collecting the resulting expressions in increasing powers of m, it is seen that the terms of ﬁrst and second order vanish, and one obtains s2 − s1 2γ (M a2 − 1) 1 = + ... R 3 (γ + 1)2
3

.

(4.67)

This results shows, that the entropy increase across a normal compression shock is small at low 3 supersonic Mach numbers. It increases with (M a2 − 1) . All changes of the state are therefore 1 almost isentropic. The relation for the pressure change across the normal compression shock p2 − p1 2γ (M a2 − 1) = (4.68) 1 p1 γ+1 reduces to p2 − p1 = p1 12 γ 2 γ+1
1 3

s2 − s1 R

1 3

.

(4.69)

The change of pressure can thereby directly be related to the increase of entropy.

4.5 Oblique Compression Shock

151

4.5 Oblique Compression Shock
In addition to normal compression shocks, also oblique shocks can occur in supersonic ﬂow. They are inclined by a certain angle, the shock angle σ, with respect to the direction of the free-stream ﬂow. 4.5.1 Jump Conditions and Turning of the Flow As for the normal shock, the jump conditions for the oblique shock can be obtained from the Euler equations for two-dimensional ﬂow. It is advantageous, to use Cartesian coordinates, oriented normally and tangentially to the shock. For uniform free-stream conditions and constant inclination of the shock, the jump conditions follow from the continuity equation, the two momentum equations, and the energy equation:

ρ1 u 1 = ρ 2 u 2 ρ1 u 1 v 1 = ρ2 u 2 v 2 ρ1 u 2 + p 1 = ρ 2 u 2 + p 2 2 1 ρ1 u1 h1 + 1 2 u2 + v1 1 2 = ρ2 u2 h2 + 1 2 u2 + v 2 2 2 (4.70)

The jump conditions for the oblique compression shock correspond to those for the normal shock. Mass, momentum, and energy ﬂuxes normal to the shock remain constant during the passage of the ﬂow through the shock. The tangential velocity component remains unchanged. The following sketch shows a part of an oblique shock, indicated by the thick black line. Shown is also the velocity upstream of the shock V1 , and downstream V2 , both decomposed into components normal and tangential to the shock u1 and v1 , and u2 and v2 . As u2 is smaller than u1 , and as v2 is equal to v1 , the direction of the oncoming velocity is changed, as the velocity triangles indicate. Thus an oblique shock always causes a turning of the ﬂow towards the shock. The ﬂow is turned by the angle β. The jump conditions can again be used to derive the relations for the density, pressure, temperature, and velocity changes across the shock. The density ratio can be expressed with the shock in1 tensity u1 = M a1 sin σ as a ρ2 (γ + 1) M a2 sin2 σ 1 = ρ1 (γ − 1) M a2 sin2 σ + 2 1 the increase of pressure as p2 − p1 2γ (M a2 sin2 σ − 1) , = 1 p1 (γ + 1) the temperature ratio as a2 (γ − 1) T2 = 2 =1+2 T1 a2 (γ + 1)2 1 the increase of entropy as γ M a2 sin2 σ + 1 1 M a2 sin2 σ 1 M a2 sin2 σ − 1 1 , (4.73) (4.72) , (4.71)

152

4. Gasdynamics
⎧
1 γ−1

⎨ s2 − s1 γ = ln 1 + 2 M a2 sin2 σ − 1 1 ⎩ R γ+1

(γ + 1) M a2 sin2 σ 1 (γ − 1) M a2 sin2 σ + 2 1

γ − γ−1 ⎬

⎫ ⎭

, (4.74)

and the Mach number downstream from the shock as M a2 sin2 (σ − β) = 2 2 + (γ − 1) M a2 sin2 σ 1 2 γ M a2 sin2 σ − (γ − 1) 1 u2 = tan(σ − β) . v . (4.75)

From the velocity diagram shown above, the normal components of the velocity u1 and u2 are u1 = tan σ v and (4.76)

The ratio of these two expressions yields, together with the jump condition for the mass ﬂux tan(σ − β) (γ − 1) M a2 sin σ 2 + 2 1 = tan σ (γ + 1) M a2 sin σ 2 1 The turning angle β can now be expressed as follows β = arctan 2 cot σ (M a2 sin2 σ − 1) 1 M a2 (γ + cos(2 σ)) + 2 1 . (4.78) . (4.77)

The equation for the increase of entropy, (4.74), with ∆s ≥ 0, indicates that the shock intensity has to satisfy the condition M a1 sin σ ≥ 1 . The shock angle σ is bounded from below by the Mach angle α α = arcsin Then σ can vary within the limits α≤σ≤ π 2 (4.81) 1 . M a1 (4.80) (4.79)

The turning angle β vanishes for both limits σ = π and σ = α; it attains an extremum value 2 in between. For M a1 → ∞ the expression for β reduces to tan β → sin 2σ γ + cos(2σ) , (4.82)

so that with γ = 1.4 the maximum turning angle results to βmax ≈ 45◦ and the corresponding shock angle to σ = 67.5◦ . Similarly, it can be shown, that a maximum turning angle exists for every ﬁnite Mach number. Another interesting relation results from (4.75), if M a2 is set equal to unity. Then for every free-stream Mach number M a1 and ﬂow turning angle β there exists a shock angle σ, which yields sonic conditions immediately downstream from the shock. It can also be expected, that the shock angle σ of ﬂows with M a2 = 1 is close to σmax as then the shock tends to be close to normal, with subsonic ﬂow conditions downstream. In the following diagram the dependence of the shock angle σ is shown for several Mach numbers as a function of β. According to the diagram, the jump conditions yield two solutions, which are characterized by diﬀerent shock angles and shock intensities. The solutions are known as the weak and strong solution. The strong solution describes a ﬂow which is subsonic downstream from the shock, the weak solution a ﬂow which is supersonic with the exception of β ∼ βmax . This region is bounded by the line M a2 = 1.

4.5 Oblique Compression Shock

153

4.5.2 Weak and Strong Solution The supersonic ﬂow over a wedge is described by the weak solution for β < βmax . The turning angle β is equal to the wedge angle. All ﬂow quantities downstream from the oblique shock can be determined with the relations obtained from the jump conditions. The weak solution is also valid for a double wedge with nose angle β.

The ﬂow quantities downstream from the shock can also be determined for a double wedge at angle of attack. With vanishing turning or nose angle of the wedge the shock intensity M a1 sinσ tends to unity, and the shock deteriorates to a sonic line.

If the nose angle of the wedge β is β > βmax , the shock is detached from the wedge and deforms into a bow shock in front of the wedge. A closed-form solution of the conservation equations does not exist for the ﬂow ﬁeld between the bow shock and the wedge. In a certain point on the shock, the tangent of the shock is normal to the direction of the freestream velocity.

154

4. Gasdynamics

The shock angle σ then attains all possible values, from σ = π to the value corresponding to the maximum turning 2 angle, further to the angle, corresponding to the speed of sound immediately downstream from the shock, until it decreases to the Mach angle of the free stream α at large distances from the wedge. The strong solution is valid for that part of the shock between the points S and S .

4.5.3 Heart-Curve Diagram and Hodograph Plane Due to the non-linearity of the jump conditions the pressure ratio cannot explicitly be formulated in terms of the turning angle. However, the pressure ratio p2 and the turning angle β can p1 be computed for all values of the shock intensity, so that a plot of p2 = f (β) can be provided. p1 The shape of the curve is similar to that of the contour of the heart, the lower part of which representing the weak solution and the upper part the strong solution. With the aid of the heart-curve diagram the pressure ratio across the oblique shock can easily be determined for every turning angle. A supplement for the description of the ﬂow in the physical plane is the so-called hodograph plane, with the velocity components as coordinates. The hodograph plane enables a simple representation of the solution of the jump conditions for the normal and the oblique compression shock.

The velocity vectors downstream from the oblique shock are ﬁrst plotted for all values of the turning angle at a constant free-stream Mach number. The resulting curve is called the shock polar. Most of the time it is plotted in the form of the velocity components, nondimensionalized with the critical speed of sound a∗ . The hodograph plane is bounded by a circle with radius 1 R = [ γ+1 ] 2 . The sonic line is represented by a circle with radius unity. The weak solution is γ−1 given by the point B, the strong solution by the point B ; the velocity of the ﬂow containing a normal shock is given by the point A . There exists a shock polar for every free-stream Mach number M a1 . v In order to determine the dimensionless downstream velocity a2 the turning angle β is plot∗ ted, intersecting the shock polar in the point B. The tangential and the two normal velocity components are obtained by constructing the normal from the origin of the coordinates to the extension of the line A − B.

4.5 Oblique Compression Shock

155

4.5.4 Weak Compression Shocks Compression shocks are termed weak, if the shock intensity M a1 sin σ approaches unity. The jump conditions can be simpliﬁed with this assumption. First the relation for the turning angle β is solved for M a2 sin2 σ: 1 M a2 sin2 σ = 1 2 tan σ (γ + 1) tan(σ − β) − (γ − 1) tan σ (4.83)

For weak shocks the shock angle σ does not diﬀer much from the Mach angle α, and the turning angle β tends to zero. One obtains for small angles β M a2 sin2 σ ≈ 1 + 1 γ+1 2 M a2 1 M a2 − 1 1 β . (4.84)

The dimensionless pressure diﬀerence is approximately p2 − p1 = p1 γ M a2 1 M a2 − 1 1
⎛

β

.

(4.85)

Also the expression for the increase of the entropy can be approximated: s2 − s1 γ + 1 ⎝ γ M a2 ⎠ 3 1 ≈ β R 12 γ 2 M a2 − 1 1 The square of the velocity ratio v2 v1
2 v2 v1

⎞3

(4.86)

is = cos2 σ tan2 (σ − β) + 1 = tan2 σ + 1 cos2 (σ − β) (4.87)

=

u2 v u1 v

2 2

+1 +1

.

With the approximation σ → α, β → 0 there results v2 v1
2

⎡

⎤2

≈ (1 − 2 tan αβ) ≈ ⎣1 −

β M a2 − 1 1

⎦

(4.88)

156

4. Gasdynamics

and ﬁnally v2 − v1 ∆v β = =− v1 v1 M a2 − 1 1 . (4.89)

The last relation shows, that the velocity is decreased, if the turning angle β is increased. The decrease of the velocity simultaneously causes an increase of the pressure and of the density. The compression can be enforced either by one or several weak oblique compression shocks.

If the compression is designed in such a way, that the turning of the ﬂow is the same in both cases, i. e. β = n ∆ β, then the pressure after the n-th turning is pn = p1 + O(n ∆ β) . The increase of entropy sn = s1 + O(n ∆ β 3 ) = O β3 n2 (4.91) (4.90)

can substantially be reduced, if the number of the turning elements is increased and approaches inﬁnity in the limiting case. Very weak shocks then reduce asymptotically to so-called Mach waves. The width of each of the compression elements is inﬁnitesimally small. The contour is continuously curved. Along the Mach lines the Mach number and the turning angle are constant. The ﬂow cannot be inﬂuenced against the main ﬂow direction, and disturbances cannot propagate upstream. A certain distance away from the contour, the Mach waves coalesce to an oblique compression shock.

4.6 The Prandtl-Meyer Flow
As just shown, the entropy remains constant if an inviscid supersonic ﬂow is continuously turned by inﬁnitesimally weak oblique compression shocks. An increase of the total turning angle causes a decrease of the velocity and an increase of the static pressure. Vice versa, the turning angle can be decreased, the velocity increased, and the pressure decreased. Then the gas is said to expand.

4.6 The Prandtl-Meyer Flow

157

In a supersonic ﬂow a continuously curved contour therefore generates an isentropic compression, if the turning angle is increased, and an expansion, if the turning angle is decreased. If a discontinuous expansion would be assumed, as indicated in the sketch, the normal velocity component would have to increase and pressure and temperature would have to decrease. The shock relations would then require a decreasee of entropy, which would contradict the second law of thermodynamics; it is for this reason, that an expansion can occur only in isentropic ﬂow with continuous turning.

4.6.1 Isentropic Change of Velocity The equation for the change of the velocity across a weak compression shock reduces to a diﬀerential equation of the following form if isentropic ﬂow is assumed √ M a2 − 1 dv dv = cot α = −dβ v v . (4.92)

The integral of this equation was ﬁrst published by Prandtl and Meyer in 1908. In order to facilitate the integration, the square of the velocity is expressed through the Mach angle with the aid of the energy equation v2 = a∗2 sin α + b2 cos2 α
2

.

(4.93)

The quantity b2 is an abbreviation of γ−1 . By diﬀerentiation of this expression the diﬀerential γ+1 equation for the velocity change can be cast into the following form: b2 − 1 dα = −dβ b2 + tan2 α (4.94)

The integration of the left-hand side of this equation yields an angle ν(M a), which became known as the Prandtl-Meyer angle. ν(M a) = α − 1 b π − arctan(b cot α) + C 2 (4.95)

The constant of integration C is so chosen, such that the Prandtl-Meyer angle vanishes for M a = 1. ν(M a) = γ+1 arctan γ−1 √ γ−1 (M a2 − 1) − arctan M a2 − 1 γ+1 (4.96)

For M a ≥ 1 there exists a uniquely deﬁned value of ν, which with increasing Mach number monotonically tends to lim ν = νmax = π 2 γ+1 −1 . γ−1 (4.97)

M a→∞

158

4. Gasdynamics

For example, for γ = 1.4 νmax = 130.5◦ . The integral of the right-hand side of the diﬀerential equation for the velocity change completes the relation between the Prandtl-Meyer angle ν(M a) and the turning angle: ν − ν1 = β1 − β (4.98) With these relations isentropic expansion and compression of the Prandtl-Meyer ﬂow can easily be determined: β < β1 , ν > ν1 ν = ν1 + | β − β1 | Compression: β > β1 , ν < ν1 ν = ν1 − | β − β1 | Expansion:

(4.99)

The Prandtl-Meyer angle ν(M a) is tabulated in the appendix as a function of the Mach number. The solution can easily be plotted in the hodograph plane. The Prandtl-Meyer integral represents an epicycloid, extending from the sonic circle (M a = 1) to the circle of the limiting velocity (M a → ∞). Expansion and compression, which are given by a certain turning of the ﬂow ∆β, can simply be entered in the hodogragh plane and the corresponding velocity change can be read oﬀ.

4.6.2 Corner Flow A uniform supersonic ﬂow with a Mach number M a1 ≥ 1 over a plane wall, by kinking of the wall at an arbitrary point P into a convex corner, can be turned by the amount ∆β. After the turn the gas moves with the Mach number M a2 parallel to the bent part of the wall. The ﬂow can be determined with the aid of the Prandtl-Meyer integral.

compression shock

4.6 The Prandtl-Meyer Flow

159

In the region bounded by the Mach lines P A and P B the ﬂow quantities are constant along each Mach line, but change from line to line. It is in this region that the ﬂow expands from M a1 to M a2 . The solution is singular in the point P , where all ﬂow conditions corresponding to M a1 ≤ M a ≤ M a2 can be found. The gas can also be thought of to ﬂow in the opposite direction. In this form the Prandtl-Meyer ﬂow provides the principle for an isentropic inlet diﬀuser, used on supersonic airplanes. The previous sketch shows the ﬂow reversal (compression by solidiﬁcation of a streamline). The ﬂow ﬁeld on the right demonstrates how a supersonic inlet diﬀuser can be constructed with the Prandtl-Meyer corner ﬂow. Expansion and compression are illustrated with two numerical examples: Assume that a plane supersonic ﬂow with a Mach number M a1 = 3 is given. The corresponding Prandtl-Meyer angle ν is ν = 49.76◦ . The oncoming ﬂow is assumed not to have been turned, so that β1 = 0◦ . If the ﬂow is to be decelerated to sonic speed, it has to be turned by βM a=1 = 49.76◦ . In the second example it is assumed, that a plane supersonic ﬂow with a free-stream Mach number M a1 = 2 (ν1 = 26.38◦ ) is turned by 10◦ . For an expansion there results a Prandtl-Meyer angle of ν2 = 36.38◦ , corresponding to a Mach number M a2 = 2.386, and for a compression, the Prandtl-Meyer angle is ν2 = 16.38◦ , and the Mach number is M a2 = 1.652. 4.6.3 Interactions Between Shock Waves and Expansions Prandtl-Meyer ﬂows and oblique compression shocks can interact with each other. To be distinguished are shock-shock and shock-expansion interactions. In the following sketch the reﬂexion of an oblique compression shock on a rigid wall is shown. If a shock impinges on a wall, it is reﬂected, as the wall aﬀects the oncoming ﬂow as a wedge. The reﬂected shock has the shock angle σ2 , which diﬀers from σ1 , although β1 and β2 have the same value. The gas can only ﬂow oﬀ parallel to the walll CD, if the wall AB is parallel to CD. The pressure ratio p3 is given by the shock intensities M a1 sin σ1 p1 and M a2 sin σ2 . If the ﬂow is mirrored at the wall CD and if the wall is replaced by a streamline, the intersection of two oblique compression shocks of equal intensity results. The shocks are bent when they intersect each other. For the ﬂow shown the gas ﬂows oﬀ parallel to the oncoming ﬂow. The intersection of two oblique shocks of diﬀerent intensity is also possible. If both shocks are assumed to be weak, the ﬂow pattern is diﬀerent from the one just discussed. The streamline through the point of intersection divides the ﬂow into two regions of diﬀerent stagnation pressures, as the shock intensities are no longer the same on both sides of the dividing streamline. As a consequence, the two regions have diﬀerent entropy levels and also the other ﬂow variables attain diﬀerent values, with the exception of the static pressure. Hence the ﬂow cannot maintain the direction of the free stream. It is found that the streamline through the point of intersection is inclined to the direction of the oncoming ﬂow by the angle δ (Page 160, a). The angle δ results from the condition, that the static pressure on both sides of the streamline through the point of intersection has to have the same value. p3 p p3 p2 p p = 3 = = 3 2 p1 p1 p2 p1 p2 p1 (4.100)

Because of the diﬀerent shock intensities the increase of entropy s3 − s1 is not equal to the increase of entropy s3 − s1 . The streamline through the point of intersection represents a line of discontinuity of the entropy, the density, the velocity, and the temperature.

160

4. Gasdynamics

a)

line of dicontinuity

weak expansion

b) If two oblique shocks are generated by two wedges, one downstream from the other (see sketch b), they coalesce and form a single shock, which approximately produces the same rise of the static pressure as the two other oblique shocks. A mutual pass through of the shocks is not possible, and a weak expansion or compression can arise downstream from the point of intersection to balance the static pressure. If a Prandtl-Meyer ﬂow interacts with an oblique compression shock, the shock is being curved. The Mach waves are reﬂected at the shock and interact with each other. If an oblique compression shock impinges on a wall at an shock angle σ, for which at M a2 β > βmax , the reﬂected shock can no longer be attached to the wall. This type of reﬂection is called Mach reﬂection. An almost normal compression shock is originated in the immediate vicinity of the wall, which joins the impinging and the reﬂected shock in the triple point.

triple point

Again a line of discontinuity of entropy is formed downstream from the triple point. The static pressure has the same value on both sides of the line, but all other ﬂow quantities attain diﬀerent values.

4.7 Lift and Wave Drag in Supersonic Flow
With the jump conditions for oblique compression shocks and the Prandtl-Meyer ﬂow pressure distributions can be determined for airfoil sections moving in supersonic ﬂow.

4.7 Lift and Wave Drag in Supersonic Flow 4.7.1 The Wave Drag

161

Airfoils in inviscid supersonic ﬂow generate drag, which is caused by the compression shocks. For example, the ﬂow over the front part of a double-wedge proﬁle is compressed by the oblique shock, attached to the leading edge, and the pressure is p2 . The gas expands at the shoulder of the proﬁle in a Prandtl-Meyer corner ﬂow, and the pressure on the rear part of the proﬁle is lowered to p3 . Finally, the pressure is increaed again at the trailing edge, when the ﬂow passes through the recompression shock, downstream of which the pressure is p4 . Caused by the overpressure on the front part and the lower pressure on the rear part of the proﬁle, a drag per unit length results, which is FD = (p2 − p3 )d . (4.101)

The quantity d is the maximum thickness of the proﬁle. The drag is called wave drag of supersonic ﬂow. Pressure drag caused by ﬂow separation and frictional drag have to be determined separately. 4.7.2 Lift of a Flat Plate at Angle of Attack A ﬂat plate, inclined in an inviscid supersonic ﬂow at the angle of attack β, generates lift in addition to the wave drag. Both, lift and wave drag can be determined with the jump conditions of oblique compression shocks and the Prandtl-Meyer ﬂow: FL = (p2 − p2 ) t cos β FD = (p2 − p2 ) t sin β

(4.102)

The quantity t is the chord of the proﬁle. Since wing sections in general are thin and since their angles of attack are small, approximate solutions can be constructed for the determination of lift and wave drag. Because of the diﬀerent shock intensities of the oblique shock at the leading edge and the recompression shock at the trailing edge, a surface of discontinuity of the entropy is generated at the trailing edge of the ﬂat plate.

4.7.3 Thin Proﬁles at Angle of Attack With the approximation for the pressure jump for weak oblique compression shocks ∆p ≈ p1 γM a2 1 M a2 − 1 1 ∆β (4.103)

the pressure coeﬃcient can be determined for small angles of attack as a function of the freestream Mach number. cp = p − p1 2 ∆p ≈ ρ1 2 = v γ M a 2 p1 1 2 1 2β M a2 − 1 1 (4.104)

162

4. Gasdynamics

Lift and wave drag of the ﬂat plate result from the pressure diﬀerence between lower and upper side of the plate. cpl − cpu = The lift coeﬃcient is cL = (cpl − cpu ) cos β = and the drag coeﬃcient cD = (cpl − cpu ) sin β = The ratio 1 cD = c2 4 L M a2 − 1 1 (4.108) 4 β2 M a2 − 1 1 . (4.107) 4β M a2 − 1 1 (4.106) 4β M a2 − 1 1 (4.105)

is independent of the angle of attack β. For the double-wedge proﬁle with nose angle 2 and vanishing angle of attack, the approximation yields for the front and rear part of the proﬁle cp = ± The drag coeﬃcient is cD = 4 M a2 − 1 1 d l
2

2 M a2 − 1 1

.

(4.109)

.

(4.110)

With the approximation for thin proﬁles, lift and wave drag can also be determined for non-zero angles of attack, variable thickness distribution, and camber.

angle of attack

camber

thickness distribution

The local angle of attack consists out of three parts: The angle of attack of the chord β0 , the angle of attack of the mean camber line ∆βc (x) with respect to β0 , and the angle of attack of the thickness distribution ∆βt (x), with respect to the mean camber line β = β0 + ∆βc + ∆βt The lift of the proﬁle is FL = ρ1 2 v 2 1 t 0

.

(4.111)

(cpu − cpo ) dx .

(4.112)

4.8 Theory of Characteristics The other integrals vanish with ∆βc (x) ≈ coeﬃcient is cL = The wave drag of the proﬁle is
2 ρ1 v 1 FD = 2 dyc dx

163

and ∆βt (x) ≈ .

dyt , dx

and the expression for the lift

4 β0 M a2 − 1 1

(4.113)

⎡ t 0

dyu ⎣ dx

2

dyl + dx

2

⎤ ⎦ dx

2 M a2 − 1 1

,

(4.114)

and the drag coeﬃcient can be written in the following form: cD = 4 M a2 − 1 1
2 2 2 (β0 + ∆βc + ∆βt )

(4.115)

2 2 The quantities ∆βc and ∆βt are the squares of the angular variations, averaged over the chord of the proﬁle. In general the wave drag consists out of three parts, which result from the angle of attack, the camber, and the thickness distribution.

4.8 Theory of Characteristics
If two Prandtl-Meyer ﬂows interact with each other, the ﬂow cannot be described by superposition of the Prandtl-Meyer solutions, since they describe only simple waves, but not their intersections. A solution for non-simple regions can be constructed with the conservation equations for two-dimensional ﬂows. The following sketch shows the simple and non-simple regions of two Prandtl-Meyer ﬂows superposed on each other. Close to the two corners the ﬂows do not interact with each other, but downstream from the ﬁrst intersection of the Mach lines the ﬂow is inﬂuenced by both corners. Non-simple regions also exist on the downstream side of curved compression shocks. Then the changes of the entropy have to be included in the determination of the pressure, density, temperature, and velocity. The interdependence between the distribution of entropy and the velocity ﬁeld is described in the Crocco vorticiy theorem.

4.8.1 The Crocco Vorticity Theorem With the second law of thermodynamics T ds = dh − dp ρ (4.116)

and the energy equation the total derivative of the entropy can be expressed through the total derivatives of velocity and pressure. If it is assumed, that the stagnation temperature is constant throughout the entire ﬂow ﬁeld, it follows that T ds = − u du + v dv + dp ρ . (4.117)

164

4. Gasdynamics

For a two-diemnsional steady ﬂow the total derivatives of the velocity components are du = ∂u ∂u dy dx + ∂y ∂x and dv = ∂v ∂v dx + dy ∂x ∂y . (4.118)

Introduction of the momentum equations leads to the following expression for T ds T ds = − dy v − u dx ∂v ∂u − ∂x ∂y udx . (4.119)

dy The ratio dx can be viewed as the slope of arbitrary unknown sets of curves. If the slope is chosen to be identical with the slope of the streamlines

dy v = dx u

,

(4.120)

it can be concluded, that the entropy remains constant along streamlines. ds = ∂s ∂s dx + dy = 0 ∂x ∂y u ∂s ∂s +v =0 ∂x ∂y (4.121)

Streamlines are therefore called the characteristic lines of the entropy. The entropy remains ∂v constant in the entire ﬂow ﬁeld, if the second bracketed term ∂x − ∂u , representing one of ∂y the components of the vorticity vector, vanishes. This is the contents of the Crocco vorticity theorem, which also holds for three-dimensional ﬂows: If in the second law of thermodynamics in the form T ∇ s = ∇h − ∇p ρ (4.122)

the gradient of the static enthalpy is eliminated with the energy equation ∇ h = −∇ v2 + ∇ h0 2 , (4.123)

and if the momentum equation is introduced in vector form ∂v ∇p + (v · ∇) v = − , ∂t ρ one obtains the Crocco vorticty theorem for three-dimensional ﬂow T ∇s + v x (∇ x v) = ∂v + ∇h0 ∂t . (4.125) (4.124)

According to this theorem the entropy remains constant along streamlines, even in threedimensional ﬂows, as long as the the velocity and the stagnation enthalpy do not change with time. This becomes clear, if the scalar product of (4.125) and the velocty vector v is obtained; then the second term of the left-hind side of the last equation vanishes. For otherwise equal conditions the entropy remains constant in irrotational ﬂows. 4.8.2 The Fundamental Equation of Gasdynamics The partial derivative of the density, pressure, and temperature can be eliminated from the conservation equations for mass, momentum, and energy, so that only the velocity components and their partial derivatives remain in the resulting equation. This equation is called the fundamental equation of gasdynamics. It will ﬁrst be derived for two-dimensional steady ﬂows, but

4.8 Theory of Characteristics

165

will then be extended to three-dimensional ﬂows. The continuity equation and the momentum equations read ∂ρ ∂u ∂v 1 ∂ρ + + +v u = 0 ∂x ∂y ρ ∂x ∂y ∂u 1 ∂p ∂u +v + = 0 u ∂x ∂y ρ ∂x ∂v 1 ∂p ∂v +v + = 0 . u ∂x ∂y ρ ∂y

(4.126)

The energy equation in the form h0 = const. and the diﬀerential of the thermal equation of state dp = yield the following relation: u up ∂p ∂p =− +v ∂y cp T ∂x u ∂v ∂u vp +v − ∂x ∂x cp T u ∂v ∂u +v −p ∂x ∂x ∂u ∂v + ∂x ∂y . (4.128) p p dρ + dT ρ T (4.127)

Multiplying the ﬁrst momentum equation with u and the second with v and adding the resulting relations yields u ∂p ∂p ∂u = − ρ u2 +ρuv +v ∂y ∂x ∂x ∂u ∂v ∂v + + ρ v2 ∂y ∂x ∂y . (4.129)

The fundamental equation of gasdynamics for two-dimensional steady ﬂow is obtained by subtracting (4.129) from (4.128) (u2 − a2 ) ∂u +vu ∂x ∂u ∂v + v 2 − a2 + ∂y ∂x ∂v =0 . ∂y (4.130)

In the above equation, a is the local speed of sound. For three-dimensional ﬂows, one obtains with the aid of the energy equation, the thermal equation of state, and the continuity equation, a relation for v · ∇ p v·∇p=− p v·∇ cp T v2 − p (∇ · v) , 2 (4.131)

which is inserted into the equation for the mechanical energy: v·∇ v2 v =− ·∇p 2 ρ (4.132)

After collecting terms, the last equation, written in Cartesian coordinates, becomes: (u2 − a2 ) +vw ∂v ∂w + ∂z ∂y ∂u ∂v ∂w + (v 2 − a2 ) + (w2 − a2 ) + ∂x ∂y ∂z ∂u ∂v ∂u ∂w + + + uv + uw =0 ∂y ∂x ∂z ∂x

(4.133)

In comparison to the equation derived previously, the fundamental equation of gasdynamics contains three additional terms for three-dimensional ﬂows.

166

4. Gasdynamics

4.8.3 Compatibility Conditions for Two-Dimensional Flows According to the Crocco vorticty theorem, in inviscid steady ﬂows the entropy remains constant along streamlines. It is therefore advantageous, to introduce streamline coordinates ξ and η, ∂s which are oriented in the direction of the streamlines and normal to them. Since ∂ξ = 0, the derivative of the entropy in the direction normal to the streamlines can be written as total derivative ds 1 = dη v The partial derivatives
∂s ∂x

u

∂s ∂s . −v ∂y ∂x

(4.134)

and

∂s ∂y

are substituted with the aid of the Crocco vorticity theorem , u ∂s = T ∂y
∂u ∂y

∂s v =− ∂x T so that the partial derivative

∂v ∂u − ∂x ∂y
∂v ∂x

∂v ∂u − ∂x ∂y

, ds dη

(4.135)

is replaced by

and the entropy gradient .

∂v ∂u T ds = + ∂x ∂y v dη

(4.136)

This expression is inserted into the fundamental equation of gasdynamics. There is obtained (u2 − a2 ) ∂u vuT ds ∂u ∂v + (v 2 − a2 ) + 2vu + =0 . ∂x ∂y ∂y v dη (4.137)

The last step consists in the elimination of the partial derivatives of the velocity components. The total derivatives du = are solved for
∂u ∂x

∂u ∂u dx + dy ∂x ∂y

and dv =

∂v ∂v dx + dy ∂x ∂y

(4.138)

and

∂v : ∂y

∂u du ∂u = −y ∂x dx ∂y

;

∂v 1 = ∂y y

∂v dv − dx ∂x

(4.139)

dy The quantity y , written as abbreviation for dx , can, as in the derivation of the Crocco vorticity theorem, be thought of as the unknown local slope of sets of curves. With the last two equations the fundamental equation of gasdynamics can be written as

(u2 − a2 )

du 1 dv + (v 2 − a2 ) − dx y dx −

y (u2 − a2 ) +

∂u 1 2 (v − a2 ) − 2 u v y ∂y 1 2 T ds =0 . (v − a2 ) − u v y v dη
∂u , ∂y

(4.140)

The slope y is now chosen in such a way, that the bracketed term multiplying y = uv ± a √ u2 + v 2 − a2 − a2

vanishes: (4.141)

u2

For supersonic ﬂows M a > 1 the fundamental equation reduces to a system of nonlinear ordinary diﬀerential equations with real coeﬃcients. If in the expression for y the inclination of the velocity vector, given by the angle β, and the Mach angle α are introduced, it is seen, that y is identical with the slope of the Mach lines.

4.8 Theory of Characteristics dy = tan(β ± α) . dx

167

(4.142)

The Mach lines are called the characteristic curves of the fundamental equation of gasdynamics. Along these curves the change of the velocity can be expressed by total derivatives. With the expressions obtained for y substituted into the fundamental equation, with u = v cos β and v = v sin β the compatibility conditions for the velocity take on the form cot α dv ds ∓ dβ + sin α cos α = 0. v γR (4.143)

These relations are valid along the Mach lines, the characteristic curves, given by the equations dy = tan(β ± α) . dx (4.144)

The diﬀerential ds must be determined with the compatibility condition for the entropy. ds = 0 along the streamlines dy = tan β dx . (4.145)

The solution of the equations describing the compatibility conditions along the characteristic curves, in general, is only possible with numerical methods for ordinary diﬀerential equations. For isentropic ﬂows, (4.142) can be integrated, since the term cot α dv represents the diﬀerential v of the Parndtl-Meyer angle ν. The compatibility conditions then reduce to the simple form d(ν ∓ β) = 0, or ν ∓ β = const. for dy = tan(β ± α) . dx (4.148) (Riemann invariants) (4.147) (4.146)

The diﬀerence or the sum of the Prandtl-Meyer angle and the local ﬂow inclination angle, respectively, are constant along the corresponding Mach lines. 4.8.4 Computation of Supersonic Flows With the Riemann invariants and the equations describing the slope of the characteristic curves, supersonic isentropic ﬂows can be determined. This is explained in the following example. Given are the ﬂow quantities in two neighboring points P1 and P2 , which are not supposed to lie on a Mach line. The Prandtl-Meyer angle ν and the ﬂow inclination angle β at the point of intersection of the characteristic curves P3 , can be determined with the integrals of the compatibiliy conditions. ν2 − β2 = ν3 − β3 ν1 + β1 = ν3 + β3 With the notation adopted, one obtains ν3 = β3 1 [(ν1 + ν2 ) + (β1 − β2 )] 2 1 [(ν1 − ν2 ) + (β1 + β2 )] = 2

(4.149)

(4.150)

168

4. Gasdynamics With the Prandtl-Meyer angle ν3 known, the Mach number M a3 and the other ﬂow quantities can be determined. The location of the point p3 is computed with a diﬀerence-solution of the characteristic equations. For the computation of a supersonic ﬂow ﬁeld it is assumed that n neighboring points (not on the same Mach line) are given in a supersonic ﬂow. At these points the ﬂow quantities are known, i. e. ν1 , ν2 ,..., νi , ...,νn and β1 , β2 , ...,βi , ... , βn . With νi , νi+1 and βi , βi+1 of two neighboring points, νj and βj of a third point Pj in the vicinity of the points Pi and Pi+1 can be determined. The position of the point Pj is obtained by discretizing the equations for the characteristic curves and solving the diﬀerence equations for small distances from the points Pi and Pi+1 . The coordinates of the new point are: xj ≈ yj (yi+1 − yi + m+ xi − m− xi+1 ) i i+1 (m+ − m− ) i i+1 ≈ yi+1 + m− (xj − xi+1 ) (4.151) i+1

The quantities m+ and m− are abbreviations of the slopes of the characteristic curves. For the determination of the coordinates of the point Pj mean values of the slopes can be employed m+ = i m− i+1 1 [tan(β + α)i + tan(β + α)j ] 2 1 [tan(β − α)i+1 + tan(β − α)j ] = 2

.

(4.152)

Only a certain part of the ﬂow ﬁeld depends on the initial data prescribed. This region is called the domain of dependence, sketched in the following ﬁgure. The ﬂow outside of the domain of dependence cannot be determined without additional information.

It is therefore necessary to prescribe boundary conditions for the computation of the other parts of the ﬂow ﬁeld. Two kinds of boundary conditions, representing streamlines of the ﬂow ﬁeld, are to be distinguished: the rigid wall and the free boundary, for example the edge of a

4.8 Theory of Characteristics

169

supersonic jet. For a rigid wall the curve of the contour of the wall yw = f (x) has to be known. If the point P1 is a point on the contour, all other points on the contour, for which the ﬂow quantities have to be determined, can be obtained by intersection of a characteristic with the curve describing the contour of the wall yw = f (x). For the points P4 and P7 the local ﬂow inclination angle is known from the relation (see sketch) dyw = f (x) = tan βw . dx (4.153)

The Prandtl-Meyer angle is obtained from the one Riemann invariant, corresponding to the Mach line intersecting the contour: ν4 = ν2 + β2 − β4 . (4.154)

After determining the ﬂow quantities at the point P4 , the quantities at all other points on the contour, for example, at the points P6 and P7 , or at all points between the points A and D are computed. ν7 = ν6 + β6 − β7 (4.155)

Along a free boundary the static pressure ps is known, but not the ﬂow inclination angle βs . Since the stagnation pressure p0 of an isentropic ﬂow is constant in the entire ﬂow ﬁeld, the ps pressure ratio p0 can be computed for the edge of the jet, also the Mach number M as , and the Prandtl-Meyer angle νs . The compatibility condition then yields the ﬂow inclination angle βs . If B is a point on the edge of the jet, it follows that βs = βn−1 + νs − νn−1 . (4.156)

The location of the point Ps is determined by letting the streamline through the point B intersect the Mach line through the point Pn−1 . With the procedure described, the ﬂow in the region B − D − E can be determined, one side of which is bounded by the edge of the jet (free boundary), and the other by the Mach line, originated in the point D. If the boundary conditions downstream from the points D and E are known, the computation of the ﬂow can be continued. It is often required, for example for nozzle ﬂows, to prescribe constant ﬂow quantities along the characteristic curve through the point D.

170

4. Gasdynamics Then the computation begins in the point F , which is assumed to lie on the characteristic curve through the point D in its vicinity. The computation is continued in the direction towards the free boundary along the characteristic D − B. The ﬂow in the region A − B − D is not inﬂuenced by the ﬂow condition prescribed. If two characteristic curves of the same kind intersect, the computation cannot be continued. In the point of intersection a compression shock is originated, causing a change of entropy, which is not included in the method of computation described here.

4.9 Compressible Potential Flows
The condition of irrotationality admits the introduction of a potential Φ for the description of compressible ﬂows. However, in contrast to incompressible ﬂows, the partial diﬀerential equation for the determination of the potential Φ is not linear. 4.9.1 Simpliﬁcation of the Potential Equation The fundamental equation of gasdynamics, in which only partial derivatives of the velocity components are contained, will be used to derive a simpliﬁed partial diﬀerential equation for the determination of the unknown potential Φ. The potential, deﬁned by u= ∂Φ ∂x v= ∂Φ ∂y w= ∂Φ ∂z , (4.157)

inserted into the fundamental equation of gasdynamics, yields the exact potential equation for steady compressible ﬂows: (u2 − a2 ) ∂2Φ ∂2Φ ∂2Φ + (v 2 − a2 ) 2 + (w2 − a2 ) 2 + 2 ∂x ∂y ∂z ∂2Φ ∂2Φ ∂2Φ +2uw +2vw = 0 +2 u v ∂x∂y ∂x∂z ∂y∂z

(4.158)

The nonlinearity of this equation is immediately evident, if the velocity components appearing in the coeﬃcients are replaced by (4.157). A closed-form solution is therefore not possible. Since in ﬂows about slender bodies the velocity components v and w are small and u deviates only little from the free-stream velocity, the exact potential equation can be simpliﬁed. With the aid of the perturbation velocities u = u − u∞ a perturbation potential can be deﬁned: u = ∂Φ ∂x
⎡

v =v

w =w

(4.159)

v =

∂Φ ∂y

w =

∂Φ ∂z

.

(4.160)

In order to simplify (4.158) the speed of sound is eliminated by employing the energy equation 2u u γ−1 a2 M a2 ⎣ =1− + ∞ a2 2 u∞ u∞ ∞
2

v + u∞

2

w + u∞

2

⎤ ⎦

.

(4.161)

4.9 Compressible Potential Flows

171

This relation is introduced together with the deﬁnition of the perturbation potential into the exact potential equation. If terms of the order O(M a2 ∞ are dropped, one obtains (1 − M a2 ) ∞ u ∂2Φ ∂2Φ ∂2Φ ∂2Φ + + 2 + 2 = (γ + 1) M a2 ∞ ∂z u∞ ∂x2 ∂x2 ∂y 2 u ∂ Φ ∂2Φ + (γ − 1) M a2 + 2 + ∞ u∞ ∂y 2 ∂z w ∂2Φ v ∂2Φ + . + 2 M a2 ∞ u∞ ∂x∂y u∞ ∂x∂z u2 ), u2 ∞ O(M a2 ∞ v2 ), u2 ∞ O(M a2 ∞ w2 ) u2 ∞ (4.162)

(4.163)

It is seen, that the simpliﬁed equation still contains nonlinear terms, and only if terms of the order O(M a2 ∞ u ), u∞ O(M a2 ∞ v ), u∞ O(M a2 ∞ w ) u∞ (4.164)

are neglected, a linear diﬀerential equation results for the determination of the perturbation potential (1 − M a2 ) ∞ ∂2Φ ∂2Φ ∂2Φ + 2 + 2 =0 . ∂x2 ∂y ∂z ∂2Φ ∂x2 (4.165)

This linearization is not valid for transonic ﬂows. For M a∞ → 1 the term (1 − M a2 ) ∞ tends to zero faster than the term (γ + 1) M a2 ∞ u ∂2Φ u∞ ∂x2 .

It is for this reason, that the last terms cannot be omitted in the computation of transonic ﬂows, since
M a∞ →1

lim

1 − M a2 − (γ + 1) M a2 ∞ ∞

u u∞

∂2Φ u ∂2Φ = −(γ + 1) ∂x2 u∞ ∂x2

.

(4.166)

For M a∞ → 1 the simpliﬁed potential equation therefore reads (1 − M a2 ) ∞ u ∂2Φ ∂2Φ ∂2Φ ∂2Φ + 2 + 2 = (γ + 1) M a2 ∞ ∂x2 ∂y ∂z u∞ ∂x2 . (4.167)

This equation must be employed for transonic ﬂows, while for subsonic and supersonic ﬂows the linearized potential equation can be used. 4.9.2 Determination of the Pressure Coeﬃcient For the linearized potential equation and also for the transonic approximation the relation, describing the pressure coeﬃcient cp = can be simpliﬁed. 2 γ M a2 ∞ p −1 p∞ (4.168)

172

4. Gasdynamics

Equation (4.168) can be rearranged for isentropic ﬂows with the aid of the energy equation:
⎨ 2 γ−1 M a2 cp = 1+ ∞ 2 ⎩ γ M a∞ 2 ⎧

u2 + v 2 + w 2 1− u2 ∞

γ γ−1

− 1⎭

⎫ ⎬

.

(4.169)

A series development of the last equation yields for the terms of ﬁrst and second order
⎡

cp = − ⎣

2u + (1 − M a2 ) ∞ u∞

u u∞

2

+

v u∞

2

+

w u∞

2

⎤ ⎦

.

(4.170)

If the second-order terms are neglected as in the derivation of the linearized potential equation, one obtains for the pressure gradient cp = −2 u 2 ∂Φ =− u∞ u∞ ∂x . (4.171)

This relation is also valid for transonic ﬂows. If the perturbation potential is known from a solution of the linearized potential equation or the transonic approximation, the pressure distribution (for example on a body) can be computed with the last equation. 4.9.3 Plane Supersonic Flows About Slender Bodies The linearized potential equation for supersonic ﬂows for determining the perturbation potential Φ in the form ∂2Φ ∂2Φ 1 − =0 2 2 − 1 ∂y 2 ∂x M a∞ can be solved with the ansatz of d’Alembert for the solution of the wave equation Φ(x,y) = f (ξ) + g(η). (4.173) (4.172)

The functions f (ξ) and g(η) must be twice diﬀerentiable functions. Their arguments are deﬁned as follows: ξ = x − λy η = x + λy (4.174)

For the solution of the linearized potential equation λ is the constant λ= For ξ = const. one obtains y= x − const . λ (4.176) M a2 − 1 . ∞ (4.175)

The curves described by this relation are straight lines with the slope dy = dx 1 M a2 − 1 ∞ (4.177)

They are identical with the Mach lines of the free stream with positive inclination. For η = const. there results x y = − + konst. λ (4.178)

4.9 Compressible Potential Flows and dy 1 =− dx M a2 − 1 ∞ .

173

(4.179)

The lines η = const. are the Mach lines of the free stream with negative inclination. The curvature of the Mach lines caused by their mutual interaction cannot be grasped by the linearized theory. The perturbations caused by a slender body propagate along the lines ξ = x−λy = const. on its upper side, and on its lower side along the lines η = x + λy = const. The solution is therefore split into two parts: 0 ≤ y 0 ≥ y Φ(x,y) = f (x − λy) Φ(x,y) = g(x + λy)

(4.180)

The still unknown functions f (ξ) and g(η) are determined from the boundary conditions. The kinematic ﬂow condition requires the velocity vector to be normal to the surface normal of the contour of the body. If the surface is described by the equation F (x,y) = 0, the kinematic ﬂow condition can be written as v · ∇ F (x,y) = 0 or (u∞ + u ) With u ∂F ∂F +v =0 . ∂x ∂y (4.182) (4.181)

u∞ , the slope of the contour of the body is approximately dy dx = tan βc ≈ c ∂F v ∂x = − ∂F u∞ ∂y

.

(4.183)

Since the body is slender, a series development for y = 0 can be employed: v (x,y) = v (x,o) + One obtains u∞ dy dx ≈ v (x,o) = c ∂v (x,o)y + ... ∂y

(4.184)

∂Φ (x,o) . ∂y

(4.185)

If the contour on the upper side is speciﬁed by Fu (x,y) = 0, the derivative of the function f (ξ) is determined by u∞ dy dx = −λ u df (x,o). dξ

(4.186)

The function f (ξ) is then obtained by integration. The pressure coeﬃcient for the upper side is cpu = − 2
∂Φ (x,o) ∂x

u∞

=−

2 df (x,o) , u∞ dξ

(4.187)

174

4. Gasdynamics

and with (4.186) cpu = dy 2 M a2 dx ∞ . u (4.188)

The pressure coeﬃcient for the lower side is obtained analogously: cpl = − 2 M a2 ∞ dy dx (4.189) l With the pressure coeﬃcient for the upper and lower side of the body known, wave drag and lift can be computed for supersonic free-stream conditions. 4.9.4 Plane Subsonic Flow About Slender Bodies Solutions of the linearized potential equation can also be constructed for compressible subsonic ﬂows. Since the partial diﬀerential equation for the perturbation potential (1 − M a2 ) ∞ ∂2Φ ∂2Φ + 2 =0 ∂x2 ∂y (4.190)

diﬀers from the exact potential equation for incompressible ﬂow only by a constant factor, solutions available for incompressible ﬂows can be used. The adaption of such a solution to compressible ﬂow is demonstrated here for a symmetric airfoil proﬁle at zero angle of attack. The thickness distribution of the proﬁle can be described by positioning sources and sinks along the axis of symmetry. For an incompressible ﬂow they are given by the following relation: Φ(x,y) = E ln (x − ξ)2 + y 2 2π (4.191)

In this equation the positions of the sources or sinks are given by the points (x = ξ, y = 0). The constant E is positive, if a source is considered, and negative if a sink is to be described. Equation (4.191) also satisﬁes the linearized potential equation for compressible ﬂow, if the coordinate y is replaced by y 1 − M a2 . Since the diﬀerential equation is linear, solutions can ∞ be superposed. With this ansatz an arbitrary contour can be generated with a distribution of sources and sinks along the line of symmetry. The potential then is of the form n Φ(x,y) = i=1 Ei ln (x − ξi )2 + m2 y 2 2π

(4.192)

2 with m2 = 1 − M∞ . Also a continuous distribution of sources and sinks can be thought of. The sum is then exchanged with an integral:

Φ(x,y) =

1 πm

t 0

f (ξ) ln (x − ξ)2 + m2 y 2 dξ

(4.193)

The source function f (ξ) has to be determined in such a way, that the contour of the proﬁle forms a streamline. It follows from the boundary condition that v (x,o) dy = u∞ dx = c 1 ∂Φ(x,o) 1 = lim y→0 π u∞ u∞ ∂y

t 0

f (ξ)

m y dξ (k − ξ)2 + m2 y 2

.

(4.194)

The limit process gives the simple relation f (x) = u∞ dyc . Then the expression for the potential dx becomes

4.9 Compressible Potential Flows Φ(x,y) = and the pressure coeﬃcient is cp (x,o) = −2 u 2 (x,0) = − u∞ πm t 0

175

u∞ πm

t 0

dyc dx

ln (x − ξ)2 + m2 y 2 dξ,

(4.195)

dyk dx

x=ξ

dξ x−ξ

.

(4.196)

The application of the solution is illustrated for the subsonic compressible ﬂow about a double parabolic wing section: yc = 4d x x − t t
2

(4.197)

In the ﬁrst step the source function f (x) is determined from the contour of the body. f (x) = u∞ dyc x d = 4 u∞ (1 − 2 )(4.198) dx t t

The perturbation velocity for y = 0 is u (x,0) = 4 d u∞ t πm t 0

1 ξ dξ (1 − 2 ) t (x − ξ)

,

(4.199)

and the expression for the pressure coeﬃcient becomes cp (x,0) = −8 d 1 t πm 2− 1−2 1− x ln x t t x t

.

(4.200)

The plot of the pressure coeﬃcient exhibits singular behavior of the solution at the leading and trailing edges. The failure of the solution at these points can be traced back to the simpliﬁcations intoduced by the linearization of the potential equation. Another point to be mentioned is, that the solution provides the same values for the product cp m, as long as the thickness ratio d retains the same value. In other t words, similar proﬁles have similar pressure distributions.

4.9.5 Flows about Slender Bodies of Revolution A slender body of revolution is an axially symmetric body, the local radius R(x) of which everywhere being much smaller than its length L. For pointed bodies solutions of the simpliﬁed potential equation for compressible ﬂow can be constructed in a manner similar to that of two-dimensional ﬂows. The velocity components in the axial, radial, and azimuthal direction are again expressed with a perturbation potential: u = ∂Φ ∂x v = ∂Φ ∂r w = 1 ∂Φ r ∂θ (4.201)

176

4. Gasdynamics The linearized potential equation for bodies of revolution, in comparison to plane ﬂows, contains an additional term, which results from the continuity equation, written in cylindrical coordinates: For a volume element of the ﬂow in cylindrical coordinates the mass balance for steady ﬂow yields the continuity equation in the following form 1 ∂ ∂ 1 ∂ (ρ v r) + (ρ w) = 0 . (4.202) (ρ u) + r ∂r ∂x r ∂θ The term containing the radial velocity component deviates in form from its counterpart for twodimensional ﬂow. This is also reﬂected in the linearized potential equation: (1 − M a2 ) ∞ 1 ∂2Φ ∂ 2 Φ ∂ 2 Φ 1 ∂Φ + 2 + 2 + =0 (4.203) 2 ∂x ∂r r ∂r r ∂θ2

The last term in the above equation vanishes for axially symmetric ﬂows ; the second-last term approaches the limit for r → 0 lim ∂2Φ 1 ∂Φ = r ∂r ∂r2 =0 . r=0 r→0

(4.204)

Equation (4.184) show, that the boundary condition for the body of revolution is given by the second derivative of the potential in comparison to the ﬁrst for two-dimensional ﬂow, see (4.185). Also the formulation of the boundary conditions for the body contour deviates from the formulation for plane ﬂows. If the contour of the body is speciﬁed by a certain distribution of the local radius over the length of the body L, the exact boundary conditon is dR = dx v u∞ + u .
R

(4.205)

If the boundary conditon is to be prescribed again in an approximation as in the case of the two-dimensional ﬂow, along the axis, then the series development for v(r,x) for r = 0 must take into account the diﬀerent variation of the radial velocity component manifested in the continuity equation. If the continuity equation is written in the form ∂ ∂ (ρ v r) = −r (ρ u) , ∂r ∂x it is seen that, if
∂ ∂x

(4.206)

(ρ u) remains ﬁnite in the vicinity of the axis, for r → 0 there results r→0 lim

∂ (ρ v r) → 0 . ∂r

(4.207)

In the neighborhood of the axis the product ρ v r is a function of the axial coordinate x, and the series development of the radial velocity component has the form: v r = a0 (x) + a1 (x)r + a2 (x)r2 + ... (4.208)

4.9 Compressible Potential Flows The approximation of the exact boundary condition then reduces to R dR = dx vR u∞ + u ≈
R

177

a0 (x) . u∞

(4.209)

Also the expression for the pressure coeﬃcient diﬀers from its counter part of plane ﬂows. Since v and u for r → 0 are of diﬀerent order of magnitude, the expression for the pressure coeﬃcient for axially symmetric ﬂows in the neighborhood of the axis is cp = −2 u v2 − 2 u∞ u∞ . (4.210)

The solution of the linearized potential equation for axially symmetric, compressible ﬂows about slender bodies can again be constructed by superposition of fundamental solutions. The thickness distribution of slender bodies of revolution is, as before, generated by a distribution of sources and sinks along the axis, given by the following relation n Φ(x,r) = i=1 Ei (x − ξi )2 + r2

(4.211)

for incompressible ﬂows, with the constants Ei having positive (sources) and negative (sinks) values. For a continuous source-sink distribution along the axis the sum is replaced by an integral, so that
L

Φ(x,r) =

f (ξ) dξ (x − ξ)2 + r2

0

.

(4.212)

The source function f (ξ) is determined with the boundary conditions, as described earlier. The solution for the compressible ﬂow is adjusted by stretching the radial coordinate by the quantity m = (1 − M a2 ). The potential is of the form ∞
L

Φ(x,r) =

f (ξ) dξ (x − ξ)2 + (mr)2

0

.

(4.213)

The solution can formally be extended to supersonic ﬂows. With λ2 = M a2 − 1 → 0 the ∞ linearized potential equation ∂ 2 Φ 1 ∂Φ ∂2Φ − λ2 + =0 2 ∂r r ∂r ∂x2 (4.214)

changes over into the form of the wave equation, the solution of which can be obtained by superposition of fundamental solutions for source-sink distributions along the axis Φ(x,r) = Ei (x − ξi )2 − (λ r)2 . (4.215)

The potential becomes imaginary, if λ2 r2 → (x − ξi )2 . Since an imaginary potential does not represent a physically meaningful solution, the integration for the continuous source-sink distribution can only be carried out up to the upper limit ξ ≤ x − λ r. The integral representation of the potential is then x−λr (4.216)

Φ(x,r) =

f (ξ) dξ (x − ξ)2 − (λ r)2

0

.

(4.217)

178

4. Gasdynamics The change of the upper limit of integration is necessitated by the ﬁnite extent of the region of inﬂuence of supersonic ﬂows. A point P (x,r) on the surface of the Mach cone, given by ξ = x − λ r, cannot be inﬂuenced by the sources and sinks further downstream.

4.10 Similarity Rules
Similar ﬂows exhibit physically similar behavior for geometrically similar conditions. Ratios of two kinds of forces or energies have to have the same value at corresponding points in ﬂows to be compared. The laws for carrying over a result obtained in one ﬂow to another are given, for example, by the similarity laws of Euler, Strouhal, Reynolds, and Mach. It is well known that not all of the similarity laws can be obeyed, if measured data are to be applied to the full-scale conﬁguration. In gasdynamics the similarity laws of Euler and Mach are extended to similarity rules for steady inviscid ﬂows in the frame of the simpliﬁed potential theory by combining the similarity laws with geometric parameters. In the following the similarity rules will brieﬂy be discussed. 4.10.1 Similarity Rules for Plane Flows After the Linearized Theory Assume a compressible subsonic ﬂow, which can be described by a solution of the linearized potential equation: ∂ 2 Φ1 ∂ 2 Φ1 1 + =0 ∂x2 1 − M a2 ∂y 2 1 1 (4.218)

The free-stream Mach number is M a1 and the perturbation potential Φ1 (x1 ,y1 ). It satisﬁes the boundary condition ∂Φ1 dy1 (x1 ,0) = u1 ∂y1 dx1 = u1 c d1 x1 f( ) . t1 1 t1

(4.219)

The quantities d1 and t1 are the maximum thickness and the length of the body, u1 is the free1 stream velocity, and f1 ( x1 ) the function describing the contour of the body, which is assumed t to be slender. The pressure coeﬃcient is cp1 = − 2 u1 ∂Φ1 ∂x1 . (4.220)

A second ﬂow with the potential Φ2 (x2 ,y2 ) satisﬁes the linearized potential equation for a diﬀerent free-stream Mach number M a2 : ∂ 2 Φ2 ∂ 2 Φ2 1 + =0 2 1 − M a2 ∂y2 ∂x2 2 2 The corresponding boundary condition is ∂ 2 Φ2 dy2 (x2 ,0) = u2 2 ∂y2 dx2 = u2 c (4.221)

d2 x2 f t2 2 t2

,

(4.222)

4.10 Similarity Rules and the pressure coeﬃcient is given by cp2 = − 2 ∂Φ2 u2 ∂x2 .

179

(4.223)

The potential Φ1 and also the pressure coeﬃcient cp1 are assumed to be known. It will be shown, that cp2 can be expressed through cp1 . First the diﬀerential equation for the potential Φ1 is transformed into the diﬀerential equation for the potential Φ2 . In order to be able to do so, it is necessary, to determine the dependencies x2 = x2 (x1 ) and y2 = y2 (y1 ). This is done by ∂2 expressing the diﬀerential operator ∂x2 by x2 :
1

∂2 = ∂x2 1

∂x2 ∂x1

2

∂2 ∂ 2 x2 ∂ + ∂x2 ∂x2 ∂x2 2 1

(4.224)

Since the ﬁrst derivative of the potential does not appear in the diﬀerential equation, the term ∂ 2 x2 has to vanish. Hence the relation connecting x2 with x1 is linear, such that x2 = a x1 , with ∂x2 1 a being a constant. the same is true for y2 and y1 , and y2 = b y1 . The diﬀerential equation for Φ1 can then be written in the form a2 ∂ 2 Φ1 ∂ 2 Φ1 b2 + =0 , 2 ∂x2 1 − M a2 ∂y2 2 2 (4.225)

and the boundary condition transforms into the relation b ∂Φ1 = u1 ∂y2 d1 t1 f1 x1 t1 . (4.226)

The pressure coeﬃcient becomes cp1 = − 2 a u1 ∂Φ1 ∂x2 . (4.227)

Since the simpliﬁed potential equation is linear and homogeneous, every solution can be multiplied by an arbitrary constant. In order to express Φ1 through Φ2 , a linear relation can therefore be used to connect the two solutions: Φ1 (x1 ,y1 ) = A Φ2 (x2 ,y2 ) (4.228)

The quantity A is an unknown constant. If Φ1 is substituted by (4.228) in the diﬀerential equation, (4.225), it is seen, that the original equation for Φ1 is transformed into the diﬀerential b equation for Φ2 , if the ratio a is given by b = a The boundary condition for Φ1 then reads u2 ∂Φ2 (x2 ,0) = bA ∂y2 d1 t1 f1 x1 t1 . d2 t2

1 − M a2 1 . 1 − M a2 2

(4.229)

(4.230) is (4.231)

The boundary condition for Φ2 with an arbitrarily assumed thickness ratio ∂Φ2 (x2 ,0) = u2 ∂y2 d2 t2 f2 x2 t2 .

180

4. Gasdynamics x1 t1

For bodies of the same geometric family f1 A=

= f2 .

x2 t2

the constant A is obtained to (4.232) d2 t2

1 u1 d1 t2 b u2 t1 d2

On the other hand the constant A can be chosen and the thickness ratio The following relation holds for the pressure coeﬃcients cp1 = − Finally there results cp1 1 − M a2 1 d1 t1

thereby be ﬁxed.

2aA u1

∂Φ2 ∂x2

=

u2 A a cp2 u1

.

(4.233)

= cp2

1 − M a2 2 d2 t2

.

(4.234)

Since the left-hand side of (4.234) was assumed to be known, the subscripts on the right-hand side can be dropped: cp 1 − M a2 1 d1 t1

= const. .

(4.235)

The last relation provides four similarity rules for bodies, described by the contour f ( x ): t 1. The pressure coeﬃcient cp remains constant, if the ratio
1 (1−M a2 )− 2 , ∞ 2 (1−M∞ ) 2 (d) t 1

remains constant.

2. The pressure coeﬃcient cp varies with if the thickness ratio d remains constant t (Prandtl-Glauert rule). 3. The pressure coeﬃcient cp varies with d , if the free-stream Mach number is kept constant. t 4. The pressure coeﬃcient cp varies with (1 − M a2 )−1 , if the thickness ratio d varies with ∞ t 1 (1 − M a2 )− 2 (G¨thert rule). o ∞ These rules are also valid for supersonic ﬂows. 4.10.2 Application of the Similarity Rules to Plane Flows If the pressure distribution on an airfoil section was measured for a certain free-stream Mach number, the results of the measurements can be converted to other free-stream Mach number with the Prandtl-Glauert rule. According to this rule the product of the pressure coeﬃcient 1 and (1 − M a2 ) 2 remains constant: ∞ cp2 = cp1 1 − M a2 1 1 − M a2 2 . (4.236)

If, for example M a1 is M a1 = 0.4, one obtains for a second Mach number, say M a2 = 0.7 ⇒ cp2 = 1.283 cp1 . Another application of the Prandtl-Glauert rule is demonstrated with the example of determining the critical free-stream Mach number for an airfoil proﬁle. If the free-stream velocity about an airfoil proﬁle is subsonic, in general the pressure on the upper side of the proﬁle is lowered, and the ﬂow can locally attain sonic or even supersonic velocity. The particular free-stream Mach number, for which locally the speed of sound is attained for the ﬁrst time, is called critical free-stream Mach number. From the expression for the pressure coeﬃcient
⎨ 2 + (γ − 1) M a2 2 ∞ cp = 2 ⎩ 2 + (γ − 1) M a2 γ M a∞ ⎧ γ γ−1

− 1⎭

⎫ ⎬

(4.237)

4.10 Similarity Rules the critical value of the pressure coeﬃcient is obtained by setting M a = 1: cpcrit
⎨ 2 + (γ − 1) M a2 2 ∞ = 2 ⎩ γ M a∞ γ+1 ⎧ γ γ−1

181

− 1⎭

⎫ ⎬

(4.238)

This relation is independent of the airfoil proﬁle considered. The critical free-stream Mach number of an airfoil proﬁle can be determined with the Prandtl-Glauert rule, if the pressure distribution is known for another Mach number, for example M a∞ = 0. Then the expression for the pressure coeﬃcient can be written as cp = cpinc 1 − M a2 ∞ . (4.239)

With (4.239) The critical Mach number can be determined by intersecting the curve obtained with (4.238) with the curve obtained with (4.239) as shown in the diagram. The application of the G¨thert rule will be shown as the third example. According to this rule o 1 the product of the thickness ratio d and (1 − M a2 ) 2 remains constant. ∞ t If the pressure distribution of a wing, known for a given angle of attack in incompressible ﬂow, is to be converted to a compressible subsonic ﬂow, the chord of the wing t is changed for constant maximum thickness to t = tinc. The corresponding angle of attack is α= αinc. 1 − M a2 ∞ . (4.241) 1 − M a2 ∞ . (4.240)

Equations (4.240) and (4.241) are graphically illustrated in the following sketch. The corresponding angle of attack α in compressible subsonic ﬂow is larger than in incompressible ﬂow and the chord is shortened as a result of the Prandtl-Glauert transformation. This result shows already, that the similarity rules play an important role in the design of wing sections and wings for compressible subsonic ﬂow, illustrated in the following sketches.

compressible ﬂow M a∞ < 1

incompressible ﬂow M a∞ = 0

For delta wings also the planform has to be changed.

182

4. Gasdynamics

compressible ﬂow

incompressible ﬂow

The sweep angle σ for compressible ﬂows is
⎛ ⎞

σ = arctan ⎝tan

σinc. 1 − M a2 ∞

⎠

.

(4.242)

The similarity rules loose their validity in the vicinity of stagnation points, since the linearized theory is not valid there. 4.10.3 Similarity Rules for Axially Symmetric Flows If Φ1 (x1 ,r1 ) and Φ2 (x2 ,r2 ) are solutions of the linearized potential equation for axially symmetric ﬂows ∂2Φ 1 + ∂x2 1 − M a2 ∞ ∂ 2 Φ 1 ∂Φ + ∂r2 r ∂r =0 , (4.243)

the relations x2 = a xr , r2 = b r1 are valid as for two-dimensional ﬂows with b = a 1 − M a2 1 and Φ1 (x1 ,r1 ) = A Φ2 (x2 ,r2 ) . 1 − M a2 2 (4.244)

Diﬀerences are encountered in the conversion of the boundary conditions. The boundary condition for the potential Φ1 (x1 ,r1 ) is with R1 = Rmax l l1 f1 x1 l1 x1 l1

1

∂Φ1 (x1 ,r1 = R1 ) = u1 ∂r1

Rmax l

1

f1

.

(4.245)

If Φ1 and r2 are replaced with the ansatz above, there is obtained bA ∂Φ2 (x2 ,r2 = bR1 ) = u1 ∂r2 Rmax l f1 x1 . l1 l2 f2 . x2 l2

1

(4.246)

The boundary condition for the potential Φ2 is with R2 = ∂Φ2 (x2 ,r2 = R2 ) = u2 ∂r2 Rmax l f2

Rmax l 2

2

x2 l2

(4.247)

4.10 Similarity Rules By comparing the last three equations it is seen, that for bodies of the same family f1 f2 x2 l2 x1 l1

183 =

the thickness ratios have to satisfy the condition Rmax l = 1 − M a2 1 1 − M a2 2 Rmax l (4.248)

2

1

.

As a consequence the constant A is ﬁxed by the boundary conditions: A= The pressure coeﬃcient cp1 cp1 = − 1 2 ∂Φ1 − 2 u1 ∂x u1
⎡

u1 u2

Rmax l 1 Rmax l 2

1 b

(4.249)

∂Φ1 ∂r1

2

,

(4.250)

after replacing Φ1 by Φ2 and r1 by r2 , takes on the following form: cp1 1 1 − M a2 ⎣ 2 ∂Φ2 1 − 2 = − 1 − M a2 u2 ∂x u2 2 ∂Φ2 ∂r2
2

⎤ ⎦

(4.251)

The content of the square bracketed term represents the pressure coeﬃcient cp2 . The similarity rule, which results from this relation, has the form cp = const. 1 − M a2 ∞ . (4.252)

The last equation can also be written in the following form with the condition for the maximum radii cp = const.
Rmax l

1 − M a2 ∞

,

(4.253)

which is equal to the similarity rules for plane ﬂows. 4.10.4 Similarity Rules for Plane Transonic Flows Instead of the linearized potential eqaution the nonlinear approximation for transonic ﬂows must be used ∂2Φ ∂2Φ 1 (γ + 1) M a2 1 ∂Φ ∂ 2 Φ ∞ + = 2 2 ∂y 2 ∂x 1 − M a∞ (1 − M a2 ) u∞ ∂x ∂x2 ∞ . (4.254)

With the same ansatz as employed previously the constant A is already ﬁxed by the right-hand side of the diﬀerential equation A= (γ2 + 1) M a2 (1 − M a2 ) 2 1 (γ1 + 1) M a2 (1 − M a2 ) 1 2 , (4.255)

and the expression for the pressure coeﬃcient becomes: cp1 (γ1 + 1) M a2 cp2 (γ2 + 1) M a2 1 2 = 2 1 − M a1 1 − M a2 2 (4.256)

Since the constant A is ﬁxed, these considerations cannot be carried over to axially symmetric ﬂows.

184

4. Gasdynamics

4.11 Selected References
Ganzer, U.: Gasdynamik Springer Verlag, 1988. Liepmann, H.W., Roshko, A.: Elements of Gasdynamics John Wiley Verlag, 1957. Oswatitsch K.: Gasdynamics Academic Press Inc., 1956. Sauer, R.: Einf¨hrung in die theoretische Gasdynamik, Springer Verlag, 1960. u Shapiro, A.H.: The Dynamics and Thermodynamics of Compressible Fluid Flow, Vol.1 John Wiley Verlag, 1953. Zierep, J.: Theoretische Gasdynamik, Band 1 Braun Verlang, 3. Auﬂage, 1976.

5. Exercises in Gasdynamics

5.1 Problems
5.1.1 One-Dimensional Steady Flows of Gases 1.1 Consider an inviscid ﬂow of gas without heat exchange through a stream tube.

a2 = (a) Formulate the ﬁrst law of thermodynamics for this system! (b) What kinds of energies interact 1. in compressible ﬂuid ﬂow 2. in incompressible ﬂuid ﬂow? 1.2 An airplane ﬂies over an observer in level ﬂight. m H = 577 m v = 680 T = 287 K; s Nm γ = 1.4 R = 287 kg K (a) Determine the Mach number! (b) What distance did the aeroplane cover, before it was heard by the observer? (c) When was the noise generated the observer heard? 1.3 Determine for an isentropic ﬂow (γ = 1.4) (a) the critical temperature ratio, (b) the critical pressure ratio, (c) the limiting value of the critical u Mach number M a∗ = a∗ for M a → ∞! How large is the change of temperature across the normal compression shock? 1.4 A small pressure disturbance moves through a gas at rest. Derive the expression for the speed of propagation!

∂p ∂ρ

s

1.5 A jet-propelled airplane passes a second jet-plane at a distance b. m m vB = 680 b = 170 m; vA = 510 s s Nm T = 287 K R = 287 γ = 1.4 kg K After what time would the pilot of the overtaken airplane have been able to hear the noise generated by the faster ﬂying airplane? 1.6 Gas ﬂows isentropically and steadily through a vertical pipe with cylindrical cross section against the direction of the gravitational acceleration.

Derive the expression describing the convective acceleration, and state, how the velocity changes for (a) M a(x = 0) < 1 (b) M a(x = 0) > 1

186

5. Exercises in Gasdynamics (a) Determine the ratio of the critical A∗ cross sections A2 ∗ 1 (b) How does the Mach number vary in the direction of the ﬂow? (c) To what value can M a1max increase? 1.11 Air ﬂows out of a large container through a nozzle into the open air. The ﬂow in the nozzle is steady and isentropic. The Mach number in the throat is smaller than unity.

1.7 Derive the expression relating M a∗ to M a for a one-dimensional isoenergetic ﬂow with the aid of the energy equation and show, that for γ = const. lim M a∗ = γ+1 . γ−1

M a→∞

1.8 Derive the relation describing the deA pendence of the area ratio A∗ on the Mach number M a and the isentropic exponent γ for an isentropic nozzle ﬂow! 1.9 The relation between the mass ﬂow Θ and the local velocity v, nondimensionalized with the speed of sound at stagnation conditions a0 , is shown in the following diagram for one-dimensional, isentropic and isoenergetic ﬂow.

(a) Derive the function f describing the retaining force F F =f A p0 γ, pu p0 .

Determine: (a) the maximum mass ﬂow, (b) the velocity, corresponding to this value of the mass ﬂow, (c) the maximum velocity, nondimensionalized with the speed of sound at stagnation conditions. 1.10 Viscous ﬂow of air passes through a thermally isolated pipe with cylindrical cross section at subsonic speed.

(b) Develop the function in a power series for the case pu ≈ p0 , and compare the result with the corresponding relation for incompressible ﬂow! 1.12 Consider an inviscid nozzle ﬂow of a perfect gas with constant speciﬁc heats. T Determine T0 , pp0 , and ρρ0 as a function of the Mach number for adiabatic conditions! 1.13 In the nozzle sketched below a ﬂow is accelerated to supersonic speed corresponding to a Mach number M a = 2.

M a1 < 1;

p01 = 2; p02

γ = 1.4

5.1 Problems (a) Sketch the pressure distribution pp0 , the variation of the Mach number M a, and the ratio of the mass ﬂow ρu along the axis of the nozzle. ρ∗ u ∗ (b) Determine for the exit cross section: A∗ 1. the cross-sectional ratio AE pE 2. the pressure ratio p0 , 3. the temperature ratio TE , T0 4. the mass-ﬂow ratio ρE uE ρ∗ u ∗ 1.14 Air is sucked through a convergent nozzle into a large container.

187

(a) the stagnation density ρ0 , (b) the quantities M a, M a∗ , p, ρ, T , u, and m in the cross section of the ˙ throat AH and in the exit cross section AE assuming isentropic ﬂow in the nozzle! 1.16 Gas ﬂows through a thermally isolated pipe with constant cross section. (a) Derive the relation h+ C1 = C2 ; C1 , C2 = const. (∗)! 2 ρ2

p0 = 105 Pa; T0 = 300 K; γ = 1.4; Nm R = 287 ; Ae = 0.02 m2 kg K Determine the mass ﬂow for (a) pc = 7 · 104 P a, (b) pc = 2 · 104 P a! (c) Sketch the variation of the massm ﬂow ratio ρ0 a˙0 Ae , as a function of pc the pressure in the container p0 ! 1.15 Consider a large pressurized vessel ﬁlled with air, that ﬂows through a nozzle sketched below into the surrounding air.

(b) With the relation (∗) the so-called “Fanno”- curve is obtained in the hs diagram (see sketch). 1. Prove that the Mach number is unity in the point P ! 2. Caused by wall friction, the stagnation pressure decreases in the direction of the ﬂow. Plot the direction of the change of state along the “Fanno”- curve for supersonic and subsonic ﬂow!

1.17 Consider the inviscid ﬂow of a gas through a thermally non-isolated pipe with constant cross section. (a) Derive the following relations for this ﬂow! P = C1 + C2 ; ρ C1 ,C2 = const.(∗)

pa = 105 Pa; T0 = 288 K; γ = 1.4; Nm R = 287 ; AH = 1 cm2 kg K Determine for p0outl. = 7.824 · 105 P a and p0 = 1.086 · 105 P a

(b) The relation (∗) yields the “Raleigh”-curve in the h-s diagram.

188

5. Exercises in Gasdynamics ﬂow is steady. Determine the quantities p, p0 , T , T0 , M a, and also u upstream of (Index “1”) and downstream (Index “2”) from the shock for: (a) the system of coordinates moving with the shock xs , (b) the coordinate system ﬁxed to the pipe xR . 1. Determine the Mach number in the point P ! 2. Mark the branches of the curve, indicating supersonic and subsonic ﬂow! 3. Indicate the direction of the change of state on the branches of the curve, if the pipe is heated! 2.4 A normal compression shock moves through quiescent air with temperature T1 and pressue p1 in a thermally isolated shock tube with velocity u. The static pressure downstream from the shock wave is p2 . time t = t
A

p
2

u model

p

1

5.1.2 Normal Compression Shock 2.1 How large can the density ratio ρ2 = f (γ) across a nomal compression ρ1 shock become? γ =const. 2.2 The free-stream velocity in front of a normal compression shock is m u1 = 300 and the critical Mach nums ∗ ber M a1 = 1.25. Compute the velocity u2 downstream from the shock, without using a diagram! 2.3 A normal compression shock moves through air resting in a thermally isolated pipe with pressure p1 and temperature T1 , with velocity u.
Ma oo p
2

T1

time t = t e

u

p

1

T1

p1 = 105 Pa; p0 = 105 Pa;

T1 = 300 K; γ = 1.4; Nm R = 287 kg K

1.) Determine the velocity of the shock u (for t = tA ). 2.) Determine for t = tB , (a) the Mach number M a∞ , (b) the stagnation temperature T0∞ , (c) the stagnation pressure p0∞ . 2.5 An airplane ﬂies at supersonic speed. A shock wave is formed, which is normal in front of the nose of the airplane.

p1 = 105 Pa; T1 = 293 K; γ = 1.4; m Nm ; u = 515 R = 287 kg K s The ﬂow is unsteady with respect to coordinates ﬁxed to the pipe xR ; for an observer moving with the shock (xs ) the

5.1 Problems Nm T = 287 K γ = 1.4 R = 287 kg K m u1 = 680 s How large is the temperature change of the air across the shock? 2.6 A normal compression shock was generated in the divergent part of a plane Laval nozzle between the cross sections AA and AB , for which the Mach numbers M aA and M aB are known.

189

M aEdesign1 . The pressure in the chamber pv can be varied.

p0 = 105 Pa; Nm ; γ = 1.4; R = 287 kg K T0 = 280 K; AH = 1 cm2 M aEdesign1 = 2.3; Assume isentropic ﬂow in the nozzle and determine: (a) pEdesign1 and AH , AE (b) the lowest pressure pv1 , for which the ﬂow in the nozzle is subsonic everywhere, and also M aE1 , (c) the pressure in the chamber pv2 , at which a normal shock is formed in the exit cross section AE , (d) the thrust Fs , for pv = pEdesign and pv = pv2 . (e) Sketch the variation of p and M a along the axis of the nozzle for pv = pv1 and pv = pEdesign . 2.9 During the test procedure of a jet engine equipped with a Laval nozzle the chamber pressure pv and the pressure pE at the exit of the nozzle are varied. The stagnation pressure at the exit of the nozzle p0E is measured with a pitot tube.

M aA = 2.2; γ = 1.4

M aB = 0.6;

AB AA

= 1.8;

Determine: (a) the ratio of the stagnation pressures p01 , p02 (b) the ratio of the static pressures p2 p1 immediately upstream of and downstream from the shock, (c) the pressure ratio pB . pA 2.7 A turbo engine sucks air out of the atmosphere. The pressure in front of the compressor is p1 .

γ = 1.4 R = 287 p0 = 105

Nm kg K

T = 287 K p0 = 105 Pa; T0 = 280 K; AH = 1 cm2 ; M aEdesign = 2.3; Nm R = 287 ; γ = 1.4; kg K Determine the position of the normal ∗ compression shock As , the mass ﬂow m, ˙ A and the Mach number at the exit of the nozzle M aE for:

N N p1 = 0.74 · 105 2 ; m2 m A = 9 · 10−3 m2

Determine the mass ﬂow through the engine! 2.8 An engine equipped with a Laval nozzle is tested on a test stand. The engine is designed for the Mach number

190

5. Exercises in Gasdynamics (a) pva = 0.645 · 105 P a ; p0Ea = 0.721 · 105 P a , (b) pvb = 0.816 · 105 P a ; p0Eb = 0.876 · 105 P a . (c) Compute the ratio ∆ sa . ∆ sb (d) Sketch the variation of p, p0 , and M a along the axis of the nozzle. (e) At the pressure pve a normal compression shock is observed at A∗ = 0.9. Determine M a∗ , ahead 1 AE of the shock, M a∗ downstream from 2 the shock, M aE , and the pressure pve . mize the testing time for a test Mach number M aM = 2.3? 2. Compute the gain in testing time for this case and compare it to the situation that the wind tunnel is operated without a diffuser! 2.11 Air ﬂows isentropically (γ = 1.4) out of a large, frictionless supported container through a well-rounded nozzle into the surroundings.

2.10 Air is sucked from the surroundings into a supersonic wind tunnel, equipped with a Laval nozzle and variable diffuser, and from there into a large container (volume Vv ) with constant temperature Tv . The Mach number in the test section for the design condition is M aM = 2.3.

(a) Determine the dimensionless thrust Fs for the pressure ratios pA pa D = 1; 0.6; 0.2; 0 ! p0 (b) How large are the corresponding values for incompressible ﬂow? 2.12 An adiabatic wind tunnel sucks air from the surroundings into an adiabatic container (volume Vc ). The Mach number M aM in the test section is controlled by adjusting the throat AL of the Laval nozzle. pa = 105 Pa; Ta = 300 K; γ = 1.4; Nm ; Vc = 400 m3 ; R = 287 kg K 2 AE = 0.508 m ; AM = 0.2 m2 ;

p0 = 105 Pa; T0 = Tv = 280 K; Nm ; Vv = 1000 m3 ; R = 287 kg K AE = AM ; AH = 0.1 m2 ; γ = 1.4; M aEdesign = 2.3 (a) A normal compression shock is observed in the exit cross section of the diﬀuser. Determine the pressure in the container pc1 ! (b) A normal compression shock is observed at As = 0.57 AE . Determine: 1. the Mach number M a∗ and the s static pressure ps downstream from the shock, 2. the pressure in the container pc2 ! (c) Sketch the variation of the Mach number M a along the axis of the nozzle up to the point P for both cases! (d) 1. How must the throat of the diffuser Av be adjusted, to maxi-

(a) the tunnel is started by opening a slide valve and at time t = t0 a steady ﬂow has built up in the test section. The Mach number M aE in the exit cross section AE of the

5.1 Problems Laval nozzle is at this time equal to the design Mach number. Determine the cross sectional area of the throat AL of the Laval nozzle, so that the Mach number in the test section M aM is M aM = 0.8. How large are then the Mach number M aE and the static pressure pE in the exit cross section AE ? (b) Determine the maximum measuring time tmax , for which the Mach number in the test section is constant M aM = 0.8. (c) A loss of stagnation pressure ∆p0 = 0.020pa is caused by installations in the test section. How large is now the Mach number M aM in the test section, if the cross section of the throat AL of the Laval nozzle is the same as before? How has the cross section of the throat AL of the Laval nozzle to be adjusted, so that the Mach number in the test section is again M aM = 0.8? 5.1.3 Oblique Compression Shock 3.1 A supersonic ﬂow is turned by the contour sketched below by the angle β. Physical plane:

191

m tangential component ut1 = 300 . s Across the shock the static temperature is increased to T2 = 1.2 T1

Determine with γ = 1.4 and R = 287 Nm/kgK: (a) the Mach number M a1 and the static temperature T1 upstream of the shock, (b) the velocity components un2 and ut2 , the Mach number M a2 , and the turning angle β downstream from the shock, (c) the velocity components un1 and ut1 for M a1 = const., so that M a2 = 1. How large is the turning angle β now? 3.3 Air is sucked out of the atmosphere through a supersonic wind tunnel into a vacuum container (see sketch). The Laval nozzle is designed for the Mach number M aE in the test section. p = 105 Pa; T = 280 K; γ = 1.4; Nm ; AH = 0.1 m2 ; R = 287 kg K Vv = 1000 m3 ; M aE = 2.3

Sketch: (a) the variation of the turning angle β as a function of the shock angle σ, and plot the state B in the physical plane, (b) the geometric construction for determining the state B in the hodograph plane (u, v velocity components in the x, y directions), (c) the ﬂow ﬁeld for β1 > βmax ! 3.2 The velocity V1 upstream of a straight compression shock is given by its norm mal component un1 = 400 and the s

During the test the temperature in the container is Tv = 280 K = const. At time t = 0 the pressure in the container is pv (t = 0) = 0.08 · 105 Pa.

192

5. Exercises in Gasdynamics (a) Determine the available testing time ∆t (undisturbed ﬂow in the test section, M aE = 2.3). (b) Determine the following quantities: Shock angle σ, turning angle β, p02 , M a2 , T2 , T02 , and the velocity V2 downstream from the shock for the pressure in the container pv = 0.16 · 105 Pa. (c) A wedge with nose angle 2 βw = 40◦ is mounted in the test section.

The ﬂow is assumed to be twodimensional and inviscid. The pressure in the atmosphere is given by: pa (z) γ−1 = 1− pa (0) γ Rocket: p0 = 15 · 105 P a; AE AH = 0.15 m2 ; surroundings: pa (0) = 105 Pa; γ γ−1

az

How large can the angle of attack be, without detachment of the compression shock? How large are the pressure difference pl − pu , and the Mach numbers M au and M al for this case? 3.4 Determine for the sketched ﬂow ﬁeld (M a∞ , h0 = const.):

= 1 m2 ; γ = 1.4 γ = 1.4 1 a = 1.1 · 10−4 ; m

M aa = 2; R = 287

M ab = 1.8; Nm kg K

γ = 1.4;

1. Determine for lift-oﬀ: (a) the Mach number M aE and the pressure pE in the exit cross section AE , (b) the shock angle σ of all shocks and the Mach numbers downstream from the shocks! (c) Sketch the ﬂow ﬁeld downstream from the exit cross section! Sketch the shocks, the streamlines, and the Mach lines! 2. What is the height of the design condition of the rocket? 3.6 The ﬂow in a two-dimensional supersonic wind tunnel is compressed by the shock conﬁguration shown in the following sketch.

(a) the entropy diﬀerence ∆ s = sb − sa , (b) the ratio of the velocities | Va | / | Vb |, (c) the density ratio ρa /ρb . 3.5 In the combustion chamber of a slender rocket a steady state with the stagnation pressure p0 is attained after ignition.

M a1 = 3;

ρ2 /ρ1 = 1.85;

γ = 1.4

5.1 Problems Determine the ratio of heights of the tunnel h3 /h1 and the ratio h1 /l. 3.7 Given is the detached shock of a blunt symmetric body as sketched below. 5.1.4 Expansions and Compression Shocks

193

Consider the shock as straight line with an averaged slope between the points A and B for the determination of ∆ x.

4.1 The radial component ur and the tangential component ut of the velocity are given for a point P inside of a PrandtlMeyer expansion.

Sketch the shock polar and the heartcurve diagram. Plot the points 1 to 4 in the diagrams. 3.8 Consider a two-dimensional curved compression shock (see sketch). The shock angles σA and σB are known for the points A and B, the verticle distance between them is ∆ y. Determine (approximately) the vorticity of the ﬂow downstream from the shock for the streamline in the middle between the points A and B.

m m ; ut = 250 ; s s Nm R = 287 ; γ = 1.4 kg K Determine: (a) the Mach number M ap for the point P (b) and the temperature Tp ! ur = 500 4.2 Derive ∂vr = f (a) for the sketched ∂Φ Prandtl-Meyer expansion; a = speed of sound.

M a1 = 2; σA = 35◦ ;

∆ y = 3 cm; T0 = 300 K; σB = 55◦ ; Nm ; γ = 1.4 R = 287 kg K Hint: Component of the vorticity in the x-y-plane: ∇xv = ≈ ∂v ∂u − k ∂x ∂y ∆v ∆u − k ∆x ∆y

Hint: ∇xv= 1 ∂vr 1 ∂ (r vΦ ) − r ∂r r ∂Φ

4.3 The ﬂow in the sketched supersonic wind tunnel is accelerated from M a1 to M a3 by two Prandtl-Meyer expansions. The ﬁrst and the last characteristic lines of both expansions are plotted in the drawing.

194

5. Exercises in Gasdynamics The ﬂow is expected to leave the tunnel parallel to the direction of the oncoming ﬂow. Determine the length of the distances AB and AC and their inclination to the horizontal line. (b) the angle β3 ! (c) Sketch the ﬂow ﬁeld for pa = 0.28 · 105 Pa! (d) At what surrounding pressure in the atmosphere is a normal shock to be observed in the exit cross section of the nozzle? 4.6 A plane inlet diﬀuser of a jet engine is so designed, that the front shock hits the lip at a free-stream Mach number M a1 = 3. The center body has a wedge angle of β1 . The ﬂow is decelerated in the diﬀuser with a Prandtl-Meyer compression to the speed of sound.

M a1 = 1; γ = 1.4

M a3 = 2.06;

h1 = 0.1 m;

4.4 Consider the plane supersonic ﬂow sketched below.

M a1 = 3; M a1 = 2.5; β1 = β2 = 12 ; p1 = 0.541·105 Pa; ∆p = 0.304·105 Pa; Nm γ = 1.4; R = 287 kg K Determine: (a) M a2 , σ, p02 , and p2 , (b) M a3 and p3 , (c) M a4 , β3 . (d) Sketch streamlines, shock waves, and expansion fans. 4.5 An oblique compression shock is observed at the exit of a plane Laval nozzle for the pressure in the surrounding atmosphere pu . The shock angle σ1 is 40◦ and the turning angle β1 = 10◦ .
◦

T1 = 250 K;

M a3 = 1;

15◦ ; l = 1 m; p1 = 0.54 · 105 Pa; width b = 0.5 m

γ = 1.4; Nm ; β1 = R = 287 kg K

(a) Determine for the design condition: 1. M a2 ; p2 ; T2 ; T02 ; p02 ; and ∆ s , R 2. β3 , p3 , and the mass ﬂow m. ˙ (b) How high would the static pressure p3is. be for isentropic compression from M a1 = 3 down to M a3 = 1? (c) At what free-stream Mach number M a1c would the compression shock still be attached to the center body? 4.7 A plane three-shock diﬀuser is designed for the free-stream Mach number M a1 . The compression of the air from “3” to “4” (see sketch below) is enforced by a normal compression shock.

pa = 0.2 · 105 Pa; γ = 1.4;

R = 287

Nm ; kg K

Determine: (a) the Mach numbers M a1 and M a3 , and the static pressure p2 ,

5.1 Problems M a1 = 3; T1 = 300 K; γ = 1.4; A = 0.1 m2 ; p1 = 0.5 · 105 Pa; Nm R = 287 ; β = 16◦ kg K Determine: (a) the static pressure p4 and the Mach number M a4 , (b) and the mass ﬂow m through the in˙ let diﬀuser. (c) How large would the mass ﬂow m ˙ and the static pressure p4is. be for isentropic compression from M a1 down to M a4 ? 4.8 The contour shown below decreases the Mach number M a1 > 1 of the free stream by a Prandtl-Meyer compression down to M a = 1. (a) Sketch the ﬂow ﬁeld for M a1 and for a Mach number M a2 , with 1 < M a2 < M a1

195

4.10 (a) Sketch the supersonic ﬂow ﬁeld about the body, shown below, including shocks, expansion fans, and streamlines.

(b) Sketch the supersonic ﬂow ﬁeld about the body, shown below, including shocks, expansion fans, and streamlines.

(b) Sketch the supersonic ﬂow ﬁeld about a ﬂat plate at angle of attack, (shocks, expansion fans, and streamlines).

4.11 (a) Sketch the supersonic ﬂow ﬁeld over the contour, shown below, including shocks, Mach lines, and streamlines!

(c) Sketch the ﬂow ﬁeld for the given wall contour with shocks, expansion fans, and streamlines for β > βmax .

(b) Sketch the supersonic ﬂow ﬁeld about a cambered plate, shown below, including shocks, Mach lines, and streamlines.

4.9 A plane supersonic jet is deﬂected with a baﬄeplate. Sketch the ﬂow ﬁeld with shocks, expansion fans and streamlines.

196

5. Exercises in Gasdynamics widened and the ﬂow be redirected, as sketched below. Use the linearized theory and compute the turning angle β2 and the Mach number M a2 !

5.1.5 Lift and Wave Drag – Small-Perturbation Theory 5.1 Consider the two-dimensional inviscid supersonic ﬂow with free-stream Mach number M a∞ about a double wedge proﬁle and a pair of proﬁles with the same thickness ratio d/l.

M a1 = 2;

δo = 10◦ ; Nm γ = 1.4; R = 287 kg K Hint (linearized theory): γ M a2 ∆p 1 ∆β = p1 M a2 − 1 1

δu = 5◦ ;

M a∞ = 2.5; β1 = 12◦ ; β2 = 15◦ ; l = 1 m; d = 0.24 m; Nm γ = 1.4; R = 287 kg K (a) Sketch a streamline of the ﬂow about the pair of proﬁles somewhere between 0 < y < h and indicate the variation of the static pressure along this streamline. (b) Sketch the ﬂow ﬁeld about the double-wedge proﬁle with shocks, expansion fans, and streamlines. (c) Sketch the static pressure distribution on the surface of the doublewedge proﬁle. (d) Determine the drag coeﬃcient cD of the double-wedge proﬁle and for the pair of proﬁles (approximately). cD = FD u2 b l ∞ b = wing span ρ 2

5.3 The contours of the upper and lower side, yu (x) and yl (x), respectively, of a thin airfoil proﬁle with h t are given. Derive the expressions for determining the lift and drag coeﬃcients cL and cD for supersonic free-stram conditions with the linearized theory! Show, that the lift coeﬃcient cL depends only on the free-stream Mach number and the angle of attack .

5.4 Air ﬂows through a two-dimensional cascade. The lower side of the blades are ﬂat, the contour of the upper side is given by the relation y = h cos( π x ). t t 2 t

5.2 Air ﬂows through a two-dimensional tunnel with a velocity twice the speed of sound. For reasons of reconstruction the cross section of the tunnel has to be

5.1 Problems

197

M a∞ =

√

2;

δ = D/l = 0.1

M a∞ = 3; p∞ = 105 Pa; γ = 1.4; h/t = 0.05 width of the blades: b = 1 m chord: t = 1 m Determine with the linearized theory: (a) the spacing s/t such that the blades do not interfere with each other, (b) the pressure coeﬃcients cpu = cpu (x/t) and cpl = cpl (x/t) for the upper and lower side of a blade and sketch their variation, (c) the forces Fx and Fy . 5.5 Use the linearized theory and compute the ratio of lift and drag coeﬃcient cL /cD for a proﬁle with angle of attack o and thickness distribution for supersonic ﬂow with M a∞ , as sketched below!

5.7 1.) Determine the drag coeﬃcient cD of the double-wedge proﬁle, sketched below, for inviscid, two-dimensional supersonic ﬂow without any approximation.

M a∞ = 3; β1 = 10◦ ; d = 0.10 m; γ = 1.4 t

β2 = 4◦ ;

2.) Determine with the linearized theory: (a) the pressure coeﬃcient cp1 on the front and rear part cp2 , and the drag coeﬃcient cD for M a∞ = 3; (b) the lift coeﬃcient cL and the drag coeﬃcient cD for an angle of attack = 5◦ (M a∞ = 3). 5.8 An airplane ﬂies in level ﬂight at a Mach number M a1 and an altitude H1 (static pressure p1 ). Lift and drag are generated by a wing with a thin doublewedge proﬁle at an angle of attack 1 and a thickness ratio δ. The airplane climbs to an altitude H2 with the static pressure p2 and ﬂies then again in level ﬂight.

Given: M a∞ ;

o;

a;

b;

t

5.6 Consider thin, two-dimensional proﬁles in supersonic ﬂow at M a∞ . Determine lift and drag coeﬃcient with the approximation for slender airfoils.

198

5. Exercises in Gasdynamics √ ◦ M a1 = 2; 1 = 2 ; p √1 ; p2 = 2 γ = 1.4 δ= d t

= 0.03;

Determine for the altitude H2 : (a) the angle of attack 2 , if the Mach number is kept constant (M a2 = M a1 ), (b) the Mach number M a2 , if the angle of attack remains unchanged ( 2 = 1 ; give reasons for the solution), (c) the ratio of drag and lift FD /FL for the wing, (d) the ratio of the necessary propulsive powers Pa /Pb (T∞ = const.) (Pa = propulsive power for M a2 = M a1 , see part a), Pb = propulsive power for 2 = 1 , see part b), (e) the ratio of the energies Wa /Wb necessary for the same route of ﬂight (notation same as under d). 5.1.6 Theory of Characteristics 6.1 (a) Under what angle γ are characteristics reﬂected at the edge of a jet? characteristic curves edge of free jet

6.3 In the ﬂow ﬁeld shown below the line OA is the ﬁrst, and the line OB the last charakteristic of an expansion fan.

p∞ = 105 Pa; | β |= 10◦ Nm γ = 1.4; R = 287 kg K Determine: (a) the pressure and the Mach number for the point C, (b) the slope of the streamline for the point E. M a∞ = 2; 6.4 In the ﬂow ﬁeld sketched below the line 01 is the ﬁrst, and the line 02 the last characteristic of the Prandtl-Meyer corner in the point 0.

α α

(b) Discuss, how the ﬂow direction is determined at the edge of a jet and how the Mach number is determined along a rigid wall. 6.2 A plane jet impinges on a concave wall (see sketch below). The ﬂow is assumed to be isentropic. The Mach number downstream from the characteristic line 4–5 is supposed to be constant. Determine the turning angle β3 and the Mach number M a3 for the point P3 !

| β |= 7◦ Nm γ = 1.4; R = 287 kg K Determine: (a) the direction of the ﬂow in the points 2 and 6, (b) the Mach numbers in the points 3 and 4, (c) the angle γ and the pressure ratio p4 /p∞ ! M a∞ = 1.7; 6.5 Air ﬂows through a plane symmetric nozzle, the contour consisting out of two circular sectors, isentropically into a large container (see the following sketch).

5.1 Problems

199

Determine for γ = 1.4 : (a) the Mach number M aA for the point A with the method of characteristics, (b) the ratio rA /r∗ (rA = distance 0−A). 6.7 Consider the ﬂow ﬁeld sketched below.

M a∞ = 1.34; p∞ = 105 Pa; ∆ p∞ = 0.98 · 105 Pa; Nm γ = 1.4; R = 287 kg K (a) State the diﬀerential equation describing the two characteristics as a function of the ﬂow angle β and the Mach angle α. (b) The region ABCEA is a simple region. Prove that all right-running characteristics in this region are straight lines! (c) Determine the Mach number at the edge of the jet and the angle γ between the edge of the jet and the wall! (d) Show that the ﬂow angle β downstream from the point D ﬁrst has to decrease! 6.6 Air ﬂows through a divergent twodimensional channel with straight walls. Along the circular segment with radius r∗ the Mach number is M a = 1. characteristic curves A 2β < 20°

M a∞ ;

β;

∆ Ma

Determine the Prandtl-Meyer angle and the ﬂow angle in the points 1 to 3. 6.8 A plane jet (M a∞ = 3) is symmetrically turned by a wedge with nose angle 2 β = 28◦ (see sketch below).

Determine β3 , M a3 , β4 , M a4 , and M ae ! 5.1.7 Compressible Potential Flows and Similarity Rules 7.1 Reﬂexion of Mach waves

0

Determine the pressure coeﬃcient cp (x,y) for 0 ≤ x ≤ 2 for the lower wall!

200

5. Exercises in Gasdynamics (a) Formulate the solution of the linearized potential equation for the regions I and II! (b) State the expression for the pressure coeﬃcient cp and the boundary condition at the edge of the jet for the region II! (c) Determine the angle δ as a funcrion of β! 7.5 A two-dimensional supersonic ﬂow is turned by a rigid wall as indicated in the sketch. The discontinuity surface, marked in the sketch, separates two ﬂow regimes of equal pressure and ﬂow direction, but with diﬀerent velocities.

7.2 Consider a supersonic ﬂow in a wind tunnel . The ﬂow is turned by the angle in the point A.

Draw the streamline through the point B, based on the linearized theory. 7.3 Supersonic ﬂow ﬂows through a cascade, consisting out of thin ﬂat plates at small angle of attack 0 .

Formulate: (a) the solutions for the linearized potential equation (M a2 − 1) ϕxx − ϕyy = 0 for the ∞ regions I, II, and III, (b) the boundary conditions for the potentials in the regions I and II! Determine: (c) the potential function ϕ(x,y) for the regions I, II, and III, (d) the magnitude of the perturbation velocities and the direction of the outﬂow in the region III! 7.4 Consider the ﬂow ﬁeld sketched below.

p1 ; δ; γ; M a1 ; M a2 ; M a1 > 1; M a2 > 1; M a 1 = M a 2 ; p4 = p6 . Determine the turning angle of the ﬂow β and the static pressure p6 with the linearized potential theory in the region 5. Hint: First formulate the solution for the perturbation potential and the boundary conditions in the regions 3, 4, and 5. 7.6 For the incompressible ﬂow about an elliptic proﬁle at zero angle of attack, the relation below describes the dependence of the maximum velocity along the contour, (Vmax ), on the free-stream velocity U∞ , and the normalized thickness δ: Vmax =1+δ U∞ .

Determine the free-stream Mach number of the compressible subsonic ﬂow, at which locally the speed of sound is attained on the surface of an ellipse with δ = 0.25; use the following diagram for the solution.

5.1 Problems

201

(c) Determine the lift coeﬃcients cLcomp. at M a∞ c = 0.4 and cainc. for incompressible ﬂow! 7.9 A model experiment is carried out in incompressible ﬂow in order to determine the pressure distribution of a compressible ﬂow over a cambered wall.

7.7 Consider a transonic ﬂow at M a1 = 0.95 about a wing section with the relative thickness δ1 = 0.05. The minimum pressure on the contour is measured to be cp1min = 0.3. Use the von K´rm´n a a similarity rule and compute the relative thickness δ2 and the pressure coeﬃcient cp2min of a second wing section for a free-stream Mach number M a2 = 0.9. 7.8 The critical Mach number of a thin wing section is M a∞crit. = 0.7. The lift coeﬃcient measured at this Mach number is cL = 0.3.

What similarity rule cp /cpinc. = f (M a∞ ) should be used for the application of the results of the measurements and what geometric dimensions Linc. /L and Hinc. /H are determined for the model? Use the solution of the linearized potential equation for compressible ﬂow and a sinosoidal wall: √ 2 π d/l cp = √ e−2 π y H 1−M a∞2 × 1 − M a∞2 2πx × sin , l with d = maximum amplitude. 7.10 1. According to the linearized theory for plane ﬂows the following similarity rule is valid for bodies of the same family: cp d/t = const., if = A A | M a2 − 1 | ∞ const., and A arbitrary.

p∞ = 105 Pa; γ = 1.4; Nm R = 287 kg K (a) Determine the lowest local pressure on the contour and the critical pressure coeﬃcient cpcrit. . (b) For what free-stream Mach number M a∞ b is the sketched pressure distribution measured with the same wing section?

Determine A in such a way, that the transformation of the body contour is identical with the coordinate transformation x = const; y | M a2 − 1 | = const. ∞

What similarity rule is obtained?

202

5. Exercises in Gasdynamics 2. An equivalent formulation of (1) is cp | M a2 − 1 | ∞ d/t = const. 4. How large is the lift coeﬃcient cL4 for the same geometry ( = const., δ = const.) in incompressible ﬂow? 5. How large do the ratios δ5inc. /δ and 5inc. / 3 have to be in incompressible ﬂow, so that cL5inc. is cL5inc. = cL3 ?

(a) Derive the corresponding similarity rule for the lift coeﬃcient cL ! (b) Show, that the relation d/t = const.

is valid for supersonic ﬂow, if designates the angle of attack! 3. Name three classes of ﬂows, fo which the above similarity rules are not valid! 7.11 (a) An airfoil with thickness ratio δa = 0.1 generates a lift coeﬃcient cL1 = 0.325 at M a∞1 = 0.6. At what free-stream Mach number M a∞2 is the lift coeﬃcient cL2 = 0.37? (b) How large is the thickness ratio δb of a wing of the same family for incompressible ﬂow, if its lift coeﬃcient cLinc. = cL1 for M a∞1 = 0.6? 7.12 An airplane generates its lift with a thin wing with a double-wedge proﬁle. It ﬂies steadily in level ﬂight.

A. Assume linearized plane supersonic ﬂow (M a > 1): 1. The lift coeﬃcient is ca1 = 0.1 at the free-stream Mach number M a∞1 = 1.5. Determine the angle of attack 1 of the wing! 2. The Mach number is increased to M a∞2 = 3. Determine the angle of attack 2 and the drag coeﬃcient cD2 ! B. Linearized plane subsonic ﬂow (M a∞ < 1): 3. How large does the lift coeﬃcient cL3 at M a∞3 = 0.75 have to be? How large is the drag coeﬃcient cD3 at this free-stream Mach number?

5.2 Solutions

203

5.2 Solutions
5.2.1 One-Dimensional Steady Flows of Gases 1.1 (a) Law of thermodynamics for steady processes h2 − h1 + u2 u2 2 − 1 2 2 + g (z2 − z1 ) = 0

(c) S a H cos α = S ∆t = 1.96 s ∆t = prior to ﬂying over the observer. 1.3 (a) cp T0 = cp T ∗ + a∗2 2

with: p1 p2 h2 − h1 = e2 − e1 + − ρ2 ρ1 (b) 1. ∆e ↔ ∆ g z + ∆ p u2 2 +∆ 2 ρ (b)

a∗2 = γ R T γR cp = γ−1 T∗ 2 = 0.833 = T0 γ+1

In compressible ﬂows internal energy is transformed into kinetic energy, potential energy, and compression work. 2. ∆e = 0 The internal energy is constant in incompressible ﬂuids, which, for thermally perfect gases (e = f (T )) implies a constant temperature. There is only an exchange between mechanical energies. 1.2 (a) Ma = √ (b) v = 2.0 γRT 1.4

p∗ T∗ = p0 T0 (c) Ma → ∞ cp T0 lim M a∗

γ γ−1

= 0.528

M a→∞

T →0 u2 max = 2 γ+1 = = 2.45 γ−1 :

L =

H tan α 1 Ma

α = arctan L = 1001 m

ρu = ⇒ p−(p + ∆ p) = − ⇒ 2 ρ u du + ⇒ u2

(ρ + ∆ ρ) (u + ∆ u) u dρ = −ρ du (ρ + ∆ ρ) (u + ∆ u)2 ρ u2 u2 dρ = −dp dp = dρ

204

5. Exercises in Gasdynamics h+ 1 u2 2 = (h + ∆h) + (u + ∆ u)2 2 2 ⇒ dh + udu = 0 dp T ds = dh − ρ dp with udu = − ρ ⇒ T ds = 0 ∂p ⇒ u2 = = a2 ∂ρ s 1.8 γ a∗2 γ u2 + RT = + RT ∗ 2 γ−1 2 γ−1 a2 1 γ + 1 ∗2 u2 + = a ⇒ 2 γ−1 2 γ−1 M a∗2 = ⇒ lim M a
M a→∞ ∗2

1+ γ+1 = γ−1

γ+1 γ−1 1 2 γ−1 M a2

1.5 b (vB − vA ) tan α M aB = 2 : α = 30◦ ∆t = 1.73 s ∆t =

A ρ∗ a∗ = ∗ A ρu

1.6

ρ∗ ρ A ⇒ ∗ A u2 cp T + 2 T∗ ⇒ T

= =

1 M a∗ a∗2 = cp T ∗ + 2 1 + γ−1 M a2 2 = 1 + γ−1 2

T∗ T T∗ T

1 γ−1

After problem 1.7 M a∗2 = A = A∗ · d(ρu2 ) = −dp − ρgdx d(ρu) = 0 1 dp du = − −g ⇒u dx ρ dx dp a2 = dρ du a2 dρ + = −g ⇒u dx ρ dx M a2 du = − g ⇒u dx M a2 − 1 with u > 0 dx > 0 it follows for a) M a < 1 du > 0 b) M a > 1 du < 0 (c) 1.7 h0 = u + cp T = const. 2 γ R cp = γ−1
2

γ+1 γ − 1 + M2a2 1 + γ−1 M a2 2 1 + γ−1 2 γ+1 γ − 1 + M2a2
1 γ−1

·
−1 2

1.9 (a) Θmax = = Θmax = (b) v(Θmax ) = a∗ = a0 T∗ = T 2 γ+1 ρ∗ a∗ ρ∗ = ρ0 a0 ρ 2 γ+1 2 γ+1
1 γ−1

T∗ T0 2 γ+1
1 2

γ+1 2(γ−1)

u2 γ max = cp T0 = RT0 2 γ−1 2 umax vmax = = a0 γ+1

5.2 Solutions 1.10 (a) ρ∗ a∗ A∗ = ρ∗ a∗ A∗ 1 1 1 2 2 2 ∗ ∗ T1 = T2 ⇒ a∗ = a∗ 1 2 A∗ p01 2 ⇒ ∗ = =2 p02 A1 (b) A∗ A∗ 2 = 2 1 ; M a1 < 1 A A ⇒ M a2 > M a1 (c) M a2max = 1 A∗ A∗ 1 1 = A max A∗ 2 1 = 2 ⇒ M a1max = 0.3 1.11 (a) Momentum: F = γpa M a2 A F pa ⇒ = γ M a2 Ap0 p0 A∗ 2 A incompressible F Ap0 ⇒ F F − Ap0 Ap0 = 2 ink 205

p0 − pa p0 F Ap0

= ∆ ink = 0( 2 ) Table of solution in the Appendix. 1.12 cp T0 = cp T + cp = max u2 2

T T0 ρ ρ0

γR γ−1 γ−1 M a2 + 1 = 2 1 T γ−1 = T0 = γ−1 M a2 + 1 2 γ T γ−1 T0 γ−1 M a2 + 1 2

−1

1 1−γ

p = p0 =

γ 1−γ

Energy: γ−1 T0 = 1+ M a2 T 2 T0 p0 = p T F 2γ ⇒ = Ap0 γ−1
⎡
γ−1 γ

1.13 (a)

pa p0
1 − γ−1

·
⎤

(b) A∗ AE pE 2. p0 TE 3. T0 ρE uE 4. ∗ ∗ ρ u 1. 1.14 ± ... m ˙ ρu = ρ 0 a0 A ρ0 a0 ρ = ρ0 p p0
1 γ

· (b)

pa ⎣ p0

− 1⎦

= 0.593; = 0.128; = 0.556; = A∗ = 0.593 AE

.

pa p0 − pa =1− =1− ; p0 p0 (1 − )− + γ−1 γ γ−1 γ

1 γ−1 γ
2

=1+ +1

γ−1 γ

2! F ⇒ = 2 + 0( 2 ) Ap0 p0 − pa + 0( 2 ) =2 p0

206

5. Exercises in Gasdynamics a2 a2 u2 0 + = 2 γ−1 γ−1 a 2 a2 0 1− ⇒ u2 = γ−1 a0 p 2 a2 ⎣ 0 = 1− γ−1 p0 m ˙ ⇒ = ρ 0 a0 A · critical state M ae = 1 ⇒ pe = 2 p∗ p0 = p0 γ−1 = 0.528 · 105 Pa γ−1 γ

⇒
2

m ˙ = const. = 0.579 ρ0 a0 Ae pc p∗ > p0 p0 m ˙ = ρ0 a0 Ae · pc p0
⎛ ⎝1 −
1 γ

⎡

γ−1 γ

⎤ ⎦

⇒

2 · γ−1 pc p0 γ−1 γ

p p0
⎛

1 γ

⎞ ⎠

2 · γ−1 p p0 γ−1 γ

⎞ ⎠

⎝1 −

with pc → p0 : M a → 0; ρ ≈ ρ0 2 ue ≈ (p0 − pc ) ρ pc m ˙ ∼ 1− ⇒ p0 ρ0 a0 Ae 1.15 (a) for: p0 = 7.824 · 105 Pa ⇒ ρ0 = (b) p0 Ma p∗ p0 ρ∗ ρ0 T∗ T0 = 7.824 · 105 P a; = M a∗ = 1 AH = A∗ ; p0 kg = 9.47 3 RT0 m

p0

(a) pc = 0.7 > 0.528 ⇒ subcritical p0 p0 m = ˙ γRT0 Ae RT0
⎛

pc p0

1 γ

·
⎞ ⎠

·

2 ⎝ pc 1− γ−1 p0

γ−1 γ

= 0.528 ⇒ pH = 4.13 · 103 P a = 0.634 ⇒ ρH = 6.00 kg m3

kg = 4.35 s (b) pc = 0.2 < 0.528 ⇒ supercritical p0 pe = p∗ m = ˙ 2 p0 γRT0 Ae RT0 γ+1 kg = 4.67 s
1 γ+1 2 γ−1

= 0.833 ⇒ TH = 240 K m s kg ∗ ∗ ∗ = ρ a A = 0.186 s = 105 Pa γRT ∗ = 310.5

u H = a∗ = ˙ m∗ ⇒

(c)

pc p∗ ≤ = 0.528 p0 p0

pE pE = 0.1278 p0 M aE = 2; M a∗ = 1.63 E ρE = 0.230 ρ0 kg ⇒ ρE = 2.18 3 m TE = 0.556 ⇒ TE = 160 K T0 m uE = M aE aE = 507.1 s kg ˙ ∗ = 0.186 mE = m ˙ s

5.2 Solutions (a) for: p0 = 1.064 · 105 Pa ⇒ ρ0 = 1.29 kg m3 (b) (T ds)P = 0 = dh − a2 dρ ρ 2 dρ dh = u ρ

207

P

(b) p0 is only slightly higher than pa ⇒ Subsonic ﬂow in the entire nozzle! The throat is no longer a special cross section! pE = 0.9398 p0 ∗ ⇒ M aE = 0.3; M aE = 0.326 AE : with isentropic relation p ργ ρE ρ0 = const., it follows:

⇒ (u2 − a2 ) p

dρ = 0 ρ P ⇒ Ma = 1

2nd and 3rd

= 0.956 ⇒ ρE = 1.23 kg m3

⇒ TE = 283 K m uE = 101.2 ; s kg m = 2.1 · 10−2 ˙ s AH = A∗ A∗ ρe u E f = = 0.492 AE ρ∗ a∗ f f A∗ f = 0.83 AH ⇒ M aH = 0.59; M a∗ = 0.62 H p0 = 1.26 pH ⇒ pH = 0.79 · 105 Pa Isentropic condition: ρH = 1.09 uH kg ; TH = 269.2 K m3 m kg ˙ = 194 ; mh = 2.1 · 10−2 s s

TE T0

= 0.982

1.17 (a) ρu = ρu2 + p = ⇒p+ (b) dp dρ =
P

const. const.

C1 = C2 ; ρ

C1 = (ρu)2

∂dp ∂dρ

= a2 P s dp = u2 P dρ P ⇒ M aP = 1 2nd and 3rd

1.16 (a) h0 = const. = h + u2 = C2 2

ρu = const. = C1 C1 ⇒ h + 2 = C2 2ρ

208

5. Exercises in Gasdynamics 2.4 1) Coordinates moving with shock: Index “s” p2 = 10.2 ⇒ M a1s = 3 p1 m u = M a1s γRT1 = 1042 s 2) (a) M a∗ = 2s ⇒ M a2s u1 u2 = a∗2 ∗2 a u1 m ⇒ u2 = 2 = = 192 u1 s M a∗ 2 1 T2 1 = 0.509 M a∗ 1s = 0.47 T2 = T1 = 795 K T1

5.2.2 Normal Compression Shock 2.1 Continuity equation: u1 ρ1 = u2 ρ2 Prandtl relation : u1 u2 = a∗2 ⇒ ⇒ ρ2 ρ1 ρ2 u1 u2 1 = = ∗2 = M a∗ 2 1 ρ1 u2 a γ+1 = M a∗ 2 = 1max γ−1 max

2.2

2.3 (a) u1 a1 p1 p01s ⇒ p01s T01s = 425 K; p2s T2s = 387 K; p02s T02s M a1 = M a∗ = 2s ⇒ M a2s = √ u1 = 1.5 γRT1
5

u2s = M a2s γRT2 m = 265.6 s u − u2s = 1.37 M a∞ = √ γRT1 (b) M a∞ = 1.37 ⇒ ⇒ T0∞ = 1104 K (c) p0∞ = p0∞ p2 = 30.9 · 105 Pa p2 T2 = 0.72 T0∞

= 0.272 = = = = 3.67 · 10 Pa 2.45 · 105 Pa 3.4 · 105 Pa T01s

1 = 0.7328 M a∗ 1s = 0.71

2.5 cp T1 + u2 u2 1 = cp T2 + 2 2 2 (γ + 1) M a2 u1 1 = u2 2 + (γ − 1) M a2 1 γR cp = γ−1 (γ − 1)(u2 − u2 ) 1 2 T2 − T1 = 2γR = 197.9 K

u2s = M a2s γRT2s m = 280 s (b) u1R = M a1R = 0 p01R = p1 ; T01R = T1 p2R = p2s ; T2R = T2s m u2R = u − u2s = 235 s ⇒ M a2R = 0.596 p2R ⇒ = 0.78 p02R ⇒ p02R = 3.14 · 105 Pa T2R = 0.94 ⇒ T02R = 399 K T02R

2.6 (a) A∗ A∗ AA AB p01 = 2 = 2 ∗ = 3.04 ∗ p02 A1 AB A1 AA (b) p01 = 3.04 p02 ⇒ Ma = 3 ⇒ p2 = 10 p1

5.2 Solutions (c) pB p01 p02 pB = = 2.76 pA p02 pA p01 2.7 m = ρ1 v1 A1 = ˙ γ p1 M a 1 A R T1 1.

209

FSdesign = AE ρE design × ×γM a2 design E = 13 N 2. Fspv2 = AE ρE2 uE2 u2 ρE2 = AE E2 u2 ρE2 u2 E1 ρE1 E1 1 Fspv2 = Fs1 = 4.2 N M a∗ 2 1 (e) for isentropic ﬂow pE = pK

2 ρ1 v1 A1 = (p0 − p1 ) A

M a1 =

1 γ

p0 −1 p1
2 v1 2

cp T0 = cp T1 + T1 = 1+ T0 γ−1 2 kg s

M a2 1

m = 1.41 ˙

2.8 (a) 2.9 (a) M aEdesign = 2.3 pE ⇒ pEdesign = p0 = 0.08 · 105 Pa p0 A∗ AH = = 0.456 AE AE (b) Subsonic ﬂow in the nozzle A∗ AE = 0.456 pE p0 = 0.95 · 105 Pa p0 = 0.28 pka p0Ea (b) see (a) AE ρE u2 = (pv − pE ) AE + Fs E AE ρE u2 = AE ρE γM a2 E E pv = pE m = 0.0242 ˙ A∗ = 0.77 Asb kg ; M aEb = 0.32; s p0 p0Ea = 1.387 A∗ = 0.593 Asa 1 = p∗ γA∗ M a∗ ∗ a p∗ = p0 γA∗ × p0 1 × ∗ γR T 0 To T = 0.0242 = 0.895 ⇒ M aEa = 0.4 kg s

⇒ M aEa = 2 ⇒ m = ρ∗ A∗ u∗ ˙

⇒ pv1 = M aE1 (c) pv2 = /d)

pv2 pEdesignl

pEdesignl = 0.48 · 105 Pa

210

5. Exercises in Gasdynamics (c) p01 ∆s ∆sa = ln ⇒ = 2.5 R p02 ∆sb (d) (c)

(d) 1. A v = A∗ = AH 2. ∆t = (e) AH As ⇒ M a∗ 2 a1 A∗ 2 AH A2 2 AE ⇒ M aE = = = = = AE AH = 0.507 As AE 1.71 ⇒ M a∗ 2 = 0.585 a2 p01 = 1.58 p02 A∗ AH 2 = 0.72 AH AE 0.47 pE p02 = 0.538 · 105 P a p02 ∆m m ˙ Vv Tv R

∆m = (pvmax − pv1 ) pkmax = 0.95 p0 ⇒ ∆m = 585 kg m = ρ∗ A∗ u∗ = ˙ kg s ⇒ ∆t = 24.22 s = 24.15 2.11

AH γ p 0 p 0 p γR T 0 To T
∗

∗

pE =

2.10 (a) 1. M aE = M aM = 2.3 pv1 pE p0 = 0.48 · 105 Pa pE p0

⇒ pv1 = (b)

AH AE A∗ = = 0.8 AS AS AE ⇒ M a∗ = 1.425 ⇒ M a∗ = 0.7 1 s ps p1 ps = p0 = 0.66 · 105 Pa p1 p0 2. A∗ 2 AE = p01 AH A∗ AH 2 = AH AE p02 AE 0.511 0.311 pk2 p02 p01 p02 p01 0.82 · 105 Pa

(a)
2 ρe ve AD = (pa − pe ) AD + Fs 2 ve = 2 cp (T0 − Te ) γR cp = γ−1

= ⇒ M a2E = pv2 = =

Fs p 0 AD

2 γ ρe γ − 1 ρ0 pa pe − + p0 p0 = ρ ρ0 γ−1 1−

Te T0 γ 1−γ

T = T0

=

p p0

5.2 Solutions Fs p0 AD pe γ 2γ γ − 1 p0 γ + 1 pe pa − − γ − 1 p0 p0 =
1

211

(c) AL AL AL AM ⇒ M aM AL AM pa − 0.02pa = 0.98 pa AL AL = = 0.944 AL AM = 0.76 = = 0.963 AL AL AM AL AM = 0.197 m2

subcritical : pe = pa supercritical : pe = p∗ = 0.528 p0 (see problem 1.3) (b)
2 Fs = ρe ve AD 2 (p0 − pa ) 2 ve = ρ Fs pa = 2 1− p0 AD p0

AL =

5.2.3 Oblique Compression Shock 3.1

pa p0

Fs p0 AD

a) compr. 0 0.66 1.07 1.27

b) incompr. 0 0.8 1.6 2

1 0.6 0.2 0

2.12 (a) A∗ AH A∗ = AL A∗ AE pE pa ⇒ pE (b) tmax = ∆m ; m ˙ Tc = T0 = Ta Vc mRTK ˙ kg s 3.2 (a) M a1 sin σ un1 σ = arctan ut1 V1 M a1 = a1 ⇒ T1 = 1.32 = 53.1◦ = 1.65 = 228.5 K = 0.963 = 0.193 m2 = 0.38 ⇒ M aE = 2.5 pE = 0.0585 p0 = 0.0585 · 105 Pa =

= (pEmax − pE (t = t0 ))

m = AL ρ∗ u∗ = 45 ˙ AL AE ⇒ M aE (t = tmax ) ⇒ pE (t = tmax ) ⇒ tmax

= 0.3793 = 0.23 = 0.96 · 105 Pa = 9.3 s

212

5. Exercises in Gasdynamics (b) ut2 = un1 M a∗ 1 M a2 m s = 1.45 ⇒ M a∗ = 1.16 2 = 1.2; β = 12◦ 3.4 (a) ut1 = 0.69; a∗ p0a p0b p0a pb = R ln pa p0b J ∆s = 88.62 kg K ∆s = R ln pu p0b − pE pE pE = 0.22 · 105 Pa ∗ M a0b = 1.55 ⇒ M a0b = 1.85 M a∗ = 0.92 ⇒ M a0b = 0.94 u pu − p0b =

M a1 sin σ = 1.32 ⇒ un2 = 259

(c) Shock polar un1 = 1.28; a∗ β = 15◦ a∗ = ⇒ un1 γR

T ∗ T0 m T1 = 344 T0 T1 s m m = 440 ; ut1 = 237.4 s s (b)

pa = pb ;

3.3 (a) ∆t = ∆m m ˙ kg s (T0a (c) ρa Tb = = 1.092 ρb Ta 3.5 1) (a) AH = 0.15 ⇒ M aEAusl. = 3.5 AE (b) pa(0) pE ⇒ M aE sin σ1 ⇒σ M a∗ E ⇒ M a∗ 1 = 5.13 = = = = | Va | / | Vb | = = M aa M ab Ta Tb Ta T0b T0a Tb

m = ρ∗ AH u∗ = 24.15 ˙ ∆m = pvmax

Vv (pkmax − pv (t = 0)) RTv = 0.48 · 105 Pa

M aa M ab = 1.063 = T0b )

(normal shock in exit cross section) ⇒ ∆m = 497.6 kg ⇒ ∆t = 20.6 s (b) pE = ⇒ pE p0 = 0.08 · 105 Pa p0

pv = 2 ⇒ M aE sin σ = 1.36 pE ⇒ σ = 36.25◦ ; β = 11.5◦ pv = 0.165 ⇒ M a2 = 1.83 p02 T2 T1 T0 = 167 K T2 = T1 T0 m V2 = M a2 a2 = 474 s

(c) βmax (M a = 2.3) ⇒ max β0b ⇒ M aE sin σ0b βu ⇒ M aE sin σu = = = = = = 27.5◦ 7.5◦ 12.5◦ ⇒ σ0b = 37◦ 1.384 27.5◦ ⇒ σu = 62◦ 2.03

2.12 37.3◦ 2.06 1.71 (β1 = 22◦ ) ⇒ M a1 = 2.2 β2 = β1 ⇒ σ2 = 53◦ M a∗ = 1.2 ⇒ M a2 = 1.25 2

5.2 Solutions (c) 3.7

213

2) (a) pa (z) = pEdesignl = 0.195 · 105 Pa z = × 1− pEdesignl pa (0)

γ−1 γ

×

3.8 ∆x = = M a∗ = 1 βB = ⇒ M a∗ A2 = a∗ = vA2 = = uA2 = = vB2 = = uB2 = = ⇒ 2∆y tan σA + tan σB 2.8 cm 1.63 ⇒ βA = 5.5◦ ; 21◦ 1.54; M a∗ = 1.13 B2 T∗ m T0 = 316.94 T0 s M a∗ a∗ sin βA A2 m 46.78 s M a∗ a∗ cos βA A2 m 485.84 s M a∗ a∗ sin βB B2 m 128.35 s M a∗ a∗ cos βB B2 m 334.35 s 1 ∇ x v = −7960 s γR

γ (γ − 1)a = 11.9 km

3.6 v1 ρ1 h1 h3 ⇒ h1 ρ2 ρ1 ⇒ M a1 sin σ1 ⇒ σ1 = 30◦ ; M a∗ 1 ⇒ β = 13◦ ; M a∗ 2 M a2 M a∗ = 1.56; σ 3 ⇒ M a2 sin σ2 ρ3 ρ2 ρ1 ρ3 h3 ⇒ h1 tan β = ⇒ h1 l = v3 ρ3 h3 M a ∗ ρ1 1 = M a ∗ ρ3 3 = 1.85 = = = = = = 1.5 1.964 1.78; 2.4 37◦ 1.44 ρ1 ρ2 = 0.309 ρ2 ρ3

= 1.75 =

= 0.389

h h1 1 − h3 h1 − h3 1 = 2l l 2 2 tan β = 0.756 = 1 − h3 h1

214

5. Exercises in Gasdynamics (b) P.-M. ﬂow = isentropic ﬂow ⇒ p02 = p03 = p04 ν3 = ν2 + | β2 | ν2 = 26.4◦ ⇒ ν3 = 38.4◦ ⇒ M a3 = 2.47 p3 p3 = p03 = 0.546 · 105 Pa p03 (c) u t = ap u2 t = 155 K ⇒ Tp = γR p4 p3 − ∆p = = 0.0273 p03 p04 ⇒ M a4 = 3; ν4 = 49.8◦ ⇒| β |= ν4 − ν3 = 11.4◦ (d)

5.2.4 Expansions and Compression Shocks 4.1 (a) r is a Mach line tan α = M ap (b) ut ⇒ α = 26.57◦ ur 1 = = 2.24 sin α

4.2 P.-M. ﬂow = plane, isentropic ﬂow ⇒ rot V ∂vr ∂Φ ∂vΦ ∂vr − ⇒ vΦ + r ∂r ∂Φ ∂vΦ = 0 and vΦ ∂r 4.3 ν3 = ν1 + 2 | β |; ν3 = 28◦ ; ν1 = 0◦ ⇒ β = 14◦ ν2 = ν1 + | β |= 14◦ ⇒ M a2 = 1.57; α2 = 39.5◦ Conti: a∗ h ρ∗ = AB u2 ρ2 sin α2 h A∗ = = 0.82 A AB sin α2 ⇒ AB = 0.192 m AC = 2 AB cos α2 = 0.296 m δAB = α2 − β = 25.5◦ δAC = β = 14◦ = 0 ∂ = (rvΦ ) ∂r = 0 = a⇒ ∂vr =a ∂Φ

4.5 (a) β1 = 10◦ ; σ1 = 40◦ ⇒ M a1 = 2 M a∗2 = 1.44; β2 = β1 ⇒ M a∗ = 1.2; M a3 = 1.26 3 p1 p1 = pa = 0.11 · 106 Pa pa

(b) β3 p4 p04 p03 M a2 p3 ⇒ p2 ⇒ p3 M a3 p3 ⇒ p03 p4 ⇒ p04 ⇒ M a4 ⇒ β3 = ν4 − ν3 p4 p3 = ; p3 p03 = p04 ; p4 = p2 = pu = 1.62; σ2 = 51◦ = 1.69 = 0.388 · 105 Pa = 1.26 = 0.38; = 0.225 = 1.63; = 10.6◦ ν4 = 15.7◦ ν3 = 5.1◦

4.4 (a) M a1 = 2.5; ⇒ σ = 34◦ ; M a2 = 2; β1 = 12◦ p02 = 0.96 p01

p02 p01 p1 = 8.87 · 105 Pa p02 = p01 p1 p2 p02 = 1.13 · 105 Pa p2 = p02

5.2 Solutions (c) pac p1 ⇒ M a1 sin σ1 ⇒ σ = 49◦ ; β1 M a2 ⇒ β1 > βmax (M a2 ) ⇒ Mach reﬂexion (d) (b) p3is = (c) M a1min (β1 = 15◦ ) = 1.62 (e) M a1 sin σ = 2 pad p1 = 0.5 · 105 Pa p1 4.7 (a) from shock polar pad =
Ma 1 2 3 4 3 2.21 1.59 0.67 M a∗ 1.96 1.72 1.42 0.71 β 16 16◦ 0◦
◦

215

2. = 2.5 = = = = 1.51 17.5◦ ; 1.32 7◦ ν3 = 0 ⇒ β2 β3 = β2 − β1 p3 p3 = p02 p03 m ˙ = ν2 − | β2 | = 33.02◦ = 18.02◦ = 9.36 · 105 Pa = u1 ρ1 b l tan σ1 kg = 232.4 s

p3is p01 = 10.48 · 105 Pa p01

σ 33.3 42.4◦ 90◦
◦

M a sin σ 1.65 1.49 1.59

p ˆ p

4.6 (a) 1. M a∗ 1 ⇒ M a∗ 2 ⇒ M a2 M a1 sin σ = = = = 1.96; β1 = 15◦ 1.73; σ = 33◦ 2.25; 1.634 p2 p1 p1 1.57 · 105 Pa T2 T1 T1 350 K T01 T1 = 700 K T1

3 2.41 2.77

p4 = (b)

p4 p3 p2 p1 = 10.03 · 105 Pa p3 p2 p1

p2 = = T2 = =

m = ρ4 u 4 A ˙ p4 = M a4 RT4 = p4 M a 4 = 169.9 1. p4is = mis ˙ kg s

γ RT4 A γ T1 T0 A RT1 T0 T4

T02 = T01 = p02 =

p02 p01 p1 p01 p1 = 17.71 · 105 Pa p01 p01 = p1 = 19.84 · 105 Pa p1 ∆s p01 = 0.1134 = ln R p02

p4 p0 p1 = 13.56 · 105 Pa p0 p1 p4is kg = m ˙ = 229.7 p4 s

216

5. Exercises in Gasdynamics 4.10 (a)

4.8 (a)

(b)

(b)

4.11 (a)

(c) (b)

4.9

5.2 Solutions 5.2.5 Lift and Wave Drag – Small-Perturbation Theory 5.1 (a) = 0.096 p1 = 2.1 p∞ p03 = p02 ; p3 p3 p01 p1 = p∞ p03 p1 p∞ = 0.345 5.2 p3 − p1 = K(−δo ); p1 γM a2 1 K = M a2 − 1 1 p2 − p3 = K(−(δo − β)) p1 p2 − p1 = K(β − 2δo ) ⇒ p1 p4 − p1 = Kδu p1 p2 − p4 = K(−(β − δu )) p1 p2 − p1 = K(2δu − β) ⇒ p1 ⇒ 2δu − β = β − 2δo ⇒ β = δo + δu = 15◦
⎡

217

(b)(c)

(d) Table of solution in the Appendix Pair of proﬁles cD = FD = ⇒ cD = 2FD γp∞ M a2 lb ∞ (p1 − p3 )db p1 p3 2 − γM a2 ty p∞ p∞ inf 0.016 2.1

p2 p1 ⎣ = 1+ p0 p0

γM a2 1 M a2 − 1 1

⎤

(δu − δo )⎦

= 0.0918 ⇒ M a2 = 2.22 5.3 d l cpu = cpl = − cL 1 = t = 2 M a2 − 1 ∞ 2 M a2 − 1 ∞ t 0

dy dx

u

= p1 = p∞ p03 = p02 ; p3 p3 p02 p2 p1 = p∞ p03 p2 p1 p∞ = 1.8 Double wedge

dy dx

l

(cpl − cpu )dx 4

M a2 − 1 ∞ cpu dy dx
⎡ ⎣ t 0

cD = =

1 t 2

t 0

+cpl − u dy dx

dx l ⎤ 2

FD = (p1 − p3 )db p1 p3 2 ⇒ cD = − γM a2 p∞ p∞ ∞

d l

1 2 −1 t M a∞

dy dx

2

+ u dy dx

⎦dx l 218

5. Exercises in Gasdynamics
⎡ ⎡ t 0 2 0

5.4 (a) h 1 s = + tan arcsin M a1 t t (b) cpu x t = = cpl x t = 2 dy M a2 − 1 dx u 1 h πx π sin 2t M a2 − 1 t 1 2 M a2 1 −1 h t = 0.4

1 × ⎣2 2 + t = 4

⎣

dy dx +

2

dy + dx u a2 + b 2

2

⎤

⎤

⎦ dx⎦ l 2

M a2 − 1 ∞ cL ⇒ = cD 4

π M a2 − 1 t ∞ 4o
2 0

+ (a2 + b2 )

π t

2

5.6 cL = cD =
⎡

4

0

M a2 − 1 ∞ 4 M a2 − 1 ∞
2 0

× d d(x) dx
2

⎤ ⎦

× ⎣

+

2 (x) s

+

cL1 = 0; 16 2 cD1 = δ 3

1 M a2 − 1 ∞

= 0.05

(c) Fy = cL b t = cL ⇒ Fy Fx ρ∞ 2 u 2 ∞

1 c L b t γ M a2 p ∞ 1 2 h 4t = 0.0707 = M a2 − 1 1 = 44.55 kN 1 c D b t γ M a2 p ∞ = 1 2 2 M a2 ∞
⎡
t 0

cL2 = 0; cD2 = 0.11 = 2cD1 cL3 = 0.21; 8 2 4 δ + cD3 = 2 −1 3 M a∞ = 0.12 5.7 (a) 1. cD = 1− × p1 p∞ p2 p1 M a1 ν2 p2 p1 ×

2 0

2 d p1 p∞ γ M a 2 t ∞

cD = ×

−1
2

×
2

= 2.08; = = = = = p2 p01 p02 p1 2.45 ν1 + | β1 + β2 | 52.9◦ 3.18

⎤

dy ⎣ dx

dy + dx u

x ⎦d t l

= 0.004 ⇒ Fx = 2.52 kN 5.5 cL = cD = 4
0

M a2 − 1 ∞ 2 × M a2 − 1 ∞

⇒ M a2 p2 = 0.024; p02 p1 = 0.0603 p01 p2 = 0.38 ⇒ cD = 0.0204 ⇒ p1

5.2 Solutions 2. cp1 = cp2 = 2 tan β1 M a2 − 1 ∞ 2 tan β2 M a2 − 1 ∞ = 0.125 = −0.0494 (e)

219

W = P t; s = u t Pa M a b Wa = = 1.11 ⇒ Wb M a a Pb 5.2.6 Theory of Characteristics 6.1 (a) characteristic curves edge of free jet

d cD = (cp1 − cp2 ) = 0.0174 t (b) cL = 4 M a2 − 1 ∞ 4 = 0.1237

α

2

α

cD = cD a) + = 0.0282 5.8 (a)

M a2 − 1 ∞ γ = 180◦ − 2α 1 α = arcsin Ma 4 M a2 ∞ −1 (b) Edge of free boundary:

FL = ρ∞ γ M a2 Aw ∞ = const. M a2 ∞

M a2 − 1 ∞ ⇒ p1 1 = p2 2 √ ⇒ 2 = 1 2 = 2.8◦ (b)
2

⇒ p∞

= const.

= ⇒ ⇒

1

8=0 √ ⇒ M a21.2 = 4 ± 8 ⇒ M a21 = 2.61 ∧ M a22 = 1.08

p2 = M a2 − 1 2 M a4 − 8 M a2 + 2 2

M a2 2

M a2 1 M a2 − 1 1

p1

ν1 − β1 = ν2 − β2 pu ⇒ ν2 po ⇒ β2 = ν2 − β1 − ν1 Rigid wall:

(transonic regime, for which √ M a2 − 1 = const. is not valid!) cL (c) FD cD δ2 + = = cL FL (d) P = u∞ FD ; δ2 + Pa = 2 Pb δ +
2 a 2 b 2

β2 is known ν2 = ν1 − β1 + β2 6.2 ν2 + β2 ν2 β2 ⇒ ν3 + β3 = = = = ν3 + β3 ν∞ β1 = 0 ν∞

FL = const.
3

M a2 − 1 M a3 b a M a2 − 1 M ab a

= 0.6

220

5. Exercises in Gasdynamics ν3 − β3 ν4 β4 = β5 ⇒ ν3 − β3 ⇒ ν3 β3 ν4 − β4 ν∞ β ν∞ − β β = ν∞ − 2 β = 2 = = = = 6.5 (a) dy dx = tan(β ± α)
Ch.

+ : left running characteristic − : right running characteristic (b) ν1 + β1 = ν2 + β2 ν∞ = ν1 − β1 ν∞ = ν2 − β2 ⇒ ν1 = ν2 ; β1 = β2 For arbitrary points 1, 2 in region ABCEA: it follows from ν1 = ν2 α1 = α2 . ⇒ dy dx = tan(β1 − α1 )
1−2

6.3 (a) ν∞ + βA = νc + βc ν∞ = 26.38◦ ; βc = 10◦ ; βA = 0 ⇒ νc = 36.38◦ ; M ac = 2.4 p3 p0 pc = p∞ = 0.535 · 105 Pa p0 p∞ (b) νD − βD = ν∞ − βE νD + βD = ν∞ + βE ν∞ + | β | −β = ν∞ − βB ⇒ βB = 2 β = −20◦ ⇒ νD = 16.38◦ ; βE = 0 dy =0 ⇒ dx E 6.4 (a) ν2 − β2 = ν∞ + | β | −β ν2 = ν∞ → β2 = −14◦ ν6 − β6 = ν4 − β4 ; β4 = β ν4 + β = ν∞ + β2 ⇒ ν4 = 10.81◦ ; β6 = 0 (b) ν3 + β3 = ν1 + β1 β1 = 0 ν1 = ν∞ , β3 = β ⇒ ν3 = 24.81◦ ⇒ M a3 = 1.94 ν4 + β4 = ν2 + β2 ⇒ ν4 = 10.81 ⇒ M a4 = 1.46 (c) γ = 180◦ − 2 α2 = 107.94◦ p4 p4 p0∞ = = 1.43 p∞ p0 4 p∞

= tan(β2 − α2 ) = const. (c) pF r. p∞ p∞ − ∆p = = 0.0066 p∞ p0 p0 ⇒ M aF r. = 4; νF r. = 65.78◦ νF r. = ν∞ + 90◦ − γ ⇒ γ = 31.5◦ (d) ν∞ = νF − βD ν2 − β2 = νF − βD2 ν2 = ν∞ + | β2 |= ν∞ − β2 −2 β2 = −βD2 + βD ⇒ βD2 = βD + 2 β2 < βD 6.6 (a) ν2 + (−10◦ ) = 0 ⇒ ν2 = 10◦ ◦ ν3 − 10 = ν2 − (−10◦ ) ⇒ ν3 = 30◦ ν4 = ν3 + 10◦ = 40◦ ⇒ M a4 = 2.54 (b) Continuity equation: A∗ u∗ ρ∗ = AA uA ρA A∗ (M aA = 2.54) = 0.36 A ⇒ rA = 2.8r∗

5.2 Solutions 6.7 ν2 = ν3 = ν∞ + | β | ν2 − β2 = ν∞ ⇒ β2 = ν3 − ν∞ =| β | ν0 = ν (M a∞ + 2∆M a) νu = ν (M a∞ + ∆M a) νu − βu = ν1 − β1 ; βu = ν∞ − νu νo + βo = ν1 + β1 ; βo = νo − ν∞ ⇒ ν1 = νo + νu − ν∞ β1 = νo − νu ν1 − β1 = ν3 − β3 ⇒ β3 =| β | +2 (ν∞ − νu ) 6.8 | β |= 14◦ ⇒ σ = 31.5◦ ⇒ M a1 sin σ = 1.57 p01 ⇒ = 1.09 p02 ∗ M a2 = 1.75; ν2 = 34◦ p3 = p1 ; p03 = p02 p3 p1 p01 = = 0.0299 p03 p01 p02 ⇒ M a3 = 2.94; ν3 = 48.5◦ ν3 + β3 = ν2 + β2 ⇒ β3 = −28.5◦ ν4 − β4 = ν3 − β3 β4 = −14◦ → ν4 = 63◦ ; M a4 = 3.8 ν5 + β5 = ν4 + β4 ν5 = ν3 → β5 = 0.5◦ νe − βe = ν5 − β5 ⇒ νe = 34◦ ; M ae = 2.29 = M a2 M a1 = 3; 5.2.7 Compressible Potential Flows and Similarity Rules 7.1 cp = −2 0 ≤ x ≤ 1; u = 0 ⇒ cp (0 ≤ x ≤ 1,y = 0) = 0 G1 : ϕ1 = g(x) + λy) = g(η) with λ = M a2 − 1 ∞ u u∞ ⇒g = = ⇒f = ⇒ ϕI = ϕII = ϕIII = B.C.: dy =− dx c dg ∂g =λ v = = − u∞ ∂y dη u∞ dg = − ⇒ dη λ v = u∞ G2 : ϕ2 = g(η) + f (ξ); B.C.: dg df ∂ϕ2 =λ −λ dξ ∂y dη dg u∞ df = =− ⇒ dξ dη λ 4 ⇒ cp2 = λ 4 cp (1 ≤ x ≤ 2; y = 0) = λ v = 0= 7.2 ξ = (x − λy)

221

7.3 (a) ϕI = g(x + λy) ϕII = f (x − λy) = f (x − λy) + g(x + λy) dy dx =

ϕII (b)

v = u∞ = (c) =

= c 1 ∂ϕII 1 ∂ϕI = u∞ ∂y u∞ ∂y

gλ u∞ u∞ (x + λy) + cI λ 1 f λ − u∞ u∞ − (x − λy) + cII λ u∞ (x + λy) + cI λ u∞ (x − λy) + cII − λ 2 u∞ y + cIII

222

5. Exercises in Gasdynamics (d) u = v = ∂ϕIII =0 ∂x ϕ4 = f4 (ξ4 ) + g4 (η4 ) ∂ϕ4 = v4 = β4 u1 ∂y dg4 df4 (−λ1 ) = λ1 + dξ4 dη4 g4 dg3 δu1 = = η4 dη3 λ1 df4 u1 ⇒ = (δ − β4 ) dξ4 λ1 cp4 = − = pF r. = p∞ ⇒ cpF r. = 0 ∂ϕII =0 ⇒ uF r = ∂x F r. (c) dfII dfI = dξ dξ vK = u∞ β = −λ vF r. = u∞ δ = −λ dgII dfII − dη dξ
F r.

∂ϕIII = 2 u∞ ∂y =2
GIII

dy dx Streamline in 7.4 (a)

ϕII (b)

ϕI = fI (x − λy) = fII (x − λy) + gII (x + λy)

= = cp5 = cp4 =

dfII dξ

cp5 = k p4 = cp4 M a2 = 1 =

∂ϕII = 0 ∂xF r. dfII dgII + ⇒ F r. = 0 dξ dη dfII ⇒ vF r. = −2 λ dξ ⇒δ = 2β 7.5 ϕ3 = g3 (η3 ); M a2 1 η3 = x + λ1 y; 7.6

β4 = ⇒β =

p5 γM a2 2 +1 = cp5 p1 2 M a2 = 1 − 1 −

2 ∂ϕ4 u1 ∂x 2 dg4 df4 − + u1 dη4 dξ4 2 δu1 u1 − + (δ − β4 ) u 1 λ1 λ1 2 − (2 δ − β4 ) λ1 2 ∂ϕ5 2 − = − β5 u2 ∂x λ2 p4 2 −1 γM a2 p1 1 p5 2 −1 γM a2 p1 2 p5 ⇒ cp4 M a2 = cp5 M a2 1 2 2M a2 1 − (2 δ − β4 ) λ1 2M a2 2 − β5 λ2 β5 = β 2 δM a2 1 λ 2 M a1 + M a2 λ1 2 2

2 δ γ M a2 M a2 2 1 M a2 M a2 − 1 + M a2 M a2 − 1 1 2 2 2

λ1 = −1 dy ∂ϕ3 u1 = v3 = ∂y dx dg3 = δu1 = λ1 dη3 ∂ϕ5 dg5 λ2 = β5 u2 = ∂y dη5

cp =

ρ∞ 2 ρ∞ V + pmin incomp.: u2 + p∞ = 2 max 2 ∞ ρ∞ 2 ρ∞ 2 u + p∞ = u (1 + δ)2 +pmin ⇒ 2 ∞ 2 ∞ ⇒ cpmin, inc. = 1 − (1 + δ)2 = −0.5626

p − p∞ ; ρ 2 u 2 ∞

Vmax = u∞ (1 + δ)

5.2 Solutions cpmin,comp. = cpmin,inc. 1 − M a2 ∞ (c) cL = → in diagram → intersection with curve for cpcrit. ⇒ M a∞,crit. = 0.7 1 t (cpl − cpu )dx t 0 ⇒ cLc 1 − M a2 ∞c

223

⇒ cLc 7.9

= cLa 1 − M a2 ∞a = 0.2246; cLinc. = 0.214

cp = f (M a∞ ) cpinc. x x = l l inc y y 1 = l l inc 1 − M a2 ∞ H 1 L = 1; = ⇒ Linc. Hinc. 1 − M a2 ∞ cp 1 = cpinc. 1 − M a2 ∞ 7.10 1. 7.7 δ1 M a2 1 (1 − M a2 ) 1
3 2

x = const. ⇒ t = const. y | M a2 − 1 | = const. ∞ (1 − M a2 ) 2 2 ⇒ δ2 = 0.15 M a2 2 = cp2 1 − M a2 2 ⇒ cp2 = 0.651 = δ2 M a2 2
3

→d

M a2 1 cp1 1 − M a2 1

| M a2 − 1 | = const. ∞ 1 ⇒A= | M a2 − 1 | ∞ ⇒ cp | M a2 − 1 |= const.; ∞

G¨thert rule o 2. (a) cL = 1 t t 0

7.8 (a) pmin = p∗ = p p0 p∞ p0 p∞ 5 = 0.733 · 10 Pa p∗ 2 = −1 γ M a∞crit. p∞ = −0.778
∗

(cpl − cpu )dt = const.

⇒ (b)

cL

| M a2 − 1 | ∞ d/t 4

cpcrit.

cL = ⇒

| M a2 − 1 | ∞ d/t = const.

(b) cpbmin = −0.65 from diagram cpbmin 1 − M a2 ∞b = cpamin 1 − M a2 ∞a ⇒ M a∞b = 0.519 7.11 (a) cL

(c) transonic ﬂows hypersonic ﬂows

| M a2 − 1 | = const. ∞ ⇒ M a∞2 = 0.711

224

5. Exercises in Gasdynamics (b) cL ⇒ δinc. = | 1 − M a2 | ∞ δ δ1 | M a2 − 1 | ∞ = const. cL1 = cLinc. = 0.125

7.12 (a) 1. cL = 4 M a2 − 1 ∞

→ = 0.028=1.6◦ ˆ 2. FL = const. (T∞ ,p∞ = const. ) ⇒ cL1 M a2 = cL2 M a2 ∞1 ∞2 ⇒ cL2 = 0.025 → 2 = 0.0177=1.013◦ ˆ 4 cD = × M a2 − 1 ∞
⎡

×⎣

d l

2

⎤

+

2⎦

= 0.0027 (b) 3. cL3 M a2 = cL1 M a2 ∞3 ∞1 ⇒ cL3 = 0.4; cD3 = 0! 4. cL4 = cL3 1 − M a2 = 0.265 ∞3 5. 1 − M a2 cL5 ∞3 = cL3 δ5 δ 1 δ5 = ⇒ = 1.512 δ 1 − M a2 ∞3 ⇒ δ5 = 0.06 1 − M a2 ∞3 = 0.042 = 2.4◦ ˆ =
3

5.3 Appendix

225

5.3 Appendix
Tables and Diagrams Table of solution problem 1.11
∆
F Ap0 F Ap0

Ma 0 0.05 0.10 0.15 0.20 0.30 0.40 0.60 0.80 1.00 0 1.8 · 10−3 7.0 · 10−3 1.6 · 10−2 2.8 · 10−2 6.1 · 10−2 1.0 · 10−1 2.2 · 10−1 3.4 · 10−1 4.7 · 10−1

2

F Ap0

F Ap0 inc

∆

F Ap0

[%]

0 3.1 · 10−6 4.9 · 10−5 2.4 · 10−4 7.6 · 10−4 3.7 · 10−3 1.1 · 10−2 4.7 · 10−2 1.2 · 10−1 2.2 · 10−1

0 3.496 · 10−3 1.390 · 10−2 3.101 · 10−2 5.446 · 10−2 1.184 · 10−1 2.006 · 10−1 3.951 · 10−1 5.878 · 10−1 7.396 · 10−1

0 3.496 · 10−3 1.394 · 10−2 3.118 · 10−2 5.501 · 10−2 1.211 · 10−1 2.088 · 10−1 4.320 · 10−1 6.879 · 10−1 9.434 · 10−1

0 −2.2 · 10−6 −3.5 · 10−5 −1.7 · 10−4 −5.5 · 10−4 −2.7 · 10−3 −8.2 · 10−3 −3.7 · 10−2 −1.0 · 10−1 −2.0 · 10−1

-0.06 -0.25 -0.56 -1.00 -2.30 -4.10 -9.30 -22.00 -28.00

Table of solution 5.1 (d) Pair of proﬁles

Ma ∞ 1 2 3 2.5 2.0 1.56 2.1

M a∗ 1.83 1.63 1.40 1.68

β 12◦ 12◦ 15◦

σ 34◦ 42◦

ν

M a sin σ 1.39 1.33

pi p0i

pi pi−1

p0i p0i−1

13◦ 28◦

0.059 0.128 2.1 0.250 1.9 0.113 0.452

0.96 0.97 1

Table of solution 5.1 (d) Double wedge proﬁle

Ma 1 3 2.0 3.2

M a∗ 1.63 2.01

β 12◦

ν 26◦ 53◦

pi p0i

pi pi−1

p0i p0i−1

0.128 2.1 0.021 0.164

0.96 1

226

5. Exercises in Gasdynamics

Gasdynamic quantities of the Laval-nozzle ﬂow (γ = 1.4)

5.3 Appendix

227

Gasdynamic quantities for compression shock

228

5. Exercises in Gasdynamics

Oblique compression shock

5.3 Appendix

229

Prandtl-Meyer-Function ν(M a∗ ) ν(M a∗ ) = γ+1 · arctan γ−1 1 γ−1 (M a∗ 2 − 1) 2 γ−1 − 2 (M a∗ 2 − 1)

1

− arctan ν(M a) =

γ+1 (M a∗ 2 − 1) 2 γ−1 − 2 (M a∗ 2 − 1)

γ+1 γ−1 · arctan (M a2 − 1) γ−1 γ+1 √ − arctan M a2 − 1

230

5. Exercises in Gasdynamics

Oblique compression shock: Heart-curve diagram

5.3 Appendix One-dimensional nozzle ﬂow Perfect gas with const. spec. heats, γ = 1.4 Ma 0.00 0.01 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00 1.10 1.20 1.30 1.40 1.50 1.60 1.70 1.80 1.90 2.00 2.10 2.20 2.30 2.40 2.50 2.60 2.70 2.80 2.90 3.00 3.50 4.00 4.50 5.00 6.00 7.00 8.00 9.00 10.00 M a∗ 0.00000 0.01095 0.10944 0.21822 0.32572 0.43133 0.53452 0.63481 0.73179 0.82514 0.91460 1.00000 1.08124 1.23114 1.29987 1.29987 1.36458 1.42539 1.48247 1.53598 1.58609 1.63299 1.67687 1.71791 1.75629 1.79218 1.82574 1.85714 1.88653 1.91404 1.93981 1.96396 2.06419 2.13809 2.19360 2.23607 2.29528 2.33333 2.35907 2.37722 2.39046 p p0 ρ ρ0 T T0 A∗ A

231

1.0000000 0.9999300 0.9930314 0.9724967 0.9394697 0.8956144 0.8430192 0.7840040 0.7209279 0.6560216 0.5912601 0.5282818 0.4683542 0.4123770 0.3609139 0.3142409 0.2724031 0.2352712 0.2025935 0.1740403 0.1490396 0.1278045 0.1093532 0.0935217 0.0799726 0.0683997 0.0585277 0.0501152 0.0429500 0.0368483 0.0316515 0.0272237 0.0131109 0.0065861 0.0034553 0.0018900 0.0006334 0.0002416 0.0001024 0.0000474 0.0000236

1.000000 0.999950 0.995017 0.980277 0.956380 0.924274 0.885170 0.840452 0.791579 0.739992 0.687044 0.633939 0.581696 0.531142 0.482903 0.437423 0.394984 0.355730 0.319693 0.286818 0.256991 0.230048 0.205803 0.184051 0.164584 0.147195 0.131687 0.117871 0.105571 0.097626 0.084889 0.076226 0.045233 0.027662 0.017446 0.011340 0.005194 0.002609 0.001414 0.000815 0.00495

1.00000 0.99998 0.99800 0.99206 0.98232 0.96899 0.95238 0.93284 0.91075 0.88652 0.86059 0.83333 0.80515 0.77640 0.74738 0.71839 0.68966 0.66138 0.63371 0.60680 0.58072 0.55556 0.53135 0.50813 0.48591 0.46468 0.44444 0.42517 0.40683 0.38941 0.37286 0.35714 0.28986 0.23810 0.19802 0.16667 0.12195 0.09259 0.07246 0.05814 0.04762

0.000000 0.017279 0.171767 0.337437 0.491385 0.628875 0.746356 0.841610 0.913766 0.963178 0.991215 1.000000 0.992137 0.910459 0.937818 0.896921 0.850219 0.799850 0.747604 0.694936 0.642981 0.592593 0.544383 0.498759 0.455969 0.416129 0.379259 0.345307 0.314168 0.285704 0.416129 0.236152 0.147284 0.093294 0.060378 0.040000 0.018804 0.009602 0.005260 0.003056 0.001866

232

5. Exercises in Gasdynamics Prandtl-Meyer angle (ν in ◦ , γ = 1.4) Ma 1.00 1.02 1.04 1.06 1.08 1.10 1.12 1.14 1.16 1.18 1.20 1.22 1.24 1.26 1.28 1.30 1.32 1.34 1.36 1.38 1.40 1.42 1.44 1.46 1.48 1.50 1.52 1.54 1.56 1.58 ν 0.00 0.13 0.35 0.64 0.97 1.34 1.73 2.16 2.61 3.07 3.56 4.03 4.57 5.09 5.63 6.17 6.76 7.28 7.84 8.41 8.99 9.56 10.15 10.73 11.32 11.90 12.49 13.09 13.68 14.27 Ma 1.60 1.62 1.64 1.66 1.68 1.70 1.72 1.74 1.78 1.80 1.82 1.84 1.86 1.88 1.90 1.92 1.94 1.96 1.98 2.00 2.05 2.10 2.15 2.20 2.25 2.30 2.35 2.40 2.45 2.50 ν 14.48 15.45 16.04 16.64 17.22 17.81 18.40 18.98 19.56 20.15 20.72 21.30 22.45 23.02 23.59 24.15 24.71 25.27 25.83 26.38 27.75 29.10 30.42 31.73 33.02 34.28 35.52 36.75 37.94 39.15 Ma 2.55 2.60 2.65 2.70 2.75 2.80 2.85 2.90 2.95 3.00 3.05 3.10 3.15 3.20 3.25 3.30 3.35 3.40 3.45 3.50 3.55 3.60 3.65 3.70 3.75 3.80 3.85 3.90 3.95 4.00 ν 40.28 41.41 42.53 43.62 44.69 45.74 49.78 47.79 48.78 49.76 50.71 51.65 52.57 53.47 54.35 55.22 56.07 56.91 57.72 58.53 59.32 30.09 60.85 61.59 62.35 63.04 63.75 64.44 65.12 65.78 Ma 4.10 4.20 4.30 4.40 4.50 4.60 4.70 4.80 4.90 5.00 5.10 5.20 5.30 5.40 5.50 5.60 5.70 5.80 5.90 6.00 6.10 6.20 6.30 6.40 6.50 6.60 6.70 6.80 6.90 7.00 ν 67.08 68.33 69.54 70.70 71.83 72.92 73.97 74.98 75.97 76.92 77.84 78.73 79.59 80.43 81.24 82.03 82.79 83.54 84.26 84.95 85.63 86.29 86.94 87.56 88.17 88.76 89.33 89.89 90.44 90.97 Ma ν

7.10 91.49 7.20 92.00 7.30 92.49 7.40 92.97 7.50 93.44 7.60 93.90 7.70 94.34 7.80 94.76 7.90 95.21 8.00 95.62 8.50 97.57 9.00 99.32 9.50 100.98 10.00 102.31 11.00 104.79 12.00 106.88 13.00 108.65 14.00 110.18 15.00 111.51 16.00 112.64 17.00 113.71 18.00 114.63 19.00 115.43 20.00 116.19 21.00 116.87 22.00 117.48 23.00 118.04 24.00 118.55 25.00 119.03 ∞ 130.45

6. Aerodynamics Laboratory

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o
Abstract The wind tunnel for low speeds of the Aerodynamisches Institut of the RWTH Aachen is a G¨ttingen-type wind tunnel. It is a continuously running closed-circuit tunnel, in which the o the air is circulated with a single-stage axial-ﬂow fan with a maximum power of 100 kW. The test section is open and has a diameter of 1.2 m. The maximum attainable air speed is 60 m/s (216 km/h). Generally a wind tunnel has to be built in such a way that: 1. an air ﬂow is generated in the test section with constant velocity distribution in time and space, 2. the losses caused by the closed-circuit circulation are as small as possible. The objective of this experiment is to examine the quality of this wind tunnel. In order to do so, the velocity distribution in the test section is measured with a Prandtl tube. The dependence of the accuracy of the measurement on the free-stream direction is tested by turning the Prandtl tube with respect to the direction of the oncoming ﬂow. Also, the quality coeﬃcient of the wind tunnel (ratio of the power of the air stream and the power of the axial-ﬂow fan) is determined as a function of the ﬂow velocity. 6.1.1 Preliminary Remarks The prerequisits for the usability of wind tunnels for ﬂuid-mechanical investigations are the following: 1. The similarity laws can be applied, so that the results obtained with models in the wind tunnel can be transfered to ﬂows about full-scale conﬁgurations, especially to ﬂows about airplanes in free ﬂight. 2. The boundary conditions for the ﬂows about the models in the wind tunnel and the full-scale conﬁgurations, especially for those in free ﬂight, have to be similar. In general, complete similarity cannot be attained in a single experiment. For this reason, it is necessary to determine the similarity parameter, which is of greatest importance for the problem to be investigated, and only obey the corresponding similarity law in the experiment. For steady low speeds, it is suﬃcient to keep Re = vlρ Inertia force = µ Frictional force (6.1)

constant. In order to obey the similarity law, it is necessary to provide similar initial and boundary conditions, e. g.: 1. the ratio of the width of the air stream in the wind tunnel and the size of the model compared to the inﬁnite width of the air stream in free ﬂight, 2. the roughness of the surface of the model, 3. the intensity of turbulence of the ﬂow in the wind tunnel, and others.

234

6. Aerodynamics Laboratory

The ﬂow velocity of the G¨ttingen-type tunnel of the Aerodynamisches Institut is small como pared to the speed of sound. The inﬂuence of the compressibility of the air on the ﬂow can therefore not be investigated. 6.1.2 Wind Tunnels for Low Speeds Wind tunnels for low speeds (with air as working medium at atmospheric temperatures: v ≤ 80 m/s) are continuously operated, i. e. a fan generates a ﬂow for an arbitrary duration of the test. Two basic types of tunnels are known: 1. The G¨ttingen-type (after Prandtl) wind tunnel: Tunnel circulating the air with either open o or closed test section. 2. The Eiﬀel-type wind tunnel : Tunnel without air circulation with open or closed test section. The Eiﬀel-type tunnel is brieﬂy described next. Eiﬀel-Type Wind Tunnel The Eiﬀel tunnel in its simplest design consists of a tube with a fan. Air is sucked from the surroundings; Eiﬀel tunnels are used for determining wind loads on buildings (low construction costs for large dimensions; the air speed is low – only up to 40 m/s). Most of the available tunnels are built at higher expenditures. In order to generate a uniform velocity proﬁle with low turbulent intensity in the test section, a nozzle with a contraction ratio of Ac /An = 5 ÷ 20 is necessary. A further improvement of the velocity distribution is possible by positioning guiding vanes and screens upstream of the test section: A high contraction yields a relatively low velocity in the settling chamber and small losses caused by installations. In order to keep the operating costs low, diﬀusers are used for pressure recovery (in a simple wind tunnel the entire dynamic pressure of the test section ρ v 2 /2 must be counted as loss). The pressure in the measuring cross section is below atmospheric. An underpressure chamber is needed for an open test section; the air ﬂows through a diﬀuser back into the open air. A recirculation is possible inside of a hangar, and testing is then independent of the weather coditions. Most of the time the fan is mounted at the end of the wind tunnel. A long diﬀuser enables a high pressure recovery, but this is also associated with higher construction costs; in very long diﬀusers the losses due to wall friction are increased. The advantages of an Eiﬀel-type tunnel are its simple construction and its low construction costs. A high quality coeﬃcient can only be obtained, if a reﬁned diﬀuser with large opening angle – up to 35o – and internal guide vanes are used, in order to avoid ﬂow separation. The guide vanes can cause fouling of the tunnel. The disadvantage of the Eiﬀel-type tunnel is given by the necessity, that with an open test section the air stream has to be enclosed by an underpressure chamber. The advantage of low construction costs is then at least in part outweighed by a lower quality coeﬃcient and higher operating costs. G¨ttingen-Type Wind Tunnel o In wind tunnels of the G¨ttingen-type after Prandtl the ﬂow medium is circulated from the o measuring cross section back to the settling chamber with the aid of diﬀusers and guide vanes. The optimal widening of the diﬀusers (αopt /2 = 3 ÷ 4◦ ) requires a minimum length for a given cross section of the test section and a prescribed contration ratio. The fan is to be installed as far away as posible from the test section, so that the disturbances caused by spiralling of the ﬂow and non-uniformities caused by the hub, have space enough to decay. The advantages of this type of tunnel are the following: If the tunnel has an open test section, the pressure in the air stream is atmospheric and the test section is easily accessible. The quality

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o

235

coeﬃcient - in particlar with a closed test section - can attain maximum values. The operation is independent of the conditions in the surroundings, as temperature and humidity of the air. The disadvantages are, that the construction costs are high, and a minimum space is needed; if foreign substances are added to the air, e. g. smoke for ﬂow visualization, fouling has to be controlled. 6.1.3 Charakteristic Data of a Wind Tunnel The power of the fan Pef f needed for a continuously operated wind tunnel is proportional to the power of the air stream PStr in the test section. 3 The power of the air stream is PStr = m v∞ /2 = ρ AStr v∞ /2 (The subscript “∞” denotes the ˙ 2 test section). The constant of proportionality is called the quality coeﬃcient. It is deﬁned as ξ= PStr Pef f . (6.2)

The quality coeﬃcient is not the same as an eﬃciency. In general it exceeds unity and is the higher the smaller the mechanical losses of the wind tunnel are. The electric power needed is usually smaller than the power of the air stream. Pef f = 1ρ 3 AStr v∞ ξ2 (6.3)

Since the power of the air stream is proportional to the third power of v∞ , the power of the fan can increase rapidly to large values. The required power is given in the following table. The values are calculated for a wind tunnel with a cross section of the test section of AStr = 1 m2 with atmospheric conditions inside (pressure, density, and temperature). v∞ [ m/s ] M a[−] PStr [ kW ] 80 0.24 314 340 1.0 2.5 104 680 2.0 2.0 106 With these data it can easily be estimated, that a G¨ttingen-type wind tunnel needs its own o power plant for high subsonic and supersonic velocities to satisfy its power requirements. The power of the air stream is also proportional to the measuring cross section AStr . It is reasonable to assume, that the diameter of the test section is approximately equal to the length of the model, as the following example shows. If the Reynolds similarity law is to be obeyed, one obtains for the length of the models considered: Original e. g. Modell e. g. Automobile Airplane Automobile Airplane l = 5 m; v = 100 km/h; Re ≈ 107 l = 20 m; v = 300 km/h; H = 3 km; Re ≈ 108 l = Re µ/ρ v = 1.85 m l = 18.5 m

For the above examples it is assumed, that these Reynolds numbers can be attained at atmospheric conditions and at most at a maximum speed of v∞ = 80 m/s in the model tests. Wind tunnels for low speeds therefore have to be as large as possible, so that the similarity laws can be obeyed. The exceptions to be mentioned here are ﬂows about bodies, of which the force coeﬃcient does not strongly depend on the Reynolds number. Power Balance of a G¨ttingen-Type Wind Tunnel o The power to be provides by the fan is: ˙ Pef f = Q ∆ptot. = AStr v∞ ∆ptot. (6.4)

236

6. Aerodynamics Laboratory

˙ Q = Volume rate of ﬂow ∆ptot. = Diﬀerence of the total pressure up- and downstream from the fan ρ 2 ∆ptot. = ∆pLoss = ζi vi . 2

(6.5)

The subscript “i” refers to the corresponding components causing the losses, e.g. diﬀuser, turning corner, etc. To be able to compare, all loss coeﬃcients are referenced to the dynamic 2 pressure ρ v∞ /2 in the test section. ∆pLoss = The quality coeﬃcient is ξ= 1 ζi∞ (6.7) ρ 2 ζi∞ v∞ 2 (6.6)

Losses of Mechanical Energy in a G¨ttingen-Type Wind Tunnel o The characteristic orders of magnitudes of the losses of mechanical energy in a G¨ttingen-type o wind tunnel are listed in the following table: Nr. compare. page 240 construction part 9 Test section open (Test section closed) 10 Collector (friction) 11 1. Diﬀuser 12 1. Turning corner 1 Fan 2 2. Turning corner 3 2. Diﬀuser 4 3. Turning corner 5 4. Turnng corner 6 Screens a. Straightener 7 Nozzle (friction) 8 Edge of nozzle Friction (λ = 0.01) ζi 0.11 (0.016) s.u. 0.06 0.15 0.05 0.15 0.06 0.15 0.15 3.0+0.2 s. b. ζi∞ 0.11 Percentage % 33

0.06 0.03 0.007 0.03 0.025 0.006 0.006 0.04

18 9 2 9 7.5 1.75 1.75 12

0.02 = 0.334

6 = 100%

9 Test section: The air stream in an open test section causes much larger losses due to the mixing at its edge, than in a closed test section, which causes only wall friction. 10 Collector: Variable Geometry or slots guarantee proper ouﬂow for diﬀerent volume ﬂow. 11, 3 Diﬀuser: To avoid separation an opening angle of α/2 = 3 ÷ 4◦ cannot be exceeded; this requirement may result in large construction lengths. Short diﬀusers with large opening angles are also built; ﬂow separation is avoided by guide vanes. 12, 2, 4, 5 Turning corners: For structural reasons, most of the time they are built as proﬁled cascades; curved plates, used as cambered surfaces can bring about substantial improvements. 1 Fan: The size of the hub, the motor can be mounted in, also depends on structural and acoustic conditions; disturbances due to swirl in the ﬂow are compensated by guide wheels. Instead of a large fan, often several smaller fans are used for better controllability. Adjustable blades are also installed, in order to attain a optimal eﬃciency at partial load.

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o

237

6 Screens and straighteners: Several screens with diﬀerent mesh size exhibit a more favorable ratio of the damping of the turbulent ﬂuctuations and the drag than a comparable single screen. Straighteners often are built out of honeycomb proﬁled cascades. Empirically determined standard values are available for the ratios of the spacing, diameter of the tunnel, and the chord of the cascade. 7 Nozzle: The contraction ratio and the contour are the most important quantities that determine the ﬂow uniformity in the test section. Small winglets, also called Seiferth wings, are mounted near the edge of the nozzle to avoid the formation of vortex rings and velocity oscillations in the air stream. Friction: The inner walls of the tunnel should be as smooth as possible; steps between the various parts of the tunnel should be avoided. For the example given in the table there results ξ= 1 1 =3 = ζi∞ 0.334

The electric power needed for this tunnel with AStr = 1 m2 and atmospheric conditions of the air in the stream would be Pel = 314 kW PStr = 140 kW = ξ ηel ηf an 3 · 0.94 · 0.8

6.1.4 Method of Test and Measuring Technique For determining the characteristic data of the tunnel the following measurements were carried out in the experiment: 1. Measurement of the velocity distribution in the measuring cross section at a ﬁxed distance from the edge of the nozzle. 2. Investigation of the directional sensitivity of the Prandtl tube. 3. Determination of the quality coeﬃcient of the wind tunnel for several wind speeds. Measurement of the Velocity with a Prandtl Tube The diﬀerence between the total pressure and the static pressure at free-stream conditions is measured at two points, the stagnation point and a point on the wall, parallel to the direction of the free stream (sketch see page 241) ρ 2 ptot. − p∞ = v∞ = q∞ 2 . (6.8)

In the experiment the pressure diﬀerence q∞ is measured with a liquid manometer. In order to obtain v∞ the density ρ has to be determined ρ= Ns2 p Ba g · 13.6 = RT RT mm Hg m3 . (6.9)

The barometric pressure Ba [mm Hg] and the temperature T [K] are measured. The gas constant of air is for 1 bar and 273 K R = 287 m2 s2 K .

238

6. Aerodynamics Laboratory

Liquid Manometer (U-Tube Prinziple) Betz-Manometer: In comparison to the simple U-tube the Betz-Manometer (sketched on page 241) has one narrow stem (measuring tube) and a wide stem (pot). Principle: ha = H A h+H = h h/h = (6.10) (6.11)

1 (6.12) 1 + a/A The data are read on a scale, which is attached to a ﬂoat and moves with the ﬂuid column in the stem. The scale division h is distorted, compared to the actual height of the ﬂuid h’. There is only one reading position. Inclined-Tube Manometer: In order to increase the accuracy of the reading, the scale stem of the inclined-tube manometer is inclined (see sketch on page 242). h ; sin α

hs =

α for large hs

Stagnation Tube – Venturi Tube The sensitivity of the stagnation tube is limited: In air: 4 m/s is equivalent to 1 mm H2 O 1 m/s is equivalent to 0.0625 mm H2 O at atmospheric conditions of the air. In comparison to the stagnation tube the reading sensitivity of the Venturi tube can be improved by changing the ratio of the cross sections of the two measuring positions (see the following sketch). ρ 2 p1 − p2 = v1 2 A1 A2
2

−1

(6.13) (6.14) (6.15)

Theor.: p1 = p∞ (v1 = v∞ ) ⇒ ρ 2 ∆p = v∞ 2 for stagnation tube A1 A2
2

−1

Nozzle-Calibration Factor The nozzle-calibration factor is deﬁned as ϑ= q∞ ∆psc . (6.16)

The quantity ∆psc is proportional to the dynamic pressure q∞ . It is important for measurements, since it can be used to determine the velocity without having a measuring probe interfere with the model. ∆psc = psc − pa , q∞ = p g − p ∞ (6.17)

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o

239

The pressure in the settling chamber psc (compare. test assembly page 242) is approximately equal to the total pressure pg (v = 0). Often the static pressure in the air stream is set equal to the ambient pressure, which is only approximately true. In reality the surrounding air is entrained into the air stream. Thereby the edge-streamline of the air stream is curved, and the following can be observed: p∞ > pa ; p∞ − pa > 0 (6.18) (6.19) (6.20)

∆psc = q∞ + p∞ − pa ∆psc > q∞

Measurement of the Quality Coeﬃcient PStr Power of air stream = =ξ Power of fan Pef f ρ 3 AStr v∞ 2

Quality Coeﬃcient =

(6.21) (6.22)

Power of air stream = PStr =

The power of the air stream can be obtained from the measurement of the velocity, the barometric pressure, and the temperature. The power of the fan, Pef f ηel ηG , is determined by measuring the voltage and the current of the armature and the exciter ﬁeld of the direct-current motor of the wind tunnel. The ratio of these two quantities yields the quality coeﬃcient. Other Methods for Measuring the Velocity With the Prandtl and the Venturi tube, the velocity is determined by measuring a pressure diﬀerence. Methods which use other principles are: 1. Measurement of a force: The drag of a body – its dependence on the free-stream velocity is assumed to be known – is measured (Example: The Robinson anemometer). 2. Heat-transfer measurement: The larger the velocity of the ﬂow over a heated body, the larger the heat transfer from the body to the cold ﬂow becomes (Example: hot ﬁlm, hot wire). 3. Optical Doppler eﬀect: The frequency shift of an incident light ray and its reﬂexion from particles, moving with the velocity of a ﬂow, can be used to determine the ﬂow velocity (Example: Laser-Doppler anemometer). Selected References: Prandtl, L.: F¨hrer durch die Str¨mungslehre, 6. Auﬂ., Vieweg-Verlag, 1965, S.301 ﬀ. u o Wuest, W.: Str¨mungsmeßtechnik, Vieweg-Verlag, 1969. o

240

6. Aerodynamics Laboratory

Wind tunnel of the Aerodynamisches Institut

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o

241

Projection-manometer after Betz

Prandtl tube

242

6. Aerodynamics Laboratory

Lever micro manometer after Betz

Test assembly

6.1 Wind Tunnel for Low Speeds (G¨ttingen-Type Wind Tunnel) o 6.1.5 Evaluation 1. Draw the following diagram of the measured data: q∞ = f (z)β=0◦ , q∞ = f (β)z=0, ξ = f (v∞ ) z 243

x=const.

Prandtl tube

const.

x

β

nozzle

test section

collector

Quantity ∆psc q∞ Ba p a ≈ p∞ t∞ ρ∞ v∞ ˙ Q∞ n u λ ηG (λ) IA UA IE UE PA PW e Pef f PStr ξ ∆psc

Formula Measurement Measurement Measurement s2 Ba 13.6 g mmNHg m3 Measurement p∞ /R T∞ 2 q∞ g/ρ∞ AStr. v∞ Measurement πnDp /60 ˙ Q∞

Dimension mm H2 O mm H2 O mm Hg N / m2
◦ C kg / m3 m/s m3 / s min−1 m/s 2

20 s. b.

40

Measured data 60 80 749.44 0.99987 105

120

160

22.4 1.18 18.2 21.7 239.5 25.1 0.37

22.5 1.179 25.8 30.8 335.6 35.14 0.37

22.8 1.178 31.6 37.7 409.1 42.8 0.37

23.1 1.178 36.5 43.6 469.7 49.2 0.38

23.4 1.176 44.7 53.4 567.9 59.5 0.38

24.2 1.172 51.75 61.8 659.2 69.0 0.38

2 u Dp π/4[1 − (dN /Dp ) ] from diagram Measurement Measurement Measurement Measurement IA UA PA ηel PW e ηG 3 AStr ρ∞ /2 v∞ PStr /Pef f ≈ q∞

−− −− A V A V W W W W −−
N m2

0.8 0.8 0.8 0.77 0.77 0.77 55 90 128 165 240 330 63 89 108 124 150 167 3 3 3 3 3 3 204 205 204 206 205 206 3465 8010 13824 20460 36000 55110 3181 7353 12690 18782 33048 50591 2544.8 5882.4 10152.0 11462 25447 38955 4247 12088 22191 34169 62705 96969 1.67 2.1 2.2 2.36 2.5 2.5 196.2 394.4 588.6 784.8 1177.2 1569.6

λ ηG

0.2 0.25 0.3 0.35 0.4 0.83 0.89 0.89 0.85 0.72

AStr = An = 1.194 m2 ; Dp = 2 m ; dN /Dp = 0.5 ; Axial velocity ηel = 0.918 ; λ = ratio of forward circumferential velocity = circumferential velocity 1 mm Hg = 13.6 mm H2 O; 1 mm H2 O = 1 kp / m2 ; 1 kp = 9.81 N; m2 Nm = 287 2 for air at 1 bar and 273 K 1 Nm = 1 H2 O = 1 J; R = 287 kg K s K

244

6. Aerodynamics Laboratory q∞ = f (z) q∞ [mm H2 O] 116.6 116.9 117.2 117.4 117.4 117.6 117.4 114.7 48.5 1.0 0 1mm H2 O = 1

z[m] 0 0.1 0.2 0.3 0.4 0.5 0.55 0.575 0.6 0.625 0.65

q∞ [N/m2 ] 1143.85 1146.79 1149.73 1151.69 1151.69 1153.66 1151.69 1125.21 475.79 9.81 0

kp kg N = 9.81 = 9.81 2 m2 sec2 m m

β[ 106 a slow increase of the drag coeﬃcient cD from 0.1 to cD ≈ 0.2 is noted. The separation point moves upstream, and periodic vortex shedding is observed downstream from the sphere, similar to the subcritical regime. If the transcritical regime Re > 106 is not accesssible with the available experimental facility, the ﬂow behavior of this regime can be simulated by mounting a “separation wire”1 on the rear part of the sphere. Thereby a clearly deﬁned separation line is provided for the turbulent boundary layer. If the transition of the boundary layer is enforced with a tripping wire on the front part of the sphere, the drag coeﬃcient does not drop to the value cD ≈ 0.1, but only to cD ≈ 0.22. Obviously the tripping wire does not only inﬂuence the laminar-turbulent transition, but also the position of the separation point of the turbulent boundary layer . The details of this behavior of the ﬂow, so far could not be suﬃciently clariﬁed and are still subject to ongoing research.

1

The separation wire is to be distinguished from the tripping wire, which, when mounted on the front part of the sphere in the laminar part of the boundary layer, causes transiton to turbulent ﬂow.

6.3 Sphere in Incompressible Flow 6.3.2 Shift of the Critical Reynolds Number by Various Factors of Inﬂuence

257

Since the abrupt drop of the drag coeﬃcient cD is observed for a certain Reynolds number regime Re, the critical Reynolds number (Recrit. ) is deﬁned as the Reynolds number, for which the drag coeﬃcient is cD = 0.3. The ﬁrst measurements in the transcritical regime were conducted in 1912 and 1914. Eiﬀel determined a drag coeﬃcient of cD = 0.176 in 1912. Prandtl repeated the measurements in 1914 and obtained a drag coeﬃcient of cD = 0.44 for the same Reynolds numbers. Only later, more accurate experiments could clarify this apparent discrepancy. It was shown that the drag coeﬃcient depends on the following factors of inﬂuence: Reynolds number, intensity of turbulence of the free stream, surface roughness, support of the sphere, and heat transfer. Intensity of Turbulence of the Free Stream An increase of intensity of turbulence of the free stream shifts the transition of the boundary layer and also the abrupt drop of the drag coeﬃcient to lower Reynolds numbers (see the following diagram).
Recrit. = 1.5 · 105 windtunnel of poor quality (high turbulence intensity) Recrit. = 3.65 · 105 windtunnel of good quality (low turbulence intensity) Recrit. = 3.85 · 105 max. value reached in free atmosphere (practically zero turbulence intensity)

The highest critical Reynolds number is obtained in measurements in the free atmosphere, it is: Recrit. = 3.85 · 105 The turbulent ﬂuctuations seem to have little inﬂuence on Recrit. , since the characteristic length scales of the turbulent eddies are one order of magnitude larger than the boundary-layer tickness of the sphere. One of the main objectives of wind tunnel testing is to carry out measurements with models of airplanes, the full-scale conﬁgurations of which are to be ﬂown in the free atmosphere. Since in many cases the separation of the boundary layer is of greatest interest, it is understandable, that the intensity of turbulence of the wind tunnel is adjusted to that of the free atmosphere. The extreme sensitivy of the critical Reynolds number of the ﬂow around the sphere with regards to the intensity of turbulence of the free stream is used to determine the quality of the wind tunnel with respect to turbulence. A turbulence factor T F is deﬁned as: TF = Recrit. (free atmosphere) > 1 (6.41) Recrit. (wind tunnel)

Surface Roughness An increase of roughness of the surface of the sphere aﬀects the ﬂow behavior in a fashion similar to an increase of the intensity of turbulence or the mounting of a tripping wire (see diagram).
K d = characteristic roughness height, referenced to diameter of the sphere

258

6. Aerodynamics Laboratory

Support of the Sphere If possible, the support of the sphere should be attached to the sphere in the dead-water region. Otherwise the boundary layer on the sphere could be disturbed (see the following sketch).

a) Spindle in the dead water b) Suspended crosswise c) Suspended by wires Bacon and Reid [1924] showed that the drag can be 2.5 times larger if a transverse spindle is used instead of a spindle in the dead-water region. If the sphere is suspended by wires, turbulent wakes are generated by the wires, which aﬀect the measurement. Special attention has to be paid to mechanical oscillations of the suspension, since they can falsify the measurements markedly. 6.3.3 Method of Test The drag of a sphere is measured for Reynolds numbers 2 · 105 < Re < 5 · 105 with the balance in the wind tunnel. The sphere is mounted in the test section with a swordlike support and a spindle, connected to the balance (see sketch). Since in the measurement of the force also the drag of the swordlike support is contained, a second experiment is necessary, in which the drag of the sword is determined. This is done by mounting a mock-up of the sphere upstream of the spindle, so that the same ﬂow is generated around the swordlike support, without connecting the mock-up with the spindle (see sketch). In a third experiment a tripping wire is attached to the front part of the sphere, and the ﬂow is kept at subcritical conditions. In this experiment the abrupt drop of the drag coeﬃcient cD can be observed. Selected References Achenbach, E.: Experiments on the ﬂow past spheres at very high Reynoldsnumbers, J. Fluid Mech., Vol. 54, 1972, Part 3, S. 565-575. Achenbach, E.: Vortex shedding from spheres, J.F.M., 62, 1974, S.209-221. Bacon, D.L., Reid , E.G.: The resistance of spheres in wind tunnels and in air, NACA Rep. 185, 1924.

6.3 Sphere in Incompressible Flow

259

Bailey, A.B.: Sphere drag coeﬃcient for subsonic speeds in continuum and free molecular ﬂows, J.F.M., 65, 1974, S.401-410 Eck, B.: Technische Str¨mungslehre, 5. Auﬂ, Springer-Verlag, 1957, S. 69, 205, 243, 255. o Fage, A.: Experiments on a sphere at critical Reynolds numbers, Aero. Res. Counc. R&M, no. 1766, 1936. Flachsbart,O.: Neuere Untersuchungen uber den Luftwiderstand von Kugeln, Phys. Z., 28, ¨ 1927, S. 461. ¨ Moller, W.: Experimentelle Untersuchungen zur Hydrodynamik der Kugel, Dissertation, Leipzig, 1937, oder Phys. Z., 39, 1938, S. 57-80. ¨ Prandtl, L.: Uber den Luftwiderstand von Kugeln, G¨ttinger Nachrichten, 177, 1914. o Schlichting, H. Grenzschichttheorie, 1. Auﬂ, Verlag Braun, Karlsruhe, 1951, S.15-19,33,81. Tietjens, O.: Str¨mungslehre, 1. Band, Springer-Verlag, 1960, S. 171, 174. o Truckenbrodt, E.: Str¨mungsmechanik, Springer-Verlag, 1968, S. 103, 353, 500-501 o 6.3.4 Evaluation 1. The drag force is to be plotted as function of the dynamic pressure q∞ of the free stream: FDs = f (q∞ ) The dynamic pressure is determined with the pressure in the settling chamber ∆psc of the wind tunnels: N N q∞ = ∆psc [mm H2 O] · g /mm H2 O · ϑ (6.42) m2 m2 g = Gravitaional acceleration ϑ = Nozzle calibration factor = 0.98[−−] The drag force is determined with the mass M placed on the balance after the formula: m FD [N] = M [kg] · g 2 · Kc (6.43) s Kc = Callibration factor of the balance = 1 m g = Gravitational acceleration = 9.81 2 s In order to obtain the drag of the sphere, the total drag of the sphere and the support must be reduced by the drag of the support (comp. Sect. 4). FDs = FDtot.. − FDsup.. (6.44)

The abrupt drop of the drag coeﬃcient cD , caused by the laminar-turbulent transition, depends on several factors of inﬂuence: Intensity of turbulence of the free stream; the model support; method of test, (distinguished are the direct and integrated measurement); the measuring probes or the boreholes for the pressure measurements can inﬂuence the ﬂow); roughness of the surface of the sphere and possible heat transfer from the sphere to the ﬂow.

260

6. Aerodynamics Laboratory

2. The drag coeﬃcient cD is to be plotted as function of the Reynolds number Re: The drag coeﬃcient is deﬁned as cD = FDs q∞ As (6.45)

with As = projection area of the sphere = 0.0163 m2 Re = with u∞ = and kg ρ m3 Ba[mm Hg] · 13.6 = mm H2 O N ·g /mm H2 O mm Hg m2 Nm R · T [K] kg K 2q∞ ρ u∞ Ds ν (6.46)

(6.47)

ρ

= density

R

= =

Ba =

barometric pressure density ratio mm H2 O Hg to H2 O = 13.6 mm Hg gravitational acceleration

T Ds ν

= = = =

gas constant of air Nm ·K 287.14 kg temperature diameter of the sphere 0.144 [m] kinematic viscosity (from table) m2 s

g

=

If measurements of diﬀerent wind tunnels are compared, the test conditions have to be considered. For example, the intensity of the free stream, which depends on the wind tunnel used, can markedly inﬂuence the result. In earlier investigations this eﬀfect was not taken into account, as the intensity of turbulence was not known yet.

6.3 Sphere in Incompressible Flow

261

3. Determine the critical Reynolds number with the curve cD = f (Re) for cD = 0.3. Recrit. = 3.6 · 105 from diagram With increasing ﬂow velocity the inﬂuence of the compressibility of the air on the ﬂow can be noted. Then the Mach number has to be introduced, so that the drag coeﬃcient does not only depend on the Reynolds number but also on the Mach number of the free stream, i. e. cD = f (M a,Re∞ ). This inﬂuence can be noted for M a∞ > 0.3. 4. Determine the intensity of turbulence of the wind tunnel: 3.85·105 T F = Recrit. (cD =0.3) > 1 TF =
3.85·105 Recrit.

= 1.07.

N ot. Dim. 1 2 3 4 5 6 7 8 9 10 11 12 13 14

∆psc mm H2 O 30 50 70 80 85 90 95 100 105 110 120 130 150 170

q∞ N/m2 288.4 480.7 673 769.1 817.2 865.2 913.3 961.4 1009.4 1057.5 1153.6 1249.8 1442 1634.3

Mtot. g 245 400 531 585 613 636 620 550 500 385 370 370 370 368

FDtot. N 2.403 3.924 5.21 5.74 6.014 6.24 6.082 5.396 4.905 3.777 3.63 3.63 3.63 3.61

Msusp. g 12 24 35 37 39 44 52 58 67 69 78 91 102 115

FDsusp. N 0.118 0.235 0.343 0.363 0.383 0.432 0.51 0.569 0.657 0.677 0.765 0.893 1.001 1.128

FDs N 2.285 3.689 4.867 5.377 5.631 5.808 5.572 4.827 4.248 3.1 2.865 2.737 2.629 2.482

cD – 0.486 0.471 0.444 0.429 0.423 0.412 0.374 0.308 0.258 0.18 0.152 0.134 0.112 0.093

u∞ m/s 22.75 29.37 34.75 37.15 38.3 39.4 40.5 41.5 42.5 43.56 45.5 47.35 50.86 54.15

Re 105 1.97 2.54 3.01 3.01 3.32 3.41 3.51 3.6 3.69 3.77 3.94 4.10 4.41 4.69

Measurements with tripping wire:

N ot. Dim. 1 2 3 4 5 6 7 8 9

∆psc mm WS 30 50 70 80 90 100 110 130 150

q∞ N/m2 288.4 480.7 673 769.1 865.2 961.4 1057.5 1249.8 1442

Mtot. g 110 160 230 265 295 325 355 430 490

FDtot. N 1.079 1.569 2.256 2.59 2.894 3.188 3.482 4.218 4.807

Msusp. g 19 31 38 47 52 57 61 68 76

FDsusp. N 0.183 0.304 0.373 0.461 0.51 0.559 0.598 0.667 0.746

FDs N 0.893 1.265 1.883 1.129 2.384 2.629 2.884 3.551 4.061

cw – 0.19 0.161 0.172 0.17 0.169 0.168 0.167 0.197 0.173

u∞ m/s 22.75 29.37 34.75 37.15 39.4 41.5 43.56 47.35 50.86

Re 105 1.97 2.54 3.01 3.22 3.41 3.6 3.77 4.10 4.41

262

6. Aerodynamics Laboratory

6.4 Flat-Plate Boundary Layer
Abstract The laminar and turbulent boundary layer generated by an air stream on a ﬂat plate at zero angle of attack is investigated. The density of the air ρ and the dynamic shear viscosity µ are assumed to be constant. The change of the ﬂow characteristics caused by the laminar-turbulent transition is discussed, and the critical Reynolds number of transition will be determined. The velocity distributions in the laminar and the turbulent part of the boundary layer will be determined in the experiment and compared with theoretical data, obtained with solutions of the boundary-layer equations. 6.4.1 Introductory Remarks For suﬃciently high Reynolds numbers, i. e. when the friction forces are small compared to the inertia forces, the ﬂow over a ﬂat plate can be considered as a potential ﬂow except for the region near the wall. Only in this thin layer of thickness δ, the boundary layer, in which the velocity increases from zero to the value uδ of the external ﬂow, are the friction forces of the same order of magnitude as the inertia forces. Characteristic Dimensions of the Boundary Layer The boundary-layer thickness δ is the distance measured from the wall, from where on the velocity u is asymptotically equal to the velocity uδ of the outer ﬂow. With experimental accuracy accepted, this requirement is generally satisﬁed for the distance measured from the wall, where the velocity is about 99% of uδ . The displacement thickness δ1 is a measure for the distance, by which the external ﬂow is displaced from the wall; with ρ = const., the displacement thickness is δ1 =
∞ 0

1−

u dy uδ

(6.48)

The momentum thickness δ2 is a measure for the loss of momentum the ﬂow is suﬀering by the friction forces; for ρ = const., there results δ2 =
∞ 0

u u 1− dy uδ uδ

.

(6.49)

The Laminar-Turbulent Transition on the Flat Plate At relatively low Reynolds numbers the ﬂow is laminar; at high Reynolds numbers the ﬂow in the boundary layer becomes turbulent after a certain laminar starting region. The laminarturbulent transition sets in at a certain position, designated as xcrit. . O. Reynolds suggested in 1883 to introduce a thin thread of dye into the ﬂow parallel to the direction of the main ﬂow in order to detect transition (see sketch). The transition from laminar to turbulent ﬂow in the boundary layer is aﬀected by several quantities, the most important ones being the pressure gradient in the external ﬂow, the roughness of the wall of the plate, and the intensity of turbulence in the ﬂow of the free stream. Small

6.4 Flat-Plate Boundary Layer

263

perturbations introduced in the ﬂow in the boundary layer can no longer be damped out anymore by the Stokes stresses, and the ﬂow changes its behavior. First it becomes intermittent, alternating in time between being laminar and turbulent, generating so-called turbulent spots, which appear in irregular intervals of time. They travel downstream. Finally the ﬂow becomes fully turbulent, and velocity ﬂuctuations occur in all three coordinate directions, causing a strong diﬀusion of the dye, which is then completely spread out in the ﬂow. While for x < xcrit. the thread remains straight in the laminar part of the boundary layer, downstream from x = xcrit. , the transition point, the thread begins to oscillate irregularly, and for x > xcrit. the dye spreads out, caused by the velocity ﬂuctuations in the turbulent part of the boundary layer. Laminar and turbulent boundary layers are characterized by diﬀerent boundary-layer thicknesses and velocity distributions. The shear stress at the wall is markedly higher in the turbulent boundary layer. The main quantity inﬂuencing the transition of the laminar boundary layer is the Reynolds number Rex = u∞ x/ν. The value at which transition occurs, is called Recrit. , usually observed between Recrit. = 105 and 3 · 106 . The actual value depends on the 1) 2) 3) 4) shape of the leading edge of the plate, alignment oﬀ the plate with respect to the direction of the oncoming ﬂow, technical quality of the surface (roughness, waviness) intensity of turbulence of the free stream.

6.4.2 Method of Test Determination of the Critical Reynolds Number of the Plate The laminar-turbulent transition is ﬁrst detected with the aid of a hot-wire signal. The hotwire anemometer is particularly apted for this measurement, since it measures the instantaneous value of the velocity and thereby also the turbulent oscillations of the velocity.

264

6. Aerodynamics Laboratory

In the experiment the hot wire is positioned shortly upstream of the trailing edge of the plate in the boundary layer. Its signal is recorded with an oscillograph. If the velocity u∞ is slowly increased, the following signals are observed:

At a certain velocity u∞ crit. ﬁrst ﬂuctuations appear clearly visuable (see the sketch in the middle). The critical Reynolds number is computed with this velocity. The transition can also be detected by overhearing the boundary layer. A Pitot tube, connected with a stethoscope, is positioned in the boundary layer. The pressure ﬂuctuations, caused by the turbulent velocity ﬂuctuations, can be heard as highfrequency crackling. In this experiment the velocity u∞ is kept constant, and the Pitot tube is held at a constant distance from the wall and moved downstream. The coordinate x, where a clear high-frequency crackling is ﬁrst heard, is used for determining the critical Reynolds number. Measurement of the Velocity Distribution The velocity distributions are measured with a Pitot tube with a diameter of 0.2 µm. It is connected with a diﬀerential pressure manometer, which measures the following diﬀerence: p+ρ u2 2 − pu ,

with pu being the static pressure outside of the air stream. This diﬀerence is approximately equal to ρ u2 2 ,

since, in a good approximation, it is p∞ = p = pu .

The Pitot tube is held by a support, its distance y from the surface of the plate is varied with a screwed spindle and read on a scale. The diameter of the Pitot tube must be considerably smaller than the thickness of the boundary layer, in order not to falsify the measurements. The dimension sizes of the Pitot tube are given in the sketch in mm.

6.4 Flat-Plate Boundary Layer 6.4.3 Prediction Methods

265

The diﬀerential equations, describing the ﬂow in the boundary layer, are obtained with an order of magnitude estimation of the various term of the Navier-Stokes equations [Schlichting 1982]. For the ﬂat plate at zero incidence they read for steady incompressible ﬂow with constant density and dynamic shear viscosity: ∂u ∂v + =0 ∂x ∂y µ ∂2u ∂u ∂u ∂ +v = − uv ∂x ∂y ρ ∂y 2 ∂y

u

(6.50)

The quantities u and v are the time-averaged velocity components in the x- and y-direction, if the ﬂow is turbulent. The product ρu v ist the only term of the Reynolds stress tensor of the turbulent ﬂuctuating motion, that is retained in the boundary-layer formulation. The boundary conditions are: a) for y = 0 : u = 0 (The Stokes’ no-slip condition) v = 0 (impermeable wall) b) for y → ∞ : u = u∞ Similar Solution for the Laminar Boundary Layer For the laminar boundary layer the velocity components u and v can be determined with a similar solution of (6.50), ﬁrst obtained by Blasius, [Schlichting 1982]. The validity of the solution is conﬁrmed by experimental data. It is tabulated, for example, in [Walz 1956]. The characteristic quantities are: δ 5 = √ x Rex 1.72 δ1 = √ x Rex τo (x) 0.332 = √ ρu2 Rex ∞ 0.664 δ2 = √ x Rex Approximate Solution for the Turbulent Boundary Layer Following Prandtl the turbulent shear stress in equ. (6.50) can be appproxinated as u v = −l2 ∂u ∂u · . ∂y ∂y (6.52) (6.51) (for y Rex = 5 it is x u = 0.99) u∞

The quantity l is the Prandtl mixing length. In the vicinity of the wall it is assumed to be proportional to y, and in the outer part of the boundary layer it is approximately constant. This assumption was experimentally conﬁrmed for turbulent pipe ﬂows, but was also shown to be valid for the turbulent boundary layer on a ﬂat plate . Equation (6.50) can be integrated with this closure, but the solution is not similar. Another possibility to determine the characteristic boundary-layer quantities is given by the approximate integral method of von K´rm´n and a a Pohlhausen.

266

6. Aerodynamics Laboratory

The von K´rm´n-Pohlhausen Method a a The simple approximate method of von K´rm´n and Pohlhausen is based on equs. (6.50), a a integrated in the y-direction for steady, inkompressible ﬂow. For the ﬂow over the ﬂat plate at zero angle of attack the so-called integral condition is (see e. g. [Schlichting 1982]): τo d = ρu2 dx ∞
∞

0

u u (1 − )dy u∞ u∞

(6.53)

The solution of (6.53) requires an assumption for the unknown velocity distribution y u(x,y)/u∞ = f (y/δ(x)). Pohlhausen replaced f by a fourth-degree polynomial of . δ Example: Laminar Boundary Layer with Linear Ansatz The velocity u/u∞ is crudely approximated by u/u∞ = y/δ. With this ansatz there are obtained τo = µ · du dy =µ y=0 u∞ δ

1

and
0

1 u u y 1− d = u∞ u∞ δ 6

√ √ After insertion of these values in (6.53), integration yields δ/x = 2 3/ Rex . The displacement thickness δ1 results to: δ δ1 =
0

1−

u dy = δ u∞

1

1−
0

δ u y d = , u∞ δ 2

and hence

√ δ1 3 =√ . x Rex

Despite of the crude approximation of for u/u∞ the displacement thickness δ1 is rather accu√ rately determined, compared with the similar solution ( 3 ≈ 1.73 versus 1.72 in the section “Similar Solution for the Laminar Boundary Layer”). Example: The Power-Law Ansatz for the Turbulent Boundary Layer Prandtl employed the experimental results obtained for the pipe ﬂow and applied them to the ﬂat-plate boundary layer. With the approximate velocity distribution u y = R U and the drag resistance law 8τo 0.316 =λ= ρu2 uDρ 4 µ after Blasius
1/7

after Nikuradse

Pipe U R u ¯ ¯ U

Plate u∞ δ
1 7

=

y R

τ0 ρ U2

u u¯ ∞

= τ0 ρ u2 ∞

y δ

1 7

6.4 Flat-Plate Boundary Layer there results for the ﬂat-plate boundary layer τo 0.023 = ρu2 u∞ δρ ∞ 4 µ and δ1 = δ
0

267

and

δ2 = δ

1

0

7 u u y 1− d = 72 u∞ u∞ δ

,

1

1−

1 u y d = u∞ δ 8

.

With (6.53) the boundary-layer thickness and the displacement thickness become δ = x 0.37 u∞ xρ 5 µ and 0.0462 δ1 = u∞ xρ x 5 µ .

The results imply that the ﬂow is turbulent from the leading edge on, which is usually not true. Selected References Kaufmann, W.: Technische Hydro- und Aeromechanik, 3. Auﬂ., Springer-Verlag Berlin, 1963, S. 232. Schlichting, H.: Grenzschichttheorie, 5. Auﬂage, Verlag Braun, Karlsruhe, 1965. Truckenbrodt, E.: Springer-Verlag, Berlin, Heidelberg, New-York, 1968, S. 450. Walz, A.: Str¨mungs- und Temperaturgrenzschichten, Verlag Braun, Karlsruhe, 1966. o

6.4.4 Evaluation Baromeric pres. Temperature Gas ckonstant Density Kinem. viscosity Ba θ R ρ ν = 735.9 mm Hg = 93849.3 N/m2 = 294.6 K = 21.6◦ C = 287 Nm/kg K = p/RT = 1.11 kg/m3 = 15.8 · 10−6 m2 /s

Determination of the Critical Reynolds Number – with hot-wire signal x = 1.31 [m], q∞crit. = 41.65 [mm H2 O], u∞crit. = 27.13 [m/s], Recrit. = 2.25 · 106 – by “overhearing” the boundary layer q∞ = 83.3 [mm H2 O], u∞ = 38.37 [m/s], xcrit. = 0.83 [m], Recrit. = 2.02 · 106 Velocity Distribution Laminar Boundary Layer: x = 1.19 [m],q∞ = 27.54 [mm H2 O] = 270.15 [N/m2 ],u∞ = 22.06 [m/s] u/u∞ = q/q∞ . Diagram: u/u∞ = f y x

√

Rex

Turbulent Boundary Layer: x = 1.19 [m],q∞ = 117.11 [mm H2 O] = 1148.85 [N/m2 ],u∞ = 45.45 [m/s] Diagram: u = f (y) (double-logarithmic). u = 2q/ρ.

268

6. Aerodynamics Laboratory Flat-Plate Boundary Layer laminar m x = 1.19 m, u∞ = 22.06 , s m2 , Rex = 1.16 · 106 ν = 15.8 · 10−6 s u y N ot. y q Rex u∞ x Dim. mm mm H2 O 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 0.2 0.4 0.7 1.2 1.7 2.2 2.7 3.2 3.7 4.2 4.7 5.2 5.7 6.2 6.7 1 1.4 2.5 6.3 10.8 15.1 20.1 22.6 25.2 26.6 27.2 27.5 27.7 27.8 27.8 0.191 0.225 0.301 0.478 0.626 0.740 0.854 0.906 0.957 0.983 0.994 0.999 1.003 1.005 1.005 0.22 0.43 0.76 1.3 1.84 2.38 2.92 3.47 4.01 4.55 5.09 5.63 6.17 6.72 7.26 turbulent m x = 1.19 m, u∞ = 45.45 , s m2 , Rex = 3.38 · 106 ν = 16 · 10−6 s y q u mm 0.2 0.7 1.2 1.7 2.2 2.7 3.7 4.7 5.7 6.7 7.7 8.7 9.7 10.7 11.7 12.7 mm H2 O 33.0 51.8 63.4 71.7 78.8 84.8 94.9 102.4 107.9 111.9 114.4 115.9 117.2 117.5 117.7 117.8 m/s 24.15 30.26 33.48 35.60 37.32 38.72 40.96 42.54 43.67 44.47 44.97 45.29 45.51 45.57 45.61 45.63

δ Rex x δ1 Rex x 6.4.5 Questions

= 5 th. δ Rex x δ1 Rex x

= 4.89 exp. = 1.72 th. = 1.36 exp. 1. State the assumptions for the validity of the boundary-layer theory. The characteristic Reynolds number has to be high, Re∞ 1. This requirement implies, that the length of the body and the wall curvature must be much larger than the thickness of the boundary layer. 2. State the reasons for the deviation of the measured velocity proﬁles from the theoretical predictions and discuss one of them in detail. It can be seen in the plot of the measured velocity distribution, that the experimental values √ y at about x Rex ≈ 0.2 fall under the curve of the theoretically predicted data, corresponding approximately to uu < 0.25. Obviously the deviation is caused by a systematic error of ∞ √ measurement. The distance from the wall ∆y ≈ 0.2 x Rex mm is of the order of magnitude of the diameter of the Pitot tube. In the vicinity of the wall the Pitot tube inﬂuences the ﬂow, so that an accurate measurement of the tangential component of the velocity is not possible.

6.4 Flat-Plate Boundary Layer

269

3. Sketch a laminar and a turbulent velocity proﬁle and discuss the diﬀerences!

laminar boundary layer on a ﬂat plate

turbulent boundary layer on a ﬂat plate

4. Assume that the boundary layer is turbulent from the leading edge of the plate on and determine the force, exerted by the plate on its the suspension. (Dimensional sizes of the plate: Length = 1.5 m , width = 0.55 m; Free-stream velocity: see data of experiment) F = 2 cD width w u oo

ρ 2 u LB 2 ∞

cD =

ρ 2

τw u2 ∞

Experiment: q∞ = 117.11 mm H2 O = 1148.85 Nm−2 u∞ = 45.45 ms−1 ; ρ = 1.11 kgm−3 L = 1.5 m; B = 0.55 m 2 −6 m ν = 1.6 · 10 s du∞ =0 dx

L

The von K´rm´n-Pohlhausen integral method a a

τ(y=0) 1 dδ2 + (2δ2 + δ1 ) + =0 dx u∞ ρ u2 ∞ yields:

τ0 d ∞ u u dδ2 = 1− dy ≈ dx 0 u∞ u∞ dx ρ u2 ∞ 1 u δ2 y u = 1− d . δ u∞ δ 0 u∞

270

6. Aerodynamics Laboratory With y u ¯ = u∞ δ there is obtained: 7 δ2 = δ 72 The Blasius approximation: u∞ δ u 0.25 −0.25 δ u∞ 0.25 x 7 δ 0.25 ν δ dδ = 0.023 dx u∞ 72 0 0 0.25 x1.25 7 1 1.25 ν δ dδ = 0.023 72 1.25 u∞ x0.25 0.377 δ = √ ⇒ 5 x Rex Local skin friction coeﬃcient: cf = − − ν u∞ x τ(y=0) ρ 2 u 2 ∞
0.2
1 7

τ0 = 0.023 ρ u2 ∞ 7 dδ = 0.023 ⇒ 72 dx

ν

0.25

τ(y=0) ρ 2 u 2 ∞ dδ2 14 dδ = 2 = = dx 72 dx
0.2

⇒ δ = 0.377 x

⇒

ν dδ = 0.377 0.8 dx u∞

0.3016 x−0.2 = √ 5 Rex

⇒ cf = cD

14 0.3016 14 0.05864 √ √ = 72 5 Rex 72 5 Rex ν 1 L 1 = cf dx = ∈L 0.05864 L 0 L 0 u∞ 0.05864 1 ν 0.2 = L 0.8 L u∞ L comp. Schlichting: cD =

0.2

x−0.2 dx

0.0733 cD = √ 5 ReL

0.455 = 3.45 · 10−3 (log ReL )2.58

u∞ L = 4.26 · 106 ⇒ cD = 3.46 · 10−3 ν ⇒ F = 2 cD q∞ L B = 6.55 N ReL = (for turbulent ﬂow and ﬂate plate of zero thickness!) 5. By what experimental measures can the value of the critical Reynolds number of the ﬂatplate boundary layer be inﬂuenced? The critical Reynolds number Recrit. can be changed by shaping the leading edge of the plate diﬀerently and by changing the surface quality of the plate. Recrit. is also aﬀected by the alignment of the plate with the ﬂow direction and by the intensity of turbulence of the free stream.

6.5 Pressure Distribution on a Wing

271

6.5 Pressure Distribution on a Wing
Abstract Pressure distributions are measured in ﬁve proﬁle sections on a wing of ﬁnite span. The aim of the experiment, carried out in the wind tunnel of the Aerodynamisches Institut for low velocities, is to demonstrate, how lift and pitching moment of the wing can be determined with the pressure measured. Also the dependence of the pressure distributions on the angle of attack is investigated, and ﬂow separation will be demonstrated. The pressure distributions are measured through 25 boreholes in every measuring cross section and recorded with multipletubed manometers. 6.5.1 Wing of Inﬁnite Span On a wing of inﬁnite span the velocity does not change in the spanwise direction. The pressure distributions are therefore the same for all wing sections (two-dimensional problem). The lift is determined by decomposing the pressure force (p − p∞ )ds · ∆y acting on a surface element of the proﬁle ds · ∆y into the components tangential (dT ) and normal (dN ) to the surface and by subsequent integration over the entire surface area of the wing.

The length element in the spanwise direction ∆y, for example, is chosen to be 1 m. The normal component is: ∆N = (p − p∞ ) ds cos ε ∆y , with ds · cos ε = dx. The resultant Nres acting on the proﬁle, is obtained by integrating dN on the upper (subscript u) and lower side (subscript l) over the wing chord and by taking the diﬀerence: l Nres = ∆y
0 l

[(pl − p∞ ) − (pu − p∞ )] dx [∆pl − ∆pu ] dx
0

Nres = ∆y

(6.54)

The resulting tangential force can be obtained in similar fashion. However, for small angles ε, i. e. for slender proﬁles, Tres is negligibly small in comparison to Nres . The dimensionless normal force coeﬃcient CN Nres CN = ρ∞ U 2 l∆y 2 ∞ (6.55)

272

6. Aerodynamics Laboratory

then becomes with the pressure coeﬃcient cpu,l = CN = 1 l l ∆pu,l ρ∞ 2 U∞ 2

(cpl − cpu ) · dx,
0

with the subscripts u and l designating again the upper and lower side of the proﬁle. If the resulting pressure coeﬃcient cpres = (cpl − cpu ) is plotted versus the wing chord (see the following diagram), it is clear, that the normal force coeﬃcient CN can be obtained by integration.

The center of gravity of the shaded area is the location of the aerodynamic center of the proﬁle, i. e. of that point, at which the resulting force acts. The product of the distance of the aerodynamic center from the leading edge of the proﬁle s and the normal force N is the pitching moment M about the leading edge of the proﬁle. ρ∞ 2 M = N · s = CN U l ∆y s (6.56) 2 ∞ With the dimensionless moment coeﬃcient CM CM = one obtains for CM CM = CN The lift coeﬃcient of the proﬁle L Cl = ρ∞ 2 U∞ l ∆y 2 can for small angles of attack α, with the exact relation Cl = CN cos α − CT sin α approximately be set equal to CN . (6.58) s l . ρ∞ 2

M 2 U∞ l2 ∆y

(6.57)

6.5 Pressure Distribution on a Wing 6.5.2 Wing of Finite Span Determination of the Downwash Distribution

273

In general the pressure on the lower side of the wing is higher than on the upper side, so that, caused by the pressure diﬀerence, a ﬂow about the wing tips is originated.

Since the lift vanishes at the wing tips, the circulation has to satisfy the condition Γ (y = ±b/2) = 0. For this reason the circulation distribution over the span is assumed to be given by a discontinuous step curve with small increments ∆Γ ; the vortex elements of strength ∆Γ leave the wing in the downstream direction as free vortices, shown in the following sketch, and form a vortex sheet, which rolls up into the two free tip vortices.

streamlines on upper and lower side y −b/2

b/2 top of view of wing

view on wing from downstream cut interface (seen from behind) top view interface

interface rolls up

Γ

free vortices

At every positon y on the wing a downwash velocity wi (y) and thereby an angle of attack αi (y) is induced; according to the Biot-Savart law they are

274

6. Aerodynamics Laboratory 1 4π b/2 wi (y) = wi (y) 1 = U∞ 2π
+1

−b/2

∂Γ (y ) ∂y

dy , (y − y ) (6.59)

αi (y) =

−1

∂γ ∂η

dη η−η

;

η=

Γ 2y , γ= . b b · U∞

The downwash velocity is superposed on the free-stream velocity, whereby the geometric angle of attack is decreased. If wi u∞ , then the approximation wi /u∞ = tan αi ≈ αi can be accepted. As a consequence of the downwash the resulting aerodynamic force of the wing is no longer normal to the free-stream direction (see the following sketch); the vortices shed from the wing generate an “induced drag” of the magnitude Di = 1 U∞ b/2 −b/2

∂L ∂y

wi (y) dy.

(6.60)

The local lift is now no longer determined by the geometric angle of attack αgeom. , but by the local “eﬀektive” angle of attack αl = αe . The expression for an element of lift dL must now be written as: ρ∞ 2 dL(y) = cl (y) U l(y)dy (6.61) 2 ∞ where dcl cl (y) = αe (y) = cl∞ αe (y). (6.62) dα ∞ If it is further assumed, that the local lift is given by the Kutta-Joukowski theorem, one obtains dcl dα αe (y)
∞

ρ∞ 2 U l(y)dy = ρ∞ U∞ Γ (y)dy 2 ∞

.

(6.63)

The last equation yields αe (y) αe (y) = with f (η) = 2b l(η) cl∞ . (6.65) 2Γ (y) = γ(η) f (η) U∞ l(y)ca∞ (6.64)

With the expression for αi the local angle of attack can, according to Prandtl, be expressed by the following integro-diﬀerential equation:

6.5 Pressure Distribution on a Wing 2Γ (y) 1 + U∞ l(y)ca∞ 4πU∞ 1 2π
+1 b/2

275

αg (y) =

−b/2

dΓ dy

dy (y − y )

(6.66)

αg (η) = f (η) γ(η) +

−1

dγ dη

dη (η − η )

(6.67)

This equation is the starting point of the lifting-line theory, for which two main problems are deﬁned: 1. Main problem: a) given: γ(η),f (η) to be determined: αg (η) distribution of the angle of attack (twist) b) given: γ(η),αg (η) to be determined: f (η) 2. Main problem: given: f (η),αg (η) to be determined: γ(η) Elliptic Circulation Distribution The results obtained for the elliptic circulation distribution as obtained in [Trunkenbrodt 1967] are repeated here. If the circulation is given by Γ (y) = Γo 1− 2y b
2

,

(6.68)

with Γo being a constant, the lift of the wing is obtained to b/2 L = ρ ∞ U∞
−b/2

Γ (y) · dy =

π ρ∞ U∞ Γo b. 4

(6.69)

The downwash velocity is constant wi (y) = Γo 2b . (6.70)

The induced drag of the elliptic circulation distribution is Di = π 2 · ρ∞ Γo 8 , (6.71)

and the coeﬃcient of the induced drag CDi can be expressed by the lift coeﬃcient CL and the aspect ratio of the wing CDi = The aspect ratio is deﬁned as Λ = b2 /A b/2 2 CL πΛ

.

(6.72)

and

A=

−b/2

l(y) · dy

.

(6.73)

The induced angle of attack is given by αi = CL πΛ . (6.74)

With the last two equations drag coeﬃcients and angles of attack can be determined for diﬀerent aspect ratios Λ.

276

6. Aerodynamics Laboratory

Solution of the Prandtl Integro-Diﬀerential Equation Following the ansatz of Glauert and Treﬀtz (6.63) can be transformed into a system of linear algebraic equations with the aid of a Fourier-series representation for the circulation. The circulation γ is represented by a truncated Fourier series m γ(ϑ) = 2 µ=1 aµ sin µϑ,

(6.75)

in which ϑ is obtained from η by the transformation ϑ = arccos η . (6.76)

The integration of (6.67) yields (special attention is to be paid to the singularity in the integral) m m

αg (ϑ) = 2 f (ϑ) µ=1 aµ sin µϑ + µ=1 µaµ

sin µϑ sin ϑ

.

(6.77)

The quantities aµ are the Fourier coeﬃcients. αµ = 1 π π γ(ϑ) sin µϑdϑ
0

m 1 γn sin µϑn m + 1 n=1

(6.78)

The trapezoidal rule is used for the solution of the integral. If the coeﬃcients aµ are replaced by the approximate expression (6.66), there results m m

αgν = fν γν + n=1 γn µ=1 µ sin µϑν sin µϑn . m + 1 sin ϑν

(6.79)

With (6.78) the Prandtl integro-diﬀerential equation is transformed into a system of m linear algebraic equations for γν (ν = 1,2....m). In matrix formulation (6.5.2) can be written as B · γ = αg (6.80)

In this equation B is the coeﬃcient matrix, which depends only on the geometry of the wing, and γ is the unknown solution vector. The model of the wing used in the experiment has a proﬁle NACA 23012; the characteristic quantities are: b = 0.9 m; l = 0.2 m; cl = 5.35; αg = 6o ; Λ = 4.5 (6.81)

The results of the calculation with the lifting-line theory are shown by the curve 1 in the following diagram. Also shown are the results of the extended lifting-line theory of [Schlichting, Truckenbrodt 1967] in curve 2. The Approximate Method of Schrenk Munk could show [Schlichting and Truckenbrodt 1967], that the induced drag attains a minimum, if the circulation distribution on the wing is elliptic. The fact, that the circulation distribution also depends on the shape of the wing (see (6.67), which contains the quantity l(y)), oﬀers the possibility to quickly estimate the deviation of the circulation distribution of a given wing from the elliptic distribution; such an estimate is advantageous for design problems. In [Schlichting and Truckenbrodt 1967] it is proposed to assume the local lift to be given by the mean value of the elliptic distribution and a distribution proportional to the local chord:

6.5 Pressure Distribution on a Wing

277

--cL = 0.39045 cDi = 0.01121 simple lifting-line theory cDi = 0.0834 extended lifting-line theory —— cL = 0.3419 Lift distribution of the rectangular wing with NACA-proﬁle 23012; α = 6◦ ⎡ ⎤ ⎦

2y 1 L(y) = ⎣a0 · l(y) + a1 1 − b 2

2

(6.82)

The constants a0 and a1 can be obtained by integration in the spanwise direction and requiring that both integrals have to have the same value. One obtains for cl (y)
⎡

cl (y) =

CL ⎣ 4lm 1+ · 2 πl(y)

1−

2y b

2

⎤ ⎦,

(6.83)

with lm being the mean chord and CL the total lift of the wing. The dimensionless circulation distribution then is: γ(y) = γ1 (y) + γ2 (y) = 2y CL l(y) CL A 1− + 4b πb2 b
2

(6.84)

The quantity γ1 (y) is the local circulation proportional to the chord, and γ2 (y) is the elliptic contribution. The following distribution is obtained for a rectangular wing:

It is to be noted, that the Schrenk approximation does not drop to zero at the end of a rectangular wing. The error is of minor importance, if one remembers the simplicity of the method.

278

6. Aerodynamics Laboratory

6.5.3 Method of Test In the experiment a model of a wing is used with a span width of 900 mm and a chord of 200 mm. The wing has zero twist and its proﬁle is the NACA 23012 proﬁle with a maximum thickness position of 30% chord, 12% relative thickness, and 1.84% relative camber at 15% chord. The pressure distribution is measured in ﬁve sections at various locations in the spanwise direction (see the following drawing of the model). In each measuring section there are 14 boreholes on the upper side of the wing, and 11 on the lower side. All pressure holes are connected with multiple-tubed manometers via ﬂexible tubes so that 125 connections have to be lead out of the model of the wing. In order to avoid disturbances of the ﬂow by the pressure tubes, they are put inside the sword-like support of the wing, which is mounted at a right angle to it in the wind tunnel (see the following drawing). A comparison of the calculated and the measured lift distribution shows, that there is some disturbance by the sword, but it is small. The pressure tubes are lead from the sword to ﬁve multiple-tubed manometers, which are ﬁlled with water as sealing liquid. The entire measuring unit consists of a storage container, ﬁlled with water and sealed oﬀ against the atmospheric pressure. 27 glas tubes are immerged in the container, and the pressure inside can be raised by letting compressed air in, which causes the water columns in the glas tubes to rise. Since the ﬂexible tubes of every measuring section are connected with 25 glas tubes, the water columns are sucked upward by underpressures and pushed downward by overpressures. The two remaining glas tubes are connected to the surrrounding atmosphere and serve for reference pressure measurements. In this manner the multiple-tubed manomaters are used as diﬀerential pressure manometers. The dimension size of the sword could be reduced and thereby also the error of the measurements, if electric pressure probes were used. The multiple-tubed manometers, however, are prefered here because of their simplicity. Also the pressure distribution can be recorded, when the ﬂow separates and stalls. The position of the boreholes is shown in the following drawing. The total lift of the wing is obtained by integrating the local lift coeﬃcients in the spanwise direction. The following page contain a perspective drawing of the model of the wing with the sword and the suspensions. Indicated are the ﬁve sections, in which the pressures are measured. Sections I and II are close to the wing tips, sections III and IV further inboard, and section V near the sword. The table below the drawing gives the positions of the boreholes for the pressure measurements on the upper and lower side of the wing. The corresponding numbering of the holes is shown in the sketch below the table. The drawing on the bottom of the page gives the spanwise positions of the ﬁve measuring sections. The various distances are measured in mm. Selected References Multhopp, H.: Die Berechnung von Auftriebsverteilung von Tragﬂ¨geln, Luftfahrtforschung, u Bd. 15, 1938 Schlichting, H., Truckenbrodt, E.: Aerodynamik des Flugzeugs, 2. Band, 2.Auﬂ., Springer-Verlag, 1967

6.5 Pressure Distribution on a Wing

279

125 connection to manometers

The wing model for measuring the pressure distribution

Position xU pperside xLowerside

Dim. 1 2 3 4 5 6 7 8 9 10 11 12 13 14 mm 0 6 15 23 32 48 62 77 92 109 122 140 158 175 mm 15 23 32 48 62 77 92 109 122 147 170 — — —

Pressure holes on the NACA proﬁle 23012

280

6. Aerodynamics Laboratory

6.5.4 Evaluation 1. Evaluation of the data: a) Draw the diagrams ∆p = f (x) for the ﬁve sections I, II, III, IV, V;

6.5 Pressure Distribution on a Wing Pressure distribution p [mm H2 O]
1 2 3 4 5 6 7 8 9 10 11 12 13 14 pu 120 -30 -75 -75 -68 -50 -40 -29 -22 -16 -13 -10 -4 5 Section I pl -2 -23 -25 -25 -21 -21 -16 -13 -9 -3 -1 — — — pu 120 -60 -94 -90 -81 -62 -47 -34 -29 -10 -13 -12 -5 -4 Section II pl -1 -5 -10 -18 -17 -16 -16 -11 -11 -7 -2 — — — Section III pu pl 121 -2 -63 -8 -94 -14 -92 -20 -85 -21 -66 -20 -53 -19 -41 -15 -33 -12 -25 -7 -20 -2 -14 — -5 — 3 — Section IV pu pl 117 9 -95 1 -120 -10 -112 -15 -102 -16 -80 -17 -65 -19 -50 -14 -41 -11 -32 -6 -27 -3 -18 — -10 — -2 — Section V pu pl 110 8 -99 0 -119 -10 -112 -21 -101 -22 -78 -23 -65 -24 -52 -20 -41 -18 -33 -11 -27 -2 -18 — -10 — 1 —

281

b) Determine the mean pressure for all sections! ∆pm = ∆pm ∆pm 1 l (∆pu − ∆po )dx l 0 1 l = ∆p dx l 0 1h (∆p0 + 4 ∆p1 + 2 ∆p2 + 4 ∆p3 + ... + 4 ∆pn−1 + ∆pn ) = l3

(Simpson’s method) h= l ; n choose: l = 19.6 cm ⇒ n = 14; h = 1.4 cm

∆pi from diagram (results see following table)
Section Nr. ∆p0 ∆p1 ∆p2 ∆p3 ∆p4 ∆p5 ∆p6 ∆p7 ∆p8 ∆p9 ∆p1 0 ∆p1 1 ∆p1 2 ∆p1 3 ∆p1 4 ∆pm l ∆pm ca γ 1 0 627.84 451.260 313.920 196.200 117.72 78.48 39.240 19.62 19.620 19.620 9.81 9.81 5.886 0 28.402 144.908 0.126 0.014 2 0 941.760 706.32 490.5 333.54 176.58 98.1 58.860 39.24 19.620 19.620 9.81 9.81 3.924 0 43.015 219.464 0.190 0.021 3 0 922.14 725.94 529.74 372.78 255.16 176.58 117.72 78.48 58.86 39.24 19.62 19.62 9.81 9.81 50.77 259.03 0.225 0.025 4 0 1314.54 961.38 725.94 549.36 392.4 294.3 235.44 176.58 137.34 98.10 78.48 39.240 19.620 9.810 74.026 377.685 0.327 0.036 5 0 1255.680 961.38 686.7 470.880 333.54 235.44 137.340 98.1 78.480 58.860 39.240 19.620 19.620 9.810 64.870 330.971 0.287 0.032

c) Calculate the local lift coeﬃcient cl = ∆pm /∆psc · δ for all sections; determine γ = ca · l/2b and plot in diagram (nozzle calibration factor δ = 0.98); cL = ∆pm ∆psc ·δ = ∆pm 1153.6 P a l γ = cl 2b

d) Determine the location of the aerodynamic center for the section III, measured from the ∆p ∆p ∆p leading edge. cpres. = cpu,l = ρ∞ v2 q∞ = 1153.6 P a
2 ∞

282

6. Aerodynamics Laboratory See diagram: Position of center of gravity of the area is approximately at s = 43 mm

extended lifting−line theory rectangular wing simple lifting−line theory γ measured

−1

−0.5

0 2y/b

0.5

1

e) Discuss the inﬂuence of the Reynolds number Re on the maximum lift of a wing! At high angles of attack the pressure minimum on the upper side of the wing is located near the leading edge. The pressure rise towards to trailing edge generates a large positive pressure gradient, which either provokes laminar-turbulent transition or separation of the laminar ﬂow. Whether transition or separation is ﬁrst, mainly depends on the Reynolds number and on the magnitude of the pressure gradient.

For low Reynolds numbers laminar separation is observed ﬁrst and leading edge stall may result. The maximum lift coeﬃcient is very low. With increasing Reynolds number transition is posssible, and the ﬂow can reattach. The maximum lift therefore increases with increasing Reynolds number.

6.6 Aerodynamic Forces Acting on a Wing

283

6.6 Aerodynamic Forces Acting on a Wing
Abstract Lift, drag, and pitching moment of a rectangular wing with the NACA 23012 proﬁle, and the drag of the model suspension are determined by measuring three aerodynamic forces with the balance of the low-speed wind tunnel of the Aerodynamisches Institut. The functioning and the utilization of the balance and the correction of the measured data are explained. Application of the experimental data to full-scale conﬁgurations will also be described. 6.6.1 Nomenclature of Proﬁles Wing proﬁles for subsonic ﬂight are well rounded at the leading edge and pointed at the trailing edge. The shape of the proﬁle is charcterized by the following parameters:

Chord of the proﬁle l: The characteristic parameter, to which all dimension sizes are referenced. Maximum relative proﬁle thickness d/l; relative thickness position xd /l; maximum relative camber f /l; relative camber position xf /l. The inﬂuence of the geometric parameters of a proﬁle on its aerodynamic characteristics has been investigated by numerous authors, see for example [Abbott, v. Doenhoﬀ, Pankhurst, Holder], but especially by NACA Research Laboratories of the United States, see for example [Ann. Reports 193, 1937]. The digits of the proﬁle, used here in the experiment, have the following meaning: For the ﬁve-digit series (e. g. NACA 23012), which begins with a 2, xd /l is xd /l = 0.3 = const.; for 230....., f /l = 0.0184. The digit 2 indicates 20/3 of the lift coeﬃcient c∗ (= 0.3) at shockless entrance, i. e. ﬂow l direction is tangent to the camber in the point on the leading edge (see sketch above). The 2nd and 3rd digit denote twice the value of xf in percent of l: xf /l = 0.15. The 4th and 5th digit denote d in percent of l: d/l = 0.12. 6.6.2 Measurement of Aerodynamic Forces The aerodynamic forces acting on a wing of inﬁnite span can be decomposed into lift, drag, and pitching moment about a suitable point. The forces, especially the drag of the proﬁle, cannot be determined from the pressure distribution, since the friction forces are not taken into account. In the ﬂow about a proﬁle the friction forces can become markedly larger than the pressure forces (e. g. the drag of the proﬁle). The accurate determination of the drag alone by integration of the normal forces is only possible for blunt bodies, since the normal forces are considerably larger than the friction forces. It is for this reason that the aerodynamic forces acting on the model are measured directly with a wind-tunnel balance.

284

6. Aerodynamics Laboratory

Wind-Tunnel Balances For force measurements with multi-components balances the model must be suspended in such a way, that individual forces do not interact with each other, if possible, i. e. a force, which acts in a certain coordinate direction (e. g. the drag) should not generate a component in another direction (e. g. lift). Mechanical and electric balances are used. The working principle of the mechanical lever balances is, that the displacement of the model by the aerodynamic forces is compensated by increasing the weight. The main advantage of this type of balance is, that the model is always brought back to its initial position. The disadvantages are: The space required and their limitation to steady measurements. The electric balances measure the displacement of the model due to the action of the aerodynamic forces indirectly with ohmic, capacitiv, or inductive displacement pickups. The suspension has to be so elastic, that the aerodynamic forces to be expected cause only a small displacement of the model and the suspension. It has to be compromized between the displacement, which generates components in other directions of force and the sensitivity of the reading of the balance. The advantage of the electric balances is mainly to be sought in their possible miniaturization. They can be mounted on models. Because of their short response times measurements of unsteady aerodynamic forces are possible. A large eﬀort is, however, necessary for the design of the suspension of the model, if the interaction of the components of forces, measured with an electric balance, is to be avoided. The balance used in the test is a mechanical three-components balance with wire suspension and transfer of force via beams of balance, suspended by knife edges. A model is held in equilibrium position by six wires, with six corresponding forces in the general case, measured with the balances A ,B,C,D,E, and F . The quantities to be measured are speciﬁed in the following drawing. The schematic arrangement of the experimental setup is shown in the sketch on the bottom of the page. Depicted is a model of an airplane with its axes of orientation. Indicated are also the six balances, the wires, and the positions, where the wires are connected with the model. The balances A and B are fastened at the wing tips, with the wires aligned in the direction of the main ﬂow. The drag is measured with these balances. The wires leading to the balances C and D are fastened at the same positions, but in the direction normal to the main ﬂow. Together with a third balance E they are used for measuring the lift. The side force is measured with the sixth balance F , with the wire connected to one of the wing tips, orientated in the spanwise direction. The rolling, yawing, and the pitching moment are obtained as indicated in the sketch from the force measurements. 1. Drag D = A + B 2. Lift L = C + D + E 3. Side force S = F 4. Rolling moment MR = (C − D) · b/2 5. Yawing moment MG = (A − B) · b/2 6. Pitching moment MN = E · t The various components to be measured are explained with the aid of the following drawing; a more detailed drawing of the balance is shown on page 296 of the section 6.6.4; the balance for measuring the lift in the rear is displaced horizontally when the model of the wing is positioned at angle of attack, so that the suspension wires in the front and at the rear remain vertical, when the distance between the suspension points is changed.

6.6 Aerodynamic Forces Acting on a Wing

285

In the diagrammatic sketch the equilibrium condition for the individual systems (beams of balance) is given: The external forces LH und LV act on the beam of balance in the middle (a), and the equilibrating forces are D and H. Ky = 0 : L H + L V = D + H (6.85)

The external force D acts on the lower beam of balance (b), and F and B are the balancing forces. MI = 0 : D · e = B · (e + d) → D = B · (e + d)/e (6.86)

On the upper beam (c) the equilibrium is given by the bearing forces in point II, B, H, and G act as external forces. MII = 0 : G · f = H · b + B · (c + b) ⇒H= (b) and (c) are inserted into (a): LV + LH = Ltot. = B · (e + d) B · (c + b) G · f − + e b b (6.88) B · (c + b) G·f − b b (6.87)

In order to ensure that the lift is proportional to the force G (weight laid on), the following condition has to hold: B · (e + d) B · (c + d) d c = , d.h. = e b e b Finally there is obtained: Ltot. = G · f = G · ZLtot. b (6.90) (6.89)

The balance is constructed according to the condition (ZLtot. = 3). The lever ratios ZLH and ZD for the lift balance in the rear and also for the drag balance can be read from the diagrammatic sketch on page 296. Determination of the Aerodynamic Coeﬃcients from the Measured Data The total lift Ltot. , the rear lift LH , and the total drag of the model, and the suspension D are determined in the measurement with the balance. For a ﬁxed geometry of the proﬁle (i. e. including the angle of attack) it follows from the similarity laws, that the dimensionless

286

6. Aerodynamics Laboratory

force coeﬃcients depend only on the Reynolds number Re: The aerodynamic coeﬃcients are presented in the form cL , cD , cm = f (α) or as polar diagrams cL , cm = f (cD ), with the Reynolds number Re as parameter. The drag of the suspension has to be measured separately and be subtracted from the total drag. The mutual interference between model and suspension has to be taken into account. The model is suspended with two swords, fastened at the wing tips and with a wire downstream from the wing, schematically depicted in the sketch. In order to keep the interference between the model and the suspension small, the drag of the swords should be as small as possible. This requirement implies, that the size of the wing should be suﬃciently large, so as to guarantee an accurate measurement of the drag. Dsusp. = Dswords + Dwire

It follows from the above sketch that the drag of the suspension consists of the drag of the two swords and of the drag of the wire. The drag of the swords is measured in a separate experiment without the model and is later on included in the evaluation of the three-components measurements. The drag of the wires is calculated with existing drag formulas for circular cylinders. The moment coeﬃcient is obtained from the following equilibrium consideration:

With the suspension included the equilibium of moments is: MP = 0 : N · x − LH · tF · cos α = 0 (6.91)

The weight of the wing drops out, since the balance is calibrated for u∞ = 0. For the free ﬂight the moment about the point E (again without the weight) is: ME = N (x − x). N x inserted, there results: ME = MY = AH tF cos α − N x The diagrammatic sketch shows the aerodynamic forces on page 296; one obtains (6.92)

6.6 Aerodynamic Forces Acting on a Wing N = L cos α + D sin α T = D cos α − L sin α . The moment coeﬃcient is: My cm,y = cm = ρ 2 u Awing lwing 2 ∞

287

(6.93)

(6.94)

For the determination of the moment the change of the rear lift (δLH ) caused by the static moment for an eccentric location of the center of gravity. The equilibrium of moments about the point P gives (for α = 0 with symmetric (•) and asymmetric ( ) location of the center of gravity (cgl)): For symmetr. (•) and eccentric ( ) cgl and α = 0 LHo tF = GF lS

for symmetr. cgl (•) LHα tF cos α = Gf lS cos α

For eccentric cgl ( ) LHα tF cos α = Gf cos α (lS + x) With x = h tan α, there is obtained δLH = LHα − LH0 h for ( ) = GF tan α tF If the wing is in the normal position the rear lift is measured too small, caused by the change of the static moment. If h = 0, (center of gravity on the chord; in general on the line connecting the suspension points), then the correction ∆LH = 0 is not needed. Estimate for the wing investigated: GF ≈ 20 kN ; tF ≈ 0.3 m; h ≈ 3.5 mm; αmax ≈ 22o → tan αmax ≈ 0.4. ∆LHmax = 2 · 104 · 3.5 · 0.4 N ≈ 100 N 300 (Reading round oﬀ to 50 N ) (6.95)

For (•) = 0

1 ∆L∗ = δLH −→ (∆L∗ )max ≈ 50 N H H 2

(6.96)

Since the maximum error is less than 1%, if δL∗ is neglected, the correction of the rear lift is H left out in the evaluation (note that ∆LH is larger for strongly cambered proﬁles). 6.6.3 Application of Measured Data to Full-Scale Conﬁgurations Inﬂuence of the Reynolds number on the Boundary Layer It the measured data are to be applied to the full-scale conﬁgurations, in incompressible ﬂow the Reynolds number has to have the same value for the ﬂight condition and the experiment; this requirement implies large wind-tunnel models. However, only very few proﬁles have been investigated in full scale in large wind tunnels.

288

6. Aerodynamics Laboratory

The dependence of the aerodynamic force coeﬃcients (e. g. cDmin and cLmax ), and of the behavior of the boundary layer on the Reynolds number Re is brieﬂy discussed for the proﬁle NACA 23012 with the following diagrams. If the laminar boundary layer separates downstream from the location of the maximum thickness (see sketch A), a dead-water region is generated; the drag coeﬃcient cD is relatively large. If the boundary layer undergoes transition downstream from the location of maximum thickness (see sketch B), because of the large momentum transport in the direction normal to the wall, the turbulent boundary layer separates only shortly upstream of the trailing edge. The drag coeﬃcient cD drops and attains a minimum value. With increasing Reynolds number Re the transition point moves upstream (see sketch C); the overall ﬂow patterns do not change much in comparison to those shown in sketch B, but the turbulent boundary layer extends over a larger region in the streamwise direction than does the laminar boundary layer; consequently the integral over the Reynolds stresses is larger, and the drag coeﬃcient cD also rises. If the boundary layer is turbulent from the leading edge on, the drag coeﬃcient decreases with increasing Reynolds number, but the drag increases further. Since the turbulent boundary layer can withstand a larger pressure increase than the laminar boundary layer, the ﬂow separates only at a larger angle of attack (i. e. high cLmax ); with increasing Reynolds number cLmax increases from 1.0 at Re ≈ 105 to 1.7 at Re ≈ 107 . The variation of cLmax clearly shows the sensitivity of the proﬁle used to changes of the Reynolds number. Correction of the Data Due to Finite Diameter of the Test Section Boundary conditions in wind tunnels. Caused by the inﬂuence of the edge of the air stream on the ﬂow near the model, a correction of the measured data is necessary prior to applying the data to a full-scale conﬁguration. In order to keep the inﬂuence of the edge small, the model must be small in comparison to the diameter of the air stream, while the similarity laws require large models, because of the high Reynolds numbers of the full-scale conﬁgurations.

6.6 Aerodynamic Forces Acting on a Wing

289

The conditions at the periphery of the test section will ﬁrst be discussed for the closed and open test section. Consider the following example: The drag coeﬃcient of a circular disk, positioned at a right angle to the oncoming ﬂow, shows the following variation with the Reynolds number cD = f (Re), indicated schematically in the sketch below.

For the range of Reynolds numbers for which cD = const., one obtains the sketched behavior of wind-tunnel measurements compared to measurements in the atmosphere (d = diameter of the circular disk, and D is the diameter of the air stream in the wind tunnel). As long as the ratio d/D is small (here d/D < 0.15), the wind tunnel measurements of the drag coeﬃcient cD agree with the values measured in the atmosphere. The comparison of the ﬂow about a blunt body in a closed test section of a wind tunnel (Boundary condition on a rigid wall: vn = 0) with the ﬂow about the same body in the free atmosphere shows, that the wall of the test section squeeses the ﬂow and narrows the stream tube of the actual ﬂow. The walls thereby cause an increase in velocity and higher drag compared to the free atmosphere (see sketch).

In an open test section the conditions are reversed; the ﬂow can stronger expand in the direction normal to the oncoming ﬂow than in the free atmosphere. The corrections of the measured data, necessitated by the false lateral boundary conditions in the wind tunnel, can either be obtained from special experiments [NACA Annual Reports 1935, 1937, Pope 1954, Wien and Harms, 1932] or with theoretical estimates [Pope 1954, Prandtl 1961]. The methods of corrections can only brieﬂy be discussed here. The boundary conditions can relatively simply be satisﬁed in potential ﬂow by superposition of singularities. In order to apply this method, the following assumptions are introduced: 1. The friction forces at the edge of air stream are not considered. 2. The velocity disturbances caused by the body at the edge of the air stream are small compared to u∞ : f (vx ,vn ,vt ) 1) the o friction coeﬃcient λ varies strongly with M a, while in subsonic ﬂows λ is almost independent of M a, so that the data measured in incompressible ﬂow remain also valid for subsonic compressible ﬂow for M a < 1. The following formula is valid for smooth pipes for Reynolds numbers 2 · 104 < Re < 2 · 106 : λ = 0.0054 + 0.396 · Re−0.3 (6.123)

308

6. Aerodynamics Laboratory

The Change of the State Variables by Skin Friction For steady, one-dimensional, adiabatic ﬂow of a perfect gas the conservation equations for mass, momentum, and energy have the following diﬀerential form: dρ du + =0 Continuity equation ρ u Momentum equation dp + ρudu + Energy equation and further Equation of state dρ dT dp = + p ρ T λdx ρu2 =0 D 2

cp dT + udu = 0

dM a2 du2 dT = 2 − 2 Ma u T The relation between the increase of entropy and the decrease of stagnation pressure is obtained from the second law of thermodynamics: and for constant speciﬁc heats T ds = cp dT − ds = cp dp ρ (6.124) (6.125)

dT dp −R T p

The stagnation conditions (T0 ,p0 ) of a ﬂowing gas are deﬁned by the assumption, that they are generated by an isentropic deceleration to zero velocity u = 0. The following relations are then valid: s = s0 , ds = ds0 and ds0 = cp dT0 dp0 −R T0 p0 (6.126)

With the adiabatic change of state assumed, i. e. (dT0 = 0), there results s02 − s01 = s2 − s1 = R ln p01 p02 . (6.127)

The changes of the state variables are obtained with the above equations [Shapiro 1953, Emmons 1958]:

6.8 Resistance and Losses in Compressible Pipe Flow Ma < 1 Ma > 1 du u dT T dp p dρ ρ dM a2 M a2 ds R dp0 p0 = λ γM a2 · dx 2 (1 − M a2 ) D λ γM a2 (γ − 1) · dx 2 (1 − M a2 ) D γM a2 λ [1 + (γ − 1)M a2 ] · dx 2 (1 − M a2 ) D λ γM a2 · dx 2 (1 − M a2 ) D >0 0

0 0 M acrit. , the increase of entropy and the decrease of the stagnation pressure are not only caused by the mixing process but also by the compression shock at the downstream end of the supersonic region. For M a1 = M amax the entropy increase and also the decrease in stagnation pressure are solely due to the terminating compression shock, since the ﬂow in front of the shock is almost isentropic. Selected References Emmons, H.W.: High Speed Aerodynamics and Jet Propulsion, Vol. II: Fundamentals of Gas Dynamics, Princeton New Jersey, 1958, S. 228 ﬀ.. ¨ ¨ Frossel, W.: Str¨mung in glatten, geraden Rohren mit Uberschall- und Untero schallgeschwindigkeit, Forsch. Ing.-Wesen, 1936, Bd. 7,S.75. ¨ Jaeneke, Ch.: Untersuchungen zur Uberschallstr¨mung in einem Rechteckkanal, Dissertation o RWTH Aachen, 1975 Naumann, A.: Druckverlust in Rohren nicht kreisf¨rmigen Querschnitts bei hohen o Geschwindigkeiten Allgem. W¨rmetechnik, 1956, Bd. 7, S.32 a Prandtl, L.: Str¨mungslehre, Vieweg-Verlag, 1669, S. 229. o Shapiro, A.: The Dynamics and Thermodynamics of Compressible Fluid Flow, Part I, New York, 1953, S.114, 155-160.

6.8.4 Evaluation Method of Test The pressure distribution on the wall of the pipe is measured in the vicinity of the oriﬁce with a multiple-tubed manometer for a prescribed Mach number M a1 .

314

6. Aerodynamics Laboratory

The following quantities are measured: pD , ∆pw , T0 , Ba,p(x) n x/D p(x) − Ba [mm Hg] p(x)[N/m2 ] · 103 n x/D p(x) − Ba [mm Hg] p(x)[N/m2 ] 1 -5 491 165 13 4.2 260 134.7 2 -4 490 165 14 4.6 261 134.8 3 -3 489 165.2 15 5.0 260 134.7 4 -2 490 165 16 5.4 261 134.8 5 0 500 166.6 17 5.8 261 134.8 6 0.1 73 109.8 18 6.2 260 134.7 7 1.8 195 126.1 19 6.6 260 134.7 8 2.2 217 129 20 7.6 259 134.6 9 2.6 234 131.3 21 8.6 257 134.3 10 3.0 244 132.6 22 9.6 257 134.3 11 3.4 251 133.5 23 10.6 255 134.1 12 3.8 259 134.6 24 11.6 254 133.9

Data Evaluation The mass ﬂow ρu can be determined with the data pD , ∆pw , T0 and the dimension sizes of the nozzle according to standard speciﬁcations (NORM DIN 1952 in Germany). ρu = mD α 2ρD ∆pw (6.136)

Inserting the quantities ρu, p, and T0 into the gasdynamic relations the state variables in the cross sections I and II can be computed, and the resistance coeﬃcient of the oriﬁce ζO is obtained as a function of M aI and ReI . If the static pressure p is known at a certain location, the Mach number can be determined with the following relations m ˙ Ap and T0 γ−1 M a2 =1+ 2 T If the abbreviation M a
T0 T

ρu RT0 = γ ρT

RT0 T0 = Ma γ T

(6.137)

.

(6.138)

= K is used, there is obtained (6.139)

1 1 + 2K 2 (γ − 2) − 1 . γ−1 If M a is known, the temperature ratio T /T0 can be computed from (6.138) and M a2 = γ T0 γ−1 p0 = . (6.140) p T For the temperature of the experiment the viscosity of air is approximated by the relation µ µ0 ◦ C The Reynolds number can be expressed as Re = ρuD ρuD = µ(T ) µ(T0 ) = T T0◦ C T0 T
0.76

.
0.76

(6.141)

0.76

= Re0

T0 T

.

(6.142)

Resistance and losses of a compressible pipe ﬂow with a standard oriﬁce (details of evaluation)

6.8 Resistance and Losses in Compressible Pipe Flow
Nr. 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 Diameter of pipe Diamter of nozzle Opening ratio of nozzle Diameter of oriﬁce Opening ratio Gas constant Viscosity at 0o C Barometer reading Stagnation temperature Nozzle excess pressure Diﬀerential pressure Pressure ahead of nozzle Density ahead of nozzle Viscosity at T0 Pressure ratio Expansion coeﬃcient Estim. discharge coeﬃcient Mass ﬂow Reynolds number Discharge coeﬃcient Mass ﬂow 1 Total mass ﬂow Pressure at I Combination Mach number ζmax (MI ) Temperature ratio Stagnation pressure Reynolds number Friction coeﬃcient Pressure at II Combination Mach number ζmax (MII ) Temperature ratio Stagnation pressure Reynolds numberl Friction coeﬃcient Total friction coeﬃcient ζF riction Resistance coeﬃcient of oriﬁce Ratio of stagnation pressures Increase of entropy Symbol D dn mD d m R µ0o C Ba T0 hn ∆pw pn ρn µ0 p2 /p1 α0 ρ u(a0 ) Re0 α ρu m ˙ pI KI MI ζmaxI T0 /TI p0I ReI λI pII KII MII ζmaxII T0 /TII poII ReII λII ζI,II ¯ λlI,II /D ζB p0I /p0II ∆s/R Numerical value 0.05 0.04 0.64 0.0353 0.5 287.3 1.71 · 10−5 753 293 555 1085 10643.85 1308 173964 2.066 1.8044 · 10−5 0.9388 0.9389 1.1732 147 4.096 · 105 1.1707 147.523 0.2896 165.2 · 103 0.2189 0.2178 13 1.0095 1.707 · 105 4.125 · 105 0.01358 134.1 · 103 0.2697 0.2678 8 1.01434 1.409 · 105 4.14 · 105 0.01357 5 0.19005 4.81 1.2115 0.192 Unit m m mn = (dn /D)2 m m = (d/D)2 Nm/kgK Ns/m2 mm Hg K mm Hg mm H2 O N/m2 mm Hg N/m2 kg/m3 Ns/m2 Equation

315

1 mm H2 O = 9.81 N/m2 pn = Ba + hn 1 mm Hg = 133 N/m2 ρn = pn /R T0 (6.141) p2 /p1 = 1 − ∆pw /pn from table from table α0 = α(Re = 105 , mn ) (6.136) Re0 = ρ u(a0 ) D/µ0 from table ρu = ρu(a0 ) a/a0 m = ρuD2 π/4 ˙ p(x) (6.137) (6.139) diagram (6.138) (6.140) (6.142) (6.123) p(x) (6.137) (6.139) diagram (6.138) (6.140) (6.142) (6.123) (6.133) ¯ λ = 1/2(λI + λII ) (6.135) (6.127)

kg/s m2

kg/s m2 kg/s N/m2

N/m2

N/m2

N/m2 N/m2

1

additional iterations do not improve the result

316

6. Aerodynamics Laboratory

6.8.5 Problems 1. The pressure distribution measured on the wall of the pipe is to be plotted in the diagram p = f (x).

2. The resistance coeﬃcient of the oriﬁce, the ratio of the stagnation pressures p0II /p0I , and the increase of entropy ∆SI,II /R are to be plotted in the corresponding diagrams.

Resistance coeﬃcient ζB , increase of entropy, and loss of stagnation pressure of a standard oriﬁce with m = 0.5 as a function of the free-stream Mach number M a1

3. What quantities must be measured for the determination of the Mach number M aI ? The quantities hn , pw , T , Ba, and pI must be measured, if the Mach number M aI is to be determined. 4. Which ﬂow is described by the Fanno curve in the T-s diagram? Do the ﬂow conditions upstream of and downstream from a normal compression shock also lie on the Fanno curve? The Fanno curve describebs the dependence of the temperature T on the entropy diﬀerence ∆S/R for an one-dimensional, compressible ﬂow with constant mass ﬂow ρ u = const. for an adiabatic change of state. The conditions upstream of and downstream from a normal compression shock lie on the Fanno curve, since ρ u and T0 remain constant. Gasdynamic quantities as a function of the pressure ratio: uρ a0 ρ0 T0 T
0.76

=

Re a0 ρ0 ·l µ0

=f

p p0

6.8 Resistance and Losses in Compressible Pipe Flow Ma RT0 p a0 m ˙ = =f p0 a pA γ u p p Ma = f ; =f p0 a0 p0

317

with the assumptions s = s0 ; µ = µ0 T0 T γ = 1.4
0.76

The quantities to be determined from the pressure measurements are summarized above and plotted in the diagram. Resistance of adiabatic pipe ﬂow (γ = 1.4) l )max = ξmax (M a1 ) D D = diameter of pipe M a1 = Mach number at pipe entrance ξmax = max. resistance coeﬃcient at M ag = 1 2 d (p + ρ u2 ) λ = x ρ u2 d D = pipe friction coeﬃcient l1,2 ξmax1 − ξmax2 = λ D = resistance coeﬃcient of a pipe (without installations) of length l1,2 , the ¯ diameter D, and the friction coeﬃcient λ. ¯ ξmax = (λ ξmax1 − ξmax2 = ξB + ¯ λ l1,2 = D

resistance coeﬃcient with installations.

318

6. Aerodynamics Laboratory

Re m 0.1000 0.1414 0.1732 0.2000 0.2236 0.2449 0.2646 0.2828 0.3000 0.3162 0.3317 0.3464 0.3606 0.3742 0.3873 0.4000 0.4123 0.4243 0.4359 0.4472 0.4583 0.4690 0.4796 0.4899 0.5000 0.5099 0.5196 0.5292 0.5385 0.5477 0.5568 0.5657 0.5745 0.5831 0.5916 0.6000 0.6083 0.6164 0.6245 0.6325 0.6403

m2 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.10 0.11 0.12 0.13 0.14 0.15 0.16 0.17 0.18 0.19 0.20 0.21 0.22 0.23 0.24 0.25 0.26 0.27 0.28 0.29 0.30 0.31 0.32 0.33 0.34 0.35 0.36 0.37 0.38 0.39 0.40 0.41

2 · 104

2.5 · 104

3 · 104

4 · 104

6 · 104 α0

7 · 104 0.9892 0.9917 0.9946 0.9973 1.0002 1.0033 1.0064 1.0096 1.0128 1.0162 1.0196 1.0230 1.0266 1.0303 1.0341 1.0380 1.0420 1.0461 1.0503 1.0546 1.0590 1.0636 1.0682 1.0730 1.0779 1.0830 1.0881 1.0934 1.0989 1.1045 1.1101 1.1159 1.1218 1.1279 1.1341 1.1406 1.1472 1.1540 1.1609 1.1680 1.1754

105 0.9895 0.9924 0.9954 0.9984 1.0015 1.0047 1.0080 1.0113 1.0147 1.0182 1.0217 1.0263 1.0290 1.0328 1.0366 1.0405 1.0445 1.0486 1.0527 1.0569 1.0612 1.0656 1.0701 1.0746 1.0793 1.0841 1.0890 1.0941 1.0993 1.1046 1.1101 1.1156 1.1214 1.1272 1.1332 1.1394 1.1457 1.1523 1.1590 1.1660 1.1732

2 · 105 0.9895 0.9927 0.9959 0.9992 1.0026 1.0059 1.0093 1.0128 1.0163 1.0199 1.0235 1.0272 1.0309 1.0347 1.0385 1.0424 1.0463 1.0504 1.0545 1.0586 1.0628 1.0671 1.0715 1.0760 1.0805 1.0852 1.0899 1.0948 1.0998 1.1049 1.1101 1.1155 1.1209 1.1266 1.1324 1.1383 1.1445 1.1508 1.1573 1.1641 1.1710

104 - 2 · 105 0.9896 0.9928 0.9960 0.9994 1.0027 1.0061 1.0095 1.0130 1.0166 1.0202 1.0238 1.0275 1.0312 1.0350 1.0388 1.0427 1.0467 1.0507 1.0547 1.0589 1.0631 1.0674 1.0718 1.0762 1.0807 1.0854 1..0901 1.0949 1.0999 1.1049 1.1101 1.1154 1.1208 1.1264 1.1321 1.1379 1.1439 1.1501 1.1565 1.1630 1.1698

0.9768 0.9822 0.9849 0.9876 0.9907 0.9939 0.9973 1.0009 1.0048 1.0088 1.0129 1.0173 1.0219 1.0266 1.0315 1.0366 1.0418 1.0472 1.0528 1.0586 1.0645 1.0706 1.0769 1.0833 1.0899 1.0966 1.1035 1.1106 1.1179 1.1253 1.1329 1.1407 1.1486 1.1567 1.1650 1.1734 1.1821 1.1909

0.9849 0.9871 0.9895 0.9921 0.9951 0.9982 1.0015 1.0050 1.0086 1.0123 1.0163 1.0206 1.0251 1.0297 1.0344 1.0393 1.0444 1.0496 1.0550 1.0606 1.0662 1.0721 1.0782 1.0844 1.0908 1.0972 1.1037 1.1106 1.1176 1.1246 1.1320 1.1394 1.1470 1.1548 1.1627 1.1709 1.1793 1.1877

0.9883 0.9906 0.9930 0.9956 0.9984 1.0014 1.0046 1.0080 1.0116 1.0153 1.0192 1.0234 1.0276 1.0321 1.0367 1.0415 1.0464 1.0515 1.0567 1.0621 1.0677 1.0734 1.0792 1.0853 1.0914 1.0976 1.1039 1.1105 1.1173 1.1241 1.1312 1.1384 1.1457 1.1532 1.1609 1.1668 1.1768 1.1861

0.9926 0.9951 0.9976 1.0002 1.0031 1.0060 1.0092 1.0126 1.0160 1.0197 1.0235 1.0274 1.0316 1.0358 1.0402 1.0447 1.0494 1.0543 1.0593 1.0644 1.0697 1.0751 1.0806 1.0864 1.0923 1.0982 1.1042 1.1104 1.1168 1.1233 1.1300 1.1368 1.1438 1.1510 1.1583 1.1658 1.1735 1.1813

0.9951 0.9977 1.0005 1.0033 1.0063 1.0093 1.0125 1.0159 1.0194 1.0230 1.0267 1.0305 1.0345 1.0386 1.0428 1.0472 1.0517 1.0563 1.0611 1.0660 1.0710 1.0763 1.0816 1.0871 1.0928 1.0985 1.1043 1.1102 1.1164 1.1225 1.1290 1.1355 1.1423 1.1493 1.1564 1.1636 1.1711 1.1788

Discharge coeﬃcients a0 = f (m2 ,Re) for standard nozzles in smooth pipes, valid for pipe diameters D between 50 tos 500 mm. The values of m2 listed (not of m) can be interpolated linearly.

6.8 Resistance and Losses in Compressible Pipe Flow p2 /p1 m m2 0 0 0.3162 0.1 0.4472 0.2 0.5477 0.3 0.6325 0.4 0.6403 0.41 0 0.3162 0.4472 0.5477 0.6325 0.6403 0 0.3162 0.4472 0.5477 0.6325 0.6403 0 0.3162 0.4472 0.5477 0.6325 0.6403 0 0.1 0.2 0.3 0.4 0.41 0 0.1 0.2 0.3 0.4 0.41 0 0.1 0.2 0.3 0.4 0.41 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 0.98 0.9874 0.9856 0.9834 0.9805 0.9767 0.9763 0.9884 0.9867 0.9846 0.9820 0.9785 0.9781 0.9892 0.9877 0.9857 0.9833 0.9800 0.9796 0.9909 0.9896 0.9879 0.9858 0.9831 0.9827 0.96 0.9748 0.9712 0.9669 0.9613 0.9541 0.9532 0.9767 0.9734 0.9693 0.9642 0.9575 0.9567 0.9783 0.9753 0.9715 0.9667 0.9604 0.9596 0.9817 0.9791 0.9759 0.9718 0.9664 0.9657 0.94 0.9620 0.9568 0.9504 0.9424 0.9320 0.9308 0.9649 0.9600 0.9541 0.9466 0.9369 0.9358 0.9673 0.9628 0.9573 0.9503 0.9412 0.9401 0.9724 0.9685 0.9637 0.9577 0.9499 0.9490 0.92 0.90 for γ = 1.2 0.9491 0.9361 0.9423 0.9278 0.9341 0.9178 0.9238 0.9053 0.9105 0.8895 0.9090 0.8877 for γ = 1.3 0.9529 0.9408 0.9466 0.9331 0.9389 0.9237 0.9292 0.9120 0.9168 0.8971 0.9154 0.8954 for γ = 1.4 0.9563 0.9449 0.9503 0.9377 0.9430 0.9288 0.9340 0.9178 0.9223 0.9038 0.9209 0.9021 for γ = 1.66 0.9629 0.9533 0.9578 0.9471 0.9516 0.9394 0.9438 0.9299 0.9336 0.9176 0.9324 0.9161 0.85 0.9029 0.8913 0.8773 0.8602 0.8390 0.8366 0.9100 0.8990 0.8859 0.8697 0.8495 0.8472 0.9162 0.9058 0.8933 0.8780 0.8588 0.8566 0.9288 0.9197 0.9088 0.8953 0.8782 0.8762 0.80 0.8089 0.8543 0.8371 0.8163 0.7909 0.7881 0.8783 0.8645 0.8481 0.8283 0.8039 0.8012 0.8865 0.8733 0.8577 0.8388 0.8154 0.8127 0.9033 0.8917 0.8778 0.8609 0.8397 0.8373 0.75 0.8340 0.8169 0.7970 0.7733 0.7448 0.7416 0.8457 0.8294 0.8102 0.7875 0.7599 0.7569 0.8558 0.8402 0.8219 0.8000 0.7733 0.7704 0.8768 0.8629 0.8464 0.8265 0.8020 0.7993

319

Expansion coeﬃcient for standard nozzles for arbitrary gases and vapors, for diﬀerent pressure ratios p2 /p1 . Listed are also the squares of the opening ratios m and Isentropic exponent γ. The numerical values m = m2 = 0 and p2 /p1 = 1 are listed only to enable the interpolation of values of for m2 < 0.1 and p2 /p1 > 0.98.

320

6. Aerodynamics Laboratory

6.9 Measuring Methods for Compressible Flows
Abstract Optical methods for measuring density ﬁelds of gas ﬂows are introduced: The shadow method, the Mach-Zehnder interferometer , the diﬀerential interferometer, and the working principle of the holographic interferometer. For measuring the local velocity the hot-wire anemometer and the Laser-Doppler anemometer are brieﬂy described. 6.9.1 Tabular Summary of Measuring Methods 1. Optical methods for measuring the density 2. Hot-wire and Laser-Doppler anemometry for measurements of velocities and turbulent ﬂuctuation velocities 3. Pressure measurement 4. Measurement of forces and moments 5. Temperature measurement 6. Skin friction measurement 7. Visualization of streamlines 8. Concentration measurement 9. Film thermometer for measuring heat transfer 6.9.2 Optical Methods for Density Measurements The propagation velocity of light c in gases depends on the gas itself and on its density ρ. It is smaller than in vacuum. The index of refraction n1 of a gas “1” is deﬁned as the ratio of the velocity of light in vacuum c0 to its value in the gas n1 = c0 c1 . (6.143)

The optical path length S is deﬁned as the distance, which a light ray would cover in vacuum in the time t, in which the ray would cover the geometric distance l1 in the gas with index of refraction n1 S = n1 l1 . (6.144)

If a light ray crosses the boundary between two gases of diﬀerent index of refraction, it is deﬂected. If two rays travel diﬀerent optical paths, a phase diﬀerence between the two results. Methods Based on the Deﬂection of the Light If a light ray enters a transparent wedge with index of refraction n2 and wedge angle α2 , the ray is deﬂected by the angle ε, when it leaves the wedge. c1 ε sin α1 sin α2 sin α1 sin α2 sin α1 > = = = c2 , n1 < n2 α1 − α2 c1 τ /AB c2 τ /AB c1 c1 τ /AB n2 = = = c2 n1 c2 τ /AB n2 = sin α2 n1

6.9 Measuring Methods for Compressible Flows From the sketch it follows for small angles α n2 −1 ε ≈ α2 n1

321

.

(6.145)

The index of refraction n of a gas or vapor can approximately be expressed in terms of the density ρ by the following relation: n−1=K ρ Gladstone-Dale (6.146)

The Gladstone-Dale constant K depends on the gas and on the wave length of the light, for example: Air at15◦ C, λ = 644 nm → K = 0.2255 · 10−3 m3 /kg λ = 447 nm → K = 0.2290 · 10−3 m3 /kg Other examples are listed below: Gladstone - Dale K [ m3 /kg ] λ [nm] 912.5 0.2239 ·10−3 0.2274 ·10−3 509.7 0.2330 ·10−3 356.2 0.190 ·10−3 589 0.229 ·10−3 589 −3 0.238 ·10 589 Index of refraction (λ = 589 nm, T = 293 K) Air: n = 1.00027 (p = 1 bar) Water: n = 1.333 Crown glass (BK 7) n = 1.519 Flint glass (SF 10) n = 1.734 constant K Gas Air (T = 288 K) O2 (T = 273 K) CO2 (T = 273 K) N2 (T = 273 K) Coherence length of light rays Sun ∼ 10−7 m Spectroscopic lamp ∼ 10−1 m Ar - Laser ∼ 102 m He - Ne - Laser ∼ 104 m

The continuous increase of the geometric path in the medium “2” in the wedge with constant index of refraction n2 causes a deﬂection of the incident ray. It can also be concluded that an increase of the index of refraction causes a deﬂection of the light if the geometric path 1 is constant. An object, which causes the product n · l to change in an arbitrary direction, is called schlieren. If the index of refraction n remains constant in the direction of the incident ray (x) and if the geometrical path does not change in the y-direction, then the schlieren is called two-dimensional. The following approximation is valid for a weak deﬂection in the schlieren ε≈ l dn n dy dρ dy (6.147)

With (6.146) the following proportionality holds for the two-dimensional schlieren: ε∼l The Shadow Method The angles ε occuring in applications are very small ( < 1/10o ). Two methods are used to make them visible. A very simple method is the shadow method. If the light after passing through a schlieren is captured on a ﬁlm, the part of the ﬁeld, from which the light is deﬂected, generates an image which is not as dark as the image of the area, which is additionally lightened by the deﬂected light rays. (6.148)

322

6. Aerodynamics Laboratory

It is clear from the sketch above, that a constant density gradient in the phase object deﬂects the entire incident light, but does not change the distribution of the light intensity on the ﬁlm. Only a change of the density gradient leads to a brightening and darkening of parts of the image. The light intensity is approximately given by ∆E ∼ d2 ρ dy 2 (6.149)

Jumps in the distribution of the density gradient, as they occur in compression shocks, are extremely well detected. The Schlieren Method In order to build up a schlieren optic a focus (focal spot) must be generated between the object and the image screen. The focal spot serves to focus the non-deﬂected rays (regular bundle of rays). A knife edge, positioned close to the focus, captures the light, deﬂected by the schlieren to it, so that the corresponding areas on the ﬁlm are not illuminated. y regular beam of rays

knife egde schlieren boundaries collector lense light screen E

Because of the ﬁnite size of the focal spot, about half of it is covered with the knife edge, so that the optical sytem is suﬃciently sensitive. The schlieren then appear in diﬀerent grey tints, depending on their direction and intensity, which is approximately given by ∆E ∼ dρ . dy (6.150)

The schlieren method is more sensitive than the shadow method and provides very clear pictures of the ﬂow ﬁeld. Methods Based on the Measurement of Phase Diﬀerences (Interferometric Methods) If two equal light rays travel along diﬀerent optical paths, the light waves of the two rays are shifted relative to each other.

6.9 Measuring Methods for Compressible Flows

323

The shift ∆f is called phase diﬀerence. If the two light rays are superposed, they cancel each other, if ∆f is an uneven, integral multiple of one half of the wave length. If ∆f is an integral multiple of the wave length, then they amplify each other. Two light rays, which are identical in all aspects except for their phases, are called coherent. They are not coherent, if they are generated by diﬀerent light sources; they therfore have to be generated by the same light source and split into two rays, if they are to be coherent. After they have travelled diﬀerent optical paths, their phase diﬀerence can be determiend by superposition of the two rays. Mach-Zehnder Interferometer In a Mach-Zehnder interferometer the splitting and superposition of the light rays is done with semitransparent mirrors. If the mirrors 1 to 4 are positioned parallel to each other, then all geometric paths of both rays are equal. If there is a density ﬁeld in the test section, constant in the direction of the ray, then the interference fringes represent lines of constant density.

This condition is very well suited for a qualitative evaluation of a ﬂow process. For quantitative evaluation so-called ﬁnite-fringe spacing is used. By turning the corresponding mirror a linearly increasing density ﬁeld, also called wedge ﬁeld is simulated, such that a ﬁeld of parallel equidistant interference fringes, generated without a ﬂow process, appears. If a real ﬂow is investigated, the known wedge ﬁeld has to be subtracted from the ﬁeld obtained with the ﬂow. This is done by measuring the shift of the fringes in the picture of the ﬂow process in comparison to its position in the wedge ﬁeld. This method of evaluation makes it possible to catch optical errors,

324

6. Aerodynamics Laboratory

resulting from the glass panes of the wind tunnel or from the mirrors and from those that are contained in the wedge ﬁeld. Diﬀerential Interferometer In the diﬀerential interferometer the light ray is split with doubly refracting crystals, for example wollaston prisms, the action principle of which is explained in the following sketch. The distance d between both rays is so samll, that both can travel through the test section (The distance is of the order of magnitude of 1 mm).

Both rays are superposed after they have passed the test section, and the superposition yields density gradients in two-dimensional and axially symmetric ﬂows. If the wollaston prism is positioned in such a way, that with constant density in the test section interference fringes do not appear (inﬁnite fringe), then the interference fringes which appear after generating a two-dimensional density ﬁeld, are lines of constant density gradients. A suitable polarisation is necessary for generating interference fringes. A ﬁeld of equidistant interference fringes can be generated by shifting the wollaston prism. The fringe shift ∆W is then a measure for the density gradient. It is [Merzkirch 1974]: ∆W d = · sin β W λ ∂n dl + cos β ∂x ∂n dl ∂y (6.151) n = Index of refractionx d = Distance between the rays l = geometrical path For the analysis of axially symmetric density ﬁelds, Abel’s integral equation has to be solved.

6.9 Measuring Methods for Compressible Flows 6.9.3 Optical Setup

325

So far only the physical principles of some of the optical measuring methods were brieﬂy explained. In order to apply them, an optical setup must be arranged, which has to depict the object to be measured and also the signal, for example the fringes of the interferometer on the ﬁlm. In addition the setup should also provide the possibility of passing parallel light through the test section. An example of an optical setup is shown below for the so-called Z-arrangement of the diﬀerential interferometer, which is often used to keep the defects of the images small.

The light of the source is focussed by a condenser in a point, which is so chosen that it coincides with the focus of a concave reﬂector. The light rays leaving the reﬂector are parallel to each other. They pass the test section and are focussed again by a second reﬂector. Special imaging optics are required to depict the test section in a suitable size on the ﬁlm. Holographic Interferometry The methods discussed so far can only be used for the analysis of two-dimensional density ﬁelds. With the development of the Laser it became possible to use the holgraphy as a measuring technique and record the wave fronts generated by a body. A typical setup is shown below. A parallel light beam, generated by a Laser, is split with a splitter plate; one part of it, the reference beam, after being deﬂected by a mirror, is directly led to a photographically sensible layer H (Hologram). The second beam is diﬀusedly scatteerd on a ground-glass plate. The object is behind the ground-glass plate in the diﬀused light. The wave fronts, passing through the object, interfere with the reference beam and are recorded on the photographically sensible layer. For reconstruction of the image the hologram H is again exposed to the reference beam. In this manner a three-dimensional picture of the phase relationship of the light penetrating the ﬂow is generated. The method of double exposure of the hologram is preferably used, i. e., the

326

6. Aerodynamics Laboratory

photographically sensible layer is exposed without the ﬂow and a second time with the ﬂow. In this manner the change of density caused by the ﬂow can be determined. 6.9.4 Measurements of Velocities and Turbulent Fluctuation Velocities The optical methods just discussed, in general provide data which represent values averaged over the width of the test section. Hot-wire and Laser-Doppler anemometer enable the measurement of the velocity at a point with a resolution of about 1 mm3 . Hot-Wire Anemometer The hot-wire and the hot-ﬁlm anemomter are measuring devices which are heated with an electric current, so that a temperature diﬀerence is generated between the device and the ﬂowing medium (Temperature diﬀerences in gases are about 200 to 300 K). The measuring devices consist of hot wires (ø 10−6 to 10−4 m) which are ﬁxed to two pins, hot ﬁlms (thickness 10−8 to 10−6 m) are fastened to glass or ceramic supporting bodies. Two methods are used to construct a relation between the heat transfer from the measuring device to the ﬂowing medium and its local velocity: Either the heating current is kept constant and temperature or electric resistance are measured (constant-current method) or the temperature or the electric resistance are kept constant and the current is readjusted (constant-temperature method). The adjustment of the heating current can be achieved in a simple manner by combining the measuring device with a Wheatstone’s bridge. The limit of the methods are given by the spatial, temporal, and amplitude resolution. It must also be remembered that the heat transfer of the anemometer is aﬀected by a change of density or temperature of the ﬂowing medium. A single hot-wire probe is shown above. The wire has a diameter of the order of 5 µm. Modern hot-wire anemometers measure all three velocity components with three wires, orthogonally arranged to each other in a measuring volume of about 1 mm3 , gold-soldered to six prongs, which lead the electric current to the bridge. Laser-Doppler Anemometer (LDA) The Laser-Doppler anemometer is an opto-elektronic measuring method. The position of the measurement is given by the volume of intersection of two intersecting Laser beams (typical size: ≈ .1 mm3 ). The velocity of small particles, travelling with the ﬂow (ø 10−7 to 10−5 m) is determined with the aid of the light scattered by the particles and is assumed to be approximately equal to the velocity of the ﬂow. Since the particles move relative to the light source, the frequencies of the stray light of the two Laser beams is distorted as a consequence of the Doppler eﬀect, by diﬀerent amounts, resulting from the diﬀerent directions of the beams. The diﬀerence in the frequencies of the Doppler shifts is called Doppler frequency f . It can be determined from the beat frequency of the two parts of the stray light with the aid of photodetectors. Of the many possible setups the “two-beam method” is shown here. With this method the frequency f can also be explained with the interference of the two Laser beams. Selecte References Durst, F., Melling, A, Whitelaw, J.H.: Principles and practice of laser-doppleranemometrie, Academic Press, 1976.

6.9 Measuring Methods for Compressible Flows Francon, M.: Holographie, Springer-Verlag, 1972. Merzkirch, W.: Flow visualition, Academic Press, 1976.

327

Schardin, H.: Die Schlierenverfahren und ihre Anwendung, Ergebnisse der exakten Naturwissenschaften, Bd.20, 1942. Strickert, H.: Hitzdraht- und Hitzﬁlmanemometrie, VEB Verlag Technik, Berlin, 1974. Wuest, W.: Str¨mungsmeßtechnik, Vieweg-Verlag, 1969. o

6.9.5 Evaluation Determination of the Mach number by measuring the pressure: Measured data: p0 − p = 696 mm Hg p0 = 747.5 mm Hg p p0 − p ⇒ = 1− = 0.068; p0 p0
⎛

p0 = 14.51 p
⎞

Ma =

2 ⎝ p0 γ−1 p

γ−1 γ

− 1⎠ = 2.39

1. Pitot-pressure measurement at supersonic speeds

Given:

p02 = g(M a∞ ,γ) p01 pg = f (M a∞ ,κ) p∞ (Give derivation!)

To be determined:

328

6. Aerodynamics Laboratory p02 2γ g(M a∞ ,γ) = = 1+ M a2 − 1 ∞ p01 γ+1 γ γ+1

⎡

⎤

γ+1 ⎦ ·⎣ γ − 1 + M2 2 a∞ p02 = pg ; p∞ = p1
1 γ−1

γ γ−1

pg p02 p02 p01 p01 γ−1 M a2 = = · = g(M a∞ ,γ) = g(M a∞ ,γ) · 1 + ∞ p∞ p1 p01 p1 p1 2

= f (M a∞ ,γ)

2. State the diﬀerence between the shadow and the schlieren method! (optical setup, measured quantities, advantages and disadvantages) The shadow method: Changes of the density gradient cause brightening or darkening in the d2 picture, ∆E ∼ d yρ . 2 The schlieren method: The density gradients visualize themselves, ∆E ∼ d ρ . dy

3. The picture shows a photograph of a candle light, obtained with a diﬀerential interferometer as ﬁnite fringe interferogram. a) How can the density proﬁle be determined from these data? b) Sketch the picture for inﬁnite fringe spacing. Explain the interference fringes generated! c) Sketch the setup of a Mach-Zehnder interferometer for ﬁnite fringe spacing, if the interference fringes in the reference ﬁeld without the ﬂame run diagonally. a) The fringe shift is the measure for the density ⇒ density proﬁle. from equ: ∆w d sin β = w λ ∂u dl + cos β ∂x ∂u dl ∂y ⇒ ∂u ∂u ∂ρ , ; n = (K · ρ) + 1 ⇒ ∂x ∂y ∂x

b)

c)

The interference fringes are lines of constant density gradients.

If the reﬂectors 2 and 4 are not turned, the light rays pass through. A phase diﬀerence is generated and thereby an interference ﬁeld.

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge

329

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge
Abstract Several types of wind tunnels are reviewed. The working principle of supersonic wind tunnels is explained. The straight oblique compression shock in steady two-dimensional ﬂow of a perfect gas with constant speciﬁc heats is experimentally investigated. Friction forces and addition and removal of heat are neglected. The pressure ratio across the shock p2 /p1 and the shock angle σ of an oblique shock generated with a wedge are measured for a certain Mach number and compared with theoretical values. 6.10.1 Introduction The objective of wind-tunnel testing is to generate spacially and temporally constant ﬂows for the investigation of ﬂuid-mechanical problems. Studies of problems in ﬂuid mechanics often have to rely on experiments, since theoretical analyses of complex ﬂow processes are not always possible. Since experimental investigations of ﬂows about full-scale models can be realized only in rare cases, wind-tunnel testing has to resort to reduced-scale models and the similarity laws have to be observed. In addition to geometric similarity the following similarity parameters are of importance: u Reynolds number Re = ρ µ l Mach number Knudsen number Prandtl number Strouhal number Schmidt number Ma = Kn = Pr = Str = Sc = u a

Flow = Speed velocity of sound Mean free = Characteristic ath length

lm l µ cp λ ω l u ν D12

=

f l u

Kinematic viscosity = Diﬀusion coeﬃcient

Lewis number Le = P r , Sc The regimes of the ﬂow characteristics to be simulated in wind tunnels can be obtained from the missions of the air- or spaceplanes to be built. It is generally not possible to cover all ﬂight conditions with a single wind tunnel (see the following Re-Ma-number diagram), and to abey all similarity laws. In particular the Reynolds similarity law is diﬃcult to implement since either very high stagnation pressures or very large wind tunnels would be required (Re ∼ p0 · lmodel ). The following diagram shows the working areas of several wind tunnels. 6.10.2 Classiﬁcation of Wind Tunnels Classiﬁcation According to Mach Number Subsonic tunnel (incompressible) 0 < M a < 0.25 Subsonic tunnel (compressible) 0.25 < M a < 0.8 Transonic tunnel 0.8 < M a < 1.2 Supersonic tunnel 1.2 < M a < 5 Hypersonic tunnel 5 < Ma The hypersonic tunnels are designed as either cold and hot tunnels. In the former only the Mach number is simulated, in the latter also the stagnation temperature. Hot wind tunnels are also used to study real-gas eﬀects.

330

6. Aerodynamics Laboratory

shock wave tunnel

1 2 3 4

DLR Hypersonic tunnel DLR Gun tunnel DFL Gun tunnel Shock-wave tunnels (RWTH Aachen u. ISL Saint Louis)

5 6 7 8

DLR Plasma tunnel PK1 DLR Vacuum tunnel AIA Vacuum tunnel 15 x 15 cm2 AIA Vacuum tunnel 40 x 40 cm2

DLR - Deutsches Zentrum f¨r Luft- und Raumfahrt u RWTH - Rheinisch Westf¨lische Technische Hochschule a ISL - Insitute Saint Louis

Classiﬁcation According to Type of Construction Closed-circuit, continuously operating wind tunnels (G¨ttingen Type) o Open wind tunnels (Eiﬀel Type) Classiﬁcation According to Time of Operation Continuously operating tunnels Wind tunnels with closed circuit and returning of the air (Transonic tunnel G¨ttingen) o Plasma tunnel MHD tunnel (Magneto-hydrodynamic tunnel) intermittently operating tunnels Vacuum storage tunnel Pressure storage tunnel Shock-wave tunnel Injector tunnel, pipe wind tunnel Kryo-shock-wave tunnel, gun tunnel

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge

331

Vacuum-storage tunnel of the AIA (above) Pressure-storage tunnel (below)

Shock-wave tannel

332

6. Aerodynamics Laboratory

6.10.3 Elements of a Supersonic Tunnel Supersonic Nozzle with Test Section The nozzle of a supersonic wind tunnel is designed with the method of characteristics [Shapiro 1953]. The purpose of the nozzle is to accelerate the ﬂow from sonic speed (M a = 1) in the throat of the nozzle to a prescribed supersonic speed in the test section (See the following sketch and diagram). The nozzle consists of the subsonic part, the transonic part, the expansion part, and the part for correction of the ﬂow to uniform conditions in the test section. The ﬂow expands in the area (7-5-3), with several reﬂections along the wall of the nozzle. The expansion waves hitting the wall of the nozzle in the correction part (3-1) are cancelled there by generating equally strong compression waves with the concave curvature of the contour of the wall. If the nozzle geometry is correctly designed, the ﬂow in the measuring rhombus is parallel and its Mach number is constant. A change of the Mach number can only be achieved by changing the entire nozzle geometry. This can be realized with a) exchangeable nozzles b) nozzles with variable contours c) combinations of a) and b).

Data of air: γ = 1.4, T0 = 288 K M a∞ A1 /A∗ p1 /p0 u [m/s] T [K] 1 1.0 0.527 310 240 1.5 1.17 0.279 424 198 2 1.68 0.1277 507 157 3 4.19 0.0273 610 102 5 24.4 0.0019 694 47.5 10 533 +2.4 · 10−5 742 13 20 15300 2.1 · 10−7 755

Diﬀuser Analogously to the generation of the supersonic ﬂow in the Laval nozzle, the deceleration of the velocity with pressure recovery is facilitated in a convergent-divergent diﬀuser. The pressure

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge

333

recovery should be as large as possible in order to keep the required compressor power as small as possible for continuously operating tunnels or to achieve optimum measuring times for intermittently operating tunnels. The largest pressure recovery is obtained, if critical ﬂow conditions prevail in the throat of the diﬀuser and if the deceleration of the ﬂow to subsonic conditions is facilitated by a weak normal compression shock in the divergent part of the diﬀuser, immediately downstream from the throat of the diﬀuser A∗ = A∗ . Since the shock is weak, the losses caused by the shock are small. The start of the supersonic tunnel , however, requires a diﬀerent cross section of the throat of the diﬀuser, as is clear from the following consideration [Pope and Goin 1960, Voigt 1960]. During the initial phase of the start of the tunnel subsonic ﬂow prevails in the entire tunnel. The speed of sound is ﬁrst attained in the throat of the Laval nozzle. The ﬂow expands to supersonic speeds in the divergent part of the nozzle, with a normal shock terminating the supersonic part. With increasing velocity the shock moves downstream. The shock intensity and the losses in the ﬂow reach their maximum, when the shock is located in the test section. In order to guarantee that the mass ﬂow can pass the diﬀuser, the cross section of the throat of the diﬀuser must then be suﬃciently large. The continuity equation yields

ρ∗ A∗ a∗ = ρ∗ A∗ a∗ T01 = T02 → a∗ = a∗ → T∗ = T∗ ∗ ∗ ∗ ∗ A ρ p T p01 = ∗ = ∗ · ∗ = >1 ρ p T p02 A∗ Several tunnels were built with an adjustable cross section of the throat of the diﬀuser, which immediately after the start of the tunnel is reduced from A∗ to A∗ , in order to garantee a maximum pressure recovery. 6.10.4 The Oblique Compression Shock Introduction The ﬂow in a concave corner is continuously turned, if the velocity is subsonic, but discontinuously by a compression shock, if the velocity is supersonic. It is justiﬁed to consider the compression shock as a discontinuity here, since the thickness of the layer, in which pressure, density, temperature, velocity and other quatities change, is of the order of magnitude of several free mean paths only. If the turning angle does not exceed a certain angle and if the Mach number in front of the shock is suﬃciently large, in two-dimensional ﬂow the shock is straight. sho ck

334

6. Aerodynamics Laboratory

If the turning angle is large, the shock is detached from the contour and curved, normal at the wall (see sketch), so that the ﬂow is subsonic in its immediate vicinity. Further away from the wall the curvature of the shock is decreased, and the ﬂow velocity downstream from it is supersonic. A sonic line w = a separates the supersonic ﬂow from the subsonic ﬂow near the wall. Computation of the Change of State Across the Oblique Compression Shock Assumptions 1. 2. 3. 4. The ﬂow is steady and two-dimensional. Gravitational and friction forces can be neglected. Heat transfer with the surroundings can be neglected. The ﬂowing medium is a perfect gas with constant speciﬁc heats.

Basic Equations With a small change in notation the jump conditions for the oblique shock are listed again in the following. They were already derived in section 4.5, but are needed here for comparison of the theoretical results with the measured data. With the assumptions listed, the conservation equations for the control surface indicated in the sketch are the following: Continuity equation: ρ1 w1n = ρ2 w2n (6.152)

Momentum equation normal to the shock: p1 − p2 = ρ2 w2n w2n − ρ1 w1n w1n (6.153) Momentum equation tangential to the shock: 0 = ρ2 w2n w2t − ρ1 w1n w1t Energy equation: γ p1 1 2 γ p2 1 2 2 2 = (w2t + w2n ) + (w + w1 n ) + 2 1t γ − 1 ρ1 2 γ − 1 ρ2 sin σ tan σ tan(σ − β) sin(σ − β) = = = = w1n /w1 w1n /w1t w2n /w2t w2n /w2 (6.154) (6.155)

The following relations can be derived from the sketch: (6.156) (6.157) (6.158) (6.159)

After introducing the Mach number in front of the shock, M a1 = w1 p1 = w1 / γ a1 ρ1 , (6.160)

with a being the local speed of sound, (6.156) can be rewritten in the following form:
2 M a2 sin2 σ = w1 n / γ 1

p1 ρ1

(6.161)

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge Solution It follows from (6.152) to (6.155) and (6.161): w1t = w2t n 1 + γ−1 γp11/ρ1 1 + γ−1 M a2 sin2 σ w2n ρ1 w2n 2 = = = γ+12 2 1 2 (γ,M a,σ) = 2 w1 n γ+1 w1n ρ2 w1n M a1 sin σ 2

335

(6.162) w2 (6.163)

2

γp1 /ρ1

p2 2γ = 1+ p1 γ+1

w1 2 2γ p2 n (M a2 sin2 σ − 1) = (γ,M a,σ) (6.164) −1 =1+ 1 γp1 /ρ1 γ+1 p1

With (6.162) there is obtained from (6.157) and (6.158): w2n tan(σ − β) , = w1n tan σ and further ⎞ ⎛ γ+1 M a2 ⎟ ⎜ 2 ⎟ cot β = tan σ ⎜ ⎝ M a2 sin2 σ − 1 − 1⎠ ; σ = σ(γ,M a,β) .

(6.165)

(6.166)

Equations (6.164) and (6.166) yield a relation for the construction of the heart-curve diagram, after elimination of σ
⎡

tan β = ⎣

2

M a2 1 1+ γ+1 2γ p2 p1

⎤ ⎡

−1

− 1⎦ · ⎣

p2 p1

−1 p2 p1

⎤2 ⎦

γM a2 − 1

−1

.

(6.167)

It follows from (6.156) and (6.159) that w2n sin σ w2 = · w1 w1n sin(σ − β)

.

(6.168)

This equation together with (6.166) and the critical speed of sound a∗ , 2γ a∗2 = R · T0 , γ+1 after elimination of σ, with R being the gas constant, and T0 the stagnation temperature, leads to the expression for the shock polar: w1 w2 cos β − a∗2 (w2 sin β)2 = (w1 − w2 cos β)2 . (6.169) 2 2 w1 − w1 w2 cos β + a∗2 γ+1 Heart-Curve Diagram and Shock Polar The heart-curve diagram and the shock polar are sketched for constant γ for a given free-stream Mach number.

heart-curve diagram

shock polar

336

6. Aerodynamics Laboratory

Only the physically meaningful regions are shown (p2 /p1 ≥ 1 and w2 /w1 ≤ 1). Two solutions exist for a prescribed value of β indicated by the symbols ◦ and •). In the actual ﬂow the weak solution (◦) with the smaller p2 and the larger w2 is observed in straigth oblique shocks. For the case thatβ = 0 the heart-curve diagram and the shock polar contain the solution for the normal compression shock (B) and the trivial solution for the unperturbed ﬂow (A). Closedform solutions for the straight shock do not exist for turning angles β > βmax ; the turning of the ﬂow takes place in a curved shock. Supplementary Remarks The Hugoniot relation: With (6.152) to (6.155) the energy equation is cast into the following form, named after Hugoniot (γ + 1) p2 + (γ − 1) ρ2 p1 = , 2 ρ1 (γ − 1) p1 + (γ + 1) p or in the form given by von K´rm´n a a p2 + p1 p2 − p1 =γ . ρ 2 − ρ1 ρ2 + ρ1 (6.171) (6.170)

It follows from (6.170) for the very strong shock (p2 /p1 → ∞), that the maximum compression ratio of an perfect gas with constant speciﬁc heats is ρ2 ρ1 = p2 /p1 →∞

γ+1 γ−1

(6.172)

Vice versa, it follows from (6.171) for the weak shock (p2 /p1 → 1) ∆p ∆ρ =γ p/p→1 ˆ

p = ρ

dp dρ

(6.173) isentrop. Very weak shocks can in a ﬁrst approximation be considered as isentropic compression waves. The maximum turning angle: For γ = 1.4 the maximum turning angle βmax. has the following values: βmax. Ma Construction of the Shock Polar All velocity components are non-dimensionalized with the critical speed of sound a∗ ,the speed of sound in the throat of the Laval nozzle, in which the ﬂow expands in the divergent part to supersonic conditions corresponding to the Mach number M a. The relation connecting the velocity non-dimensionalized with a∗ and M a is w w a ≡ M a∗ = · ∗ = M a a∗ a a γ+1 2 γ−1 M a2 1+ 2 0o 1 23o 2 34o 3 42o 5 44o 8 45o ∞

.

(6.174)

The shock polar was explained in Sect. 4.5.3, where it was introduced in conjunction with the transformation of the ﬂow quantities from the physical plane into the hodograph plane. The shock polar oﬀers a convenient means of determining the velocity immediately downstream from an oblique shock.

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge If the velocity components in front and downstream from the shock are plotted as indicated in the sketch, it is seen that for given values of M a,γ,β und w1 the quantities w2 ,wn2 ,wt2 , and σ can be determined from the shock polar.

337

6.10.5 Description of the Experiment The experiment is carried out in the intermittently operating supersonic wind tunnel of the Aerodynamisches Institut. The air is sucked from the surroundings (p0 , T0 ) through the Laval nozzle and the test section into the vacuum storage.

A supersonic ﬂow with a Mach number of M a ≈ .2.0 in the test section is generated by expansion of the air in the Laval nozzle. The Laval nozzle is a continuously adjustable nozzle and admits a variation of the test Mach number from about M a = 1 to M a = 5. The contour of the nozzle is generated with ﬂexible steel sheets, which are fastened at one end of the nozzle frame. The sheets are held in position by six supporting jacks. The jacks are adjusted in such a way, that the metal sheets take on the form of the nozzle contour required for the test Mach number. The model support enables a change of the angle of attack of the model. It can be varied in steps of one half, one, and two degrees. The maximum angle of attack the model can be held at is 39.5◦ . In order to obtain a symmetric ﬂow about the wedge, it is turned with the aid of its support until the pressure diﬀerence p2 2 − p2 3 measured at the two symmetrically positioned pressure holes vanishes. After this directional adjustment of the wedge, the pressure diﬀerences p0 − p, p0 − p2 , and p0 − pexp are measured with U-tubes. The shock angle σ and the Mach angle α of a Mach line of the free stream are visualized with a schlieren picture:

338

6. Aerodynamics Laboratory

The Mach line, inclined by the angle α, can be recognized in the photograph. In the actual ﬂow it is a very weak oblique compression shock, generated by a small perturbation of the ﬂow, caused by a piece of Scotch tape, pasted to the wall.

The sketch shown above depicts the arrangement of the pressure measurement on the wedge. The U-tube on the left measures the pressure ratio p∞ /p0 . The U-tube in the middle of the sketch measures the pressure ratio p21 /p0 , i. e. the ratio of the static pressure on the wedge to the stagnation pressure. The third U-tube measures the pressures p22 and p23 , which serve to align the model with the direction of the free stream. The angle of attack of the wedge is changed until p22 = p23 . The U-tube on the right measures the pressure downstream from the shoulder, after the expansion, pexp. , referenced to the stagnation pressure p0 . Also indicated is the Mach angle a of the free stream, as made visible by a roughness element of the wall of the wind tunnel. The lines at the nose of the wedge indicate the shock angle σ for the weak solution. Selected References Wuest, W.: Str¨mungsmeßtechnik, Vieweg-Verlag, 1969. o Pope, A., Goin, K.L. High-Speed Wind Tunnel Testing, Jihn Wiley, 1965. Oertel, H.: Stoßrohre, Springer-Verlag, 1966. Naumann, A.: Moderne Entwicklung bei Windkan¨len, Die Umschau in Wissenschaft und a Technik, Heft 20, 1964, S. 613. Naumann, A.: Aerodynamische Gesichtspunkte der Windkanalentwicklung, Jahrbuch WGL, 1954. ¨ Heyser, A.: Vorlesung: Versuchsanlagen der Str¨mungstechnik f¨r Unter-, Uber- und Hypero u schallgeschwindigkeiten, RWTH Aachen, 1972. Shapiro, A.H.: The Dynamics and Thermodynamics of Compressible Fluid Flow, Ronald Press, Vol. I, 1953, S. 507-512. Cummings, J.C.: Development of a high-performance cryogenic shock tube, Journal of Fluid Mechanics, Vol. 66, Part 1, 1874, S. 177-187. ¨ Voigt, F.: Vorg¨nge beim Start einer Uberschallstr¨mung, Dissertation, RWTH Aachen, 1960 a o Prandtl, L. F¨hrer durch die Str¨mungslehre, Vieweg-Verlag, 1965, S.302. u o ¨ Flugge, S.: Handbuch der Physik: Str¨mungsmechanik, Springer-Verlag, Band IX, 1959. o

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge 6.10.6 Evaluation Measured Data p0 = 747 mm Hg p0 − p1 = 690 mm Hg p0 − p2 = 632 mm Hg β = 5.75o γ = 1.4 p0 − pexp = 700 mm Hg Determination of the Mach Number 1. Ma = 1 = 2.13 sinα
⎡

339

p0 1 = 13.1 = p 1 − (p0 − p1 )/p0 p2 1 − (p0 − p2 )/p0 = = 2.017 p1 1 − (p0 − p1 )/p0

α is determined from photograph γ−1 κ
⎤

2.

Ma =

2 ⎢ p0 ⎢ ⎢ γ − 1 ⎣ p1

− 1⎥ = 2.33

⎥ ⎥ ⎦

Determination of σ 1. 2. σ = 33.0◦ σ = 39.9◦ from photograph from sin σ = sin α · γ+1 2γ
∗

p2 −1 +1 p1

3.

σ = 32.5

◦

γ+1 2 from shock polar (M a = M a = 1.767) γ−1 M a2 1+ 2

Determination of the Pressure Ratio p2 /p1 1. 2. p2 = 2.017 p1 p2 = 1.4 p1 from measurement from heart-curve diagram

Questions 1. Sketch the contour of the shock in the sections (1, 2, 3) parallel to the side walls of the tunnel.

340

6. Aerodynamics Laboratory The sketch shown above depicts the top view of the wedge in the test section of the wind tunnel. The model extends to both side walls of the tunnel. As indicated in the sketch by the dashed curved lines a turbulent boundary layer develops on the side walls. The thickness of the boundary layer is strongly magniﬁed. Also indicated in the sketch are 3 sections, one in the middle of the tunnel (section 1), one closer to the side wall (section 2), and the third near the wall (section 3). Case 1: M a = 2.13 Case 2: Ma > 1 σ > 33◦ or detached shock for β > βmax

Case 3: M a < 1, no shock, ﬂow decelerated by wall friction or M > 1 ⇒ detached shock

2. The pressures in front of and downstream from a normal compression shock are measured with two Pitot tubes. What pressures are measured and how do the readings diﬀer from each other? (It is assumed that the shock is not aﬀected by the Pitot tube positioned in front.) The total pressure downstream from a normal compression shock is measured. The reading is the same in both cases. 3. Given is a wedge at an angle of attack with β = 20o at a free-stream Mach number M a∞ = 2.91.

How large can the angle of attack α become, without having the compression shock detach from the wedge? βmax = 33.8◦ from heart-curve diagram here βmax = β + αmax for lower side ⇒ αmax = 13.8◦ pl − pu How large is then the pressure diﬀerence ? p∞ pl −pu = 6.43 p∞

6.10 Supersonic Wind Tunnel and Compression Shock at the Wedge

341

Heart-curve diagram for γ = 1.4

342

6. Aerodynamics Laboratory

6.11 Sphere in Compressible Flow
Abstract The drag of a sphere is investigated experimentally in compressible subsonic ﬂow. The drag coeﬃcient cD depends on the Reynolds number Re and on the Mach number M a. The inﬂuence of the critical Mach number on the separation of the boundary layer is studied with two spheres of diﬀerent diameter. The drag of the spheres is measured for several free-stream velocities. The diﬀerent ﬂow patterns of sub- and supercritical Mach numbers M a are visualized with the schlieren method. 6.11.1 Introduction In this experiment the inﬂuence of the compressibility of the air on the ﬂow ﬁeld about a sphere is studied in an experiment. The compressibility has to be taken into account in steady ﬂows, if the density changes are larger than about one percent. The drag coeﬃcient cd then also depends on the Mach number in addition to the similarity parameters mentioned already in the experiment sphere in incompressible ﬂow. The drag coeﬃcient of the sphere measured in the wind tunnel cD = with D 2 q∞ = ρ w∞ 2 DS drag, dynamic pressure of the free stream, diameter of the sphere, D 2 q∞ π DS 4 (6.175)

depends on the following similarity and other parameters: cD = f (Re,M a∞ ,γ, ks AW T ,T u, , support,Kn,...) DS AS ρ∞ u∞ DS µ∞

(6.176)

The quantities mentioned in the above equation are deﬁned as follows: Re∞ M a∞ γ ks DS

Reynolds number

Tu
AW T AS

Kn

∞ Mach number w∞ a ratio of the speciﬁc heats relative roughness of the surface Turbulence intensity of the free stream ratio of the cross-sections tunnel-sphere Knudsen number λ∞ with λ∞ = mean free path. DS

In (6.179) the function f is unknown. In the experiment only the Reynolds number Re∞ and the Mach number M a∞ are varied; the other parameters, which are given by the experimental facility, are kept constant. 6.11.2 The Experiment Experimental Facility The experiment is carried out in the intermittently operating vaccum-storage tunnel of the Aerodynamisches Institut, with a test section of 15 · 15 cm2 . Air is sucked out of the atmosphere through a well-rounded intake and the test section into the vacumm tank. A convergentdivergent diﬀuser with an adjustable throat is positioned between the test section and the tank. If the ratio of the pressures in the tank and in the atmosphere is suﬃciently small, the velocity

6.11 Sphere in Compressible Flow

343

of the air in the throat of the diﬀuser is equal to the speed of sound. With this arrangement it is guaranteed, that despite of the increasing pressure in the tank the ﬂow conditions in the test section remain constant for a certain time. The similarity parameters Re and M a are coupled by the following relation, if it is assumed, that the viscosity µ = µ(T ) is konwn: Re = ρ0 a0 µ0 ρ u DS ρ u a∞ µ0 1 ρ0 · a0 · = · ρ0 · · · a0 · · DS = f (M a∞ ) · · DS a∞ a0 µ µ0 µ0 µ ρ0 (6.177)

Stagnation density of the air Speed of sound at stagnation conditions viscosity

In a vacuum-storage tunnel with constant ρ0 ,a0 , and µ0 , with the Mach number M a held constant, the Reynolds number Re∞ can only be changed by changing the diameter of the sphere. Two spheres with diameters of 25 mm and 50 mm were used in the experiments. The spheres were mounted on the model support, which was positioned in the dead-water region of the spheres. The drag was measured with a strain-gauge balance mounted in the support. Balance for Measuring the Drag The balance mounted in the support measures the drag of the sphere by recording the change of the electric resistance, caused by the elongation of an electric conductor (strain gauge). The installation of the support and the balance are sketched below.

The spring elements of the balance are deformed when they are exposed to strain. The deformation is measured with two strain gauges connected to a bridge circuit. The strain gauges are arranged in such a way that suﬃcient temperature compensation is guaranteed. The balance was calibrated in an extra experiment prior to the drag measurement by taking the balance out of the support and loading it with weights. The inﬂuence of the support on the drag of the sphere could not be taken into account.

344

6. Aerodynamics Laboratory

Description of the Measurement The drag of two spheres with 25 mm and 50 mm diameters is measured for several free-stream velocities w. The free-stream velocity is varied by varying the cross section of the throat of the diﬀuser, where sonic conditions prevail. The unperturbed free-stream conditions are determined with a measurement of the pressure. A sketch of the pressure measurements is given on the data sheet. The ﬂow ﬁeld about the sphere is visualized with the schlieren method. 6.11.3 Fundamentals of the Compressible Flow About a Sphere Potential Flow According to potential theory, the ﬂow of an incompressible medium about a sphere is obtained by superposing a spacial dipole with a parallel ﬂow. Some of the streamlines are shown below.

The velocity and the pressure on the surface of the sphere in incompressible ﬂow are described by the following relations: w 3 = sin θ w∞ 2 9 = 1 − sin2 θ on r = Rs 4 (6.178) (6.179)

ps − p∞
∞ ρ u2 2

Lamla [1939] investigated the potential ﬂow of a compressible gas about a sphere. The following diagram shows the ratio of the velocities wcompr. /wincompr. for the surface of the sphere. In the vicinity of the upstream stagnation point the deceleration of the compressible ﬂow is more pronounced, and near the equator the acceleration is stronger than in the incompressible ﬂow. According to Lamla [1939] the critical Mach number of the free stream M a is given by the condition, that the velocity at the equator becomes sonic, which with γ = 1.4 yields a value of M acrit. = 0.57 M a∗ = 0.6 crit. (6.180)

The velocity wcompr. at the equator for M acrit. is about ten percent higher than the velocity computed for incompressible ﬂow.

6.11 Sphere in Compressible Flow

345

Deﬁnition of the Critical Reynolds Number The location of the separation line and the size of the dead-water region are strongly inﬂuenced by the nature of the ﬂow in the boundary layer, which can either be laminar or turbulent. If the laminar-turbulent transition occurs upstream of the separation line, it is shifted from the upstream to the downstream side of the equator and reduces the dead-water region. The Reynolds number Re, at which the transition occurs, is called critical Reynolds number. Deﬁnition of the Critical Mach Number An increase of the Mach number M a in subsonic ﬂow leads to an increase of the velocity in the vicinity of the equator, and, aided by the displacement eﬀect of the sphere, eventually to the formation of a local supersonic pocket, which is terminated by a shock. The lowest free-stream Mach number, at which the local velocity is equal to the local speed of sound, is called critical Mach number. The Dependence of the Flow Field on Reynolds and Mach Number As previously mentioned, the magnitude of the drag coeﬃcient depends on the location of the separation line, which in turn is inﬂuenced by the nature of the ﬂow in the boundary layer. It was, however, shown in experiments, that for supercritical Mach numbers separation in the vicinity of the equator is also inﬂuenced by the shock, terminating the supersonic pocket. The following diagrams and sketches on the left show the ﬂow patterns, if the critical Reynolds number Recrit. is attained prior to the critical Mach number M acrit. . When the critical Reynolds number Recrit. is reached, the drag coeﬃcient cD decreases abruptly. The ﬂow separates downstream from the equator. When the critical Mach number is reached, a shock is generated at the equator, which causes the boundary layer to separate. The drag coeﬃcient cD rises again. If M acrit. is reached prior to Recrit. (diagram and sketches on the right below), the shock prevents the drag coeﬃcient

346

6. Aerodynamics Laboratory

cd from dropping abruptly at the critical Reynolds number, since the separation line, usually observed to move downstream, when Recrit. is reached, is now ﬁxed by the shock at the equator. The diﬀerent cD -distributions can be explained by using spheres with diﬀerent diameters; the results for the sphere with the larger diameter are shown on the left, those for the sphere with the smaller on the right.

Schematic diagram of the drag coeﬃcient and the ﬂow patterns as a function of the free-stream velocity for large diameters of the sphere.

Schematic diagram of the drag coeﬃcient and the ﬂow patterns as a function of the free-stream velocity for small diameters of the sphere.

The free-stream velocities, at which Recrit. and M acrit. are observed, are wRe,crit. = wM a,crit. Recrit. µ∞ ρ∞ DS = M acrit. a∞ . (6.181) (6.182)

If the free-stream conditions are kept constant, then wM acrit. = const., and wRecrit. ≈ 1/DS . Therefore for large diameters wRecrit. < wM acrit. (see diagram on the left), and for smaller diameters wRecrit. > wM acrit. (see diagram on the right). In the limiting case wRecrit. = wM acrit. the critical Reynolds number and the critical Mach number are reached at the same free-stream velocity. The diameter of the sphere, correponding to this condition is somwhere between 30 and 50 mm. Experimental Results The experimental results cD = f (Re,M a) of Naumann and Walchner, shown in the following two diagram conﬁrm these considerations, [Naumann 1953]. The curves in the ﬁrst of the following diagrams show the inﬂuence of the free-stream Mach number on the drag coeﬃcient cD for diﬀerent diameters of the sphere. It can be seen, that for small diameters up to approximately 30 mm M acrit. is reached before Recrit. . After M acrit. is exceeded, the drag coeﬃcient increases slightly, as the shock intensity and also the ﬂuid mechanical losses are increased. For spheres with diameters larger than 50 mm, Recrit. is reached before M acrit. . The abrupt drop of the drag coeﬃcient cd is observed as in incompressible ﬂow. If the critical Mach number is exceeded, the drag coeﬃcient cD increases sharply. In the second diagram the drag coeﬃcients measured by Naumann [1953] are shown as a function of the Reynolds number Re and of the Mach number M a.

6.11 Sphere in Compressible Flow

347

348

6. Aerodynamics Laboratory

Selected References Bailey: Sphere drag coeﬃcient for subsonic speeds in continuum and free molecular ﬂows, J. Fluid Mech., Vol. 65, 1974, pp. 401-410. Lamla, E.: Symmetrische Potentialstr¨mung eines kompressiblen Gases Kreiszylinder und o Kugel im unterkritischen Gebiet, F.B. 1014, DVl, 1939. Naumann, A.: Luftwiderstand der Kugel bei hohen Unterschallgeschwindigkeiten, Allgemeine W¨rmetechnick 4/10, 1953. a

6.11.4 Evaluation The drag coeﬃcients cD = f (Re,M a) are to be computed from the measured data given on page 350: All equations, needed for the computation are listed on the data sheet. The drag coeﬃcients cD = f (M a,Ds ) and cD = f (Re,Ds ) are to be plotted and to be discussed. The experiment and the results of the measurements are to be commented critically.

For Ds = 50 mm the drag coeﬃcient cD drops abruptly at Re ≈ 3.5·105 ≈ Recrit. and M a = 0.3, caused by the sudden shift of the separation line further downstream from the equator, reducing the dead-water region and also the drag coeﬃcient. At M a = M acrit. ≈ 0.55 the drag coeﬃcient cD increases sharply, since the ﬂow velocity exceeds the local speed of sound just upstream of the equator and becomes supersonic, with a shock terminating the supersonic pocket near the equator. The sudden pressure rise enforces ﬂow separation just downstream from the shock, and the dead-water region is immediately increased. The drop of the drag coeﬃcient cD is not observed in the experiment with the sphere with a diameter of 25 mm, as in this case M acrit. is reached before Recrit. and the shock prevents the shift of the separation line. In principle, measurements of the drag are rather inaccurate, which explains the deviation of the data from each other. 6.11.5 Questions 1. Sketch the ﬂow ﬁeld about a sphere for the following free-stream conditins: a.) M a = 0, Re > Recrit. b.) M a > M acrit. , Re > Recrit.

6.11 Sphere in Compressible Flow

349

2. Why does the drag coeﬃcient cD increase with increasing Ma for M acrit. < M a < 1? The shock moves upstream, which causes an enlargement of the dead-water region. 3. How large is the density change at the equator in a potential ﬂow about the sphere (ρ∞ − ρ)/ρ∞ , if the free-stream Mach number M a∞ = M acrit. ? Assume isentropic change of state (γ = 1.4,air). ρ0 = 1.17 ρ∞ ρ∞ − ρ ρ ρ ρ0 = 1− · =1− ρ∞ ρ∞ ρ 0 ρ0 M∞ = 0.57 →

ρ0 ρ∞

= 1 − 0.634 · 1.17 = 0.258

4. Up to what ﬂow velocity can the inﬂuence of the compressibility of dry air on the ﬂow be neglected (T0 = 293K)? ρ∞ − ρ > 0.01 ⇒ compressible ﬂow ρ∞ ρ < 0.99 ⇒ M a > 0.14 ⇒ ρ0 m u∞ = M a γ R T0 ⇒ u∞ ≥ 48 s The inﬂuence of the compressibility can be neglected up to a ﬂow velocity of about 48 m , since s then the relative density changes are below 1%. Calibration values of the strain gauge balance loading F [N] 0 0.71 2.71 4.71 6.71 8.71 10.71 15.71 20.71 [Units] 0 1.4 4.7 8.1 11.6 15.0 18.5 27.1 35.7 Reading loading Reading unloading [Units] 0.2 1.5 4.9 8.4 11.8 15.1 18.5 27.1 35.7

Calibration curve

350

1 DK 7 p1 p0 u a0

2 ∆p0 [ms−1 ] DIA 0.160 0.230 0.305 0.370 0.445 0.480 0.520 0.555 54.9 79.0 106.5 133.9 166.6 183.7 202.6 221.5 54.9 65.2 79.0 91.0 106.5 0.987 0.982 0.974 0.964 0.952 1.17 1.16 1.15 1.14 1.13 1763 2466 3589 4720 6408 1.7 2.6 3.1 4.7 4.5 0.49 0.54 0.43 0.51 0.36 1.79 2.12 2.57 2.91 3.41 0.987 0.974 0.952 0.926 0.885 0.863 0.836 0.804 1.17 1.15 1.13 1.10 1.05 1.02 0.99 0.95 1763 3589 6408 9861 14572 17210 20318 23305 0.58 0.45 0.51 0.48 0.55 0.54 0.55 0.61 0.89 1.29 1.70 2.09 2.49 2.68 2.91 3.10 0.160 0.190 0.230 0.260 0.305 [Skt] Me 1.1 1.5 2.9 4.2 7.0 8.2 9.7 12.4 [N] Calibr. 0.5 0.8 1.6 2.3 3.9 4.6 5.5 7.0 [ms−1 ] [kgm−3 ] ρ1 ρ0

3 p0 [kgm−3 ] [Nm−2 ] a0 ρ0 DK µ0

[mm] Me 25 0.984 0.964 0.934 0.897 0.844 0.813 0.777 0.737 0.984 0.975 0.964 0.951 0.934 0.160 0.190 0.230 0.270 0.315 0.160 0.190 0.230 0.265 0.310 DIA 0.160 0.230 0.315 0.395 0.500 0.550 0.610 0.675 DIA 0.160 0.230 0.310 0.390 0.485 0.535 0.590 0.645

[mmH2 O] Me

[Nm−2 ]

4 ∆p1

5 ∆p1

6 p1

8 Ma

9

10 R

11 a0

12 u1

13 ρ0

14

15 ρ1

16 q1

17 W

18 W

19 cD

20

21 Re · 10−5

[mmHg] Me 12 27 49 77 117 140 167 197

[Nm−2 ]

[Nm−2 ]

1601 3602 6537 10273 15610 18678 22280 26283

98194 96193 93258 89522 84185 81117 77515 73512

6. Aerodynamics Laboratory

50

12 19 27 37 49 77 94 117 128 149 0.844 0.829 0.801 0.500 0.525 0.575 0.485 0.510 0.555 0.445 0.465 0.495 166.6 175.1 190.6 0.885 0.874 0.852 1.05 1.04 1.01 14572 15943 18346 2.1 2.8 16.1

1601 2535 3602 4936 6537

98194 97260 96163 94859 93258

15610 17077 19878

84185 82718 79917

3.0 4.7 5.5 8.4 8.0 2.5 3.1 3.8 5.0 28.3

0.07 0.08 0.45

4.98 5.20 5.53

(3) (5) (6) (11) (12) (13) (14)

p0 ∆p1 p1 a0 u1 ρ0 ρ1 ρ0

= = = = = = =

pa − ∆p0 ∆p1 · 133.3 pa − ∆p1 √ γ · R · T0 u1 a0

(15) (16) (19)

ρ1 q1 cD

= = =

ρ1 ρ0 1 2

ρ0 ρ1 u 2 1
D q1

748

20.5

π D2 S 4

Ba pa ta Ta R γ µ0

= = = = = = =

[mmHg] [Nm−2 ] [◦ C] [K] [m2 s−2 K−1 ]

(21)

Re

=

κ

a0 ρ0 DS µ0

287 1.4 1.82 · 10−5

[Nsm−2 ]

ρa ·

p0 pa 1 1+ 2 γ−1 1 γ−1

M a2 1

Index

aﬃne velocity proﬁles 59 airfoil 54, 65, 174, 180, 181 d’Alembert’s paradox 52 d’Alembert’s solution 172 apparent shear stresses 26, 62 atmospheric pressure 1, 4, 15 axially symmetric ﬂow 176 barometer 4 barometric height formula 3 bent pipe 13 Bernoulli equation 35, 48, 53, 57, 143, 145, 146 Bingham model 22 Blasius law 29 Blasius solution 60 boundary layer 55, 56, 91, 132, 262, 265–267 boundary layer separation 64, 254, 256, 342 boundary-layer equations 56–59, 61, 65 boundary-layer thickness 56, 57, 61 Buckingham’s Π theorem 38 calorically perfect gas 140, 141, 143 capillary viscometers 25 capillary waves 300 Carnot’s equation 15 Cauchy-Riemann diﬀerential equations 47 cavitation 11 characteristic curves 167, 168, 170 circular cylinder 51, 52, 54, 65 circulation 44, 45 communicating vessels 3, 4 compatibility conditions 166, 167, 169 complex stream function 48 compressible potential ﬂow 139, 170, 199, 221 compression 147, 151, 155 compression shock 147, 149, 151, 188, 191, 193, 208, 211, 214, 329, 333 computation of supersonic ﬂow 139 continuity equation 7, 13, 14, 31, 32, 44, 46, 61, 145 continuum 1, 2, 12 contraction 9, 10, 33 contraction ratio 15

control surface 12–16, 53 control volume 16, 32 convective acceleration 8 corner ﬂow 158 creeping motion 41 critical depth 17 critical free-stream Mach number 180 critical Mach number 144, 145, 181, 346 critical Reynolds number 61, 261, 345, 346 critical speed of sound 144 critical state 144 Crocco vorticity theorem 163 dead water region 14 diﬀerential pressure 10 dimensional analysis 38, 39 dipole 50 dipole moment 50 discharge coeﬃcient 10 discontinuous widening of a pipe 14 dispersion 55 displacement thickness 58 dissipation 150 dissipation function 37 drag 53, 65, 66, 86, 91, 92, 134 drag coeﬃcient 162, 196 drag of a ship 42 dyadic product 67 dynamic ﬂow condition 54 dynamic pressure 8, 9 dynamic shear viscosity 21, 22, 25, 31 Eiﬀel-type wind tunnel 234, 245 energy equation 8, 11, 20, 35–38, 40, 143, 144, 148, 150, 151 energy height 17 enthalpy 37, 140, 141, 143, 150, 164 entropy 139–142, 147–150, 160, 161, 163, 164, 166, 167 equilibium of moments of momentum 12 equipotential lines 47 Euler equations 41, 47, 48, 57 Euler number 39, 41 Euler’s turbine equation 19

352

Index isentropic exponent 41, 141 isothermal atmosphere 3 jump conditions 147, 151 von K´rm´n vortex street 66 a a von K´rm´n’s constant 27 a a von K´rm´n’s integral relation 58, 59 a a von K´rm´n’s velocity defect law 63 a a kinematic ﬂow condition 47, 54 kinematic viscosity 21 Kutta condition 54 Kutta-Joukowski theorem 53 Lagrange’s method 5 Lagrangian particle path 5 laminar boundary layers 55, 262 laminar ﬂow 20, 24, 30, 57 Langrange’s method 6 Laplace equation 47–49, 53, 54 Laser-Doppler anemometer 320 Laval nozzle 146 lift 54, 65 limiting velocity 145 linearized potential equation 139 liquid manometer 3 local acceleration 6, 8, 11 Mach angle 152, 154, 155, 157, 166 Mach cone 143, 178 Mach line 159, 167, 169 Mach number 41, 142–146, 148, 150, 152, 154, 156–159, 161, 168, 178, 180, 181 Mach-Zehnder interferometer 320 Magnus eﬀect 52 manometer 4 mechanical energy 8, 35, 37 method of diﬀerential equations 31 moment of momentum theorem 12, 18, 76, 106 moment of reaction 19 momentum 12 momentum equation 26, 40, 43, 53, 57–60, 64 momentum theorem 12–16, 18, 19, 26 momentum thickness 59 Navier-Stokes equations 32, 35, 36, 45, 46 Newton’s law 7 Newtonian ﬂuid 21, 24 Non-Newtonian ﬂuids 22 normal compression shock 139, 147, 148, 150, 160 normal stress 22, 26 nozzle ﬂow 148, 169, 226

Eulerian method 5, 6 expansion 139, 157–159 explosion 142 Fanno curve 310, 316 ﬂat plate 56, 59, 61, 62 ﬂat plate turbulent boundary layer 62, 265, 269 ﬂow-measurement regulation 10 ﬂuid coordinates 5 ﬂux of vorticity 44, 45 form drag 66 Fourier heat ﬂux 36 friction coeﬃcient 56, 59, 60, 63 friction velocity 27 Froude number 17, 39, 41 Froude’s Theorem 16 fully developed pipe ﬂow 25 fundamental gasdynamic equation 139 fundamental hydrostatic equation 3 gas constant 41, 139 Gauss integral theorem 31 Goethert rule 180, 181 G¨ttingen-type wind tunnel 233–235, 246 o gravity waves 17, 41, 54, 300 Hagen-Poiseuille law 24, 25 half body 51, 247, 248, 251 heart-curve diagram 154, 193, 230, 335, 339, 341 heat ﬂux 36, 37 heat source 35 Hele-Shaw ﬂow 247 hodograph plane 154 Hooke’s law 21, 33 hot-wire anemometer 326 Hugoniot relation 336 hydraulic diameter 29, 30 hydraulic jump 18 hydraulic press 4 hydrodynamics 5, 71, 99 hydrostatic lift 5 hydrostatic paradox 4 hydrostatics 2, 69, 96 incompressible potential ﬂows 49 inertia forces 42, 56, 62, 64 inﬂuence quantity 38, 40 intake pipe 11 intake region 25 interaction 139 internal energy 35, 140 irrotationality 44, 46

Index oblique compression shock 191, 194, 211, 228, 230 one-dimensional isentropic ﬂows 145 open channel ﬂow 17, 18, 29, 41 oriﬁce 10, 307, 311 Ostwald-de Waele model 22 outﬂow velocity 9 parallel ﬂow 49, 51 path line 5 phase velocity 55 pipe ﬂow 10, 17, 20, 25–27, 39, 61, 62, 80, 83, 119, 307 pipe friction coeﬃcient 28, 29 Pitot tube 8 plane ﬂow 180 plane subsonic ﬂow 174 plunger pump 11 potential equation 47 potential ﬂow 44, 46–48, 54, 126, 344 potential function 46 potential theory 46, 57 potential vortex 50 Prandtl boundary-layer hypothesis 55 Prandtl integro-diﬀerential equation 276 Prandtl mixing-length hypothesis 62 Prandtl number 41 Prandtl relation 208 Prandtl tube 9 Prandtl’s mixing-length hypothesis 26 Prandtl-Glauert rule 180 Prandtl-Meyer angle 157, 232 Prandtl-Meyer ﬂow 156 Prandtl-Meyer-function 229 pressure coeﬃcient 48, 51, 161, 171, 173 pressure loss coeﬃcient 14, 15 principle of solidiﬁcation 3 propeller of a ship 15 quasi-steady ﬂow 11, 41 radial turbine 19 Rankine’s slip-stream theory 15 recirculation region 65 relative roughness 29 resistance law 27, 29 Reynolds hypothesis 61 Reynolds number 1, 9, 24, 39, 41, 55, 56 Reynolds stress tensor 62 Riemann invariants 167 rotation 44 rotation of the ﬂow 15 rotational ﬂow 44 rough pipes 29 roughness of the wall 29

353

sand roughness 29 scaling function 59, 60 schlieren method 322 Schrenk’s approximate method 276 Segner’s water wheel 19 separation points 65 Ser’s disc 9 shadow method 321 shallow water waves 55 shear action 21 shear experiment 21, 22 shear ﬂow 22 shear stress 21, 22, 24–28, 56 shock angle 151 shock intensity 151 shock polar 139 similar ﬂows 38, 86, 123 similar solution 59 similarity law 41 similarity parameter 38 similarity rules 139, 178, 180–182, 183, 199, 221 similarity transformation 60 single-stem manometer 3 slender axially symmetric body 175 slender body 139, 170, 172–175, 177 sonic conditions 144 source 49, 50 speciﬁc heat 31, 40, 41, 140, 143 speed of sound 142 sphere 253, 258 stagnation enthalpy 37, 143 stagnation point 8, 49, 51, 54, 65 standard nozzle 10 standard oriﬁce 10 starting moment 20 starting vortex 54 steady ﬂow 6, 12, 13, 17, 19 steady gas ﬂow 141, 185, 203 Stevin’s principle of solidiﬁcation 3 Stokes hypothesis 33 Stokes’ no-slip condition 20, 23, 24, 27 stream ﬁlament 7, 8 stream function 46–50, 59, 60 stream tube 7, 10, 15 streamline 6, 7, 37, 44, 46, 47, 49–51 stress tensor 13, 32, 34–37 Strouhal number 39, 41, 66 supersonic tunnel 329, 332, 333 surface forces 2, 32, 35

354

Index

surface roughness 66 surface waves 299, 300 tangential stress 1, 20, 21, 34 theory of characteristics 139, 163, 198, 219 theory of lift-generating bodies 54 theory of similitude 38 thermal conductivity 31 thermal energy 35 thermal equation of state 3, 31, 139 thermally perfect 139 thin proﬁle 161 Thomson’s theorem 45 total energy 35 transonic ﬂow 139, 142, 146, 150, 171, 172, 183 transport properties 40 turbulent ﬂow 61, 64 turbulent pipe ﬂow 25 turbulent shear stress 27 U-tube manometer 3 universal law of the wall 27 vapor pressure 11 velocity correlation 26 velocity ﬂuctuation 26, 27, 61, 62 velocity height 17 Ventury nozzle 10 viscosity law 20–22, 26 viscous sublayer 28 volume dilatation 34 volume force 2, 3, 12, 32, 35, 37, 45, 46, 48, 61 volume rate of ﬂow 7, 10, 11, 17, 23–25, 43 volume viscosity 34 vortex line 44 vortex theorem 44 vortex tube 44 vorticity transport equation 45 vorticity vector 44, 45 wake 65 wall shear stress 23–25, 27, 62, 64, 65 wall velocity 23 water analogy 299, 304, 306 wave drag 87, 93, 139, 160–163, 174, 196, 217 weak and strong solution 152, 153 weak compression shock 155, 157 wind tunnel 190–193, 200 wind tunnel balance 283 wing 271, 273, 278

Fluid Mechanics

You May Also Find These Documents Helpful

Science Study Notes: Fluids

Science Study Notes: Fluids

Computational Fluid Dynamics

Computational Fluid Dynamics

Fluids Mechanics

Fluids Mechanics

Assignment 08: Business Consulting

Assignment 08: Business Consulting

Mech Eng Yr1 Fluid Dynamics Lab Report

Mech Eng Yr1 Fluid Dynamics Lab Report

Hot Air Balloon and Wood Block Float

Hot Air Balloon and Wood Block Float

Annotated Bibliography On Graffiti And Vandalism

Annotated Bibliography On Graffiti And Vandalism

h7 experiment

h7 experiment

Fluids

Fluids

Channel Flow of a Compressible Fluid

Channel Flow of a Compressible Fluid

Critique of a Research Article

Critique of a Research Article

The Greatest Story Never Told

The Greatest Story Never Told

Fluid Dynamics and Flow Rate

Fluid Dynamics and Flow Rate

fluid

fluid

conduct by fluid

conduct by fluid

Related Topics