Pressure Distribution Around Symmetrical Aerofoilc
Pressure Distribution around a Symmetric Aerofoil
Abstract:
The following report is based on an experiment conducted to calculate the lift curve slope for a symmetrical aerofoil subjected to varying angles of attack. Pressure readings were taken at different points on the upper and lower surface of the aerofoil. The report concludes that maximum lift is generated between 12 º -15º, which is also the stall point. It also states that region close to the leading edge contributes most to the lift force.
Introduction:
This experiment is designed to measure the static pressure distribution around a symmetric aerofoil, find the normal force and hence to determine the lift- curve slope.
For zero angle of attack the pressure distribution is symmetrical around the aerofoil. Increasing the angle of attack (lifting the leading edge) increases the velocity of airflow hence decreases the air pressure on the upper-surface. The opposite happens on the lower-surface where high pressure is created. This difference in pressure creates a force normal to the chord line in the direction of lower pressure, this force is called lift. As the angle of attack increases so does the lift until at a particular angle the airflow on the upper-surface is cut-off. This dramatically increases the drag and decreases the lift.
The Experiment:
Aerofoil of chord length 3.5” is mounted inside a wind tunnel running at a suitable at a suitable wind speed.
Pressure at different points on the surface of the aerofoil is measured using wall tappings. These tappings are connected to a multi-tube manometer.
The dynamic pressure is measure using the tunnel reference pressure (hs) and atmospheric pressure (ha). Pressure readings will be taken for angles of attack from -1° to 16° at intervals of 5°.
Theory:
The Pressure coefficient can be calculated from the manometer readings as follows:
[pic]
Where h is the reading for the tapping being considered, ha is the atmospheric pressure reading and hs is the
Abstract:
The following report is based on an experiment conducted to calculate the lift curve slope for a symmetrical aerofoil subjected to varying angles of attack. Pressure readings were taken at different points on the upper and lower surface of the aerofoil. The report concludes that maximum lift is generated between 12 º -15º, which is also the stall point. It also states that region close to the leading edge contributes most to the lift force.
Introduction:
This experiment is designed to measure the static pressure distribution around a symmetric aerofoil, find the normal force and hence to determine the lift- curve slope.
For zero angle of attack the pressure distribution is symmetrical around the aerofoil. Increasing the angle of attack (lifting the leading edge) increases the velocity of airflow hence decreases the air pressure on the upper-surface. The opposite happens on the lower-surface where high pressure is created. This difference in pressure creates a force normal to the chord line in the direction of lower pressure, this force is called lift. As the angle of attack increases so does the lift until at a particular angle the airflow on the upper-surface is cut-off. This dramatically increases the drag and decreases the lift.
The Experiment:
Aerofoil of chord length 3.5” is mounted inside a wind tunnel running at a suitable at a suitable wind speed.
Pressure at different points on the surface of the aerofoil is measured using wall tappings. These tappings are connected to a multi-tube manometer.
The dynamic pressure is measure using the tunnel reference pressure (hs) and atmospheric pressure (ha). Pressure readings will be taken for angles of attack from -1° to 16° at intervals of 5°.
Theory:
The Pressure coefficient can be calculated from the manometer readings as follows:
[pic]
Where h is the reading for the tapping being considered, ha is the atmospheric pressure reading and hs is the
References: “Engineering Laboratory Instruction Booklet”, pp 19-27. 2009.