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 static pressure in the tunnel working section.

The tunnel speed can be determined using:

[pic]

Where θ is the angle of inclination of the manometer to the horizontal, ρm is the density if the manometer fluid (usually about 830 kg/m3) and ρ is the density of air. Density of air can be calculated as follows:

[pic]

P is the pressure in the lab, T is the lab temperature, and R is the gas constant for air (given above). Reynolds number for an aerofoil can be calculated:

[pic]

Results:

Graphs below show the Coefficients of Pressure CP at different points on the aerofoil at different Angles of attack α. The CP values for Hole 2 have been changed using linear interpolation.

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

[pic]

The Lift Coefficient CL is calculated using counting squares method. The values are shown in Table: 5; and the graph below shows the variation of lift coefficient CL with the angle of attack, α.

[pic]

[pic]

Initial Gradient of the Curve (CL/α) = 0.071

Discussion:

Graphs 1-6 show how the coefficients of pressure CP vary at different positions on the surface of aerofoil as the Angle of attack, α is altered. The values for Hole 2 have been linearly interpolated as they seem to give abnormal values consistently for all angles of attack. These flier values can be clearly observed on graphs 8-13 in the appendices. These values could have resulted due to a blockage in tube 2. Linear interpolation method corrects these values by taking into consideration the previous and the next value to the value with error, and finds the average of the two.

Stagnation point is where CP value is +1. For the experiment, it occurs only when the angle of attack is -4º. It exists on the upper surface close to the leading edge. The Flow separation point is where the pressure distribution on the upper surface becomes constant. In this experiment this occurs when the angle of attack is 16º. This...