Pressure Distribution Around Symmetrical Aerofoilc free essay sample

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 ? -15? , 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. We will write a custom essay sample on Pressure Distribution Around Symmetrical Aerofoilc or any similar topic specifically for you Do Not WasteYour Time HIRE WRITER Only 13.90 / page 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 ? is the angle of inclination of the manometer to the horizontal, ? m is the density if the manometer fluid (usually about 830 kg/m3) and ? is the density of air. Density of air can be calculated as follows: [pic]Initial Gradient of the Curve (CL/? ) = 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, ? 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 -4?. 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 16?. This means that the flow cuts off and thus the pressure is almost the same as the atmospheric pressure, which means there in so suction. Another observation made during the test was that the wall tappings started swirling rigorously as the angle was increased; this uggests that the flow became turbulent on the upper surface. From the graphs above it can be seen that maximum suction for all the angles (except -4? ) exists at Hole 1, which is 1. 27 mm from the leading edge on the upper surface. The angle of attack at which most suction is achieved is 11? where the CP value reaches -3. 0; and it increases as the angle of attack is increased or decreased from 11?. The CP values at the trailing edge vary for different angles. They all range between 0 and -0. 5. It is maximum at -4? and 0? while it is minimum at 16?.