Measurement of Pressure. The aerofoil shape used in wing is to. Distribution and Lift for an Aerofoil. generate lift due to the difference

Transcription

1 Measurement of Pressure Distribution and Lift for an Aerofoil. Objective The objective of this experiment is to investigate the pressure distribution around the surface of aerofoil NACA 4415 and to determine the graph of lift coefficient versus angle of attack. Introduction In 1871, the history began with the inventing, designing, and operating the first wind tunnel by Francis Herbert Wenham by installing a 12 feet long and 18 inches square trunk in the test section to test the aerodynamics of the trunk. The history of aerofoil started in 1884 where H. F. Phillips invents a series of aerofoil shapes. During World War Two, largest wind tunnel was built in 1941, located in Ohio by United States to test full scale of large aircraft model. Post World War Two, United States construct wind tunnels at universities and military sites because of lagging to The aerofoil shape used in wing is to generate lift due to the difference velocity of wind on top of the aerofoil and the below part of the aerofoil. The difference of velocity produced different pressure where according to Bernoulli, the velocity of wind on top is higher than below thus produced lower pressure while the below part experienced higher pressure and generate lift. As the flight move in the air, the wing is inclining to the direction of flight direction at some angle where the angle between the chord line and the flight direction is called angle of attack resulting high effect on the lift generated by the wing. Drag is a force that exerted on a moving solid body in the direction of free stream flow, the drag comes from the forces due to the pressure distribution over the body surface. Besides, pitching moment is produced by the aerodynamic force on the aerofoil, it is part of the total moment that must be balanced by horizontal stabilizer. those wind tunnel built by German.

2 In this low speed wind tunnel testing located in Aerolab UTM, I measured pressure distribution at various point along the surface of NACA 4415 aerofoil, manually change the angle of attack of the aerofoil and also change the velocity of the wind. Materials and Method. To conduct this experiment, I used Aerofoil NACA 4415 (Figure 1) and low speed wind tunnel (Figure 2) in UTM Aerolab which has 0.457m x 0.457m x 1.27m(L) in size and speeds up to 33 m/s. This wind tunnel using axial fan drive system driven by 15 kw DC motor. Measurement system of pressure for this testing is using multitube manometer and pressure transducer system. To gauge pressure distribution along the chord, I used Electronic Auto-Zeroing Pressure Aerofoil NACA 4415 is placed into the test section and 17 pressure sensor tubes are connected from the aerofoil to the pressure scanner in chordwise direction. The experiment began with the measurement of the pressure at zero velocity of air at zero angle of attack and measures the pressure along the surface of aerofoil using pressure scanner and data is recorded using Data Acquisition software. Then the angle of attack was set up to -5 degree and velocity 10m/s. After 10 seconds, the pressure is recorded with 1000 reading per seconds. This process is repeated for angle of attack -3, 0, 3, 6, 9, 12, 15, 18 for velocity of wind 15m/s and 20m/s. The total pressure is equal to the sum of static pressure and dynamic pressure. Scanner and to record the pressure, I used Data Acquisition program in LabView software. To change the angle of attack, knob is used where there stated the scale of the angle of attack. To measure the static and total pressure, I used anemometer. Where is the static pressure, is the density of air (1.225 ), is the free stream velocity and is the total pressure.

3 the upper surface pressure. In figure 4, The equation becomes Where Cp is the pressure coefficient. Results and Discussion. Figure 3 illustrated the graphs of coefficient of pressure (Cp) against x/c for velocity of wind 10 m/s at various angles of attack. The first graphs (angle of attack -5 0 ) shows that the lower part of aerofoil has lower pressure (red line) at the beginning of x/c then it gradually increase until the pressure line of upper (blue line) and lower surface intercept. The difference of Cp after the interception is smaller even though x/c increases. This patterns almost the same for the angle of attack However, no interception is detected for angle of attack 3 0 until 12 0 which means that the lower pressure is higher than the upper pressure. Further increment of angle of attack cause the lower surface pressure to be decreased until it is lower than the velocity of free stream is changed to 15m/s however, the trend of the graphs remains almost the same as in Figure 3 for every angle of attack, just a small increment in area between the two curves. By referring to Figure 5, the velocity of free stream used is 20m/s and the graphs plotted shows that there has only small difference as compared to graphs in Figure 3 and Figure 4. The trend is the same but the difference only the area between the curves which has larger area. In figure 6, the graph is about the coefficient of lift against the angle of attack. It can be seen that 20m/s velocity of wind gives bigger lift coefficient than other velocities. They reach maximum lift coefficient between angle of attack 5 0 to 10 0 before the lift coefficient decreases even though the angle of attack increases. Based on Figure 3, 4, and 5, they show the same trend for every graph. Angle of attack higher than 15 0 will have lower surface pressure greater than upper surface pressure. This is because

4 when the angle of attack is greater, the air has tangential velocity that is almost to zero since the air is near zero velocity and wind away from the aerofoil has velocity, it produces thin boundary layer and the boundary layer is turbulent as shown in figure 7. Due to near zero velocity at the upper surface of the aerofoil, it gives high pressure than the lower surface. In Figure 6, higher velocity produce bigger lift coefficient. This is because the area between the curves in Figure 5 is bigger than the area between curves in figure 3 and 4. However, at near to 10 0 angle of attack, the lift coefficient for the three velocities dropped due to the upper surface pressure is higher than the lower surface. This angle is called stall angle. We assume that when the angle of attack is greater, circulation is happening on the upper surface and the circulation push the aerofoil to go down. Conclusion This wind turbine testing has demonstrated the measurement of NACA 4415 and the results was plotted as shown in figure 3, 4, 5, and 6. This testing shows that when the angle of attack is negative, there is only small lift coefficient and the value is negative. When the angle of attack is near 10 0, the maximum lift is produced consequently when the angle is bigger than that, the lift starting to decrease. Acknowledgement. I would like to thank Dr. Nazri Mohd Nasir (Lecturer) for guiding and instructing me on how to run this experiment. I also like to thank En. Abd Basid Bin Abd Rahman (Engineer) and Mohd Mahathir Mohmad (Technicians) for demonstrating and helping me do this experiment. References Donald J. Graham, et al. Report No. 832 A Systematic Investigation of Pressure Distribution at High Speed Over Five Representative NACA Low Drag and Conventional Airfoil Sections. pressure along the surface of aerofoil

5 airfoilhistory.html e) moment nnel 12/WindTunnel/history.html John D. Anderson, JR. (2011), Fundamentals of Aerodynamics 5 th Edition, Mc Graw Hill, p Joseph Katz, Allen Plotkin (2001), Low-Speed Aerodynamics Vol. 13, Cambridge University Press, p M.J. Hoffmann, et al. (1996), Effects of Grit Roughness and Pitch Oscillations on the NACA 4415 Airfoil, The Ohio State University.

6 Appendix Figure 1: Aerofoil NACA Figure 2: Schematic diagram of Low Speed Wind Tunnel in UTM Aerolab.

7 Figure 3: Graph of Cp against x/c for 10m/s velocity of free stream for various angles of attack.

8 Figure 4: Graph of Cp against x/c for 15m/s velocity of free stream for various angles of attack.

9 Figure 5: Graph of Cp against x/c for 20m/s velocity of free stream for various angles of attack.

10 Figure 6: Graph of lift coefficient against angle of attack at various speed of free stream.

11 Figure 7: Flow separation over an aerofoil.