SKMA394 Measurement of Pressure Distribution around an Airfoil NACA445 OBJECTIVE The objective of this experiment is to study about the pressure distribution around the surface of aerofoil NACA 445 starting from before angle of attack(α CL = 0 ) until after static angle and plot the graph of lift coefficient C l and drag coefficient C D against angle of attack α, for the aerofoil. INTRODUCTION Wind Tunnel is a tool to simulate flight of an aircraft as a real flight. It is used to test models of proposed aircraft. (Benson, 2009). Wind tunnel will provide the aerodynamic characteristics of the aircraft for further analysis. The air flow also can be seen clearly. This will help aeronautic s engineers and researchers to modified or in designing the aircraft that have specific characteristics such as minimum drag for and maximum lift force. There are various types of wind tunnel such as low subsonic, subsonic closed return, water tunnel, subsonic open return full scale, supersonic closed return propulsion, subsonic open return smoke tunnel, etc. From wind tunnel test, we can determine aerodynamic s characteristics such as lift coefficient(cl), drag coefficient(cd), and polar drag lift curve as well. METHODOLOGY This experiment required a low subsonic wind tunnel, aerofoil NACA 445, 7 pressure vessels, speed detector, and manometer. All 7 pressure vessels need to be connected to the aerofoil and be placed inside the test section of the wind tunnel. Manometer and speed detector need to be connected to the wind tunnel as well in order to get the reading for the speed inside the wind tunnel and the pressure reading. After all the apparatus were set up, the technician will switch on the wind tunnel and let the wind tunnel on for few minutes with 0 o angle of attack. Pressure reading(p total ) on the manometer will be taken after the speed on the speed detector reach low speed range().
This step will be repeated for different angle of attack (-6 o, -3 o, 0 o, 3 o, 6 o, 9 o, 2 o, 5 o and 8 o ) by changing the angle of the aerofoil. After done with low velocity, we do the same thing but for different velocity which is for 0,5 and 20 m/s. Cn = [Area Cp x for upper surface c Area Cp x for lower surface] c Cn = Normal coefficient All data recorded will be use to calculate the Cp(Pressure coefficient). The calculation of Cp can be determine using this formula; Ca = [Area Cp y c for upper surface Pstatic P cp = ( ) 2 ρ V 2 Pstatic = Pressure taken from experiment P experiment = Dynamic Pressure, from All the data we get from the experiment is a raw data which mean not considering the blockage effect inside the wind tunnel. Here is the formula to calculte the corrected velocity which will be used to calculate the correct Cp, Cn, Ca, Cl, and Cd. Area Cp y for lower surface ] c Ca = Axial coefficient Cl = [Cn cos(α) Ca sin(α)] Cl = Lift coefficient α = Angle of Attack in radian Cd = [Cn sin(α) Ca cos(α)] Cd = Drag coefficient α = Angle of Attack in radian
RESULTS Figure shows the graph of C p versus angle of attack(α) at (-6 o, -3 o, 0 o, 3 o, 6 o, 9 o, 2 o, 5 o and 8 o ) at various speed. By this graphs we can clearly see that the pressure distribution along the airfoil is increasing with angle of attack and speed as well. Meanwhile, in figure 2, it shows the graph of coefficient of lift and drag against angle of attack. From this graph it is found that as the angle of attack increases, the both coefficient of drag and lift keep on increasing. The process continues until the coeficient of lift reaches at some point which to be peaked at (5, 4 and 2) angle of attack. This proves the theoretical part of this experiment when the airfoil tend to stall when the angle of attack reached at peak. Inversely to that, coefficient of drag keep on increasing as the angle of attack incease because mainly due to the parasite and induced drag. It is known that the C l and C d graph doesn t produce a linear graph. It continues for different speed of velocity which might be caused by some error during handling the experiment plus error from the human and the equipment itself. Moreover, some assumption in experimental data(raw data) has been made in calculating the coefficient of pressure so that most of the error can be ignored due to the blockage in the pressure tube vessels. When conducting the experiement, it is found out that the pressure detector is having a slight error for detecting the experimental pressure. DISCUSSION As in this experiment, we used airfoil shaped NACA 445, which is a cambered airfoil that can produce lift even at zero angle of attack. As in the graphs, which shows coefficient of lift and drag increases as angle of attack increases. As in the draglift polar graph, it shows an identical graph when it is compared with theoretical ideas. Can t be denied that this experimental value is having certain failures of data due to the defects happen in the pressure vessels. But, it has been resolved by taking an assumption that the defected pressure vessel are working fine. The data that has been defected were recorrected by taking average values between two other data from pressure vessel beside the defected pressure vessel.
CONCLUSION From this experiment we can conclude that the various pressure distributions does really affect the coefficient of drag and lift of NACA 445 airfoil. As the angle of attack increases, it does help to increase the Cl and Cd of the airfoil. The objective of the experiment succeeded as the coefficient of drag and lift been plotted and studied based on pressure distribution along the airfoil. REFERENCES J.D. Anderson: Fundamental of Aerodynamics, 5th Ed, New York. Steven C.Chapra, Raymond P.Canale:Numierical Methods for Engineers,Sixth Edition, McGraw Hill.
Figure.: graphs of pressure distribution for upper surface at various angles of attack at speed of 0 m/s(angle of attack increasing starting from left to right towards bottom) Figure.2: graphs of pressure distribution for upper surface at various angles of attack at speed of 5 m/s(angle of attack increasing starting from left to right towards bottom)
Figure.3: graphs of pressure distribution for upper surface at various angles of attack at speed of 20 m/s(angle of attack increasing starting from left to right towards bottom) Figure.4: graphs of pressure distribution for upper surface at various angles of attack at speed of 0 m/s(angle of attack increasing starting from left to right towards bottom)
Figure.5: graphs of pressure distribution for lower surface at various angles of attack at speed of 5 m/s(angle of attack increasing starting from left to right towards bottom) Figure.6: graphs of pressure distribution for upper surface at various angles of attack at speed of 20 m/s(angle of attack increasing starting from left to right towards bottom)
Coeffcient Cl and Cd versus angle of attack.4.2 0.8 0.6 0.4 0.2 0-0 -5 0 5 0 5 20-0.2 Angle of Attack Cd v=0m/s Cl v=0m/s cl v=20m/s cd v=20m/s cl v=5m/s cd v=5m/s Figure 2.: Coefficient of Lift and Drag for airfoil NACA 445.4 Drag-Lift polar curve.2 0.8 0.6 V=0m/s 0.4 0.2 0-0. -0.2 0 0. 0.2 0.3 0.4 0.5 0.6 Figure 2.2: Drag polar curve for airfoil NACA 445 at velocity of 0 m/s
Cl Cl Drag-Lift polar curve.4.2 0.8 0.6 0.4 0.2 0-0.05 0 0.05 0. 0.5 0.2 0.25-0.2 Cd V=5m/s Figure 2.3: Drag polar curve for airfoil NACA 445 at velocity of 5 m/s.2 Drag-Lift polar curve 0.8 0.6 0.4 0.2 v=20m/s 0-0.2 0 0.0 0.02 0.03 0.04 0.05 Cd Figure 2.4: Drag polar curve for airfoil NACA 445 at velocity of 20 m/s