A1 Lab Chordwise Pressure Distribution on an Aerofoil with a Flap
Student Name: REUBEN LEW WAI CHOON
Tutor: COLLIN GREATWOOD
Date of Lab: 7TH NOVEMBER 2014
Summary
A 23015 NACA aerofoil was tested in a wind tunnel with a Re number of 4.28 x 105 and Mach number of 0.07. The pressure distribution over the aerofoil was recorded and used to derive the coefficient of lift, pressure drag and pitching moment. Variations of these parameters with incidence and flap deflection were investigated. Thin aerofoil theoretical prediction of a 2π/radian lift curve, aerodynamic hysteresis and theoretical aerodynamic centre at 0.25C were examined.
1 Introduction
The experiment investigates the characteristics of an aerofoil in a 2D incompressible viscous …show more content…
This is unusual as an asymmetrical aerofoil (NACA 23015) should have generated lift at 0° incidence. This is anomalous result is excluded from the analysis.
Another reason for the error is for the implementation of a thin aerofoil theory, the thickness and camber of the aerofoil must be relatively small. Therefore, the thin aerofoil assumptions is not applicable to NACA 23015.
Another observation during the experiment was once the boundary layer was separated from the aerofoil at a stall angle, it would require a much lower incidence angle for the boundary layer to reattach itself back to the aerofoil. This phenomenon is known as aerodynamic hysteresis. Graph 1 shows the exhibits the phenomenon, where the aerofoil before stall generated more lift at the same incidence angle. [1]
4.2 Effects of Flap Deployment
Flaps increases lift by increasing the area of the wing and altering the camber of the wing. Increasing camber of the wing causes an increase in local flow velocity which decreases local surface static pressure. This would cause a greater pressure difference between the upper and lower surface of the wing, resulting in an overall increase in lift. …show more content…
A comparison made with various flap angles and CLmax in Graph 19 showing CLmax increasing as the angle of attack of the flap increases. However, CLmax stop increasing with the increase in flap angles from 40° to 55°, this is due to separation of flow at the flaps as there is not enough momentum to overcome the increased adverse pressure gradient due to high flap angles.
Graph 16 shows the relationship between flap angles and drag coefficient. Above wing incidence angle of 10°, drag coefficient is proportionate to flap angles. This is due to the fact that flaps would increase local flow velocity and wing area, however, following the drag equation, drag would increase as well.
Camber also helps with reducing stalling speed and this is useful for take-off and landing where maximum lift at low speed is needed. However, there would only be a substantiate decrease in stall speed with a large increases in CLmax. If there is a large increase in CLmax, there would be a large increase in drag as well. A balance would need to be strike to ensure that flaps are aerodynamically efficient. A typical landing and take-off flap configuration for a commercial aircraft would be 40° and 20° respectively.
4.3 Pressure
Student Name: REUBEN LEW WAI CHOON
Tutor: COLLIN GREATWOOD
Date of Lab: 7TH NOVEMBER 2014
Summary
A 23015 NACA aerofoil was tested in a wind tunnel with a Re number of 4.28 x 105 and Mach number of 0.07. The pressure distribution over the aerofoil was recorded and used to derive the coefficient of lift, pressure drag and pitching moment. Variations of these parameters with incidence and flap deflection were investigated. Thin aerofoil theoretical prediction of a 2π/radian lift curve, aerodynamic hysteresis and theoretical aerodynamic centre at 0.25C were examined.
1 Introduction
The experiment investigates the characteristics of an aerofoil in a 2D incompressible viscous …show more content…
This is unusual as an asymmetrical aerofoil (NACA 23015) should have generated lift at 0° incidence. This is anomalous result is excluded from the analysis.
Another reason for the error is for the implementation of a thin aerofoil theory, the thickness and camber of the aerofoil must be relatively small. Therefore, the thin aerofoil assumptions is not applicable to NACA 23015.
Another observation during the experiment was once the boundary layer was separated from the aerofoil at a stall angle, it would require a much lower incidence angle for the boundary layer to reattach itself back to the aerofoil. This phenomenon is known as aerodynamic hysteresis. Graph 1 shows the exhibits the phenomenon, where the aerofoil before stall generated more lift at the same incidence angle. [1]
4.2 Effects of Flap Deployment
Flaps increases lift by increasing the area of the wing and altering the camber of the wing. Increasing camber of the wing causes an increase in local flow velocity which decreases local surface static pressure. This would cause a greater pressure difference between the upper and lower surface of the wing, resulting in an overall increase in lift. …show more content…
A comparison made with various flap angles and CLmax in Graph 19 showing CLmax increasing as the angle of attack of the flap increases. However, CLmax stop increasing with the increase in flap angles from 40° to 55°, this is due to separation of flow at the flaps as there is not enough momentum to overcome the increased adverse pressure gradient due to high flap angles.
Graph 16 shows the relationship between flap angles and drag coefficient. Above wing incidence angle of 10°, drag coefficient is proportionate to flap angles. This is due to the fact that flaps would increase local flow velocity and wing area, however, following the drag equation, drag would increase as well.
Camber also helps with reducing stalling speed and this is useful for take-off and landing where maximum lift at low speed is needed. However, there would only be a substantiate decrease in stall speed with a large increases in CLmax. If there is a large increase in CLmax, there would be a large increase in drag as well. A balance would need to be strike to ensure that flaps are aerodynamically efficient. A typical landing and take-off flap configuration for a commercial aircraft would be 40° and 20° respectively.
4.3 Pressure