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All aerodynamics about the airfoil is determined by the pressure distribution developed around the airfoil in flight.
For convenience we resolve the net effect of the pressure distribution as concentrated loads called the lift, drag, and the pitching moment resolved at the center of pressure. By definition the drag is parallel to the relative velocity at the airfoil. The lift is perpendicular to the relative velocity. Nose up pitching moment is positive
The aerodynamic loads (or their non dimensional coefficients) depend on
airfoil shape
Reynolds number
Mach number
angle of attack
For two very different airfoils, their pressure coefficients and their aerodynamic characteristics are compared below (Pablo - a MATLAB program for airfoil analysis)
Airfoil NACA 0012
symmetric airfoil , is one of the earlier airfoils used for helicopter blades, the horizon tail etc.
Airfoil GAW1
Specially designed at NASA after considerable research. Targeted at small general aviation aircraft. Has high aerodynamic efficiency.
Angle of attack = 4
Re = 10,000,000
Cl = 0.4729
Cd = 0.0059
Cm = -0.0032
xcp/c= 0.25
E = Cl/Cd = 80
Angle of attack = 4
Re = 10,000,000
Cl = 1.0872
Cd = 0.0077
Cm = -0.1334
xcp/c= 0.37
E = Cl/Cd = 141
It is instructive
- to note the effect of thickness, maximum thickness, area under the curve
- scale and sign for the pressure coefficient (y - direction)
- scale for x (x/c)
- sharp and smooth variation
Wing Pressure Distribution
Wing Pressure Distribution
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