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Supersonic Flow (Mo > 1, в = JM( — 1)
The same profile equips the wing of an airplane cruising at Mach number M0 > 1 in a uniform atmosphere.
Pressure Distribution and Flow Features
Using Ackeret formula one finds C + — ±20/в and C- — 0 . This is shown in Fig. 15.32 at a — 0. The flow features are displayed in Fig. 15.33 (shocks, characteristic lines, expansion shocks).
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Fig. 15.32 Cp distribution on the half double wedge at a — 0 |
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Aerodynamic Coefficients |
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At a = 0 , the drag coefficient (Cd)a=0 is given by
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and moment coefficient (Cm, o) 0 by
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2в
Static Equilibrium About an Axis
If an axis is located at the leading edge, x = 0, the equilibrium angle aeq corresponds to a zero moment coefficient, i. e.
Cm, o(aeq) = (Cmo)a=0 – 2в = 0,
The equilibrium is stable because dCm, o/da < 0.
