2-D Inviscid, Linearized, Thin Airfoil Theories
220.127.116.11 Incompressible Flow (M0 = 0)
Consider a half-double wedge profile of chord c of equation
where © is the wedge angle, see Fig. 14.3.
Calculate the distributions of camber d(x) and thickness e(x) for this profile. Check your result.
Fig. 14.3 Half double-wedge geometry
Give the expressions of the Fourier coefficients A0 and An in the expansion of the vorticity for an arbitrary profile.
Calculate the Fourier coefficients A0, A1 and A2 for this thin profile (Hint: you
need to split the integral into two pieces J02 + Jc).
Give the incidence of adaptation aadapt.
Sketch the flow at the incidence of adaptation, showing in particular the streamlines near the leading and trailing edges.
Definition of Aerodynamic Center
Give the definition of the aerodynamic center.
Give the aerodynamic coefficients C;(a) and Cm, o (a).
18.104.22.168 Supersonic Flow (M0 > 1, в = yjM( — 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
Calculate and plot —C + and – C— versus x for this airfoil at a = 0. Sketch the flow at a = 0 (shocks, characteristic lines, expansion shocks).
At a = 0 , calculate the drag and moment coefficients (Cd)a=0 and (Cm, o) 0
Static Equilibrium About an Axis
If an axis is located at the leading edge, x = 0, find the equilibrium angle aeq if there are no forces other than the aerodynamic forces. Is the equilibrium stable, unstable, neutral? (Hint: calculate Cm, o (a)).