11)

Consider the NACA 2410 airfoil (shown below) at M=0.3. (40 pts.)

Note:

Chord length = 10 feet, Alt. sea level;

Max thickness 10% at 29.9 % chord; Max camber 2% at 40% chord.

a) Compute Reynolds number: Re

b) Sketch (on the graph below) what the notional C, vsa

between -20 degrees and +20 degrees should look like:

a

c) Estimate the zero lift a in degrees:

d) Estimate the lift curve slope at a = 3 degrees:

e) Estimate maximum G:

f) Sketch (on the graph below) how C, vsa (for-5

change with the addition of a trailing edge flap deflection of 20 degrees.

(show the baseline-undeflected case and the flap deflected case):/ng) What is C, (pressure coefficient) at the stagnation point?

h) Estimate C at 10 degrees angle of attack:

i) Estimate C at 0 degrees angle of attack:

Show any work for prob. 11: