Question

8. (8) An airfoil is tested with a trailing edge flap using XFLR5. Which curve below belongs to the plain airfoil and which to the airfoil with flap deflected downward. Annotate the figure or describe. 9. (8) The plot right shows the lift curve for a wing with a NACA 2412 airfoil. Add a second curve representing the same wing but now with a leading edge slat deflected 20 degrees down. I'm just looking for the trends -- no calculations. 10. An airfoil camber line is analyzed and Fourier coefficients are computed with results given as: $A_0 = \alpha = -0.004$ $A_1 = 0.075$ $A_2 = -0.019$ If the chord $c = 1$, and the angle of attack in radians is 0.004, iii) (6) Find the section lift coefficient, $C_l$ iv) (8) Find the center of pressure, $x_{cp}$

          8. (8) An airfoil is tested with a trailing edge flap using XFLR5. Which curve below belongs to
the plain airfoil and which to the airfoil with flap deflected downward. Annotate the figure or
describe.
9. (8) The plot right shows the lift curve for a wing with
a NACA 2412 airfoil. Add a second curve representing
the same wing but now with a leading edge slat deflected
20 degrees down. I'm just looking for the trends -- no
calculations.
10. An airfoil camber line is analyzed and Fourier
coefficients are computed with results given as:
$A_0 = \alpha = -0.004$
$A_1 = 0.075$
$A_2 = -0.019$
If the chord $c = 1$, and the angle of attack in radians is 0.004,
iii) (6) Find the section lift coefficient, $C_l$
iv) (8) Find the center of pressure, $x_{cp}$

8. (8) An airfoil is tested with a trailing edge flap using XFLR5. Which curve below belongs to
the plain airfoil and which to the airfoil with flap deflected downward. Annotate the figure or
describe.
9. (8) The plot right shows the lift curve for a wing with
a NACA 2412 airfoil. Add a second curve representing
the same wing but now with a leading edge slat deflected
20 degrees down. I'm just looking for the trends – no
calculations.
10. An airfoil camber line is analyzed and Fourier
coefficients are computed with results given as:
A0 = α = -0.004
A1 = 0.075
A2 = -0.019
If the chord c = 1, and the angle of attack in radians is 0.004,
iii) (6) Find the section lift coefficient, Cl
iv) (8) Find the center of pressure, xcp

Added by Manuel W.

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