Consider a rectangular wing with span b = 15 m and chord c = 1,77 m, flying horizontally in steady atmosphere with velocity U∞ = 50 m/s, and air density ρ = 1.225 kg/m3. We can assume that Re >> 1, M << 1, and that we know the circulation distribution along the wing Γ(y):

where A = 0,73, B = 2,53, and C = 14,27 are dimensionless constants, and Γo = 17,7 m²/s. You are asked to compute, for this wing, the global lift L in [N]:

Note: For the computed results, provide at least 6 decimals, and use the coma "," as decimal separator.

The

image shows a vortex line with horseshoe shape. The vortex line has different

intensity per unit length depending on the considered segment/leg of the

horseshoe, where

Say that we have a 3D wing which is entirely built using only a single type of airfoil. The aspect ratio of the wing is AR=10,8. As per the relationship between the vs polar curve of the 3D wing and the vs polar curves of the 2D airfoils it is made of, what is the value of in [º] at which 0,81 for the 3D wing, if the value of at which 0,81 for the 2D airfoils is 4,3º?

Consider a rectangular wing with span b = 31,3 m and chord c = 1,49 m, flying horizontally in steady atmosphere with velocity U∞ = 73 m/s, and air density ρ = 1.225 kg/m3. We can assume that Re >> 1, M << 1, and that we know the circulation distribution along the wing Γ(y):

where A = 0,63, B = 3,75, and C = 16,6 are dimensionless constants, and Γo = 29,3 m²/s. You are asked to compute, for this wing, the global induced drag Di in [N]:

Note: For the computed results, provide at least 6 decimals, and use the coma "," as decimal separator.

Say we have a doublet with intensity 16 [m³/s], a source with intensity 38 [m²/s], and a vortex with intensity -17 [m²/s], all three located in the origin of the wind reference frame. You are asked to compute the axis component of the flow velocity in units [m/s], in the position 17,3 m and 9,6 m.

Note: Please, provide at least 4 decimals and use the comma "," as decimal separator.

The perturbation velocity component in the -axis, in [m/s], in x = -0.5 m, for a thickness problem with

Say we have an airfoil with leading edge (LE) and trailing edge (TE) located at 0 and 1.84 m, respectively. The pressure coefficient distribution in the upper surface of the airfoil is trapezoidal, with -0.32 constant for all values of between and 63% of the chord, and then ranging linearly from -0.32 at 63% of the chord to 0 at . On the other hand, the pressure coefficient distribution in the lower surface fits well a cosenoidal function with period equal to 0.57 times the chord, and amplitude 0.67, crossing the positive part of the ordinate axis in the leading edge with null slope. Please, compute the value of the airfoil’s global lift coefficient (with 3 decimals; and please, use the comma "," as decimal separator).

Consider

a 2D airfoil of chord

The measured component of 

the

flow velocity is

where

For

this purpose, you can assume that the

Bear

in mind that the velocity cannot be assumed to be horizontal far

upwind and far downwind, while

and the following unknown axis velocity far downwind:

Finally, a

ssume

an unknown velocity in the form

Note: The results are in SI units with 2 decimals and the comma "," is used as decimal separator