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I have been trying to estimate the lift calculation for a straight wing. The previous project I worked on was a swept wing. But this is a straight wing. I would like to to get the stall speed so I need the Clmax formula. What would the lift formula for this type of wing be?

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    $\begingroup$ Change the cos$\varphi$ terms to 1 and the sin$\varphi$ terms to 0. Then the method for swept wings works for a straight one. $\endgroup$ – Peter Kämpf Apr 2 '16 at 17:58
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Regardless of the geometrical shape of the wing, you need to know the wing area to determine the lift force as it is one of the factors that affect lift.

Calculate lift using this formula: $$Lift = \frac{1}{2} (C_L * P * S * V^2)$$.

  • $C_L$: coefficient of lift
  • $P$: density of air
  • $S$: surface area of the wing
  • $V$: velocity of air
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Yo have the lift given by : $$ Lift = \frac{1}{2}\rho S V^2CL $$

The linear part of the lift curve can be approximated by formulae coming from the lifting-line theory :

$$CL = \frac{dCL}{dAoA}\left(AoA-AoA_0\right)$$ With the lift slope given by : $$\frac{dCL}{dAoA} =\frac{2\pi}{\sqrt{1-M^2+\left(\frac{2}{AR}\right)^2}+\frac{2}{AR}} $$

Still you cannot compute the $CL_{max}$ from this and this stands for a straight low aspect ratio wing in the subsonic domain.

If the overall span of your wing is $2b$ and the chord is $c$ you have :

  • The wing surface : $$S=2bc$$
  • Its aspect ratio : $$AR = \frac{(2b)^2}{S}=\frac{4c}{b}$$

For the zero-lift angle of attack ($AoA_0$) it's approximately the one of the 2D-wing section if the wing is untwisted and if the wing section has a constant shape.

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