I understand that the critical angle of attack is mostly affected by the shape of the airfoil. However does the airspeed of the airfoil or the density of the air affect the critical angle of attack? I would think that the airspeed or air density would affect the laminar flow properties and could change the critical angle of attack of that airfoil, perhaps only slightly.

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    $\begingroup$ Could you perhaps wait with accepting the answer to attract more answers? I've been wondering about this exact question, but I'm hoping there will be some more quantitative answers than the current one. $\endgroup$
    – Sanchises
    Commented May 4, 2018 at 12:43

2 Answers 2


Yes, through the Reynolds number

The Reynolds Number is a function of airspeed and is used to predict the transition from laminar to turbuluent flow:

$$ {Re} = \frac{\rho v L}{\mu} = \frac{v L}{\nu} $$

The critical angle of attack is affected by the separation of the boundary layer, which, in turn, depends on the laminar or turbulent qualities of said layer.

Do note that the Reynolds number of the layer is not the same as the Reynolds of the wing, although they are related.

EDIT: to visualize the efect, see for example the following graph (thanks @MikeY) Image illustrating Reynolds number effect on lift coefficiant curves, from Kishore, Ravi & Coudron, Thibaud & Priya, S.Jeba. (2013). Small-scale wind energy portable turbine (SWEPT). Journal of Wind Engineering and Industrial Aerodynamics. 116. 21–31. 10.1016/j.jweia.2013.01.010.

Quite a bit of experimental work on low airspeed performance was done by Selig et al. at the UIUC, and their website maintains a significant catalogue of airfoil datapoints.

Another resource worth checking is Airfoiltools, which provides graphed Xfoil polars for an extensive seleciton of airfoils. Do keep in mind Xfoil´s prediction of flow separation points is numerical and therefore may not converge with reality at all points.

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    $\begingroup$ Nice diagram...researchgate.net/profile/Ravi_Kishore16/publication/257011613/… $\endgroup$
    – MikeY
    Commented Apr 24, 2019 at 21:55
  • $\begingroup$ If this diagram represents something about velocity, why are neither the X or Y axes labeled with anything having to do with velocity? $\endgroup$ Commented Apr 24, 2019 at 22:27
  • $\begingroup$ @RyanMortensen velocity is subsumed into the Reynolds number, which appears as the coefficient generating the curve family. $\endgroup$ Commented Apr 24, 2019 at 22:29
  • $\begingroup$ Although an increase of velocity 2 orders in magnitude is a bit much, there are differences in doubling speed for a given Reynolds number. Good graph. +1 $\endgroup$ Commented Apr 25, 2019 at 13:26
  • $\begingroup$ @RobertDiGiovanni No need to increase velocity by a factor of 100, as the chord also plays a role; this is another reason why wingtips often have different airfoils. The effect is most pronounced in small model planes, which fly close to the flow transition $Re$ anyway. $\endgroup$ Commented Apr 25, 2019 at 13:42

From a pilot’s perspective, no. Stalling will occur at critical AoA at any airspeed or altitude.

Planes fly because of Lift being equal or greater than weight. Lift is the result of this formula:

$$ L = \frac{1}{2} \cdot \rho \cdot v^2 \cdot C_L \cdot S $$

(ρ=density, v=airspeed, CL=lift coefficient, S=surface area)

CL varies directly with AOA. The more AoA, the more CL... until the point where the boundary layer separates, and no Lift is generated. At that point the maximum CL, or CLmax is attained. The relationship between AoA and CL is obtained empirically for each aerodynamic surface.


The more amount of airspeed, the more Lift.

And the more AoA, more Cl, and therefore, more Lift.

But beyond the critical angle of attack, regardless of the airspeed, you stall.

You may think that you can go beyond the AoA and still be flying and climbing, but that is going to be a product of the thrust, that, with the acceleration is going to modify the direction of the relative wind. Remember that the AoA is measured against the relative wind, so when you apply thrust, you are modifying the "point of view" of your angle.

Regarding density, or altitude, you have to consider the difference between the True Airspeed (TAS) and the Indicated Airspeed (IAS). With a same IAS, at a higher altitude you are getting more TAS. This is because IAS is a function of the dynamic pressure (q), that depends on the density (ρ):

$$ q = \frac{1}{2} \cdot \rho \cdot v^2 $$

As you see from the equation, if you lower ρ, or increase altitude and still keep your IAS, you will be obtaining more v (airspeed generating the same amount of Lift).

But CL is not depending on density, and the CLmax is obtained at the same angle of attack at different altitudes.

  • $\begingroup$ Welcome to Aviation.SE! Could you explain further how you came to this conclusion? Just as with drag,, the lift coefficient is dependent on Reynolds number so the Cl/α curve is valid for a certain Reynolds number (and therefore speed). $\endgroup$
    – fooot
    Commented Apr 24, 2019 at 19:20
  • $\begingroup$ "Stalling occurs at a specific AoA at any airspeed or altitude" is simply not correct, I´m afraid, see my answer. $\endgroup$ Commented Apr 24, 2019 at 22:27
  • $\begingroup$ @fooot: I interpreted the question from a pilot's perspective, where the pilot has no control over the Reynold's number, but on the AoA, surface configuration (ailerons, spoilers...) and speed. My point is valid from this perspective, where the pilot needs to know that he can stall the plane at any speed or altitude at a specific AoA. That's why planes have an AoA sensor. But of course, from a design perspective of the airfoil, you are correct. $\endgroup$ Commented Apr 26, 2019 at 8:24

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