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A symmetric airfoil is operating with flow velocity of 350m/s. The lift produced by the airfoil is 21 N at 0.008 rad AOA. If chord is c m then, what will be the pressure difference across the airfoil?

A symmetric airfoil is operating with flow velocity of 350m/s. The lift produced by the airfoil is 21 N at 0.008 rad AOA. If chord is c m then, what will be the pressure difference across the airfoil? Correct Answer 21/c Pa

Given, symmetric airfoil Flow velocity V=350m/s, Lift L=21N, AOA = 0.008rad Chord = c m, Span of airfoil = 1 unit Pressure difference = Lift/Area of airfoil = 21/c*1 = 21/c Pa.

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