A NACA airfoil has a mean camber line given by z/c 0.600 [ 0.5 (x/c) - (x/c)² ] for 0< x /c< 0.25; z/c 0.111 [0.3 + 0.4 (x/c) – (x/c)² ] for 0.25 < x /c < 1.0. 8°. Using thin airfoil theory, find: (a) angle of attack at zero lift, and (b) lift coefficient when a =
A NACA airfoil has a mean camber line given by z/c 0.600 [ 0.5 (x/c) - (x/c)² ] for 0< x /c< 0.25; z/c 0.111 [0.3 + 0.4 (x/c) – (x/c)² ] for 0.25 < x /c < 1.0. 8°. Using thin airfoil theory, find: (a) angle of attack at zero lift, and (b) lift coefficient when a =
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![A NACA airfoil has a mean camber line given by
z/c
0.600 [ 0.5 (x/c) – (x/c)? ] for 0 < x /c< 0.25;
z/c = 0.111 [ 0.3 + 0.4 (x/c) – (x/c)² ] for 0.25 < x/c < 1.0.
Using thin airfoil theory, find: (a) angle of attack at zero lift, and (b) lift coefficient when a = 8°.](/v2/_next/image?url=https%3A%2F%2Fcontent.bartleby.com%2Fqna-images%2Fquestion%2F5718771a-ab2c-4c23-86aa-27244a553195%2F4d65a08e-e221-4472-b7c3-59814bb4ff63%2Fm813pcq_processed.png&w=3840&q=75)
Transcribed Image Text:A NACA airfoil has a mean camber line given by
z/c
0.600 [ 0.5 (x/c) – (x/c)? ] for 0 < x /c< 0.25;
z/c = 0.111 [ 0.3 + 0.4 (x/c) – (x/c)² ] for 0.25 < x/c < 1.0.
Using thin airfoil theory, find: (a) angle of attack at zero lift, and (b) lift coefficient when a = 8°.
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