Consider an infinite wing with a NACA 1412 airfoil section and a chord length of 3...

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Consider an infinite wing with a NACA 1412 airfoil section and a chord length of 3 ft. The wing is at an angle of attack of 5° in an airflow velocity of 100 ft/s at standard sea-level conditions. Calculate the lift, drag, and moment about the quarter-chord per unit span. Refer to the Appendix graphs given below for the standard values. al -4 Section lift coefficient, c 36 3.2 2.8 24 2.0 1.6 1.2 -4 -8 -1.2 -1.6 -20 32 -24 B 198 -16 2 The lift per unit span is The drag per unit span is The moment about the quarter-chord per unit span is Be -8 0 8 16 Section angle of attack, , deg NACA 1412 Wing Section lb. lb. 676 ib. SO₂ Home 24 32 024 8.020 E Section drag coefficien 016 012 008 004 16 .2 E1 -1.2 4 ayc 6 -8 .8 1.0 a.c. position 2/C++w/ 250 0 250 0 252 026 ..0280 0 0932 R LO 30x10 06.0 09.0 6.0 Standard roughness 1 020c simulated split flap deflected 60° 06.0 60 028 H Standard roughness -4 4 Section lift coefficient, NACA 1412 Wing Section (continued) .8 1.2 1.6 Consider an infinite wing with a NACA 1412 airfoil section and a chord length of 3 ft. The wing is at an angle of attack of 5° in an airflow velocity of 100 ft/s at standard sea-level conditions. Calculate the lift, drag, and moment about the quarter-chord per unit span. Refer to the Appendix graphs given below for the standard values. al -4 Section lift coefficient, c 36 3.2 2.8 24 2.0 1.6 1.2 -4 -8 -1.2 -1.6 -20 32 -24 B 198 -16 2 The lift per unit span is The drag per unit span is The moment about the quarter-chord per unit span is Be -8 0 8 16 Section angle of attack, , deg NACA 1412 Wing Section lb. lb. 676 ib. SO₂ Home 24 32 024 8.020 E Section drag coefficien 016 012 008 004 16 .2 E1 -1.2 4 ayc 6 -8 .8 1.0 a.c. position 2/C++w/ 250 0 250 0 252 026 ..0280 0 0932 R LO 30x10 06.0 09.0 6.0 Standard roughness 1 020c simulated split flap deflected 60° 06.0 60 028 H Standard roughness -4 4 Section lift coefficient, NACA 1412 Wing Section (continued) .8 1.2 1.6