279.4 230.4 9.6 0.6 {metric} {26.0 {21.4} = 2.0 0.088 1.768 0.5 1.9 12.0 Wing Sref-wing Sexp-wing A true Lambda Ct/Cr Sweep-LE t/c average Delta Y Cl-design CLmx-airfoil X (quarter-chord) Horizontal Tail S-tail Sexp-tail A true A effective Lambda 5Ct/Cr Sweep-LE t/c average Delta Y X (quarter-chord) {3.7} {6.5} {4.4} 70.0 47.5 6.0 6.0 0.57 4.6 0.08 1.629 34.0 {10.4} Fuselage Swet length diam-effctiv Q (interfer) Upsweep deg X (nose) = 516.9 = 35.5 - 6.1 1.1 = 10.0 = 0.0 {48.0} {10.8} {1.9) {0} | 12.1 For the aircraft just defined, prepare the wing lift curve at Mach 0 and at Mach 0.6. Also graph lift-curve slope vs Mach up to Mach 2. 12.3 Calculate drag due to lift at C1 = 0.5 and Cz = 0.9 using the LE suction method. Compare to result via the Oswald method. Which is better? Easier? 12.1 For the aircraft just defined, prepare the wing lift curve at Mach 0 and at Mach 0.6. Also graph lift-curve slope vs Mach up to Mach 2. 279.4 230.4 9.6 0.6 {metric} {26.0 {21.4} = 2.0 0.088 1.768 0.5 1.9 12.0 Wing Sref-wing Sexp-wing A true Lambda Ct/Cr Sweep-LE t/c average Delta Y Cl-design CLmx-airfoil X (quarter-chord) Horizontal Tail S-tail Sexp-tail A true A effective Lambda 5Ct/Cr Sweep-LE t/c average Delta Y X (quarter-chord) {3.7} {6.5} {4.4} 70.0 47.5 6.0 6.0 0.57 4.6 0.08 1.629 34.0 {10.4} Fuselage Swet length diam-effctiv Q (interfer) Upsweep deg X (nose) = 516.9 = 35.5 - 6.1 1.1 = 10.0 = 0.0 {48.0} {10.8} {1.9) {0} | 12.1 For the aircraft just defined, prepare the wing lift curve at Mach 0 and at Mach 0.6. Also graph lift-curve slope vs Mach up to Mach 2. 12.3 Calculate drag due to lift at C1 = 0.5 and Cz = 0.9 using the LE suction method. Compare to result via the Oswald method. Which is better? Easier? 12.1 For the aircraft just defined, prepare the wing lift curve at Mach 0 and at Mach 0.6. Also graph lift-curve slope vs Mach up to Mach 2