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+//Example 11_5
+clc;clear;
+//Properties
+rho_ag=1.20;// kg/m^3
+rho_ac=0.312;// kg/m^3
+C_Lmax1=1.52;// The maximum lift coefficient of the wing with flaps
+C_Lmax2=3.48;// The maximum lift coefficient of the wing without flaps
+//Given values
+m=70000;// kg
+A=150;// m^2
+V=558;/// km/h
+g=9.81;// m/s^2
+
+// Calculation
+//(a)
+W=m*g;// N
+V=V/3.6;// m/s
+V_min1=sqrt((2*W)/(rho_ag*C_Lmax1*A));// m/s
+V_min2=sqrt((2*W)/(rho_ag*C_Lmax2*A));// m/s
+V_1s=1.2*V_min1*3.6;// 1 m/s=3.6 km/h
+printf('(a)Without flaps:V_min1,safe =%0.0f km/h\n',V_1s);
+V_2s=1.2*V_min2*3.6;// 1 m/s=3.6 km/h
+printf('    With flaps:V_min2,safe =%0.0f km/h\n',V_2s);
+//(b)
+F_l=W;// N
+C_l=F_l/(1/2*rho_ac*V^2*A);// The lift coefficient
+//For the case with no flaps, the angle of attack corresponding to this value of C_L is determined from Fig. 11–45 to be
+alpha=10;// The angle of attack in degree
+printf('(b)The angle of attack,alpha~=%0.0f degree\n',alpha);
+//(c)
+// From Fig.11-45,C_d~=0.03
+C_d=0.03;// The drag coefficient
+F_d=(C_d*A*rho_ac*(V^2/2))/1000;//kN
+P=F_d*V;// kW
+printf('(c)The power that needs to be supplied to provide enough thrust to overcome wing drag,P=%0.0f kW\n',P);
+// The answer vary due to round off error