Try this:
W = 73000; %Airplane weight in lbs
Alt = 30000; %altitude given at 30000 feet
S = 950; %Area of the wing given in ft^2
AR = 5.92; %Aspect Ratio
Cd_o = 0.015; %coefficient of drag at 0 The value of 0.015 chosen for C0 _o is based on a generic value typical of streamlined, multiengine jet aircraft.
e = .9; %span efficiency factor
k_3 = 1/(pi*e*AR);
k_2 = 0;
k_1 = (1/3)*k_3;
K = k_1+k_2+k_3; %I'm not sure
V = linspace(400, 2400, 1000); %Velocity feet per second
rho = 8.9068*10^(-4); %Density at 30000 ft
C_l = ((2*W)./(rho*V.^2*S)); %coefficient of lift
C_d = Cd_o+K*(C_l).^(2); %Coefficient of drag
T_r = (1/2*rho.*V.^2.*S.*C_d); %Thrust given in lbs
D = T_r; %Thrust = Drag page 205 last paragragh
P_r = V.*T_r;
T_a = 13850*2; %lb/f
P_a = V.*T_a;
intx_V = interp1(P_r-P_a, V, 0);
intx_Pr = interp1(V, P_r, intx_V);
figure
hold on
plot(V, P_r)
plot(V, P_a)
plot(intx_V, intx_Pr, 'sr')
hold off
text(intx_V,intx_Pr, sprintf('(%6.1f, %9.1f) \\rightarrow', intx_V,intx_Pr), 'HorizontalAlignment','right', 'VerticalAlignment','middle')
producing:
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