% Typical Range Calculation for Boeing 737-300
% flying at Mach 0.8 at an altitude of 30,000ft
%
%velocity
v=240;
%Mach number
M=0.8;
%Wing planform area
S=91.0;
%wing span
b=28.35;
%Aspect ratio
AR=b*b/S
%wing planform efficiency factor
eff=0.7;
%base drag coefficient
cdo=0.019;
%lift induced drag constant
K=1.0/pi/AR/eff
%
%Thrust Specific fuel compumption (TSFC) (lbm/hr)/lbf --> /sec or (N/s)/N
tsfc=0.6/3600.0
%Altitude in meters
alt=30000*.3048
%air density calculated as a function of altitude
[rho]=alt_dens(alt)
%initial weight at start of cruise (=MTOW rough estimate)
winit=56470.0;
%lift coefficient required for L=W
cl=winit*9.8/(0.5*rho*v*v*S)
%drag coefficient
cd = cdo + K *cl*cl;
%corrected for compressible flow effects
cd=cd/sqrt(1-M*M) 
%lift to drag ratio
lond=cl/cd
%final weight = initial weight - 15083 Kg fuel (Fuel tank capacity=16000kg)
wfinal=56470-15083
% range = V/TSFC * L/D * ln(Winitial/Wfinal) in meters
range=v/tsfc*lond*log(winit/wfinal)
% final lift coefficient required
cl=wfinal/(0.5*rho*v*v*S)
% final drag coefficient
cd=cdo+K*cl*cl
cd=cd /sqrt(1-M*M)
% drag of vehicle at final weight
drag=cd*0.5*rho*v*v*S
% thrust required
thrust=drag
% approximate fuel flow rate in Kg/s
fuelflow=tsfc*thrust/9.8
%  max reserve (Kg)
fuel=1000
% reserve endurance at M=0.8 in mins
time=fuel/fuelflow/60.0