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A MATLAB Script for Propagating Interplanetary Trajectories from Earth to Mars

A MATLAB Script for Propagating Interplanetary Trajectories from Earth to Mars

by

David Eagle (view profile)

 

Numerically integrate the orbital equations of motion of an Earth to Mars interplanetary trajectory.

oeprint1(mu, oev)
function oeprint1(mu, oev)

% print six classical orbital elements,
% argument of latitude and orbital period in days

% input

%  mu      = gravitational constant (km**3/sec**2)
%  oev(1)  = semimajor axis (kilometers)
%  oev(2)  = orbital eccentricity (non-dimensional)
%            (0 <= eccentricity < 1)
%  oev(3)  = orbital inclination (radians)
%            (0 <= inclination <= pi)
%  oev(4)  = argument of perigee (radians)
%            (0 <= argument of perigee <= 2 pi)
%  oev(5)  = right ascension of ascending node (radians)
%            (0 <= raan <= 2 pi)
%  oev(6)  = true anomaly (radians)
%            (0 <= true anomaly <= 2 pi)

% Orbital Mechanics with MATLAB

%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%%

rtd = 180 / pi;

% unload orbital elements array

sma = oev(1);
ecc = oev(2);
inc = oev(3);
argper = oev(4);
raan = oev(5);
tanom = oev(6);

arglat = mod(tanom + argper, 2.0 * pi);

if (sma > 0.0)
    
    period = 2.0 * pi * sma * sqrt(sma / mu) / 86400.0d0;
    
else
    
    period = 9.9999999999e99;
    
end

% print orbital elements

fprintf ('\n        sma (km)              eccentricity         inclination (deg)          argper (deg)');

fprintf ('\n %+16.14e  %+16.14e  %+16.14e  %+16.14e \n', sma, ecc, inc * rtd, argper * rtd);

if (sma > 0.0)
    
    fprintf ('\n       raan (deg)          true anomaly (deg)         arglat (deg)           period (days)');
    
    fprintf ('\n %+16.14e  %+16.14e  %+16.14e  %+16.14e \n', raan * rtd, tanom * rtd, arglat * rtd, period);

else
    
    fprintf ('\n       raan (deg)          true anomaly (deg)         arglat (deg)');
 
    fprintf ('\n %+16.14e  %+16.14e  %+16.14e \n', raan * rtd, tanom * rtd, arglat * rtd);
    
end




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