Hello everyone, I'm new here so If I'm writing anything wrong tell me. I'm an aerospace engineering student, for a Uni project we're designing a Mars lander and I'm trying to model the transfer orbit from Earth to Mars on GMAT in order to have the orbit and position for the landing and departure phase. My orbital mechanics knowledge is very limited since my bachelor is focused on aeronautics rather than space so I'm having some trouble. The launch date from Earth should be around 2040 so I've used a matlab script with a lambert solver that I found in the book Orbital Mechanics for Engineering Students, Exercise 8.08 for reference, to have the data about the transfer orbit.
This is the output of the script
Example 8.8
Departure:
Planet: Earth
Year : 2040
Month : May
Day : 7
Hour : 0
Minute: 0
Second: 0
Julian day: 2466281.500
Planet position vector (km) = [-1.03902e+08 -1.09492e+08 11864.8]
Magnitude = 1.50944e+08
Planet velocity (km/s) = [21.1239 -20.6176 0.00138267]
Magnitude = 29.5179
Spacecraft velocity (km/s) = [19.3726 -28.0145 -0.931206]
Magnitude = 34.0731
v-infinity at departure (km/s) = [-1.75132 -7.39691 -0.932589]
Magnitude = 7.6584
Time of flight = 493 days
Arrival:
Planet: Mars
Year : 2041
Month : September
Day : 12
Hour : 0
Minute: 0
Second: 0
Julian day: 2466774.500
Planet position vector (km) = [8.87105e+07 2.08065e+08 2.18782e+06]
Magnitude = 1.50944e+08
Planet velocity (km/s) = [-21.3728 11.5668 0.766488]
Magnitude = 24.3141
Spacecraft Velocity (km/s) = [-23.9898 0.456169 0.496433]
Magnitude = 23.9992
v-infinity at arrival (km/s) = [-2.61699 -11.1106 -0.270056]
Magnitude = 11.4179
Orbital elements of flight trajectory:
Angular momentum (km^2/s) = 5.03387e+09
Eccentricity = 0.374728
Right ascension of the ascending node (deg) = 46.6618
Inclination to the ecliptic (deg) = 1.60102
Argument of perihelion (deg) = 134.835
True anomaly at departure (deg) = 45.0034
True anomaly at arrival (deg) = 245.419
Semimajor axis (km) = 2.22129e+08
Period (days) = 660.875
Mind that I'm still trying to reproduce the orbit on GMAT so optimization will come after.
The problem is the GMAT modelling, I've watched the GMAT tutorial "Mars B-Plane Targeting" and I'm trying to adapt it but I don't know where to start. Any help would be appreciated
This is the output of the script
Example 8.8
Departure:
Planet: Earth
Year : 2040
Month : May
Day : 7
Hour : 0
Minute: 0
Second: 0
Julian day: 2466281.500
Planet position vector (km) = [-1.03902e+08 -1.09492e+08 11864.8]
Magnitude = 1.50944e+08
Planet velocity (km/s) = [21.1239 -20.6176 0.00138267]
Magnitude = 29.5179
Spacecraft velocity (km/s) = [19.3726 -28.0145 -0.931206]
Magnitude = 34.0731
v-infinity at departure (km/s) = [-1.75132 -7.39691 -0.932589]
Magnitude = 7.6584
Time of flight = 493 days
Arrival:
Planet: Mars
Year : 2041
Month : September
Day : 12
Hour : 0
Minute: 0
Second: 0
Julian day: 2466774.500
Planet position vector (km) = [8.87105e+07 2.08065e+08 2.18782e+06]
Magnitude = 1.50944e+08
Planet velocity (km/s) = [-21.3728 11.5668 0.766488]
Magnitude = 24.3141
Spacecraft Velocity (km/s) = [-23.9898 0.456169 0.496433]
Magnitude = 23.9992
v-infinity at arrival (km/s) = [-2.61699 -11.1106 -0.270056]
Magnitude = 11.4179
Orbital elements of flight trajectory:
Angular momentum (km^2/s) = 5.03387e+09
Eccentricity = 0.374728
Right ascension of the ascending node (deg) = 46.6618
Inclination to the ecliptic (deg) = 1.60102
Argument of perihelion (deg) = 134.835
True anomaly at departure (deg) = 45.0034
True anomaly at arrival (deg) = 245.419
Semimajor axis (km) = 2.22129e+08
Period (days) = 660.875
Mind that I'm still trying to reproduce the orbit on GMAT so optimization will come after.
The problem is the GMAT modelling, I've watched the GMAT tutorial "Mars B-Plane Targeting" and I'm trying to adapt it but I don't know where to start. Any help would be appreciated