Hashir
New member
Dear all,
I have a small doubt in the plane change manoeuvre calculation between two circular orbits in GMAT.
My initial circular orbit has Semi major axis = 8865 KM , inclination = 50 deg and final orbit has same a but i = 0 deg.
Satellite’s initial position is set in a such a way that, RAAN is 90 deg, Argument of perigee as 0 deg and TA as 0 deg in the first circular orbit.
So the satellite will be at inertial Y axis at its initial position, where the ascending node is present. Now after the first propagate, I want to bring satellite to the ascending node position (the initial position) before applying the delta V for the plane change.
So In GMAT, I set timeElapsed property of first propagate as 1200 second and now I want to set Earth.TA Parameter for the second propagate to bring satellite to ascending node, where TA is true anomaly. I could not find any proper calculation to find the true anomaly for this problem.
Could anyone please give me any insight on this matter?
Thanks in advance
I have a small doubt in the plane change manoeuvre calculation between two circular orbits in GMAT.
My initial circular orbit has Semi major axis = 8865 KM , inclination = 50 deg and final orbit has same a but i = 0 deg.
Satellite’s initial position is set in a such a way that, RAAN is 90 deg, Argument of perigee as 0 deg and TA as 0 deg in the first circular orbit.
So the satellite will be at inertial Y axis at its initial position, where the ascending node is present. Now after the first propagate, I want to bring satellite to the ascending node position (the initial position) before applying the delta V for the plane change.
So In GMAT, I set timeElapsed property of first propagate as 1200 second and now I want to set Earth.TA Parameter for the second propagate to bring satellite to ascending node, where TA is true anomaly. I could not find any proper calculation to find the true anomaly for this problem.
Could anyone please give me any insight on this matter?
Thanks in advance
