[SOLVED] Orbit Question 1. The problem statement, all variables and given/known data A satellite is in a circular earth orbit of altitude 400km. Determine the new perigee and apogee altitudes if the satellite on-board engine gives the satellite a radial (outward) component of velocity of 240m/s. Answers: Z_apogee = 621km, Z_perigee = 196km 2. Relevant equations mu = 398600km^3/sec^2 h (specific angular momentum) = r*v_perpendicular a = semimajor axis of ellipse 2*a = perigee radius + apogee radius v^2/2 + mu/r = -mu/(2*a) ( specific energy equation ) e= eccentricity of the orbit theta = the true anomaly angle ( angle between eccentricity vector and the position vector) r_perigee = a(1-e) radius of the earth = 6378km r = h^2/mu * (1+e*cos(theta)) V_cir = sqrt(mu/r) 3. The attempt at a solution I started off by getting the magnitude of the velocity (v_perp^2 + v_radial^2)^.5 then calculated the escape velocity to see if it would be a closed orbit. I found that it was still an ellipse (i also calculated the specific energy and found that it was negative... thus being an ellipse) from the specific energy i got 'a' ( -mu/(2*a) ) and using rp = a(1-e) = h^2/mu*(1+e)^-1 (at theta = 0 for rp) however... when i did this I used the same h that i got from the satellite being a circular orbit.. which i am unsure if that was incorrect to do.. finally after solving for e, i plugged it back into rp = a(1-e) to get ~ 6543km... then the altitude zp = 6543- 6378 = 165km.. which is incorrect... can anyone help?? Thanks.