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Here are the steps to solve these orbital mechanics problems: 1) Given: Launch altitude = 250 km, Velocity = 7900 m/s, Zenith angle = 89 deg Using the orbital equation: h = r (1 - cosθ) r = h/sin^2θ r = 250,000 m / sin^2(89) = 7,500,000 m = 7,500 km Semi-major axis a = r Eccentricity e = 0 Perigee (Rp) = Apogee (Ra) = a = 7,500 km 2) Given: Semi-major axis a = 7,500 km, Eccentricity e = 0.1

Author

M S Prasad

Language

EN