Moved to Kepler Orbit Simulator
The ballistic orbit page has been consolidated into the Kepler orbit simulator.
How to Use
- Input orbital parameters: semi-major axis (km), eccentricity (0–1), inclination (degrees), and argument of perigee
- Select central body (Earth, Moon, Mars) to auto-populate gravitational parameter μ
- Click "Propagate Orbit" to compute Kepler elements and visualize the trajectory; adjust time step for precision
- Toggle periapsis/apoapsis markers and velocity vectors to verify ballistic properties
Worked Example
For a Low Earth Orbit: semi-major axis a = 6,678 km, eccentricity e = 0.001, inclination i = 51.6°. Using Earth's μ = 398,600 km³/s², the orbital period T = 90.4 minutes and velocity at perigee v_p = 7.81 km/s. At apogee (6,687 km altitude), v_a = 7.73 km/s. The simulator computes true anomaly progression and predicts ground track coverage for communications planning.
Practical Notes
- Geostationary orbits require a ≈ 42,164 km and e ≈ 0; verify station-keeping delta-v budgets by comparing circular vs. elliptical transfers
- Hohmann transfers between orbits use initial and final semi-major axes; input both to calculate burn timing and fuel requirements
- High-eccentricity trajectories (e > 0.7) are useful for lunar flybys; check periapsis altitude does not intersect planetary atmosphere
- Inclination changes are expensive; validate launch site latitude constraints before designing polar missions