Orbital Mission & Ground Track Planner
Simulate orbital trajectories, Earth rotation ground tracks, eclipse daylight/shadow windows, and ground station telemetry passes over hours of sustained spaceflight.
| Orbit # | Ground Station | Acquisition (AOS) | Loss of Signal (LOS) | Duration | Max Elevation |
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| Orbit # | Daylight Entry | Umbra Entry (Eclipse) | Sunlight Fraction | Solar Array Efficiency |
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| MET (HH:MM:SS) | Latitude | Longitude | Altitude (km) | Speed (km/s) | Sunlit? |
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The Physics of Multi-Orbit Endurance Tests
For Starship’s Flight 14, remaining in orbit for 10 hours represents a massive operational leap beyond early suborbital test trajectories. A standard Low Earth Orbit (LEO) at ~200 km altitude takes approximately 88.5 minutes to complete one full revolution.
Over a 10-hour flight, the spacecraft executes roughly 6.8 complete orbits. Because the Earth rotates eastward beneath the inclined orbital plane at 15.04° per hour (plus nodal precession caused by the Earth’s oblateness J₂), each successive equator crossing shifts westward by ~22.5° longitude. This ground-track shift exposes the ship to differing telemetry tracking stations and dynamic thermal sun/shadow cycles.
Mission Critical Checkpoints in a 10-Hour Orbit
- Thermal Soaking & Boil-off: Continuous alternating 54-minute sunlight / 35-minute shadow periods test vacuum cryogenic propellant insulation.
- Ground Telemetry Handoffs: Tracking relays switch between Starbase (TX), Bermuda, Hawaii, Guam, and maritime tracking vessels.
- Microgravity RCS & Relight: Settling propellants and executing in-space Raptor vacuum engine relight tests.
- Deorbit Burn Timing: A ~95–110 m/s retrograde burn lowers perigee into the atmosphere for targeted reentry over the Indian or Pacific Ocean.
How does the simulator calculate ground tracks and J₂ nodal regression?
The orbital period is derived from Kepler's third law: T = 2π √(a³ / μ) where μ = 398600.4418 km³/s² and a = R_earth + (r_perigee + r_apogee)/2. As the satellite progresses along its true anomaly, Earth rotates at ωearth = 7.292115 × 10⁻⁵ rad/s. Furthermore, the oblateness of the Earth (J₂ perturbation = 1.08263 × 10⁻³) causes a secular regression of the ascending node ΔΩJ2 ≈ -9.9639 × (R_E / a)3.5 × cos(i) deg/day, shifting the ground track on every subsequent orbit.
What is required for a safe deorbit burn?
To safely deorbit from a 200 km circular orbit into an atmospheric entry interface (nominally ~80–100 km), the spacecraft fires its engines opposite the velocity vector (retrograde). A delta-V (ΔV) of ~85 to 110 m/s reduces perigee below 50 km, ensuring atmospheric drag will capture the vehicle on the planned ocean corridor.