Satellites
Add satellites to a scenario — import from TLE/OMM or design from orbital elements — and propagate the orbit using Two-Body, J2 Analytical, SGP4, or Numerical propagators.
Satellites are the core objects in a scenario. You can import real catalog satellites or design your own, then propagate their orbit to get position and velocity over time.
Adding a satellite
When you add a satellite, you choose:
- Central body — Earth or Moon. This determines the available propagators and input formats.
- Propagator — the mathematical model used to compute the orbit over time.
- Input format — how you define the initial orbit.
Central body
| Central body | Available propagators |
|---|---|
| Earth | Two-Body, J2 Analytical, SGP4, Numerical |
| Moon | Two-Body, Numerical |
Moon orbits use the Moon inertial frame with lunar gravity. SGP4 and J2 Analytical are Earth-only.
Propagators
Two-Body · Keplerian
Keplerian motion — a two-body problem solved from classical or circular element sets. No perturbations.
- Use for: quick studies, lunar transfer approximations, sanity checks.
- Limitations: no atmospheric drag, no J2, no higher-order perturbations. The orbit is a perfect ellipse.
- Available for: Earth, Moon.
J2 Analytical
Secular J2 perturbations for LEO orbits. Includes nodal regression and apsidal precession.
- Use for: LEO orbit design where you need RAAN drift and perigee precession but not full numerical fidelity.
- Limitations: Earth-only. Uses Brouwer-Lyddane first-order J2 secular rates, derived via a series expansion that assumes small eccentricity, so accuracy degrades at high eccentricity — best for near-circular orbits. No atmospheric drag, no solar radiation pressure.
- Available for: Earth.
SGP4 · Simplified Perturbations
The standard analytical propagator for TLE-based satellites. Fast and tuned for catalog objects.
- Use for: importing and propagating real satellites from Space-Track or public TLE catalogs.
- Limitations: Earth-only, requires TLE input, omits higher-order force models.
- Available for: Earth.
Numerical
High-fidelity numerical integration using Orekit, with an adaptive step-size Dormand-Prince 853 (DP853) integrator and configurable force models.
- Use for: high-fidelity design studies where accuracy matters — drag, solar radiation pressure, third-body perturbations, gravity field harmonics, and relativistic effects.
- Limitations: computationally intensive. Longer propagation times than analytical methods.
- Available for: Earth, Moon.
Gravity models
| Central body | Model | Description |
|---|---|---|
| Earth | EGM2008 | NGA combined spherical harmonics gravity model. In high-fidelity mode the harmonic expansion is truncated to the selected accuracy tier: Fast (40×40), Balanced (100×100), or High (150×150). In simplified mode only zonal harmonics are used: J2 (degree 2) or J2–J6 (degree 6). |
| Moon | GL0660B | GRAIL + LOLA combined lunar gravity model. Same accuracy tiers: Fast (5×5), Balanced (80×80), or High (120×120). The Moon’s slow rotation means a time-frozen gravity field is an excellent approximation, so a constant provider is used. |
Numerical force model options
When the Numerical propagator is selected, the following force models can be toggled and configured:
| Option | Description | Default |
|---|---|---|
| Gravity field | Enable spherical harmonics gravity model instead of point-mass | On |
| Gravity mode | Simplified (J2 or J2–J6) or High fidelity (full spherical harmonics) | Simplified |
| Simplified gravity | When in simplified mode: J2 or J2-J6 | J2 |
| Gravity model accuracy | When in high-fidelity mode: Fast, Balanced, or High (controls degree/order of the harmonic expansion) | Fast |
| Atmospheric drag | Model drag using the NRLMSISE-00 empirical atmosphere model | Off |
| Solar radiation pressure | Model SRP using the satellite’s area and reflectivity coefficient | Off |
| Third-body Moon | Include lunar gravity perturbation | Off |
| Third-body Sun | Include solar gravity perturbation | Off |
| Relativity | Include relativistic correction | Off |
| Mass (kg) | Satellite mass (used for drag/SRP) | 100 |
| Area (m²) | Cross-sectional area (used for drag/SRP) | 1 |
| Drag coefficient (Cd) | Drag coefficient | 2.2 |
| Reflectivity coefficient (Cr) | Radiation pressure coefficient | 1.8 |
For Moon orbits, Earth and Sun third-body perturbations are included automatically (Earth’s gravity is the dominant perturbation for lunar orbits), and drag is not applicable (the Moon has no atmosphere).
Atmospheric drag model
When drag is enabled, atmospheric density is computed using the NRLMSISE-00 empirical model — the naval research laboratory standard for upper-atmosphere density from ground to 1000+ km altitude.
NRLMSISE-00 is not a static density table — it accounts for actual solar and geomagnetic activity using historical and predicted space weather data:
- Solar flux (F10.7): the 81-day centered average is used, which smooths short-term fluctuations while tracking the solar cycle.
- Geomagnetic activity (Ap): 3-hour resolution Ap values are used (8 values per day), capturing geomagnetic storms and daily variations.
The space weather data comes from the CSSI Space Weather dataset (SpaceWeather-All-v1.2.txt), which is bundled with SMD and covers from 1947 through October 2096. This means:
- For past dates, measured historical solar and geomagnetic data is used.
- For future dates (up to October 2096), predicted values from the CSSI dataset are used.
- You do not need to configure or upload any space weather data — it is included.
Input formats
The available input formats depend on the selected propagator and central body.
For SGP4
SGP4 requires TLE input. Three entry modes:
| Mode | Description |
|---|---|
| TLE Search | Search by NORAD ID to fetch the latest TLE from the catalog |
| TLE Paste | Paste a raw two-line TLE manually |
| OMM Parameters | Enter CCSDS Orbit Mean-elements Message parameters directly |
SMD handles TLE and OMM seamlessly: when OMM is available, elements are converted internally without lossy TLE string synthesis.
For Two-Body, J2 Analytical, and Numerical
| Format | Fields | Available for |
|---|---|---|
| Circular | Altitude, inclination, RAAN | Earth, Moon |
| Keplerian | Semi-major axis, eccentricity, inclination, RAAN, argument of perigee, true anomaly | Earth, Moon |
| Cartesian | Position (x, y, z) and velocity (vx, vy, vz) in km and km/s | Earth, Moon |
| Geosynchronous (GSO) | Longitude, inclination | Earth only |
| Sun-synchronous (SSO) | Altitude, Local Time of Descending Node (LTDN), true anomaly | Earth only, J2 Analytical or Numerical |
SSO is not available for Two-Body (it requires J2 to maintain the sun-synchronous condition).
Viewing results
Propagated data renders on the 3D globe (or Moon viewer) and as time series: position, velocity, ground track, and classical elements along the trajectory.
Next steps
- Ground Stations — define sites for access analysis
- Access — compute visibility windows
- Sensor — define satellite sensors for field-of-view access analysis
- Antenna — define satellite antennas for link budget computations
- Target Pointing — override attitude during access windows
- Model — design the satellite 3D model for power analysis
- Power — analyze power generation and battery state