NASA · ARTEMIS PROGRAM
GATEWAY ORBITAL STABILITY EXPLORER
NRHO · THREE-BODY SIMULATION ENGINE
INITIALIZING TELEMETRY LINK…
CLASSIFIED · INTERNAL · MISSION-CRITICAL OPS
Mission Elapsed
T+ 00d 00h 00m
Awaiting Burn
Control Room
CRIT-C
Insertion Velocity 0.900 km/s
0.8000.9001.000
Satellite Mass LOCKED 40,000 kg
10k40k80k
Insertion Angle LOCKED 90.000°
85°90°95°
±0.01° sensor noise applied per trial.
Orbital Sandbox · Earth–Moon System
CRIT-D
Frame
Moon-centered, J2000
Playback Speed
Stable orbit
Warning drift
Critical decay
Current Perilune
— km
Current Apolune
— km
100%
Telemetry
CRIT-C
Velocity
—km/s
Perilune Δ
—km
Stability
AWAITING
Thruster Risk
—
Sim Day
0/ 30 days

Research Log — Trial Results

Each "Initiate Orbital Burn" run appends to this table. Per-condition averages and standard deviations appear once a condition has two or more recorded trials.

# Condition Velocity (km/s) Perilune Dev. (km) Stability Thruster Risk Recorded
No trials recorded. Initiate an orbital burn to begin.

Methodology

This simulation models the Lunar Gateway's Near-Rectilinear Halo Orbit (NRHO) using a three-body gravitational framework (Earth, Moon, Gateway) with verified NASA and NIST constants. Below are the hard-coded values used by the physics engine.

NASA / NIST Physics Constants

Lunar mass
7.342 × 10²² kg
NASA Lunar Fact Sheet
Earth–Moon distance
384,400 km
NASA Planetary Fact Sheet
NRHO perilune altitude
~3,000 km
NASA Gateway Overview 2022
NRHO apolune altitude
~70,000 km
ESA Gateway Portal
Baseline insertion velocity
~0.900 km/s
NASA NRHO parameters
Gravitational constant G
6.674 × 10⁻¹¹ N·m²/kg²
NIST CODATA

Independent Variable

Initial orbital insertion velocity, tested at three levels: 0.855 km/s (−5%), 0.900 km/s (baseline), 0.945 km/s (+5%).

Dependent Variable

Perilune altitude deviation (km) from the ideal 3,000 km perilune, measured over 30 simulated days.

Control Variables

  • Gateway satellite mass: 40,000 kg (locked)
  • Orbital insertion angle: 90° (locked)
  • Lunar mass: 7.342 × 10²² kg (constant)
  • Earth-Moon distance: 384,400 km (constant)
  • Simulation duration: 30 days per trial (constant)

Trial Noise Model

Each simulation run introduces ±0.01° insertion-angle noise plus ±0.2% velocity-magnitude jitter, modelling combined sensor margin-of-error and thruster impulse precision. These produce the trial-to-trial variance that emerges in the recorded perilune deviations.

Three-Body Physics Note

The engine models Moon gravity directly plus a scaled Earth perturbation (5% of true Earth gravity) representing third-body effects. The reduced scale is a deliberate simplification that preserves the qualitative three-body character (non-Keplerian orbit shape, slow drift) while keeping the velocity → perilune-deviation signal visible across the 30-day window. A full N-body simulation would require station-keeping logic beyond the scope of this engine.

Sources (MLA, 9th edition)

All references cited in the Methodology, Hypothesis, and Design Rationale tabs are listed below.

NASA. “Gateway: About.” NASA, 2022, www.nasa.gov/reference/gateway-about/.
European Space Agency. “Gateway Overview.” ESA, 2022, www.esa.int/Science_Exploration/Human_and_Robotic_Exploration/Exploration/Gateway.
NASA. “Artemis: Moon to Mars.” NASA, 2023, www3.nasa.gov/specials/artemis/.
National Institute of Standards and Technology. “CODATA Internationally Recommended Values of the Fundamental Physical Constants.” NIST, 2023.
NASA. “Lunar Fact Sheet.” NASA Goddard Space Flight Center, 2022.
University of Colorado Boulder. “Gravity and Orbits.” PhET Interactive Simulations, 2023, phet.colorado.edu.
AZoBuild. “Lunar Regolith Radiation Shielding.” AZoBuild, 2022, www.azobuild.com/article.aspx?ArticleID=8794.

Design Rationale

Every interface decision in this simulation is documented and justified below.

Why sliders, not text inputs?
Sliders constrain user inputs to physically realistic ranges, preventing invalid entries and mirroring how real mission control systems restrict operator input to safe operational envelopes. This is a deliberate design choice, not a convenience.
Why a live graph, not a static table?
Real-time plotting mirrors NASA telemetry dashboards and lets the user observe orbital decay as it happens, making the causal relationship between insertion velocity and NRHO stability immediately visible and intuitive.
Why NRHO specifically?
Circular orbits are the standard introductory case. NRHO is a real three-body, highly elliptical orbit representing the actual engineering challenge NASA faces — a more authentic system to study and visualise.
Why hard-coded NASA/NIST constants?
Using verified, cited physical constants ensures the simulation produces scientifically defensible results rather than abstract approximations, directly bridging the digital model to real Artemis mission parameters.
Why ±0.01° insertion-angle noise (with velocity jitter) per trial?
A deterministic simulation produces identical results on every run, making repeated trials meaningless. Modelling sensor and thruster precision creates realistic trial-to-trial variance, so the dataset has a legitimate mean and standard deviation.