RPO Suite

TypeScript library for spacecraft relative motion propagation implementing the Yamanaka-Ankersen and Clohessy-Wiltshire algorithms for Rendezvous and Proximity Operations.

Overview

This library provides analytical solutions for propagating the relative state between two spacecraft in orbit. The implementation is based on the analytical state transition matrix derived in Yamanaka & Ankersen (2002), which extends the classical Clohessy-Wiltshire equations to handle elliptical reference orbits with exact analytical solutions.

Two propagation methods are provided:

• Yamanaka-Ankersen State Transition Matrix for elliptical orbits (0 <= e < 1)
• Clohessy-Wiltshire equations for circular orbits (e = 0)

Demo

View Demo on StackBlitz

Installation

npm install rpo-suite

Usage

import {
  propagateYA,
  trueAnomalyAtTime,
  type OrbitalElements,
  type RelativeState,
} from "rpo-suite";

const elements: OrbitalElements = {
  eccentricity: 0.1,
  gravitationalParameter: 3.986004418e14,
  angularMomentum: 5.409e10,
};

const initialState: RelativeState = {
  position: [100, 200, 50] as const,
  velocity: [0.5, -0.2, 0.1] as const,
};

const theta0 = 0;
const deltaTime = 1000;
const thetaF = trueAnomalyAtTime(elements, theta0, deltaTime);

const finalState = propagateYA(
  initialState,
  elements,
  theta0,
  thetaF,
  deltaTime,
  "RIC"
);

API Reference

Type Definitions

type Vector3 = readonly [number, number, number];

type RelativeState = {
  readonly position: Vector3;
  readonly velocity: Vector3;
};

type OrbitalElements = {
  readonly eccentricity: number;
  readonly angularMomentum: number;
  readonly gravitationalParameter: number;
};

type TrueAnomaly = number;

type Frame = "RIC" | "LVLH";

type InPlaneState = {
  readonly x: number;
  readonly z: number;
  readonly vx: number;
  readonly vz: number;
};

type OutOfPlaneState = {
  readonly y: number;
  readonly vy: number;
};

type DeriveAngularMomentum = (
  eccentricity: number,
  meanMotionRevPerDay: number,
  mu: number
) => number;

Core Functions

propagateYA

function propagateYA(
  initialState: RelativeState,
  elements: OrbitalElements,
  theta0: TrueAnomaly,
  thetaF: TrueAnomaly,
  deltaTime: number,
  frame: Frame
): RelativeState;

Propagates relative state using the Yamanaka-Ankersen State Transition Matrix. Implements Equations 80-84 from Yamanaka & Ankersen (2002) for elliptical reference orbits.

propagateHCW

function propagateHCW(
  initialState: RelativeState,
  orbitalRate: number,
  deltaTime: number,
  frame: Frame
): RelativeState;

Propagates relative state using the Clohessy-Wiltshire equations for circular reference orbits.

Utilities

trueAnomalyAtTime

function trueAnomalyAtTime(
  elements: OrbitalElements,
  theta0: TrueAnomaly,
  deltaTime: number
): TrueAnomaly;

Computes true anomaly at a future time using Kepler propagation.

trueAnomalyFromMean

function trueAnomalyFromMean(
  meanAnomaly: number,
  eccentricity: number,
  tolerance?: number
): TrueAnomaly;

Converts mean anomaly to true anomaly by solving Kepler equation using Newton-Raphson iteration.

orbitalPeriod

function orbitalPeriod(elements: OrbitalElements): number;

Calculates orbital period from orbital elements.

deriveAngularMomentum

function deriveAngularMomentum(
  eccentricity: number,
  meanMotionRevPerDay: number,
  mu: number
): number;

Derives specific angular momentum from TLE mean motion and eccentricity.

Coordinate Transformations

function toModifiedCoordinates(
  state: RelativeState,
  elements: OrbitalElements,
  theta: TrueAnomaly
): RelativeState;

function fromModifiedCoordinates(
  modifiedState: RelativeState,
  elements: OrbitalElements,
  theta: TrueAnomaly
): RelativeState;

Transforms between true and modified coordinates as defined in Yamanaka & Ankersen (2002).

Auxiliary Functions

function kSquared(elements: OrbitalElements): number;
function rho(eccentricity: number, theta: TrueAnomaly): number;
function s(eccentricity: number, theta: TrueAnomaly): number;
function c(eccentricity: number, theta: TrueAnomaly): number;
function sPrime(eccentricity: number, theta: TrueAnomaly): number;
function cPrime(eccentricity: number, theta: TrueAnomaly): number;
function J(elements: OrbitalElements, deltaTime: number): number;

Low-level functions corresponding to auxiliary variables in Yamanaka & Ankersen (2002).

Reference Frames

Two local-orbital reference frames are supported:

RIC: Radial, In-track, Cross-track

• R: Radial (away from Earth center)
• I: In-track (along velocity)
• C: Cross-track (normal to orbital plane)

LVLH: Local Vertical Local Horizontal (ordered as I, C, R)

Units

All quantities use SI units: meters (m), meters per second (m/s), seconds (s), and radians (rad). Gravitational parameter mu is in m^3/s^2 and angular momentum h is in m^2/s.

Advanced Usage

Computing Orbital Elements from TLE

import { deriveAngularMomentum } from "rpo-suite";

const elements = {
  eccentricity: 0.0001084,
  angularMomentum: deriveAngularMomentum(
    0.0001084,
    15.54225995,
    3.986004418e14
  ),
  gravitationalParameter: 3.986004418e14,
};

Propagating Over Multiple Orbits

import { propagateYA, orbitalPeriod, trueAnomalyAtTime } from "rpo-suite";

const period = orbitalPeriod(elements);
let state = initialState;
let theta = 0;

for (let i = 0; i < 3; i++) {
  const thetaF = trueAnomalyAtTime(elements, theta, period);
  state = propagateYA(state, elements, theta, thetaF, period, "RIC");
  theta = thetaF;
}

Development

bun install
bun run build
bun test

References

1. Yamanaka, K., & Ankersen, F. (2002). "New State Transition Matrix for Relative Motion on an Arbitrary Elliptical Orbit." Journal of Guidance, Control, and Dynamics, 25(1), 60-66.

2. Clohessy, W. H., & Wiltshire, R. S. (1960). "Terminal Guidance System for Satellite Rendezvous." Journal of the Aerospace Sciences, 27(9), 653-658.

3. Vallado, D. A. (2013). Fundamentals of Astrodynamics and Applications (4th ed.). Microcosm Press.

License

MIT