Orbit Module

Directory List

Constellations

CrossScaleConstellation	Create a cross scale constellation
WalkerConstellation	Compute orbital elements for a Walker constellation.

DemoFuns

PlotZermelo	Propagate and plot the Zermelo trajectory.
RHSZermelo	Zermelo's differential equation right hand side.
ZermeloAnalyticalCost	Computes the analytical cost for Zermelo's problem.
ZermeloAnalyticalLambda	Computes the analytical lambda for Zermelo's problem.
ZermeloCost	Cost function of Zermelo's differential equations.

Demos/GravityModels

GravityCompare

Compare gravity models

Demos/OrbitControl

RelativeOrbitControl	Simulates two orbits and applies a relative controller.
TargetDemo	Demonstrates the Lambert targeting function DVTarget.

Demos/OrbitCoord

TestComputeTLE

Test ComputeTLE

Demos/OrbitManeuver

ComStarAKMDV	Plan the AKM burn by looking at the AKM thrust variations as a function of temperature.
DragCompensationDemo	Demonstrates drag compensation maneuvers.
DragCompensationDemo.mat	OP Data for drag compensation maneuvers.
GEOStationKeeping	GEO Stationkeeping Example
GeoOrbitSim	Low thrust geo transfer
GeoTransferSim	Low thrust geo transfer simulation
HohmannDemo	Demonstrate a Hohmann Transfer in simulation
InsertionDeltaV	Computes the insertion delta-V for a Hohmann transfer.
IspLimits	DV limits from fuel tank mass ratio
LowThrustOrbitRaising	Simulate a low-thrust orbit raising from an ISS orbit
OrbMnvrDemo	Runs demonstrations of selected orbit maneuver functions.
OrbitPlaneChange	Inclination Change
OrbitSeparationSimulation	Orbit separation simulation with discrete delta-Vs
OrbitSeparationTradeStudy	Orbit separation demo
TDV	Demonstrate the various orbit change functions.
TDVDisp	Compute delta-v dispersions given declination and velocity errors.
TwoBurnGeoTransfer	Delta-V for a Hohmann transfer between LEO and GEO.
TwoStageDesign	Computes a preliminary design of a two stage to orbit vehicle.

Demos/OrbitMechanics

ComStarFBudget	Demonstrate a fuel budget for the ComStar satellite.
J2EffectsDemo	Demo of J2 Orbit Effects in Simulation
SunSynchElements	Sun-Synchronous Orbit Demo
TNORAD	Demo two of the five NORAD element propagators.
TwoOrbSim	Generates two orbits and plots their relative positions

Demos/OrbitPropagator

DragDemo	Demonstrate drag reentry.
DragDemo.mat	Drag reentry demo OP data.
GravityModelDemo	Demonstrate using an external gravity model with PropagateOrbitPlugIn.
IntegrationAccuracyDemo	Integration Accuracy for PropagateOrbitPlugin.
LowThrustDemo	Demonstrate low thrust transfer using PropagateOrbitPlugIn.
LowThrustDemo.mat	Low thrust demo OP data.
LunarModelDemo.mat	OP Data for propagating with lunar gravity model.
OPBatchDemo	Demonstrate using PropgateOrbitPlugin in batch
OPDemo.mat	OP Data for batch mode demo.
OrbitIntegrationAccuracy.mat	OP Data for accuracy demo.
PropagatorComparison	Integration accuracy study comparing RK4, RK45, and ode113.
TopexDemo	Compare various orbit propagators with a Topex ephemeris.

Demos/OrbitSim

EarthOrbitDragDemo	Simple demo to compute drag over one orbit
EarthOrbitSim	Simulates a spacecraft in earth orbit with spherical harmonics.
GeoSim	Using a spherical harmonic gravity model
LEOPAnalysis	Perform LEOP (Launch and Early Analysis Phase) analysis.
LinOrbitDemo	Linear orbit simulation. This compares nonlinear with linear.
LowThrustEscapeSim	Simulates a low thrust escape from an earth orbit.
RelativeOrbitSim	Simulates two orbits and plots their relative positions.
TOrbit	Investigate orbit coordinate systems.
TVarEq	Simulate an orbit using the variational equations.

Demos/Visualization

MarsOrb	Generates a Mars orbit plot.
OrbViewDemo	Demonstrate the OrbView function.
RelMotionVis	Visualize relative orbital motion in the inertial frame.
SphericalHarmonicPlot	Demonstrate a high order gravity model.

Glideslope

CWSimAndPlot	Visualize a set of maneuvers propagated with Clohessy-Wiltshire equations.
ClohessyWiltshire	A special formulation of the Clohessy-Wiltshire equations.
DVGlideslope	Calculate delta-Vs for glideslope rendezvous.
GlideSlope	Compute the glide slope angle.
GlideslopeCircumnav	Calculate delta-Vs for a pulsed circumnavigation.
GlideslopeGuidance	Guidance trajectory where range rate is proportional to range.
GlideslopePlanner	Calculate sequential glideslope maneuvers for planning purposes.
SolveT	Cost function for glideslope period solver
SolveTD	Derivative for glideslope period solver

GravityModels

EGM100x100.mat	First 100 coefficients of Earth EGM2008 model (normalized).
EGM50x50.mat	First 50 coefficients of Earth EGM2008 model (normalized).
EarthGravityModel.mat	Earth GEMT1 model in a mat file (unnormalized).
GEMT1.geo	GEMT1 gravity model circa 1987.
GravityMap	Create a gravity map from a gravity model.
JGM2.geo	Earth JGM2 70-term gravity model, circa 1993.
JGM3.geo	Earth JGM3 70-term gravity model, circa 1995.
LoadGEM	Load the GEMT1 data.
LoadGravityModel	Load a spherical harmonic gravity model (.geo) from GSFC/U. of Texas at Austin.
NormalizationMatrix	Generate a normalization matrix for spherical harmonics
ProcessASCIIGravityModel	Process an ascii file of spherical harmonic coefficients.
UnnormalizeGravity	Remove normalization from a gravity model
WGS84.geo	Earth WGS84 gravity model, 1984.

LowThrust

AFLTSpiral	Low thrust delta v between two circular orbits.
DVLowThrust	Delta-V from a circular to noncoplanar circular orbit.
LTPlaneChange	Low-thrust transfer between two circular, nonplanar orbits
LTSpiral	Delta v of a low thrust spiral between two circular orbits.
LowThrustEscape	Compute the delta-V required for a low thrust escape.
LowThrustSim	Simulate low thrust spacecraft.

Optimization

CostLowThrust2D	Cost function for 2D low thrust trajectory optimization.
EqPartials	Computes sail partials using a first difference approximation
FLowThrust2D	Derivatives for the low-thrust planar orbit problem.
Plot2DTrajectory	Propagate and plot the planar trajectory from costates (indirect optimization).
SpacecraftFromAccel	Size a spacecraft from acceleration time history
TrajectoryOptimization	Performs indirect trajectory optimization on costates with fminsearch.

OrbitControl

ApogeeRaisingDeltaV	Compute two delta-v's to raise the apogee and circularize the orbit.
DVTarget	Perform targeting in Earth orbit using the Lambert algorithm.
DiscreteTrajectoryToDeltaV	Compute accelerations for a point to point trajectory.
DoDeltaV	Updates orbital elements based on an impulsive delta-V
LQOrbitController	Creates a linear quadratic relative orbit controller.
LambertOpt	Use Lambert with optimization of start and transfer time
LowThrustControl	Compute control acceleration for low thrust orbit raising
RendezvousPlanner	Compute delta-Vs for rendezvous.
TransferTimeToDV	Converts the transfer time to delta-V for a rendezvous.

OrbitCoord

A2Apogee	Computes the apogee radius from a and e.
A2Perige	Computes the perigee radius from a.
AB2RADec	Convert angles to right ascension and declination
AE2RPRA	Converts semi-major axis and eccentricity to right perigee and apogee.
AE2VR	Computes v and r given a, e and nu.
AENuToVelAltFpa	Convert semi-major axis & ecc. to velocity, altitude, flight path angle
AERToNu	Converts semi-major axis, eccentricy and r to true anomaly.
Alfriend2El	Convert an Alfriend orbital element set into the standard orbital element set
CEcl2SER	Transformation matrix from ecliptic to Sun-Earth rotating frame.
CEq2C	Transformation matrix from equinoctial to cartesian coordinates.
CP2I	Transformation matrix from the perifocal frame to the inertial frame.
CT2R	Converts tangential accelerations into radial using the true anomaly.
CT2REl	Converts tangential accelerations into radial using the elements vector.
CT2REq	Converts tangential accelerations into radial using equinoctial elements.
CartOrb2Cyl	Convert a cartesian orbit state to cylindrical orbit coordinates.
CircularOrbitState	Compute inertial position and velocity given circular orbit elements
ComputeTLE	Compute two-line elements from osculating elements
ConvertNORAD	Convert NORAD TLE string into a data structure.
CylOrb2Cart	Convert a cylindrical orbit state to a cartesian orbit.
ECEFToECI	Compute the transformation matrix that rotates ECEF to ECI coordinates.
ECI2OrbitPlane	This function takes a position vector and orbit normal vector in ECI
ECIToRLP	Transform from ECI to Rotating Libration Point Coordinates
ENUToECEF	Compute the transformation matrix that rotates ENU to ECEF coordinates.
El2Alfriend	Convert the standard orbital element set into an Alfriend orbital element set
El2Eq	Converts conventional elements to equinoctial elements.
El2LKl	Converts orbital elements to commonly used angular quantities
El2RV	Converts orbital elements to r and v for an elliptic orbit.
ElToFPA	Planar elements to flight path states
ElToMEq	Transforms Keplerian elements to modified equinoctial elements.
Eq2El	Converts equinoctial elements to Keplerian elements.
Eq2RV	Converts equinoctial elements to r and v for an elliptic orbit.
InitialOrientation	Compute the initial quaternion and body rates for a circular orbit,
LatLonToRAz	Compute range and azimuth angle between two points on a sphere.
LocusOfVisibility	Determine the locus of visibility for a satellite above a planet.
MEQToECI	Transform to ECI frame from tangential coordinates.
MEqToEl	Transforms modified equinoctial elements to Keplerian elements.
MEqToRV	Transforms modified equinoctial elements to r and v.
OrbElemDiff	Computes the differences between orbital element vectors.
OrbNEq	Computes the orbit normal from equinoctial elements.
OrbitClosestPoint	Computes closest orbit point to points on the earth.
OrbitalPlane	Display the orbital plane with the Earth.
OverArea	Determine whether r is in a area on the surface of an Earth sphere
PtchCnc	Computes planet centered orbital elements.
RAzToLatLon	Transform (range,azimuth) to (latitude,longitude) coordinates.
REq	Computes the orbit radius and the parameter from equinoctial elements.
RHAzToECEF	Transform (range,azimuth) to (latitude,longitude) coordinates
RLagrange	Computes a collinear Lagrange point.
RPQ	Computes the orbit position vector in the perifocal plane.
RPV2AE	Converts perigee and velocity at perigee to a and e
RV2A	Computes the semimajor axis given position and velocity magnitudes.
RV2AE	Converts R and V to a, e, rP and rA.
RV2El	Converts Cartesian state to Keplerian orbital elements.
RV2Eq	Converts R and V to Equinoctial elements
RV2RADec	Right ascension and declination with their rates from position and velocity vectors.
RVSet2El	Batch form of RV2El.
RVToMEq	Transforms elements r and v to modified equinoctial.
RelState2Absolute	Given the reference absolute state, converts a relative orbit state to absolute.
RelativeCoord	Computes relative coordinates between two orbits.
RelativeOrbitState	Find the relative coordinate state vector between x2 and x1.
TEMEToPEF	Computes the matrix from TEME to the pseudo-earth fixed frame.
VR2AE	Computes sma and eccentricity given v and r.
VelAltFpaToAENu	Converts flight path angle, velocity and r to orbital elements.
VelAltFpaToHAE	Compute (VF,gammaF) at hF from (V0,gamma0) at h0.

OrbitData

Earth.txt	HORIZONS ephemeris data for Earth.
LEOData.mat	Sample LEO orbit data.
Messenger.txt	HORIZONS ephemeris data for Messenger.
MirNORAD.txt	Mir NORAD two-line elements data.
NewHorizons.txt	HORIZONS ephemeris data for New Horizons.
ReadHorizons	Reads in JPL Horizons data.
earthmap.fig	3D Earth map in Normalized Earth-Fixed Coordinates.
mir.txt	Mir NORAD elements.
sdp.txt	SDP NORAD 2 line elements test data
sgp.txt	SGP NORAD 2 line elements test data

OrbitManeuver

AccelToThrustAndQ	Convert an acceleration vector into a thrust quaternion.
ApogeeDV	Delta-V to change apogee.
ChangePerigee	Compute the delta V required to change perigee
CircularManeuver	Compute maneuver profile for a circular maneuver.
CoplanarOrbit	Find a coplanar orbit with a different altitude
DVAKM	Computes the temperature required to achieve the desired delta-V.
DVDisp	Computes the velocity and inclination errors for a delta V burn
DVGeoTrx	Delta-V needed to control the longitudinal acceleration in geo.
DVHoh	Computes the delta-v for a coplanar Hohmann transfer
DVHohE	Compute Hohmann transfer between two elliptical coplanar orbits.
DVHohInc	Compute impulses for an eccentric Hohmann transfer with a plane change.
DVIDrift	Computes inclination drift for geosynchronous orbits.
DVOCCapt	Computes the optimal circular capture radius and delta-v.
DVOHohmn	Delta-v for injection into an interplanetary orbit using Olberth's method.
DVReentry	Computes the delta-V for reentry from a near-circular orbit.
HohmannTransferRendezvous	Compute delta-Vs for rendezvous using Hohmann transfers.
LTVCon	Solve the Linear Terminal Velocity Constraint problem.
ManeuverDuration	Compute maneuver duration given constant thrust.
ManeuverEnvelope	Compute the maneuver envelope for a spacecraft rotating around another.
MaximumDeltaV	Returns the maximum velocity change fraction
OptDIH	Optimal inclination corrections for a Hohmann transfer with a plane change.
OrbMnvrBielliptic	Delta-V for a bielliptic orbit maneuver between circular orbits.
OrbMnvrCircularize	Computes the delta-V for a 2-burn transfer to a circular orbit at rF.
OrbMnvrHohmann	Computes the delta-V for a Hohmann transfer between circular orbits.
OrbMnvrInclAndHohm	Delta-V for a Hohmann transfer between circular non coplanar orbits.
OrbMnvrInclination	Computes the delta-V for an inclination correction.
OrbMnvrInsertion	Computes the delta-V to insert into a planetary or lunar orbit.
OrbMnvrLongAndIncl	Delta-V for an longitude and inclination change maneuver for a circular orbit.
OrbMnvrLongitude	Computes the delta-V for a longitude only maneuver for a circular orbit.
OrbMnvrLowerApogee	Computes the delta-V to lower apogee.
OrbMnvrOneTangent	Computes the delta-V for a one-tangent transfer between coplanar orbits.
OrbMnvrPerigee	Compute the delta V required to change perigee.
OrbMnvrPhaseChange	Computes the delta-V for a phase change.
OrbMnvrSemimajor	Computes the delta-V for a change in semi-major axis.
SemimajorAxisRevisit	Computes semi-major axis from revisit and swath angle.
Tfr2Imp	Two impulse optimal orbit transfer.
Tfr3Imp	Three impulse orbit transfer.
TransferAndPhaseAngle	Computes the transfer and phase angle given transfer time.
TriangularManeuver	Compute maneuver profile for a triangular maneuver.
TrxCntrl	Plans delta-v maneuvers to correct triaxial motion.
VDA	Computes the heliocentric departure vector for a spacecraft.

OrbitMechanics

AGravity	Compute the gravitational acceleration in spherical coordinates.
AGravityC	Compute the gravitational acceleration in cartesian coordinates.
APlanet	Perturbing acceleration due to a planet on a spacecraft.
AZLaunch	Launch azimuth to achieve a desired orbit inclination.
AdjustMeanAnomaly	Adjusts mean anomaly in an elements vector for a different start time.
BallisticCoefficient	Compute the ballistic coefficient from a CAD model.
DVDragFPA	Drag loss along the arc of a transfer orbit determined from flight path angle
DVOrbitDrag	Compute the drag loss over an orbit arc in planetary orbit.
DVUpperStageTwoImpulse	Delta-V for an upper stage with two impulses to orbit
E2M	Converts eccentric anomaly to mean anomaly.
E2Nu	Computes the true anomaly from the eccentric or hyperbolic anomaly.
E2RAORP	Ratio of apogee radius to perigee radius given eccentricity for an ellipse
Energy	Compute the energy for any orbit
FBudget	Computes a fuel budget from a list of events.
FPA	Computes the flight path angle from eccentricity and true anomaly.
FlightPathAngle	Compute the flight path angle for a Cartesian state.
FlightPathAngleFromCartEl	Compute the flight path angle from elements
Flyby	Calculates orbital elements and perigee radius for a planetary flyby.
GVEDynamics	Compute continuous-time relative dynamics for Gauss' variational equations
GetFuelBudgetDataFromCAD	Retrieve fuel budget information from a CAD model.
GroundCoverage	Compute ground coverage for an orbit.
HEl	Computes orbit angular momentum from the elements vector.
HEq	Calculates the orbit angular momentum given the input in equinoctial elements
IDrift	Computes the inclination rate of an earth orbit due to the sun.
InclinationRate	Finds the inclination rate and orbit normal vector.
J2OrbitEffects	Compute J2 effects on the orbital elements - analytic model.
K2L	Converts eccentric longitude to true longitude
L2K	Converts eccentric longitude to true longitude
LTAN	Compute the local time of the ascending node.
LambertTOF	Solves the Lambert time of flight problem using Battin's method.
LonMot	Computes the longitudinal motion due to an acceleration a
M2E	Computes the eccentric anomaly
M2EApp	Approximate root to Kepler's Equation for elliptical and hyperbolic orbits.
M2EEl	Eccentric anomaly for an ellipse.
M2EHy	Eccentric anomaly for a hyperbola.
M2Nu	Computes the true anomaly from the mean anomaly.
M2NuAbs	Computes the true anomaly from the mean anomaly without wrapping.
M2NuPb	Generate the true anomaly from the mean anomaly for a parabola.
M2NuS	Computes true anomaly from a series expansion
ML2K	Eccentric longitude from the mean longitude by solving Kepler's equation.
ML2KApp	Approximate eccentric longitude from the mean longitude.
MajorMinorToAE	Convert semi-major and semi-minor lengths to a and e
MeanAnom	Computes the mean anomaly from change in time
MeanMotn	Computes the mean orbit rate for an elliptic orbit
N2A	Computes the semi-major axis given the mean orbit rate.
NORADToEl	Convert NORAD to Keplerian elements.
NonDimensionalOrbit	Dimensionalize or non-dimensionalize an orbit.
Nu2E	Converts true anomaly to eccentric or hyperbolic anomaly.
Nu2M	Converts true anomaly to mean anomaly.
Nu2MAbs	Computes the mean anomaly from the true anomaly without wrapping btwn -pi / pi
NuInf	Computes the asymptote angle for a hyperbola.
OrbDecay	Approximate time for an orbit to decay based on an exponential atmosphere.
OrbNEl	Conmputes the orbit normal from orbital elements.
OrbRate	Compute the orbital rate from distance and semi-major axis.
OrbitDrag	Approximate the acceleration due to drag over an orbit arc, h0 to hF.
OrbitFromSynodic	Computes the elements of an orbit with the desired synodic period.
OrbitInsertion	Compute orbit insertion velocity and flight path angle.
OrbitJacobian	Computes the Jacobian for a spherical planet gravity model.
OrbitVel2D	Compute (VF,gammaF) at hF from (v0,gamma0) at h0.
PAHohmn	Computes the phase angles for a Hohmann intercept
PDrag	Dynamic pressure on a satellite based on an exponential atmospheric model.
PToSMA	Compute the semi-major axis from an orbit period or mean motion.
Parameter	Computes the parameter for any orbit.
Period	Compute the period for an orbit.
PeriodToSMA	Compute the semi major axis from the period
PlanetaryAccelerations	Compute perturbations due to other planets aside from the center.
PropEph	Propagates the ephemeris one step.
RAEToM	Computes m, nu and E from r, a and e.
RARP2A	Computes the semi major axis from apogee and perigee radii
RARP2E	Computes the eccentricity from apogee and perigee radii
REl	Computes the orbit radius and the parameter from Keplerian elements.
ROrb	Computes the orbit radius.
RPRA2AE	Converts perigee and apogee to semi-major axis and eccentricity
RToNu	Computes true anomaly from semi-major axis, eccentricity and radius.
RandomOrbit	Monte-Carlo orbit simulation.
SLR	Computes the semi-latus rectum.
SphOI	Computes the sphere of influence ratio
SunSyncInclination	Compute the sun-synchronous inclination for an orbit.
SunSyncNode	Compute the ascending node for the desired local time.
SynPer	Computes the synodic period between two orbits.
SynodicFromSemimajorAxes	Computes the synodic period from semi-major axis.
TARes	Resolves the transfer angle ambiguity.
THohmn	Computes all times for a Hohmann Transfer.
TLETime	Time vector for a TLE
TTrnsHoh	Computes the Hohman transfer time between two orbits
TWaitHoh	Computes the wait time at the target.
TimeOfFlightHyperbola	Time of flight in a hyperbola
TrueAnomalySpacing	Generates an array of true anomalies.
UGravity	Compute the earth's gravitational potential.
VEscape	Computes the escape velocity.
VInf	Computes the asymptotic velocity for a hyperbola (v infinity)
VInfRPToRV	Generate RV from parameters v infinity and periapsis
VOrbit	Computes the orbital velocity from radius.
VPQ	Computes the orbit velocity in the perifocal plane.
VarEqP	Computes the right hand side of the equinoctial variational equations (polar).
VarEqT	Computes the right hand side of the variational equations (tangential)
VarGausP	Computes the right hand side of the Gauss variational equations.
VarGausT	Variational Keplerian equations with the acceleration in tangential coordinates.

OrbitPropagatorGUI

FAero	The aero force model. This is a pure drag model.
FAtmDensity	Atmospheric density model function that calls the Jacchia J70 model.
FEngine	Computes the engine force vector in the ECI frame.
FOrbHFOP	High fidelity orbit model designed to work with ode113.
FPlantNoise	The plant noise model.
FSolar	The solar force model for OrbitPropagator.
PropagateOrbitPlugIn	Main GUI for propagating an orbit using ode113 and user-selected models.
ReadTopexPOE	Reads a Topex Precision Orbit Ephemerides file.
StoppingConditions	Stopping conditions for an orbit.
StoppingPlugIn	Create a stopping conditions plug in for OrbitPropagator.

OrbitSim

J2Prop	Propagate an ECI state into the future with the J2 perturbation.
LoadNORAD	Load a set of NORAD elements (TLEs) and convert to Cartesian states
NORAD	Propagates the NORAD two line elements (TLE), ex. SGP, SGP4, SGP8.
OrbitProp	Propagate an ECI state into the future with point mass orbit.
PropagateLowThrust	Propagate a low thrust trajectory with constant thrust
PropagateTLE	Propagates the NORAD two line elements, ex. SGP, SGP4, SGP8.
PropagateTLECommonTime	Propagates a set of NORAD two line elements with a common time frame.
RVFromKepler	Generate an orbit by propagating Keplerian elements.
RVOrbGen	Generate an orbit by propagating Keplerian elements.
RVOrbHF	High fidelity orbit simulation with drag and planetary perturbations.

RHSOrbit

CylOrb	Right-hand-side of the cylindrical orbit equations about a mass point.
DLinOrb	Computes the linearized orbit equations for a discrete system.
FOrb	Computes the right-hand-side of the orbit equations about a mass point
FOrbCart	Computes the right-hand-side of the orbit equations about a mass point.
FOrbCartH	Computes the right-hand-side of a spherical harmonic orbit (Cartesian).
FOrbHF	High fidelity orbit model right-hand-side.
FOrbHF2	High fidelity orbit model designed to work with ode113.
FOrbLEOP	Orbit model for Launch and Early Orbit Operations.
FOrbLEOPStopping	Events function for FOrbLEOP
FOrbLTE	Orbit model for low thrust escapes.
FOrbLowThrust	Right hand side of the planar orbit equations assuming constant thrust
FVarEq	Computes the right hand side of the variational equations.
LinOrb	Computes the linearized orbit equations.
RHS2DPolarOrbit	RHS for 2D polar orbit
RHSEquinoctial	Right-hand-side for equinoctial elements.
RHSGeoHarm	Computes the right hand side for a gravity model.
RHSGeoJ2	Computes the right hand side for Earth gravity with J2.
RHSLinOrb	Continuous form of linearized orbit (Hills equations).
RHSOrbit	Right hand side for a simple Cartesian orbit.
RHSOrbit2D	Two dimensional orbit dynamics.
RHSOrbitPlanetFixed	Cartesian orbital equations in a planet fixed (rotating) frame.
SphOrb	Point mass orbit equations in equatorial spherical coordinates.

Visualization

AttitudeFrames	Draw attitude frames. You can add these to any 3D picture.
DrawOrbitPlane	Draw an orbit plane. You can add this to any 3D plot.
OrbTrack	Plot a planet-fixed orbit track in 2D or 3D.
OrbTrackECI3D	Plot an orbit track in the ECI frame.
OrbView	View orbit related parameters on a quad chart.
PlotLambertTransfer	Plot a Lambert transfer trajectory as computed by DVTarget
PlotOrbitPage	Create a plot page with 2D and 3D views of an orbit.
PlotOrbitTransfer	Plot two initial orbits and a transfer orbit.
PlotPlanetaryOrbit	Plot the state in 3D with an planet map.
SeeTLE	View a set of satellite TLEs in Earth orbit, using TLEs from celestrak.
SelectTLEGroup	A GUI to view / select from a list of TLE groups at www.celestrak.com.
TrajectoryPlot	Draws a 2D or 3D trajectory plot.
TwoObjectTrajectoryViewer	Displays spacecraft and moon in a trajectory about the Earth in real time.

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