Orbit Module

Directory List

Constellations

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

Demos/GravityModels

GravityCompare

Compare gravity models

Demos/Interplanetary

HelioTrajectorySimulation	Set up and run a heliocentric trajectory simulation with two objects.
SpiralToHohmann	Spiral to a Hohmann Transfer to Apophis.

Demos/LowEnergyManeuver

CRTBPExampleTrajectory	CRTBP Example Trajectory
EnergyMin4Body	Energy minimization in the 4 body problem.
LissajousOrbit	Compute the lissajous trajectory about a collinear libration point.
LoadFamilies	Periodic Orbit Families
LowEnergyMission	Low Energy Mission
PeriodicOrbitFamilies	Periodic Orbit Families
PlanetIDDemo	Planet id number from name.
PropagateDemos	Demonstrate propagation functions.
RefineXAxisInterceptDemo	Demonstrates RefineXAxisIntercept
Targeting	Demonstrate targeting
TwoDimensionalEscape	Simulate a low thrust departure from the Earth

Demos/Optimization

TrajectoryOptimizationDemo	Demonstrate the 2D trajectory optimization function for low thrust.
ZermeloCostDemo	Compute the Zermelo cost function as a function of costate.
ZermeloOptimization	Demonstrate the Trajectory optimization using the Zermelo problem

Demos/OrbitControl

ApophisTransfer	Compute a transfer from earth orbit to the asteroid Apophis.
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.
EarthDeparture	Simulate a departure from the Earth
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.
JIMOMissionAnalysis	Compute the total delta-V for the JIMO mission
LowThrustMarsMission	Compute the delta-V for a low thrust mission to Mars synchronous orbit.
LowThrustMarsOptimalTransfer	Perform an optimal transfer from earth to mars orbits.
LowThrustOrbitRaising	Simulate an orbit raising from an ISS orbit
MarsHohmannTransfer	Hohmann transfer to Mars.
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
LunarOrbitSim	Propagate an orbit about the moon. Show the visibility sun, earth, moon.
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

EarthMoonSunSim	Simulate a spacecraft in orbit near the earth and moon.
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
InterstellarSim	Interstellar mission simulation with a sun gravity assist.
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.
LowThrustHelicocentricSim	Perform a helicocentric simulation from one circular orbit to another.
NBodySolarSystem	Simulate the solar system using an n-body model
RelativeOrbitSim	Simulates two orbits and plots their relative positions.
SolarSailSim	Orbit simulation of a solar sail.
TOrbit	Investigate orbit coordinate systems.
TVarEq	Simulate an orbit using the variational equations.

Demos/StraightLine

InterstellarAsteroid	Set up and run the trajectory simulation for an asteroid intercept.
Planet9	Planet 9 orbit
PowerLimitedRocket	Power-Limited Rocket Demo
SLPLComparisonDemo	Compare a constant-thrust mission to optimal linear acceleration

Demos/ThreeBody

EarthSpiralDemo

Simulates a spiral out from LEO.

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	Form:
GlideslopeCircumnav	Form:
GlideslopeGuidance	Form:
GlideslopePlanner	Calculate sequential glideslope maneuvers for planning purposes.
SolveT	Function to pass to NewtRaph for solving for the period of a glideslope
SolveTD	Derivative function to pass to NewtRaph for solving the period

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.
LP150Q.sh	Lunar data file for LoadLP150Q
LoadGEM	Load the GEMT1 data.
LoadGravityModel	Load a spherical harmonic gravity model (.geo) from GSFC/U. of Texas at Austin.
LoadLP150Q	Load the LP150Q Lunar gravity model.
LoadSGM150	Load the SGM150 Lunar gravity model.
LunarGravityModel.mat	Gravity model based on Lunar Prospector, 75x75
NormalizationMatrix	Generate a normalization matrix for spherical harmonics
ProcessASCIIGravityModel	Process an ascii file of spherical harmonic coefficients.
SEMGravAcc	Computes the acceleration due to the Sun and the Earth/Moon.
SGM150.geo	Lunar model file for LoadSGM150
UnnormalizeGravity	Remove normalization from a gravity model
WGS84.geo	Earth WGS84 gravity model, 1984.
jgl075g1.sha	Lunar spherical harmonic ASCII file

Interplanetary

ApophisOrbit	Generate elements and orbit state for the asteroid Apophis.
EarthHelioState	Compute the position and velocity of Earth in the heliocentric frame.
EarthOrbToHelioOrb	Computes the heliocentric orbital elements after an Earth escape spiral.
ForceSimple	Simple thruster force model for orbit simulations
GravityHelio	Calculate the heliocentric gravity for the position and date.
HelioFromPlanetInit	Initialize a heliocentric orbit from a planet centered orbit.
HelioToBPlane	Transformation matrix from the heliocentric to B-plane frame.
JDStartTransfer	Find the Julian Date for a straight line trajectory to the target
LambertDoubleRendezvous	Double rendezvous problem between two planets.
NoForce	Thruster force model producing zero force.
PlanetAlignments	Plots planet alignments fur successive synodic periods.
PlanetHelioPhase	Compute the phase angle in the ecliptic plane of a given planet at jD.
PlanetOpposition	Compute Julian Dates for times when two planets are closest.
PlanetPhasing	Compute and plot positions of planet1 and planet2 at the specified times.
PlanetTransferLambert	Generate a Lambert transfer between two planets.
PlotBPlane	B-plane plots using HelioToBPlane.
RHSTrajectory	Right-Hand-Side function for solar system object trajectories
SolarSysState	Computes planet heliocentric states for a range of dates.
SynodicPlanets	Computes the synodic period from planets.

LowEnergyManeuver

CRTBP2kms	Scales CRTBP Jacobi coordinates to km and sec
CRTBPJacobiConstant	Calculate from Jacobi constant from a scaled state vector
CheckForOptimizationToolbox	Check to see if MATLAB contains the Optimization Toolbox
EnergyWRTMoon	Keplerian Energy with respect to the Moon.
IteratePeriodicOrbit	Compute a periodic orbit for a LET transfer.
J20002RotPuls	Transform from J2000 frame to rotating-pulsating
J20002RotPulsPlanetMoon	Transform a state vector from the J2000 reference frame to a rotating
Kms2CRTBP	Transform from km and sec to Jacobi coordinates for circular restricted
LETGuess	Compute a periodic orbit for a LET transfer.
LETPhasing	Calculate the nominal transfer time for a low energy transfer
LibrationCoeff	Compute the nth coefficient for the about a collinear libration point
LibrationData	Compute libration point data for the restricted three body problem.
LibrationRHS	Computes the time derivative of a state at a collinear libration point
LibrationRHSODE45	Compute the time derivative of a state at a collinear libration point,
Lissajous	Compute the lissajous trajectory about a collinear libration point
LowEnergyEarthToMoonTransfer	Compute a ballistic lunar capture trajectory from a circular Earth orbit
LowEnergyTransferInCRTBP	Calculate a low energy transfer in a CRTBP.
MinE4BP	Minimize the keplerian energy of the final state of a transfer orbit
PeriodicOrbitFromGuess	Find a second perpendicular crossing of the x-axis
PlanetId	Planet id number from name.
PlotLET	Plot the Low Energy Transfer
PlotLET3BP	Plot the Low Energy Transfer, 3 Body Problem
Propagate3BP	Propagate a Sun-Earth 3BP with origin at the Earth system barycenter.
Propagate4BP	Propagate a Sun-Earth-Moon 4BP with origin at the Earth-Moon barycenter.
PropagateCRTBP	Propagate an arbitrary CRTBP with origin at the secondary body.
PropagateCRTBPBackwards	Propagate an arbitrary CRTBP with origin at the secondary body
PropagateCRTBPToPeriapse	Propagate the CRTBP initial state to the next perigee.
PropagateCRTBPToXAxis	Propagate the CRTBP equations of motion for arbitrary mu
RHSCRTBP	Compute the RHS of the CRTBP equations of motion.
RefineXAxisIntercept	Gain a precise X axis intercept with a good initial guess
RotPuls2J2000	Transform a state vector from a rotating pulsating reference frame to J2000
Targeting3BP2	Transfer a CRTBP LET to the restricted 3 body, non-planar problem.
Targeting4BP	Transfer a LET from the restricted 3BP to the Sun/Planet/Moon 4BP
TwoDEscape	Simulate a low thrust departure from a planet
V0FromJacobiConstant	Calculate an initial velocity from the jacobi constant
f16Data.mat	Markellos f16 family orbit data
f16Ref.mat	Markellos f16 family reference data
f16pData.mat	Markellos f16' family orbit data
f16pRef.mat	Markellos f16' family reference data
f17Data.mat	Markellos f17 family orbit data
f17Ref.mat	Markellos f17 family reference data
f17pData.mat	Markellos f17' family orbit data
f17pRef.mat	Markellos f17' family reference data
f18Data.mat	Markellos f18 family orbit data
f18Ref.mat	Markellos f18 family reference data
f18pData.mat	Markellos f18' family orbit data
f18pRef.mat	Markellos f18' family reference data
f19Data.mat	Markellos f19 family orbit data
f19Ref.mat	Markellos f19 family reference data
f19pData.mat	Markellos f19' family orbit data
f19pRef.mat	Markellos f19' family reference data
f20Data.mat	Markellos f20 family orbit data
f20Ref.mat	Markellos f20 family reference data
f20pData.mat	Markellos f20' family orbit data
f20pRef.mat	Markellos f20' family reference data
f26Data.mat	Markellos f26 family orbit data
f26Ref.mat	Markellos f26 family reference data
f26pData.mat	Markellos f26' family orbit data
f26pRef.mat	Markellos f26' family reference data

LowEnergyManeuverData

f16Data.mat	Markellos f16 family orbit data
f16Ref.mat	Markellos f16 family reference data
f16pData.mat	Markellos f16' family orbit data
f16pRef.mat	Markellos f16' family reference data
f17Data.mat	Markellos f17 family orbit data
f17Ref.mat	Markellos f17 family reference data
f17pData.mat	Markellos f17' family orbit data
f17pRef.mat	Markellos f17' family reference data
f18Data.mat	Markellos f18 family orbit data
f18Ref.mat	Markellos f18 family reference data
f18pData.mat	Markellos f18' family orbit data
f18pRef.mat	Markellos f18' family reference data
f19Data.mat	Markellos f19 family orbit data
f19Ref.mat	Markellos f19 family reference data
f19pData.mat	Markellos f19' family orbit data
f19pRef.mat	Markellos f19' family reference data
f20Data.mat	Markellos f20 family orbit data
f20Ref.mat	Markellos f20 family reference data
f20pData.mat	Markellos f20' family orbit data
f20pRef.mat	Markellos f20' family reference data
f26Data.mat	Markellos f26 family orbit data
f26Ref.mat	Markellos f26 family reference data
f26pData.mat	Markellos f26' family orbit data
f26pRef.mat	Markellos f26' family reference data

Optimization

CostLowThrust2D	Cost function for 2D low thrust trajectory optimization.
CostLowThrust3D	Cost function for 3D low thrust trajectory optimization.
FLowThrust2D	Derivatives for the low-thrust planar orbit problem.
FLowThrust3D	This function is 3D orbit problem in equinoctial coordinates.
Plot2DTrajectory	Propagate and plot the planar trajectory from costates.
Plot3DTrajectory	Propagate and plot a 3D trajectory using costate outputs from optimization.
PlotZermelo	Propagate and plot the Zermelo trajectory.
RHSZermelo	Zermelo's differential equation right hand side.
TrajectoryOptimization	Performs indirect trajectory optimization.
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.

OrbitControl

ApogeeRaisingDeltaV	Compute two delta-v's to raise the apogee and circularize the orbit.
DVTarget	Perform targeting using the Lambert algorithm.
DiscreteTrajectoryToDeltaV	Compute accelerations for a point to point trajectory.
DoDeltaV	Updates orbital elements based on an impulsive delta-V
FLambert	Right hand side for locally optimizing the Lambert time of flight problem.
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	-------------------------------------------------------------------------
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	[a,th,i,q1,q2,W] -> [a,i,W,w,e,M]
BarycenterEarthMoon	Computes the ECI position of the Earth-Moon Barycenter.
CEcl2SER	Transformation matrix from ecliptic to Sun-Earth rotating frame.
CEq2C	Transformation matrix from equinoctial to cartesian coordinates.
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.
E2RAORP	Ratio of apogee radius to perigee radius given eccentricity for an ellipsew
ECEFToECI	Compute the transformation matrix that rotates ECEF to ECI coordinates.
ECI2OrbitPlane	This function takes a position vector and orbit normal vector in ECI
ECIToEarthMoonRot	Convert ECI to Earth Moon rotating coordinates about the barycenter.
ECIToRLP	Transform from ECI to Rotating Libration Point Coordinates
ECIToSunEarthRot	Convert ECI to Sun Earth rotating coordinates about the barycenter.
EMBToEMBND	Scale EMB system to be non-dimensional
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
Eq2El	Converts equinoctial elements to Keplerian elements.
Eq2RV	Converts equinoctial elements to r and v for an elliptic orbit.
FPA	Computes the flight path angle from eccentricity and true anomaly.
HelioToPlanet	Computes the transformation matrix from heliocentric to planet fixed frame.
InitialOrientation	Compute the initial quaternion and body rates for a circular orbit,
K2L	Converts eccentric longitude to true longitude
L2K	Converts eccentric longitude to true longitude
LatLonToRAz	Compute range and azimuth angle between two points on a sphere.
LocusOfVisibility	Determine the locus of visibility for a satellite above a planet.
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.
OrbElemDiff	Computes the differences between orbital element vectors.
OrbNEl	Conmputes the orbit normal from orbital elements.
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
Parameter	Computes the parameter for any orbit.
PtchCnc	Computes planet centered orbital elements.
RARP2A	Computes the semi major axis from apogee and perigee radii
RARP2E	Computes the eccentricity from apogee and perigee radii
RAzToLatLon	Transform (range,azimuth) to (latitude,longitude) coordinates.
REl	Computes the orbit radius and the parameter from Keplerian elements.
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.
ROrb	Computes the orbit radius.
RPQ	Computes the orbit position vector in the perifocal plane.
RPRA2AE	Converts perigee and apogee to semi-major axis and eccentricity
RPV2AE	Converts perigee and velocity at perigee to a and e
RToNu	Computes true anomaly from semi-major axis, eccentricity and radius.
RV2AE	Converts R and V to a, e, rP and rA.
RV2Eq	Converts R and V to Equinoctial elements
RV2RADec	Right ascension and declination with their rates from position and velocity vectors.
RelativeCoord	Computes relative coordinates between two orbits.
RelativeOrbitState	Find the relative coordinate state vector between x2 and x1.
SEMIToSEMR	Transform pos. and vel. from inertial heliocentric frame to
SEMRToSEMI	Transform from Sun-Earth/Moon rotating to inertial frame
SEMToSEMND	Scale position and velocity in SEM system (rotating or inertial) by
SLR	Computes the semi-latus rectum.
TEMEToPEF	Computes the matrix from TEME to the pseudo-earth fixed frame.
TransformECIToEMBR	Transform ECI pos. and vel. to Earth Moon barycenter rotating frame.
TransformECIToSEMR	Transform ECI pos. and vel. to Sun-Earth-Moon rotating (SEMR) frame.
TransformEMBRToECI	Transform pos. and vel. from Earth Moon barycenter rotating frame to ECI
TransformSEMRToECI	Transform Sun-Earth-Moon rotating (SEMR) state to ECI frame.
TransformSEMRToHelio	Transform Sun-Earth-Moon rotating (SEMR) state to the heliocentric frame.
TrueAnomalySpacing	Generates an array of true anomalies.
VPQ	Computes the orbit velocity in the perifocal plane.
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.
EarthMoonSystemConstants	Specifies Earth-Moon system constants
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.
SEML2OrbitData.mat	Reference HALO trajectory data
SunEarthMoonSystemConstants	Specifies Sun-Earth-Moon system constants.
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	Computes two impulses needed to 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.
DVLowThrust	Delta-V from a circular to noncoplanar circular orbit.
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.
LTPlaneChange	Low-thrust transfer between two circular, nonplanar orbits
LTSpiral	Low thrust delta v between two circular orbits.
LTVCon	Solve the Linear Terminal Velocity Constraint problem.
LowThrustEscape	Compute the delta-V required for a low thrust escape.
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.
OrbitLoweringManeuvers	Lower orbits starting from a hyperbolic orbit
PatchedConicPlanner	Helps design patched conic trajectories.
SemimajorAxisRevisit	Computes semi-major axis from revisit and swath angle.
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

AFLTSpiral	Low thrust delta v between two circular orbits.
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.
BallisticCoefficient	Compute the ballistic coefficient from a CAD model.
C3ToV	Convert c3 to orbital velocity
CylOrb	Right-hand-side of the cylindrical orbit equations about a mass point.
DVOrbitDrag	Compute the drag loss over an orbit arc in planetary orbit.
EclipticInjectionPlane	Computes the parking orbit plane for a heliocentric injection.
Energy	Compute the energy for any orbit
FBudget	Computes a fuel budget from a list of events.
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.
FlybyHyperbola	Compute the required orbital elements for a flyby
GVEDynamics	Compute continuous-time relative dynamics for Gauss' variational equations
GetFuelBudgetDataFromCAD	Compute the power balance in a spacecraft.
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.
J2Prop	Propagate an ECI state into the future with the J2 perturbation.
LTAN	Compute the local time of the ascending node.
LambertTOF	Solves the Lambert time of flight problem using Battin's method.
LoadNORAD	Load a set of historical NORAD elements and convert to Cartesian states.
LonMot	Computes the longitudinal motion due to an acceleration a
LunarPolarOrbit	Computes a polar lunar orbit
LunarRelativeState	Position and velocity relative to the target
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.
NORAD	Propagates the NORAD two line elements (TLE), ex. SGP, SGP4, SGP8.
NORADToEl	Convert NORAD to Keplerian elements.
NonDimensionalOrbit	Dimensionalize or non-dimensionalize an orbit.
NuInf	Computes the asymptote angle for a hyperbola. If e is < 1 it substitutes 1.
OrbDecay	Approximate time for an orbit to decay based on an exponential atmosphere.
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.
PatchedConicElements	Computes patched conic elements.
PlanetaryAccelerations	Compute perturbations due to other planets aside from the center.
PropEph	Propagates the ephemeris one step.
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.
RVOrbGen	Generate an orbit by propagating Keplerian elements.
RVSet2El	Batch form of RV2El.
RandomOrbit	Monte-Carlo orbit simulation.
SphOI	Computes the sphere of influence ratio
SphOrb	Point mass orbit equations in equatorial spherical coordinates.
SphereOfInfluenceHelio	Computes the ratio of the sphere of influence for interplanetary
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.
Tfr2Imp	Two impulse optimal orbit transfer.
Tfr3Imp	Three impulse orbit transfer.
TimeOfFlightHyperbola	Time of flight in a hyperbola
UGravity	Compute the earth's gravitational potential.
VEscape	Computes the escape velocity.
VInf	Computes the asymptotic velocity for a hyperbola
VInfRPToRV	Generate RV from parameters
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.

OrbitPropagator

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

LowThrustSim	Simulate low thrust spacecraft.
PropagateLowThrust	Propagate a low thrust trajectory assuming that the thrust is constant
PropagateNBody	Propagate n-bodies in an n-body problem.
RVOrbHF	High fidelity orbit simulation with drag and planetary perturbations.

RHSOrbit

DLinOrb	Computes the linearized orbit equations for a discrete system.
FEarthMoon	Gravity acceleration in the earth moon system
FEarthMoonSun	Orbit dynamics with spherical harmonic models of the Earth and Moon gravity.
FHel	Orbit equations for InterstellarSim.
FOrb	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	Its companion function is FOrbLEOP
FOrbLTE	Orbit model for low thrust escapes.
FOrbLowThrust	Right hand side of the planar orbit equations assuming constant thrust
FSolarSail	Solar sail simulation right hand side.
FVarEq	Computes the right hand side of the variational equations.
LinOrb	Computes the linearized orbit equations.
RHS2DPolarOrbit	RHS for 2D polar orbit
RHSGeoHarm	Computes the right hand side for a gravity model.
RHSGeoJ2	Computes the right hand side for Earth gravity with J2.
RHSHelioWithPlanets	Right hand side for a heliocentric system with a thruster.
RHSOrbit2D	Two dimensional orbit dynamics.
RHSOrbitPlanetFixed	Cartesian orbital equations in a planet fixed (rotating) frame.

StraightLine

DVConstantThrust	Velocity change with constant thrust given fuel and time
MassFuelElectricConstantPower	Fuel mass for a constant accel/power electric propulsion system
MassFuelElectricConstantThrust	Fuel masss for a constant thrust electric propulsion system.
MassFuelElectricConstantUE	Fuel mass for a constant accel/exhaust velocity electric propulsion system.
RoundTripTime	Round trip time for a straight-line, constant thrust (Kammash)
SLPLDataStructure	Straight-line, power-limited data structure
SLPLDeltaV	Compute delta-V of an ideal power-limited rocket (straight-line)
SLPLFindDistance	Finds duration, ideal power-limited rocket, straight-line trajectory
SLPLFindDuration	Finds duration, ideal power-limited rocket, straight-line trajectory
SLPLFindMass	Finds final mass, ideal power-limited rocket, straight-line trajectory
SLPLFindPower	Finds required power, ideal power-limited rocket, straight-line traj.
SLPLReport	Print a summary of an SLPL problem
SLPLSolver	Solve a straight-line power-limited rocket problem.
SLPLTrajectory	Returns trajectory for an ideal power-limited rocket (straight-line)
SimulateStraightLineTrajectory	Simulate a straight line, constant-thrust trajectory
SpacecraftFromAccel	Size a spacecraft from acceleration data
StraightLineConstantAccel	Constant acceleration simulation
StraightLineConstantThrust	Compute straight line bang-bang trajectory with constant thrust
StraightLineDataStructure	Returns data structure for straight line functions
StraightLineOptimalFixedUE	Compute thrust for a minimum mass trajectory, with a fixed uE
SwitchTime	Calculate the switch time for a straight line trajectory
ThrustElectric	Computes power and thrust as a function of specific power.
TrajectoryBetweenTwoPlanets	fmincon solution for a trajectory between two planets
UEVsSigma	Compute exhaust velocity vs specific power
UToMF	Calculate masses for a spacecraft with an electric propulsion system.

ThreeBody

CRTBPJacobi	Compute the Jacobi constant for a state in the CRTBP system.
CRTBPLinearSystem	Compute linear system matrix for the circular restricted 3 body problem
CRTBPRHS	Right-hand-side for the circular restricted three body problem.
EarthEscape	Simulate an escape trajectory from Earth using CRTBPRHS.
FCRTBPRHS	Right-hand-side for the forced circular restricted three body problem.
HALOFamily	Compute and plot a family of stable HALO orbits in the given system
HALOInit	Compute a stable initial state for a HALO orbit
HALOState	Approximate an initial state for a HALO orbit given the orbit size
L2OrbitSim	Simulate the Earth orbit to L2 point low thrust transfer.
LEOToSunEarthL2Design	Design a low thrust transfer from Earth orbit to Sun/Earth-Moon L2 point
LowThrustCRTBP_StopFcn	Stoppping Function for the circular restricted three body problem.
LyapunovFamily	Generate a family of planar Lyapunov orbits
LyapunovInit	Compute a stable initial state for a Lyapunov orbit

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.
EarthMoon	Plot an orbit track in the Earth/Moon System.
EarthMoonRotFrame	Plot an orbit track in the Earth/Moon Rotating System.
HALOPlot	Generate a plot of a HALO trajectory
L2OrbitPlots	Simulate and plot an Earth orbit to L2 point low thrust transfer.
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.
PlanetTrajectoryPlot	Plots a spacecraft between two planets
PlotEMTraj	Plot a trajectory in the Earth-Moon system
PlotOrbitPage	Create a plot page with 2D and 3D views of an orbit.
PlotOrbitTransfer	Plot the two initial orbits and the transfer orbit.
PlotPlanetaryOrbit	Plot the state in 3D with an planet map.
PlotSEMTraj	Plot an orbit track in the Sun/Earth-Moon Rotating System.
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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