Name

Description

A

The Keplerian semimajor axis (a). Half the length of the long axis of the orbit ellipse. Can not be set to 0.

Acceleration

Returns the vector of the Spacecraft's instantaneous acceleration referenced to the Mean of J2000 EarthEquator coordinate frame. This vector provides the Cartesian components of the change in the Spacecraft's velocity with respect to time. The components of this vector are populated using the ForceModel associated with the Spacecraft's Propagator and will be zero if an analytic propagator is used.

Accelerations

The accelerations due to various forces on the spacecraft, referenced to the Mean of J2000 EarthEquator coordinate frame.

AHF

AHF for advancing the attitude of the Spacecraft.

AlphaRE

Returns the angle between the Spacecraft's +Y body axis and the projection of the spacecraftearth line on the spacecraft's YZ body plane.

AlphaRS

Returns the angle between the Spacecraft's +Y body axis and the projection of the sunspacecraft line on spacecraft's YZ body plane.

AlphaTE

Returns the angle between the spacecraftearth line and the Spacecraft's +X body axis.

AlphaTS

Returns the angle between the sunspacecraft line and the Spacecraft's +X body axis.

AngularAcceleration

The angular acceleration vector that is associated with the quaternion attitude system. When using a Spinner attitude system the angular acceleration is populated with values; which will be nonzero if you have a nonzero nutation angle, and the Spacecraft moment of inertia Ixx != Iyy != Izz.

AngularMomentumDeclination

The declination of the angular momentum vector with respect to the MJ2000 EarthEquatorial frame for a spinning Spacecraft. AngularMomentumRightAscension and AngularMomentumDeclination define the orientation of the "angular momentum frame" with respect to the MJ2000 EarthEquatorial frame. The zaxis of the angular momentum frame is coincident with the spinning Spacecraft's angular momentum vector.

AngularMomentumRightAscension

The right ascension of the angular momentum vector with respect to the MJ2000 EarthEquatorial frame for a spinning Spacecraft.

AngularVelocity

The angular velocity vector associated with the quaternion attitude system. The angular velocity will be applied with respect to the attitude frame defined by the Spacecraft.AttitudeRefFrame property.

Annotation

When displayed in a Mission View the specified text will be drawn at each tail point. The value may be modified between updates to change the text drawn at subsequent points.

Apoapsis

Returns an instantaneous osculating estimate of the apoapsis distance from the center of the CentralBody to the Spacecraft, based on the current orbital state. The Spacecraft.Apoapsis property is not based on a force model and is computed directly from the instantaneous Keplerian elements where the instantaneous Keplerian elements are computed from the active state of the Spacecraft.

Apogee

Returns an instantaneous osculating estimate of the apoapsis distance from the center of its central body to the Spacecraft, based on the current orbital state. The Spacecraft.Apogee property is not based on a force model and is computed directly from the instantaneous Keplerian elements where the instantaneous Keplerian elements are computed from the active state of the Spacecraft.

ApogeeHeight

Returns the instantaneous osculating apoapsis height of the Spacecraft above its central body, measured along a vector normal to the surface of the body at the Spacecraft's subsatellite point. The oblateness of the central body is taken into account when calculating the height.

ArgumentOfLatitude

Returns the angle between the vector pointing to the ascending node and the vector pointing to the Spacecraft's current position. This angle varies from 0 to 360 degrees.

AscendingNode

This method can be used with the Step command to advance the Spacecraft to the point in its orbit at which the Spacecraft crosses its central body's instantaneous equatorial plane, moving from the south to the north. It serves as the conditional statement for the "to" statement that can be used with the Step command.

AttitudeCoordinateSystem

CoordinateSystem that serves as the attitude reference frame for the Spacecraft object.

AttitudeMatrix

The DirectionCosineMatrix (DCM) that defines the rotation from the specified Attitude Reference Frame to the Spacecraft Body Coordinate System.

AttitudeRefFrame

This is the Spacecraft's attitude reference frame. It is necessary to define the AttitudeRefFrame property of the spacecraft prior to defining the attitude system properties.

AttitudeSystem

Specifies which attitude system set defines the Spacecraft state.

BetaAngle

Returns the elevation angle measured between the orbital plane and the EarthSun vector.

BL_A

The mean semimajor axis in the BrouwerLyddane Mean Element set. Half the length of the long axis of the orbit ellipse.

BL_E

The mean eccentricity in the BrouwerLyddane Mean Element set. The distance from the center of the ellipse to one of the foci divided by the semimajor axis.

BL_Elements

Represents the instantaneous Spacecraft orbital state as BrouwerLyddane Mean Keplerian elements. The Spacecraft state is reported as the following mean element set semimajor axis, eccentricity, inclination, right ascension of ascending node, argument of perigee, and mean anomaly.

BL_I

The mean inclination in the BrouwerLyddane Mean Element set. The angle between the orbit plane and the fundamental plane of the Mean of J2000 EarthEquator coordinate frame.

BL_MA

Mean Anomaly of the Spacecraft in Mean of J2000 EarthEquator coordinates in the BrouwerLyddane Mean Element set. The Mean Anomaly is defined as the angle from the Eccentricity Vector to a position vector corresponding to the location of the satellite if it always moved at a constant angular rate.

BL_RAAN

The mean right ascension of the ascending node in the BrouwerLyddane Mean Element set. The angle, measured at the center of the Earth, from the vernal equinox to the ascending node. The ascending node is the point where the satellite crosses the fundamental plane of the Mean of J2000 EarthEquator coordinate frame going from south to north.

BL_W

The mean argument of perigee in the BrouwerLyddane Mean Element set. The angle, measured at the center of the Earth, from the ascending node. The point in the satellites orbit closest to the Earth.

BLJ2A

Mean SemiMajor Axis in the J2 BrouwerLyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion.

BLJ2E

Mean Eccentricity, in the J2 BrouwerLyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion. The Eccentricity is a dimensionless value that defines the shape of the Spacecraft's orbit.

BLJ2Elements

Represents the instantaneous Spacecraft orbital state as BrouwerLyddane Mean Keplerian elements, where ONLY the J2 zonal harmonic coefficient is used in the conversion. The Spacecraft state is reported as the following mean element set semimajor axis, eccentricity, inclination, right ascension of ascending node, argument of perigee, and mean anomaly.

BLJ2I

Mean Inclination in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion. The Inclination is defined by the angle between the orbit plane and the fundamental plane of the Mean of J2000 EarthEquator coordinate frame.

BLJ2MA

Mean Anomaly in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion, and referenced to the Mean of J2000 EarthEquator coordinate frame. The Mean Anomaly is defined as the angle from the eccentricity vector to a position vector corresponding to the location of the satellite if it is always moved at a constant angular rate.

BLJ2RAAN

Mean Right Ascension of the Ascending Node in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion, and referenced to the Mean of J2000 EarthEquator coordinate frame.

BLJ2W

Mean Argument of Perigee in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the BrouwerLyddane conversion, and referenced to the Mean of J2000 EarthEquator coordinate frame. The Argument of Perigee is defined by the angle measured at the center of the Earth in the orbit plane, from the ascending node to perigee.

BodyScale

A scale factor to increase or decrease the size of the Spacecraft model in a Mission View. The BodyScale property assumes the model dimensions are in units of meters, similar to the ThreeDModel.Scale property. If, for example, the ThreeDModel.Scale is set to 1, the Spacecraft.BodyScale will have to be set to 1 in order to make bounding boxes coincide with the model.

C3

Returns the characteristic energy of the calling Spacecraft relative to its CentralBody. The characteristic energy is a measure of the excess specific energy over that required to escape from the CentralBody. For a Spacecraft in a nonescape trajectory, the characteristic energy will be less than 0. For a Spacecraft in a parabolic trajectory, the characteristic energy will be equal to 0. A Spacecraft in a hyperbolic trajectory will have a characteristic energy that is equal to the Hyperbolic Excess Velocity squared.

Cd

The Coefficient of Drag is a property that is a function of the properties and shape of the spacecraft's area incident to the velocity direction. It is used in the calculation of the force on the spacecraft due to atmospheric Drag.

CenterOfMass

Returns the X, Y, and Z components of the center of mass of the dry Spacecraft and all attached Tanks. The Spacecraft's center of mass is updated during maneuvers as fuel is depleted from the tanks. The center of mass location is relative to the origin of the Spacecraft body fixed coordinate frame.

CentralBody

Name of the body that defines the origin of the Spacecraft's position and velocity. This can be a userdefined CelestialObject or one of the following predefined bodies:
Mercury
Venus
Earth
Mars
Jupiter
Saturn
Uranus
Neptune
Pluto
Moon
Sun
In order to translate the origin of a Spacecraft's state to a different central body, simply change the Spacecraft.CentralBody property.

CentralBodyObject

CelestialObject that defines the origin for the Spacecraft's position and velocity.

Cl

The Coefficient of Lift is a property that is a function of the properties and shape of the spacecraft's area incident to the velocity direction. It is used in the calculation of the force on the spacecraft due to aerodynamic Lift.

CoElevation

The initial right ascension (not coelevation) of a spinning Spacecraft's spin axis with respect to the angular momentum frame. NutationAngle and CoElevation define the orientation of the "spin frame" with respect to the "angular momentum frame".
The zaxis of the spin frame is coincident with the spinning Spacecraft's spin axis.

Color

The color of the Spacecraft tail when displayed in a Mission View. The color can be overridden for an individual view using the ViewWindow.SetObjectColor() method.

Cr

The Coefficient of Reflectivity is a property that is a function of the properties of the spacecraft's area incident to the SunEarth line. It is used in the calculation of the force on the spacecraft due to solar radiation pressure.
Note that this value is only used when ForceModel.SRPForceGeometry is set to spherical. If using a plate model, the coefficients of reflectivity should be set for each individual plate.

DayOfMonth

Returns the numerical value representing the calendar day of the month for the current spacecraft Epoch. The day is given as an integer number ranging from 1 to 31.

DayOfWeek

Returns the numerical value representing a day of the week for the current spacecraft Epoch. The day is given as an integer number ranging from 1 to 7, where 1 corresponds to Sunday.

DayOfYear

Returns the numerical value representing the current day of year counted from January 1 of the current year for the current spacecraft Epoch. The day is given as an integer number ranging from 1 to 365 (366 for leap years).

DEC

The declination of the satellite in the Spherical Element set. It is defined as the angle between the satellite position vector and the fundamental plane of the Mean of J2000 EarthEquator coordinate frame.

DeclaredName

The name of the object as declared.

Density

Represents the density of the atmosphere at the Spacecraft's altitude calculated using the defined atmospheric density model associated with the Spacecraft's ForceModel.

DescendingNode

This method can be used with the Step command to advance the Spacecraft to the point in its orbit at which the Spacecraft crosses its central body's instantaneous equatorial plane, moving from the north to the south. It serves as the conditional statement for the "to" statement that can be used with the Step command.

DisplayName

The name displayed for this object in output windows such as views, plots, and reports.

Drag

Represents the instantaneous force on the Spacecraft due to atmospheric drag modeling, calculated using the defined atmospheric density model associated with the Spacecraft's ForceModel. Returns zero if Drag is not enabled.

DragArea

The Drag Area is the incident cross sectional area to the velocity direction. It is used in the calculation of the force on the spacecraft due to atmospheric Drag.

E

The Keplerian eccentricity (e). A unitless property that describes the curvature of an ellipse. The distance from the center of the ellipse to one of the foci divided by the semimajor axis.

EA

Represents the instantaneous Eccentric Anomaly of the Spacecraft. The Eccentric Anomaly is an angle related to the position of the spacecraft and the ideal orbit associated with its orbit (ie the circular orbit which inscribes the true elliptical orbit)

EDR

Represents the instantaneous Energy Dissipation Rate of the Spacecraft due to atmospheric drag modeling, calculated using the defined atmospheric density model associated with the Spacecraft's ForceModel. Returns zero if Drag is not enabled.

ElapsedTime

Returns the time passed since the first time this property is read on a specific line of script.

ElectronDensity

Returns the number of Electrons per m^3 at the Spacecraft location. The Electron Density is calculated using the Ionosphere model type and properties specified in the Ionosphere Options within the Solar System.

ElementType

Determines which element set defines the Spacecraft state. The chosen element set will determine the properties that are written to an external file when using the Put Command. A FreeFlyer ephemeris always contains Cartesian elements despite the value of this property.

Energy

Represents the instantaneous specific energy (energy per unit mass) of the Spacecraft orbit. This is calculated based on the Keplerian gravitational potential model as Energy = mu/(2*a).

Epoch

The current Spacecraft date and time in International Atomic Time (TAI) in the Modified Julian Date format. The returned value has units of days, measured from January 5, 1941 at 12:00:00.000 TAI.
Changing the value of this property does not cause the Spacecraft's state to be propagated  the Spacecraft's state will not be changed. In order to propagate a Spacecraft to a desired epoch, use the Step command.

EpochTAI

Returns the current spacecraft date and time in International Atomic Time (TAI) in the Modified Julian Date format as measured from January 5, 1941 at 12:00:00.000 TAI.

EpochText

The current Spacecraft date and time in Coordinated Universal Time (UTC). The returned time is in the Gregorian time format of "Mmm DD YYYY HH:MM:SS.SSS", where Mmm stands for the first three (3) letters of the month; DD stands for the day of month; YYYY stands for the 4digit year; and HH:MM:SS.SSS refers to hours, minutes, and seconds to milliseconds.

EpochUSNO

Returns the current spacecraft date and time in International Atomic Time (TAI) in the Modified Julian Date format as measured by the United States Naval Observatory from November 17, 1858 00:00:00.000 TAI.

EpochUT1

Returns the current Spacecraft date and time in Universal Time (UT1) in the Modified Julian Date format.

EpochUTC

Returns the current Spacecraft date and time in Coordinated Universal Time (UTC) in the Modified Julian Date format.

EquinoctialA

The a component of the Equinoctial element set. Half the length of the long axis of the orbit ellipse.

EquinoctialH

The h component of the Equinoctial element set, defined by: h = e*sin(RAAN + w), where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

EquinoctialK

The k component of the Equinoctial element set, defined by: k = e*cos(RAAN + w), where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

EquinoctialLongitude

Equinoctial Longitude defined by the mean longitude, defined as: RAAN + w + MA, where RAAN and w are the Right Ascension of the Ascending Node and the Argument of Perigee, respectively, and MA is the Keplerian mean anomaly. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

EquinoctialP

The p component of the Equinoctial element set, defined by: p = tan(i/2)sin(RAAN), where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

EquinoctialQ

The q component of the Equinoctial element set, defined by: q = tan(i/2)cos(RAAN) where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

EulerAngles

These are the angles associated with the Euler attitude system. These angles will be applied as a set of three rotations to the Spacecraft's attitude about the axes set in the Spacecraft.EulerSequence property.

EulerAnglesLVLH

Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the LVLH frame of the spacecraft. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence.

EulerAnglesMJ2000

Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the Mean of J2000 EarthEquator coordinate frame. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence.

EulerAnglesRIC

Returns the three Euler Angles that define the orientation of the Spacecraft BodyCoordinateSystem with respect to the RIC frame of the spacecraft. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence.

EulerAnglesUVW

Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the UVW frame of the spacecraft. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence.

EulerRates

Rate of change of the Euler angles.

EulerSequence

Axes of rotation that the EulerAngles are around.

EVector

Represents the instantaneous eccentricity vector referenced to the Mean of J2000 EarthEquator coordinate system. The eccentricity vector points from the center of the central body toward perigee with a magnitude equal to the eccentricity of the orbit.

FixedPosition

Reports the object's position components (X, Y, Z) referenced to the Body Fixed coordinate system of its central body.

FixedVelocity

Reports the Spacecraft's velocity components (VX, VY, VZ) referenced to the Body Fixed coordinate system of its central body.

ForceCoefficientType

Flag indicating whether Cd and Cr are used in the physical model, or B* and AGOM from the SGP4 model is used.

FPAngle

Returns the flight path angle, which is measured between the perpendicular of the instantaneous position and the instantaneous velocity vectors of the spacecraft.

FPAzimuth

Returns the azimuth angle of the flight path, which is measured between the local North vector and the velocity vector.

GHA

Represents the Greenwich Hour Angle at the Spacecraft's current epoch. The Greenwich Hour Angle is the angle measured in the fundamental plane between the xaxis of the Mean of J2000 EarthEquator coordinate frame at the spacecraft's current epoch and the meridian passing through the Greenwich observatory in England.

GNSSReceivers

The spacecraft's Globlal Navigation Satellite System (GNSS) receivers.

Height

Returns the height of the Spacecraft above its central body, measured along a vector normal to the surface of the body at the Spacecraft's subsatellite point. The oblateness of the central body is taken into account when calculating the height. If Earth is set as the central body using the default radius and flattening coefficient, regardless of gravitational potential model, FreeFlyer will use the WGS 84 reference ellipsoid.

HistoryMode

Determines if the visualized Spacecraft tail is unlimited in length, or whether new points replace the oldest points once the tail exceeds the Spacecraft tail length. The history mode can be overridden for an individual view using ViewWindow.SetObjectHistoryMode() method.

HorizontalFPA

The horizontal flight path angle in the Spherical Lat/Long Element Set. It is measured from the velocity vector to the plane perpendicular to the position vector.

HyperbolicExcessVelocity

Returns the hyperbolic excess velocity of the calling Spacecraft at its current semimajor axis. This property is defined for hyperbolic orbits, and will return undefined for trajectories which have a semimajor axis greater than 0.

I

The Keplerian inclination (i). The angle between the orbit plane and the fundamental plane of the Mean of J2000 EarthEquator coordinate frame.

IsUsingAHF

Boolean value indicating whether or not the attitude motion model is based on an AHF.

KSCEarthSunAngles

Returns a fiveelement Array containing the KSC EarthSun angles. These angles can be accessed individually using the AlphaTS, AlphaRS, AlphaTE, AlphaRE, and LVBeta properties.

Latitude

The Latitude property is the bodydetic latitude of the Spacecraft subsatellite point in the Spherical Lat/Long Element Set. It is defined as the angle between the fundamental plane of the BodyFixed coordinate system of the central body and the vector normal to the central body's surface passing through the satellite's position. This property returns a scalar value.

LatLongAzimuth

The velocity azimuth angle in the Spherical Lat/Long Element set. It is defined as the angle, measured eastward in the plane perpendicular to the position vector, from the projection of the central body's inertial coordinate frame's zaxis onto this plane to the projection of the velocity vector onto this plane.

LatLongRadius

The magnitude of the position vector in the Spherical Lat/Long Element set.

LatLongVi

The magnitude of the velocity vector in the inertial frame in the Spherical Lat/Long Element set.

LCSEpoch

The reference epoch associated with the Launch Coordinate reference frame. This is usually set at the time of liftoff. The format can be either in days or a calendar format, depending on the LCSEpochCoordSys.

LCSLatitude

The latitude of the launch site for the Launch Coordinate System. The LCSLatitude is for reference only and has no effect on conversion to the Launch Coordinate System.

LCSLongitude

The LCSLongitude is the associated reference longitude and is usually set at the longitude of the launch site. The Launch Coordinate System was designed to facilitate analysis of satellite launches. Launch coordinates have advantages such that the right ascension of the ascending node may be specified independently from the launch time and date. Launch Coordinate System Definition is as follows:
Origin  center of the Earth.
Zaxis  northpointing vector normal to the true equatorial plane at the reference epoch.
Xaxis  lies in the true equatorial plane at the reference epoch, and points from the center of the Earth to the reference longitude on the Earth's surface.
Yaxis  forms a righthanded coordinate system.

Lift

Represents the instantaneous force on the Spacecraft due to atmospheric lift modeling.

LiftArea

The Lift Area is the incident cross sectional area to the velocity direction. It is used in the calculation of the force on the spacecraft due to atmospheric Lift.

LongAscendingNode

Returns the geographic East longitude of the subsatellite point at the ascending node, when the Spacecraft crosses the equator from south to north. This method will return a null data value (999) until the first nodal crossing.

LongDescendingNode

Returns the geographic East longitude of the subsatellite point at the descending node, when the Spacecraft crosses the equator from north to south. This method will return a null data value (999) until the first nodal crossing.

Longitude

The Longitude property is the longitude of the Spacecraft subsatellite point in the Spherical Lat/Long Element set. It is defined as the angle between the prime meridian and the projection of the position vector onto the fundamental plane of the BodyFixed coordinate system of the central body. It is measured as increasing in a counterclockwise sense when viewed from the North Pole, thus producing the longitude east from the prime meridian. This property returns a scalar value.

LongitudeOfPeriapsis

Returns the angle measured between the vernal equinox and periapsis.

LongitudeRate

Calculates the rate of change in longitude of the orbital plane in a BodyFixed frame of reference. Computes the difference between the Spacecraft orbital period and the CentralBody's rotational period, and then maps the result to degrees of longitudinal drift per second. Typically this property is used to examine the rate of drift within a GEO stationkeeping box. Note that this property is only supported for builtin CelestialObjects.

LunarBetaAngle

Returns the declination of the EarthMoon vector with respect to the EarthSun ecliptic plane.

LVBeta

Returns the angle measured between the orbit angular momentum vector and the sunspacecraft line.

MA

Represents the instantaneous Mean Anomaly of the Spacecraft.

Mass

Returns the total mass of the Spacecraft, including all attached Tanks.

MassFlowRate

The rate at which the mass of the Spacecraft is currently being depleted. This value will be nonzero only when the Spacecraft is maneuvering.

MassTotal

The total mass of the Spacecraft including all attached Tanks.
If you set this property, VehicleDryMass is set to (MassTotal  total hardware mass).

MeanMotion

Returns the mean motion, which is a measure of how fast a spacecraft progresses through its elliptical orbit.

MES

Returns the Moon  Earth  Sun angle. The angle has the vertex as the center of the Earth with rays to the center of the Moon and the center of the Sun.

MEV

Returns the Moon  Earth  Vehicle angle. The angle has the center of the Earth as the vertex with rays to the center of the Moon and Spacecraft position.

MGL

Returns the sum of the Mean Anomaly and the longitude of periapsis minus the current GHA.

Mlt

Returns the Mean Local Time of the Spacecraft's subsatellite point, which is the current UTC time of the Spacecraft's epoch adjusted by accounting for the Spacecraft's longitude.

MltAscendingNode

Returns the Mean Local time (MLT) of the Spacecraft's subsatellite point at each Ascending Node crossing. This method will return a null data value (999) until the Spacecraft crosses an ascending node. See the help file for a description on how FreeFlyer calculates the MLTAN.

MltDescendingNode

Returns the Mean Local time (MLT) of the Spacecraft's subsatellite point at each Descending Node crossing. This method will return a null data value (999) until the Spacecraft crosses a descending node. See the help file for a description on how FreeFlyer calculates the MLTDN.

ModifiedEquinoctialF

The f component of the Modified Equinoctial element set, defined by: f = e*cos(w + RAAN) where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

ModifiedEquinoctialG

The g component of the Modified Equinoctial element set, defined by: g = e*sin(w + RAAN) where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

ModifiedEquinoctialH

The h component of the Modified Equinoctial element set, defined by: h = tan(i/2)*cos(RAAN) where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

ModifiedEquinoctialK

The k component of the Modified Equinoctial element set, defined by: k = tan(I/2)*sin(RAAN) where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

ModifiedEquinoctialL

The True Longitude, L, is defined by: L = RAAN + w + TA, where RAAN and w are the Right Ascension of the Ascending Node and the Argument of Perigee, respectively, and TA is the Keplerian true anomaly. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

ModifiedEquinoctialP

The semiparameter p is defined by: p = a*(1e^2) where a and e are the Keplerian SemiMajor Axis and Eccentricity.

MOI

Returns the components of the moment of inertia matrix of the dry Spacecraft and all attached Tanks. The Spacecraft's moment of inertia is updated during maneuvers as fuel is depleted from the tanks. The MOI is also used as an input with the Spinner attitude system.

Momentum

Represents the instantaneous specific orbital angular momentum (angular momentum per unit mass) vector referenced in the Mean of J2000 EarthEquator coordinate system.

MonthOfYear

Returns a numerical value (integer) representing the current month of the year. The month is given as an integer number ranging from 1 to 12, where 1 corresponds to January.

MoonBAngle

Returns the angle, in the moon Bplane, of the piercing point measured from the T vector.

MoonBDotR

Returns the projection of the piercing point in the moon Bplane onto the R vector.

MoonBDotT

Returns the projection of the piercing point in the moon Bplane onto the T vector.

MoonBMagnitude

Returns the distance of the piercing point from the origin of the moon Bplane. In terms of the Cartesian coordinates.
MoonBMagnitude = SQRT(MoonBdotT^2 + MoonBdotR^2)

MoonVelocity

Returns a three dimensional vector representing the velocity of the moon in the Mean of J2000 EarthEquator coordinate system.

MVE

Returns the Moon  Vehicle  Earth angle. The angle has the vertex as the Spacecraft position with rays to the center of the Moon and to the center of the Earth.

NodeRate

Returns an estimate of the rate of change of the ascending node with respect to time. This estimate is based on the Gauss VOP equation, where:
J2= geopotential coefficient for the oblateness of the Earth
n = mean motion of the Spacecraft
Re= radius of the Earth
a = Spacecraft's osculating semimajor axis
e = Spacecraft's osculating eccentricity
i = Spacecraft's osculating inclination

NonSingularA

The semimajor axis (a) property for the NonSingular Keplerian element set. Half the length of the long axis of the orbit ellipse.

NonSingularE1

The e1 component of the Nonsingular Keplerian element set, defined by e1 = e*cos(RAAN + w), where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

NonSingularE2

The e2 component of the Nonsingular Keplerian element set, defined by e2 = e*sin(RAAN + w), where e, w, and RAAN are the Keplerian Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

NonSingularE3

The e3 component of the NonSingular Keplerian element set, defined by: e3 = sin(i/2) sin(RAAN), where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

NonSingularE4

The e4 component of the NonSingular Keplerian element set, defined by: e4 = sin(i/2) cos(RAAN), where i is the orbit inclination and RAAN is the Keplerian Right Ascension of the Ascending Node. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

NonSingularE5

The e5 component of the NonSingular Keplerian element set, defined by: e5 = RAAN + w + MA, where RAAN is the Right Ascension of the Ascending Node, w is the Right Ascension of the Ascending Node, and MA is the Mean Anomaly. These orbit properties are with respect to the Mean of J2000 EarthEquator coordinate frame.

NutationAngle

The nutation angle of a spinning Spacecraft is the coelevation of a spinning Spacecraft's spin axis with respect the zaxis of the angular momentum frame.

ObjectId

The unique identifier for the object.

ObjectType

The type of the object.

OD

This property contains the OD settings for all the Spacecraft properties that can be estimated.

OrbitNumberBounded

Returns the number of times a Spacecraft has orbited the central body (revolutions). It is incremented each time the Spacecraft crosses the ascending node, and decremented similarly for backwards propagation. This property resets to orbit 1 the orbit after the orbit number reaches Spacecraft.OrbitNumberResetNumber. Note, you must enable orbit number calculations before propagating this Spacecraft by setting Spacecraft.OrbitNumberCalculationEnabled = 1.

OrbitNumberCalculationsEnabled

This property enables calculations that are required in order to compute the Spacecraft's orbit number. Enable this property at the beginning of your propagation if you plan to report the Spacecraft's orbit number in this Mission Plan. If this property is disabled, an error will be reported if any of the Spacecraft's orbit number properties are accessed. Orbit number calculations incur an overhead for every propagation step, and therefore should be disabled unless needed.

OrbitNumberCumulative

Returns the number of times a Spacecraft has orbited the central body (revolutions). It is incremented each time the Spacecraft crosses the ascending node, and decremented similarly for backwards propagation. Note, you must enable orbit number calculations before propagating this Spacecraft by setting Spacecraft.OrbitNumberCalculationEnabled = 1.

OrbitNumberDaily

Returns the number of times a Spacecraft has orbited the central body (revolutions) and resets to orbit 1 for the first ascending node that crosses Spacecraft.OrbitNumberResetLongitude degrees East longitude. It is incremented each time the spacecraft crosses the ascending node, and decremented similarly for backwards propagation. Note, you must enable orbit number calculations before propagating this Spacecraft by setting Spacecraft.OrbitNumberCalculationEnabled = 1.

OrbitNumberResetLongitude

The Spacecraft.OrbitNumberDaily property will reset to 1 if the Spacecraft's longitude of the ascending node crosses this boundary in either direction (i.e. from East to West or from West to East). If the OrbitNumberResetLongitude is changed, the RefreshDailyOrbitNumber() method must be called before the OrbitNumberDaily property will update. Note, you must enable orbit number calculations before propagating this Spacecraft by setting Spacecraft.OrbitNumberCalculationEnabled = 1.

OrbitNumberResetNumber

The orbit number value after which Spacecraft.OrbitNumberBounded will reset back to 1. If set to 0, the bounded orbit number will never reset back to 1. If set to something other than 0 and the bounded orbit number is outside of bounds, the OrbitNumberBounded property will be reset back 1. Note, you must enable orbit number calculations before propagating this Spacecraft by setting Spacecraft.OrbitNumberCalculationEnabled = 1.

P

Represents the osculating value of the semilatus rectum of the Spacecraft's orbit. The value of the semilatus rectum is equal to the square of the specific angular momentum of the orbit (angular momentum per unit mass) divided by the central body's gravity constant m.

PercentEarthShadow

Returns a numerical value representing the percentageofshadow the Spacecraft encounters in its orbit with respect to Earth's shadow. The calculation is based on percentage of shadow the Spacecraft encounters during penetration into the penumbral region. Once into the umbral region the percentage of shadow is 100%.

PercentMoonShadow

Returns a numerical value representing the percentageofshadow the Spacecraft encounters in its orbit with respect to the Moon's shadow. The calculation is based on percentage of shadow the Spacecraft encounters during penetration into the penumbral region. Once into the umbral region the percentage of shadow is 100%.

Periapsis

Returns an instantaneous osculating estimate of the periapsis distance from the center of the CentralBody to the Spacecraft, based on the current orbital state. The Spacecraft.Periapsis property is not based on a force model and is computed directly from the instantaneous Keplerian elements where the instantaneous Keplerian elements are computed from the active state of the Spacecraft.

Perigee

Returns an instantaneous osculating estimate of the periapsis distance from the center of its central body to the Spacecraft, based on the current orbital state. The Spacecraft.Perigee property is not based on a force model and is computed directly from the instantaneous Keplerian elements where the instantaneous Keplerian elements are computed from the active state of the Spacecraft.

PerigeeHeight

Returns the instantaneous osculating periapsis height of the Spacecraft above its central body, measured along a vector normal to the surface of the body at the Spacecraft's subsatellite point. The oblateness of the central body is taken into account when calculating the height.

Period

Returns the osculating period, which corresponds to the instantaneous unperturbed twobody Keplerian orbit.

Pitch

Returns the angle about the yaxis of the 312 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.

PlateModel

The child object which contains all of the properties and methods associated with the Spacecraft's PlateModel. The PlateModel can be used to model higherfidelity solar radiation pressure (SRP).

Position

Returns the Spacecraft's X, Y, Z position referenced to the Mean of J2000 EarthEquator coordinate frame.

Propagator

The propagator for updating the orbital state of the Spacecraft.

PropagatorType

The object type of the propagator that is associated with the Spacecraft.

ProximityZones

The Spacecraft's proximity zones.

Quaternion

Returns the four components of the attitude quaternion of a Spacecraft. The first three components represent the vector part of the quaternion, the last component represents the scalar part. FreeFlyer uses the righthanded convention when applying the quaternion rotation.

RA

The right ascension of the satellite in the Spherical Element Set. This is the angle from the xaxis of the Mean of J2000 EarthEquator coordinate frame to the projection of the satellite position vector onto the fundamental plane of the coordinate frame.

RAAN

The Keplerian right ascension of the ascending node (RAAN). The angle, measured at the center of the central body, from the vernal equinox to the ascending node. The ascending node is the point where the satellite crosses the fundamental plane of the Mean of J2000 EarthEquator coordinate frame going from south to north.

RADescendingNode

Represents the osculating Right Ascension of the Descending Node (RADN) in the Mean of J2000 EarthEquator coordinate system. The RADN is defined as the angle, measured at the center of the central body, from the vernal equinox to the Descending Node.

Radius

The magnitude of the Cartesian position vector of the Spacecraft with respect to the central body.

Roll

Returns the angle about the xaxis of the 312 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.

SatelliteId

Used by orbit determination algorithms to identify the Spacecraft.

Sensors

The Spacecraft's sensors.

SEV

Returns the Sun  Earth  Vehicle angle. The angle has the vertex as the center of the Earth with rays to the Sun center and Spacecraft position.

SGP4

Object holding options for managing the SGP4 propagation model and orbital element set.

SolarFlux

Returns the instantaneous value of the Solar Flux at the spacecraft, measured in W/m^2. The SolarFlux property is not associated with the environmental information of any drag model, and is not the same as the F10.7 cm solar flux that is used in some atmospheric density models.

SphericalAzimuth

The velocity azimuth angle in the Spherical Element Set. The angle, measured eastward in the plane perpendicular to the position vector, from the projection of the Mean of J2000 EarthEquator coordinate frame's zaxis onto this plane to the projection of the velocity vector onto this plane.

SphericalRadius

The magnitude of the position vector in the Spherical Element set.

SpinPhase

Phase angle of the xaxis of the Body Coordinate System with respect to the projection of the xaxis of the MJ2000 EarthEquatorial frame into the spin plane. If SpinPhase = 0, the initial xaxis of the Body Coordinate System will be coincident with the projection of the xaxis of the MJ2000 EarthEquatorial frame into the spin plane.

SpinRate

The angular rotation rate of a spinning Spacecraft about its spin axis, i.e, the zaxis of the "spin frame".

SRPArea

The SRP Area is the incident cross sectional area to the SunEarth line. It is used in the calculation of the force on the spacecraft due to solar radiation pressure.
If the Spacecraft's ForceModel.SRPForceGeometry is set to use a PlateModel, then this property will reflect the incident SRP area of the PlateModel, and an error will be reported if you try to assign a value to this property.

SunVelocity

Returns a threedimensional vector representing the velocity of the Sun in the Mean of J2000 EarthEquator coordinate system.

SVE

Returns the Sun  Vehicle  Earth angle. The angle has the vertex as the Spacecraft position with rays to the center of the Sun and to the center of the Earth.

SVM

Returns the Sun  Vehicle  Moon angle. The angle has the vertex as the Spacecraft position with rays to the center of the Sun and to the center of the Moon.

TA

The Keplerian true anomaly. The angle measured at the center of the central body, from perigee to the current position of the satellite.

TailLength

The number of historical position points to save and draw in visualizations generated using the View command. A new point will be added each time the View command is called. If you use a ViewWindow object, the tail length set there will override this tail length.
In millisecond timing precision mode only, the deprecated "as Global" syntax can be used to update the View when any Vehicle is Stepped or Maneuvered.
Use caution in setting tail lengths longer than 10,000  30,000 points. Tail lengths exceeding this range can cause memory problems due to the large amount of relative position data that must be maintained. The maximum tail length that can be used safely depends on the amount of memory in the system and the number of objects managed by the Mission View.

Tanks

The spacecraft's fuel tanks.

ThreeDModelFile

The file name of the file containing the 3D model that will represent the Spacecraft in a Mission View.

ThreeDModelObject

The name of the ThreeDModel object that holds information about the physical geometry of the Spacecraft.

Thrusters

The spacecraft's thrusters.

TickType

The tick mark shape to draw at each point of the tail in a Mission View.

TimeOfDay

Returns the current Spacecraft time referenced to Greenwich Mean Time (GMT or UTC).

TotalElectronContent

Returns the Total Vertical Electron Content for a 1 m^2 column that extends radially through the entire Ionosphere. The location of the column is based on the Geodetic Latitude and Longitude of the Spacecraft. The Total Electron Content is calculated using the Ionosphere model type and properties specified in the Ionosphere Options within the Solar System.

Transponders

The spacecraft's transponders.

TrueLongitude

Returns the value of the angle measured from the vernal equinox to the spacecraft's current position with the center of the central body as its vertex.

VehicleDryCenterOfGravity

The vector representing the dry vehicle center of mass (i.e. no attached Tanks are included in this property) referenced from the origin of the Spacecraft's body fixed frame.. This property is used in calculating the CenterOfMass, which includes all attached Tanks in its calculation.

VehicleDryMass

The Spacecraft's dry vehicle mass. The fuel mass is input through the tank and updated as mass is depleted.

VehicleDryMOI

The dry vehicle moment of inertia matrix (i.e. no attached Tanks are included in this property). This property is used in calculating the Spacecraft MomentOfInertia, which includes all attached Tanks in its calculation. The MOI is referenced from the origin of the Spacecraft BCS.

Velocity

Returns the instantaneous Spacecraft velocity vector in Cartesian components (VX, VY, VZ) referenced to the Mean of J2000 EarthEquator coordinate frame.

VerticalFPA

The vertical flight path angle in the Spherical Element set. The angle between the position and velocity vectors of the satellite.

Vi

The magnitude of the velocity vector in the inertial frame in the Spherical Element set.

VMag

Represents the instantaneous magnitude of the Spacecraft's velocity.

VX

The Cartesian Xcomponent of the Spacecraft's velocity vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.

VY

The Cartesian Ycomponent of the Spacecraft's velocity vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.

VZ

The Cartesian Zcomponent of the Spacecraft's velocity vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.

W

The Keplerian argument of perigee (w). The angle measured at the center of the central body, from the ascending node to the point in the satellite's orbit closest to the central body.

WeekOfMonth

Returns a numerical value representing the current week of the month. The week is given as an integer number ranging from 1 to 5, where 1 is equivalent to the first week of the month.

X

The Cartesian Xcomponent of the Spacecraft's position vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.

Y

The Cartesian Ycomponent of the Spacecraft's position vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.

Yaw

Returns the angle about the zaxis of the 312 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.

Year

Returns the current fourdigit year.

Z

The Cartesian Zcomponent of the Spacecraft's position vector, referenced to the Mean of J2000 EarthEquator Coordinate frame.
