Spacecraft Object

Description

The Spacecraft object is the heart of FreeFlyer. You can define the orbit and attitude states of a Spacecraft in a variety of element sets and reference frames, define physical attributes such as mass and drag area, add various subsystems, and control the motion model. See The Spacecraft Object Guide for more information. You can assign Spacecraft objects using the = operator. This will equalize the Spacecrafts' properties, but not the properties of any attached objects (such as Propagators, ForceModels, Tanks, Sensors, and Thrusters).

Inheritance Hierarchy: Object->Vehicle->Spacecraft

Available In Editions:

Engineer

Mission

Timing Precision Mode

This page describes functionality in nanosecond timing precision mode.

Click here to see the documentation for this object in millisecond timing precision mode.

Name	Description	Attributes
A	The Keplerian semi-major axis (a). Half the length of the long axis of the orbit ellipse. Can not be set to 0.	Type: Variable Access: Read/Write Units: km
Acceleration	Returns the vector of the Spacecraft's instantaneous acceleration referenced to the ICRF. 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.	Type: Array Access: Read-Only Units: km/s^2
Accelerations	The accelerations due to various forces on the spacecraft, referenced to the ICRF.	Type: Array of AccelerationTerm Access: Read-Only
AHF	AHF for advancing the attitude of the Spacecraft.	Type: Ephemeris Access: Read-Only
AlphaRE	Returns the angle between the Spacecraft's +Y body axis and the projection of the spacecraft-earth line on the spacecraft's Y-Z body plane.	Type: Variable Access: Read-Only Units: deg
AlphaRS	Returns the angle between the Spacecraft's +Y body axis and the projection of the sun-spacecraft line on spacecraft's Y-Z body plane.	Type: Variable Access: Read-Only Units: deg
AlphaTE	Returns the angle between the spacecraft-earth line and the Spacecraft's +X body axis.	Type: Variable Access: Read-Only Units: deg
AlphaTS	Returns the angle between the sun-spacecraft line and the Spacecraft's +X body axis.	Type: Variable Access: Read-Only Units: deg
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 non-zero if you have a non-zero nutation angle, and the Spacecraft moment of inertia Ixx != Iyy != Izz.	Type: Array Access: Read-Only Units: deg/s^2
AngularMomentumDeclination	The declination of the angular momentum vector with respect to the ICRF for a spinning Spacecraft. AngularMomentumRightAscension and AngularMomentumDeclination define the orientation of the "angular momentum frame" with respect to the ICRF. The z-axis of the angular momentum frame is coincident with the spinning Spacecraft's angular momentum vector.	Type: Variable Access: Read/Write Units: deg
AngularMomentumRightAscension	The right ascension of the angular momentum vector with respect to the ICRF for a spinning Spacecraft.	Type: Variable Access: Read/Write Units: deg
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.	Type: Array Access: Read/Write Units: deg/s
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.	Type: String Access: Read/Write
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.	Type: Variable Access: Read-Only Units: km
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.	Type: Variable Access: Read-Only Units: km
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 sub-satellite point. The oblateness of the central body is taken into account when calculating the height.	Type: Variable Access: Read-Only Units: km
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.	Type: Variable Access: Read-Only Units: deg
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.	Type: Variable Access: Read-Only Units: km
AttitudeCoordinateSystem	CoordinateSystem that serves as the attitude reference frame for the Spacecraft object.	Type: CoordinateSystem Access: Read-Only
AttitudeMatrix	The Direction-Cosine-Matrix (DCM) that defines the rotation from the specified Attitude Reference Frame to the Spacecraft Body Coordinate System.	Type: Matrix Access: Read/Write
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.	Type: String Access: Read/Write
AttitudeSystem	Specifies which attitude system set defines the Spacecraft state.	Type: String Access: Read/Write
BetaAngle	Returns the elevation angle measured between the orbital plane and the CentralBody-Sun vector. Note that the orbital plane is defined by the orbital momentum vector evaluated in the ICRF.	Type: Variable Access: Read-Only Units: deg
BL_A	The mean semi-major axis in the Brouwer-Lyddane Mean Element set. Half the length of the long axis of the orbit ellipse.	Type: Variable Access: Read/Write Units: km
BL_E	The mean eccentricity in the Brouwer-Lyddane Mean Element set. The distance from the center of the ellipse to one of the foci divided by the semi-major axis.	Type: Variable Access: Read/Write
BL_Elements	Represents the instantaneous Spacecraft orbital state as Brouwer-Lyddane Mean Keplerian elements. The Spacecraft state is reported as the following mean element set semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee, and mean anomaly.	Type: Array Access: Read/Write
BL_I	The mean inclination in the Brouwer-Lyddane Mean Element set. The angle between the orbit plane and the fundamental plane of the ICRF.	Type: Variable Access: Read/Write Units: deg
BL_MA	Mean Anomaly of the Spacecraft referenced to the ICRF in the Brouwer-Lyddane 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.	Type: Variable Access: Read/Write Units: deg
BL_RAAN	The mean right ascension of the ascending node in the Brouwer-Lyddane 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 ICRF going from south to north.	Type: Variable Access: Read/Write Units: deg
BL_W	The mean argument of perigee in the Brouwer-Lyddane 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.	Type: Variable Access: Read/Write Units: deg
BLJ2A	Mean Semi-Major Axis in the J2 Brouwer-Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the Brouwer-Lyddane conversion.	Type: Variable Access: Read/Write Units: km
BLJ2E	Mean Eccentricity, in the J2 Brouwer-Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the Brouwer-Lyddane conversion. The Eccentricity is a dimensionless value that defines the shape of the Spacecraft's orbit.	Type: Variable Access: Read/Write
BLJ2Elements	Represents the instantaneous Spacecraft orbital state as Brouwer-Lyddane 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 semi-major axis, eccentricity, inclination, right ascension of ascending node, argument of perigee, and mean anomaly.	Type: Array Access: Read/Write
BLJ2I	Mean Inclination in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the Brouwer-Lyddane conversion. The Inclination is defined by the angle between the orbit plane and the fundamental plane of the ICRF.	Type: Variable Access: Read/Write Units: deg
BLJ2MA	Mean Anomaly in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the Brouwer-Lyddane conversion, and referenced to the ICRF. 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.	Type: Variable Access: Read/Write Units: deg
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 Brouwer-Lyddane conversion, and referenced to the ICRF.	Type: Variable Access: Read/Write Units: deg
BLJ2W	Mean Argument of Perigee in the J2 Brouwer Lyddane Element set, which is calculated using only the J2 zonal harmonic coefficient in the Brouwer-Lyddane conversion, and referenced to the ICRF. 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.	Type: Variable Access: Read/Write Units: deg
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.	Type: Variable Access: Read/Write
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 non-escape 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.	Type: Variable Access: Read-Only Units: km^2/s^2
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.	Type: Variable Access: Read/Write
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.	Type: Array Access: Read-Only Units: m
CentralBody	Name of the body that defines the origin of the Spacecraft's position and velocity. This can be a user-defined CelestialObject or one of the following pre-defined 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.	Type: String Access: Read/Write
CentralBodyObject	CelestialObject that defines the origin for the Spacecraft's position and velocity.	Type: CelestialObject Access: Read-Only
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.	Type: Variable Access: Read/Write
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 z-axis of the spin frame is coincident with the spinning Spacecraft's spin axis.	Type: Variable Access: Read/Write Units: deg
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.	Type: Variable Access: Read/Write
Cr	The Coefficient of Reflectivity is a property that is a function of the properties of the spacecraft's area incident to the Sun-Earth 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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read-Only
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.	Type: Variable Access: Read-Only
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).	Type: Variable Access: Read-Only
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 ICRF.	Type: Variable Access: Read/Write Units: deg
DeclaredName	The name of the object as declared.	Type: String Access: Read-Only
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.	Type: Variable Access: Read-Only Units: kg/km^3
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.	Type: Variable Access: Read-Only Units: km
DisplayName	The name displayed for this object in output windows such as views, plots, and reports.	Type: String Access: Read/Write
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.	Type: Array Access: Read-Only Units: N
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.	Type: Variable Access: Read/Write Units: m^2
E	The Keplerian eccentricity (e). A unit-less property that describes the curvature of an ellipse. The distance from the center of the ellipse to one of the foci divided by the semi-major axis.	Type: Variable Access: Read/Write
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)	Type: Variable Access: Read-Only Units: deg
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.	Type: Variable Access: Read-Only Units: W/kg
ElapsedTime	Returns the time passed since the first time this property is read on a specific line of script.	Type: TimeSpan Access: Read-Only Units: day
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.	Type: Variable Access: Read-Only Units: 1/m^3
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.	Type: String Access: Read/Write
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).	Type: Variable Access: Read-Only Units: km^2/s^2
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.	Type: TimeSpan Access: Read/Write Units: day
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.	Type: TimeSpan Access: Read-Only Units: day
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 4-digit year; for millisecond mode, HH:MM:SS.SSS refers to hours, minutes, and seconds to milliseconds; and for nanosecond mode HH:MM:SS.SSSSSSSSS refers to hours, minutes, and seconds to nanoseconds.	Type: String Access: Read-Only
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.	Type: TimeSpan Access: Read-Only Units: day
EpochUT1	Returns the current Spacecraft date and time in Universal Time (UT1) in the Modified Julian Date format.	Type: TimeSpan Access: Read-Only Units: day
EpochUTC	Returns the current Spacecraft date and time in Coordinated Universal Time (UTC) in the Modified Julian Date format.	Type: TimeSpan Access: Read-Only Units: day
EquinoctialA	The a component of the Equinoctial element set. Half the length of the long axis of the orbit ellipse.	Type: Variable Access: Read/Write Units: km
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write Units: deg
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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.	Type: Array Access: Read/Write Units: deg
EulerAnglesICRF	Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the ICRF. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence.	Type: Array Access: Read-Only Units: deg
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.	Type: Array Access: Read-Only Units: deg
EulerAnglesRIC	Returns the three Euler Angles that define the orientation of the Spacecraft Body-Coordinate-System 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.	Type: Array Access: Read-Only Units: deg
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.	Type: Array Access: Read-Only Units: deg
EulerRates	Rate of change of the Euler angles.	Type: Array Access: Read/Write Units: deg/s
EulerSequence	Axes of rotation that the EulerAngles are around.	Type: Array Access: Read/Write
EVector	Represents the instantaneous eccentricity vector referenced to the ICRF. The eccentricity vector points from the center of the central body toward perigee with a magnitude equal to the eccentricity of the orbit.	Type: Array Access: Read-Only
FixedPosition	Reports the object's position components (X, Y, Z) referenced to the Body Fixed coordinate system of its central body.	Type: Array Access: Read-Only Units: km
FixedVelocity	Reports the Spacecraft's velocity components (VX, VY, VZ) referenced to the Body Fixed coordinate system of its central body.	Type: Array Access: Read-Only Units: km/s
ForceCoefficientType	Flag indicating whether Cd and Cr are used in the physical model, or B* and AGOM from the SGP4 model is used.	Type: Variable Access: Read/Write
FPAngle	Returns the flight path angle, which is measured between the perpendicular of the instantaneous position and the instantaneous velocity vectors of the spacecraft.	Type: Variable Access: Read-Only Units: deg
FPAzimuth	Returns the azimuth angle of the flight path, which is measured between the local North vector and the velocity vector.	Type: Variable Access: Read-Only Units: deg
GNSSReceivers	The spacecraft's Globlal Navigation Satellite System (GNSS) receivers.	Type: Array of GNSSReceiver Access: Read-Only
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 sub-satellite 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.	Type: Variable Access: Read-Only Units: km
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write Units: deg
HyperbolicExcessVelocity	Returns the hyperbolic excess velocity of the calling Spacecraft at its current semi-major axis. This property is defined for hyperbolic orbits, and will return undefined for trajectories which have a semi-major axis greater than 0.	Type: Variable Access: Read-Only Units: km/s
I	The Keplerian inclination (i). The angle between the orbit plane and the fundamental plane of the ICRF.	Type: Variable Access: Read/Write Units: deg
IsUsingAHF	Boolean value indicating whether or not the attitude motion model is based on an AHF.	Type: Variable Access: Read-Only
KSCEarthSunAngles	Returns a five-element Array containing the KSC Earth-Sun angles. These angles can be accessed individually using the AlphaTS, AlphaRS, AlphaTE, AlphaRE, and LVBeta properties.	Type: Array Access: Read-Only Units: deg
Latitude	The Latitude property is the bodydetic latitude of the Spacecraft sub-satellite point in the Spherical Lat/Long Element Set. It is defined as the angle between the fundamental plane of the Body-Fixed 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.	Type: Variable Access: Read/Write Units: deg
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 z-axis onto this plane to the projection of the velocity vector onto this plane.	Type: Variable Access: Read/Write Units: deg
LatLongRadius	The magnitude of the position vector in the Spherical Lat/Long Element set.	Type: Variable Access: Read/Write Units: km
LatLongVi	The magnitude of the velocity vector in the inertial frame in the Spherical Lat/Long Element set.	Type: Variable Access: Read/Write Units: km/s
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.	Type: TimeSpan Access: Read/Write Units: day
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.	Type: Variable Access: Read/Write Units: deg
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. Z-axis - north-pointing vector normal to the true equatorial plane at the reference epoch. X-axis - 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. Y-axis - forms a right-handed coordinate system.	Type: Variable Access: Read/Write Units: deg
Lift	Represents the instantaneous force on the Spacecraft due to atmospheric lift modeling.	Type: Array Access: Read-Only Units: N
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.	Type: Variable Access: Read/Write Units: m^2
LongAscendingNode	Returns the geographic East longitude of the sub-satellite 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.	Type: Variable Access: Read-Only Units: deg
LongDescendingNode	Returns the geographic East longitude of the sub-satellite 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.	Type: Variable Access: Read-Only Units: deg
Longitude	The Longitude property is the longitude of the Spacecraft sub-satellite 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 Body-Fixed 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.	Type: Variable Access: Read/Write Units: deg
LongitudeOfPeriapsis	Returns the angle measured between the vernal equinox and periapsis.	Type: Variable Access: Read-Only Units: deg
LongitudeRate	Calculates the rate of change in longitude of the orbital plane in a Body-Fixed 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 built-in CelestialObjects.	Type: Variable Access: Read-Only Units: deg/s
LunarBetaAngle	Returns the declination of the Earth-Moon vector with respect to the Earth-Sun ecliptic plane.	Type: Variable Access: Read-Only Units: deg
LVBeta	Returns the angle measured between the orbit angular momentum vector and the sun-spacecraft line.	Type: Variable Access: Read-Only Units: deg
MA	Represents the instantaneous Mean Anomaly of the Spacecraft.	Type: Variable Access: Read-Only Units: deg
Mass	Returns the total mass of the Spacecraft, including all attached Tanks.	Type: Variable Access: Read-Only Units: kg
MassFlowRate	The rate at which the mass of the Spacecraft is currently being depleted. This value will be non-zero only when the Spacecraft is maneuvering.	Type: Variable Access: Read-Only Units: kg/s
MassTotal	The total mass of the Spacecraft including all attached Tanks. If you set this property, VehicleDryMass is set to (MassTotal - total hardware mass).	Type: Variable Access: Read/Write Units: kg
MeanMotion	Returns the mean motion, which is a measure of how fast a spacecraft progresses through its elliptical orbit.	Type: Variable Access: Read-Only Units: rad/s
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.	Type: Variable Access: Read-Only Units: deg
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.	Type: Variable Access: Read-Only Units: deg
MGL	Returns the sum of the Mean Anomaly and the longitude of periapsis minus the Spacecraft.CentralBodyObject's prime meridian apparent sidereal time (see the CelestialObject.GetPrimeMeridianApparentSiderealTime for more information).	Type: Variable Access: Read-Only Units: deg
Mlt	Returns the Mean Local Time of the Spacecraft's sub-satellite point, which is the current UTC time of the Spacecraft's epoch adjusted by accounting for the Spacecraft's longitude.	Type: Variable Access: Read-Only Units: hr
MltAscendingNode	Returns the Mean Local time (MLT) of the Spacecraft's sub-satellite 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.	Type: Variable Access: Read-Only Units: hr
MltDescendingNode	Returns the Mean Local time (MLT) of the Spacecraft's sub-satellite 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.	Type: Variable Access: Read-Only Units: hr
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write Units: deg
ModifiedEquinoctialP	The semiparameter p is defined by: p = a*(1-e^2) where a and e are the Keplerian Semi-Major Axis and Eccentricity.	Type: Variable Access: Read/Write Units: km
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.	Type: Matrix Access: Read-Only Units: kg m^2
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.	Type: Variable Access: Read-Only
MoonBAngle	Returns the angle, in the moon B-plane, of the piercing point measured from the T vector.	Type: Variable Access: Read-Only Units: deg
MoonBDotR	Returns the projection of the piercing point in the moon B-plane onto the R vector.	Type: Variable Access: Read-Only Units: km
MoonBDotT	Returns the projection of the piercing point in the moon B-plane onto the T vector.	Type: Variable Access: Read-Only Units: km
MoonBMagnitude	Returns the distance of the piercing point from the origin of the moon B-plane. In terms of the Cartesian coordinates. MoonBMagnitude = SQRT(MoonBdotT^2 + MoonBdotR^2)	Type: Variable Access: Read-Only Units: km
MoonVelocity	Returns a three dimensional vector representing the velocity of the moon in the ICRF.	Type: Array Access: Read-Only Units: km/s
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.	Type: Variable Access: Read-Only Units: deg
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= equatorial radius of the Earth a = Spacecraft's osculating semi-major axis e = Spacecraft's osculating eccentricity i = Spacecraft's osculating inclination	Type: Variable Access: Read-Only Units: deg/day
NonSingularA	The semi-major axis (a) property for the NonSingular Keplerian element set. Half the length of the long axis of the orbit ellipse.	Type: Variable Access: Read/Write Units: km
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write
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 ICRF.	Type: Variable Access: Read/Write Units: deg
NutationAngle	The nutation angle of a spinning Spacecraft is the coelevation of a spinning Spacecraft's spin axis with respect the z-axis of the angular momentum frame.	Type: Variable Access: Read/Write Units: deg
ObjectId	The unique identifier for the object.	Type: Variable Access: Read-Only
ObjectType	The type of the object.	Type: String Access: Read-Only
OD	This property contains the OD settings for all the Spacecraft properties that can be estimated.	Type: SpacecraftODProperties Access: Read-Only
OrbitalAngularAcceleration	Represents the instantaneous orbital angular acceleration about the central body referenced in the ICRF.	Type: Array Access: Read-Only Units: rad/s^2
OrbitalAngularMomentum	Represents the instantaneous specific orbital angular momentum (angular momentum per unit mass) vector about the central body referenced in the ICRF.	Type: Array Access: Read-Only Units: km^2/s
OrbitalAngularVelocity	Represents the instantaneous orbital angular velocity vector about the central body referenced in the ICRF.	Type: Array Access: Read-Only Units: rad/s
OrbitNormal	Represents the instantaneous orbit normal vector about the central body referenced in the ICRF.	Type: Array Access: Read-Only
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write
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.	Type: Variable Access: Read/Write
P	Represents the osculating value of the semi-latus rectum of the Spacecraft's orbit. The value of the semi-latus 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.	Type: Variable Access: Read-Only Units: km
PercentEarthShadow	Returns a numerical value representing the percentage-of-shadow 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%.	Type: Variable Access: Read-Only
PercentMoonShadow	Returns a numerical value representing the percentage-of-shadow 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%.	Type: Variable Access: Read-Only
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.	Type: Variable Access: Read-Only Units: km
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.	Type: Variable Access: Read-Only Units: km
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 sub-satellite point. The oblateness of the central body is taken into account when calculating the height.	Type: Variable Access: Read-Only Units: km
Period	Returns the osculating period, which corresponds to the instantaneous unperturbed two-body Keplerian orbit.	Type: Variable Access: Read-Only Units: min
Pitch	Returns the angle about the y-axis of the 3-1-2 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.	Type: Variable Access: Read-Only Units: deg
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 higher-fidelity solar radiation pressure (SRP).	Type: PlateModel Access: Read-Only
Position	Returns the Spacecraft's X, Y, Z position referenced to the ICRF.	Type: Array Access: Read/Write Units: km
Propagator	The propagator for updating the orbital state of the Spacecraft.	Type: Propagator Access: Read-Only
PropagatorType	The object type of the propagator that is associated with the Spacecraft.	Type: String Access: Read-Only
ProximityZones	The Spacecraft's proximity zones.	Type: Array of ProximityZone Access: Read-Only
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 right-handed convention when applying the quaternion rotation.	Type: Array Access: Read/Write
RA	The right ascension of the satellite in the Spherical Element Set. This is the angle from the x-axis of the ICRF to the projection of the satellite position vector onto the fundamental plane of the coordinate frame.	Type: Variable Access: Read/Write Units: deg
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 ICRF going from south to north.	Type: Variable Access: Read/Write Units: deg
RADescendingNode	Represents the osculating Right Ascension of the Descending Node (RADN) in the ICRF. The RADN is defined as the angle, measured at the center of the central body, from the vernal equinox to the Descending Node.	Type: Variable Access: Read-Only Units: deg
Radius	The magnitude of the Cartesian position vector of the Spacecraft with respect to the central body.	Type: Variable Access: Read-Only Units: km
Roll	Returns the angle about the x-axis of the 3-1-2 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.	Type: Variable Access: Read-Only Units: deg
SatelliteId	Used by orbit determination algorithms to identify the Spacecraft.	Type: String Access: Read/Write
Sensors	The Spacecraft's sensors.	Type: Array of Sensor Access: Read-Only
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.	Type: Variable Access: Read-Only Units: deg
SGP4	Object holding options for managing the SGP4 propagation model and orbital element set.	Type: SGP4Properties Access: Read-Only
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.	Type: Variable Access: Read-Only
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 ICRF's z-axis onto this plane to the projection of the velocity vector onto this plane.	Type: Variable Access: Read/Write Units: deg
SphericalRadius	The magnitude of the position vector in the Spherical Element set.	Type: Variable Access: Read/Write Units: km
SpinPhase	Phase angle of the x-axis of the Body Coordinate System with respect to the projection of the x-axis of the ICRF into the spin plane. If SpinPhase = 0, the initial x-axis of the Body Coordinate System will be coincident with the projection of the x-axis of the ICRF into the spin plane.	Type: Variable Access: Read/Write Units: deg
SpinRate	The angular rotation rate of a spinning Spacecraft about its spin axis, i.e, the z-axis of the "spin frame".	Type: Variable Access: Read/Write Units: deg/s
SRPArea	The SRP Area is the incident cross sectional area to the Sun-Earth 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.	Type: Variable Access: Read/Write Units: m^2
SunVelocity	Returns a three-dimensional vector representing the velocity of the Sun in the ICRF.	Type: Array Access: Read-Only Units: km/s
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.	Type: Variable Access: Read-Only Units: deg
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.	Type: Variable Access: Read-Only Units: deg
TA	The Keplerian true anomaly. The angle measured at the center of the central body, from perigee to the current position of the satellite.	Type: Variable Access: Read/Write Units: deg
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.	Type: Variable Access: Read/Write
Tanks	The spacecraft's fuel tanks.	Type: Array of Tank Access: Read-Only
ThreeDModelFile	The file name of the file containing the 3D model that will represent the Spacecraft in a Mission View.	Type: String Access: Read/Write
ThreeDModelObject	The name of the ThreeDModel object that holds information about the physical geometry of the Spacecraft.	Type: ThreeDModel Access: Read-Only
Thrusters	The spacecraft's thrusters.	Type: Array of Thruster Access: Read-Only
TickType	The tick mark shape to draw at each point of the tail in a Mission View.	Type: String Access: Read/Write
TimeOfDay	Returns the current Spacecraft time referenced to Greenwich Mean Time (GMT or UTC).	Type: TimeSpan Access: Read-Only Units: hr
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.	Type: Variable Access: Read-Only Units: 1/m^2
Transponders	The spacecraft's transponders.	Type: Array of Transponder Access: Read-Only
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.	Type: Variable Access: Read-Only Units: deg
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.	Type: Array Access: Read/Write Units: m
VehicleDryMass	The Spacecraft's dry vehicle mass. The fuel mass is input through the tank and updated as mass is depleted.	Type: Variable Access: Read/Write Units: kg
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.	Type: Matrix Access: Read/Write Units: kg m^2
Velocity	Returns the instantaneous Spacecraft velocity vector in Cartesian components (VX, VY, VZ) referenced to the ICRF.	Type: Array Access: Read/Write Units: km/s
VerticalFPA	The vertical flight path angle in the Spherical Element set. The angle between the position and velocity vectors of the satellite.	Type: Variable Access: Read/Write Units: deg
Vi	The magnitude of the velocity vector in the inertial frame in the Spherical Element set.	Type: Variable Access: Read/Write Units: km/s
VMag	Represents the instantaneous magnitude of the Spacecraft's velocity.	Type: Variable Access: Read-Only Units: km/s
VX	The Cartesian X-component of the Spacecraft's velocity vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km/s
VY	The Cartesian Y-component of the Spacecraft's velocity vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km/s
VZ	The Cartesian Z-component of the Spacecraft's velocity vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km/s
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.	Type: Variable Access: Read/Write Units: deg
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.	Type: Variable Access: Read-Only
X	The Cartesian X-component of the Spacecraft's position vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km
Y	The Cartesian Y-component of the Spacecraft's position vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km
Yaw	Returns the angle about the z-axis of the 3-1-2 Euler sequence that represents the orientation of the Spacecraft Body Coordinate System from its reference attitude frame.	Type: Variable Access: Read-Only Units: deg
Year	Returns the current four-digit year.	Type: Variable Access: Read-Only
Z	The Cartesian Z-component of the Spacecraft's position vector, referenced to the ICRF.	Type: Variable Access: Read/Write Units: km

Name	Description
Access	Returns the start time, end time, and duration of a Spacecraft Sensor's view of a Region, another Spacecraft, a GroundStation, a Star, or a celestial object. The Access time can be based on any active Sensors on the calling Spacecraft or a user-specified subset of Sensors.
AddGNSSReceiver	Creates a new GNSSReceiver object and adds it as hardware to the Spacecraft.
AddProximityZone	Creates a new ProximityZone object and adds it as hardware to the Spacecraft.
AddSensor	Creates a new Sensor object and adds it as hardware to the Vehicle.
AddTank	Creates a new Tank object and adds it as hardware to the Spacecraft.
AddTDRSTransponder	Creates a new TDRSTransponder object and adds it as hardware to the Spacecraft.
AddThruster	Creates a new Thruster object and adds it as hardware to the Spacecraft.
AddTransponder	Creates a new Transponder object and adds it as hardware to the Spacecraft.
AlongArcTrackSeparation	Returns the distance along a circular arc centered around the central body. The arc is in the orbit plane of the calling Spacecraft and passes through the calling Spacecraft to the projection of the argument Spacecraft position vector on the calling Spacecraft orbit plane. The distance returned is measured from the calling Spacecraft.
AlongTrackSeparation	Returns the distance along the calling Spacecraft velocity vector between the projection of the argument Spacecraft position on to the velocity vector and calling Spacecraft. The distance returned is measured from the calling Spacecraft.
AnnularShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the shadow of a central body. Events are reported for annular shadow conditions only. If the specified body is a member of the Earth-Moon system, the counterpart not specified will also be considered in shadow evaluation.
AvgLongAscendingNodeErrorKm	Returns a moving average of the Spacecraft's longitude of the ascending node error. The longitude of ascending node error is the offset from a user-defined global reference longitude grid measured at the ascending node, also referred to as a repeating groundtrack grid. The moving average will use a maximum of 288 data points to determine the average. A simple average is returned when the number of data points available is less than 288 points. Once at least 288 data points are available, the returned average value is an average over the most recently calculated set of data points, where the number of points in the set equals 288. (Assuming a 300 second integration step size, 288 points generates an average calculation over a day). This property requires two arguments by which the user defines the repeating groundtrack grid. The two arguments are the base longitude, used to initialize the repeat groundtrack grid, and the repeat cycle, the number of orbits before the Spacecraft repeats the repeat groundtrack grid. This property returns a scalar value. Note that this property is Earth-specific.
AvgLongDescendingNodeErrorKm	Returns a moving average of the Spacecraft's longitude of the descending node error. The longitude of descending node error is the offset from a user-defined global reference longitude grid measured at the descending node, also referred to as a repeating groundtrack grid. The moving average will use a maximum of 288 data points to determine the average. A simple average is returned when the number of data points available is less than 288 points. Once at least 288 data points are available, the returned average value is an average over the most recently calculated set of data points, where the number of points in the set equals 288. (Assuming a 300 second integration step size, 288 points generates an average calculation over a day). This property requires two arguments by which the user defines the repeating groundtrack grid. The two arguments are the base longitude, used to initialize the repeat groundtrack grid, and the repeat cycle, the number of orbits before the Spacecraft repeats the repeat groundtrack grid. This property returns a scalar value. Note that this property is Earth-specific.
Azimuth	Returns the azimuth angle of the argument as seen from the reference object. The definition of azimuth depends on the reference object and the passed arguments.
BPlaneBdotR	Returns the projection of the piercing point in the B-plane onto the R vector.
BPlaneBdotT	Returns the projection of the piercing point in the B-plane onto the T vector.
BPlaneBMag	Returns the distance of the piercing point from the origin of the B-plane. In terms of the Cartesian coordinates. BPlaneBMag = SQRT(BPlaneBdotT^2 + BPlaneBdotR^2)
BPlaneTheta	Returns the angle in the B-plane of the piercing point measured from the T vector.
CardinalPoints	Returns the times when the Spacecraft is at each of the cardinal points. The cardinal point calculations are based on the alpha angle, which is the angle between the Spacecraft's position and the target CelestialObject's position with respect to the Spacecraft's central body when both vectors are projected onto the Spacecraft's orbit plane. Note that the target CelestialObject is modeled as a point.
CelestialObjectWidth	Returns the angular width of the provided CelestialObject as observed from the Spacecraft position.
ClearSavedStates	Clears previous saved states for this object.
ClearThreeDModel	Clears any ThreeDModel object that has been associated with the Vehicle using the SetThreeDModel method.
CloseApproachRange	Returns a sequential listing of events related to the proximity of two objects. It provides the entry time and exit time of one object moving within the user specified range of another object, as well as the time and range of closest approach, and the time and range of greatest separation. The times and ranges are computed through one of two user-selectable formulations of the Alfano Negron Close Approach Software (ANCAS) interpolation algorithms. Note that the ANCAS algorithms rely on crossing a range threshold in order to kick off the algorithm, so if two objects are always within the threshold or outside the threshold the no events will be reported. The ANCAS algorithms are designed to work well for high-relative velocity conjunctions and the results won't be as accurate for low-relative velocity conjunctions. For more information on the ANCAS algorithms, please refer to: Alfano, Salvatore. Determining Satellite Close Approaches, Part II. Journal of the Astronautical Sciences 42.2. Apr-Jun 1994, 143-152.
CommZAspectAngle	Returns the aspect angle of the argument as seen from the reference object.
Contact	Returns a value of 1 when the argument is in contact and a value of 0 if no contact is available with the Spacecraft.
ContactTimes	Returns the acquisition of signal time (AOS), loss of signal time (LOS), and duration of the argument's contact with the Spacecraft.
CrossTrackSeparation	Returns the distance along the Spacecraft orbit normal vector between the calling Spacecraft and the projection of the argument Spacecraft position on to the orbit normal. The distance returned is measured from the calling Spacecraft.
DutyCycle	Returns the duty cycle of a Thruster Object. Duty cycle is the percentage on-time of a given thruster during a burn. This property scales the mass depletion and scales the acceleration due to thrust. This property ranges between 0 and 1.
EarthShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the Earth's shadow. Events are reported for Penumbra, Umbra, or Annular shadow condition geometries.
ElapsedTimeFromEpoch	Returns the time between a user-specified epoch and the current Spacecraft epoch.
Elevation	Returns the elevation angle of the argument as seen from the reference object.
EPV	Extended Precision Vectors give spacecraft position and velocity information in the Geocentric True of Date Rotating reference frame.
ExclusionZone	Returns the start time, end time, and duration of a pass based on a Spacecraft's location over a Region. The ExclusionZone time is based on the Spacecraft's sub-satellite point within a Region.
FixedVectorToObject	Returns an Array containing the components of the vector originating at the location of the prefix object and terminating at the location of the argument object, in the body fixed coordinate frame of the prefix object.
FOVInterference	Returns the time of entry into an area of interference, the time of exit from the area of interference, and the duration of interference. Interference is defined to be the period when both an object of interest and an interfering object are within the field-of-view of a Spacecraft sensor, and the interfering object lies within a user-specified cone about the Spacecraft to object-of-interest vector.
FOVTimes	Returns the entry and exit times when the argument Spacecraft, GroundStation, Star, or celestial body enters and leaves the Spacecraft's field of view. See each overload for a summary of the information returned.
FrozenMAGuess	Calculates an estimated mean anomaly of a one-burn maneuver that restores frozen orbit conditions.
GetApoapsisDeclination	Returns the declination of the apoapsis vector (DAV). The DAV is the latitude of the apoapsis vector referenced to the chosen coordinate frame. This value is only valid for elliptical cases (e < 1), it is invalid for hyperbolic orbits (e > 1) and will report an error.
GetApoapsisRightAscension	Returns the right ascension of the apoapsis vector (RAV). The RAV is the equatorial east longitude of the apoapsis vector, referenced to the chosen coordinate frame. This value is only valid for elliptical cases (e < 1), it is invalid for hyperbolic orbits (e > 1) and will report an error.
GetApoapsisVector	Returns the X, Y, Z unit vector of the Spacecraft's apoapsis vector referenced to the chosen coordinate frame. This value is only valid for elliptical cases (e < 1), it is invalid for hyperbolic orbits (e > 1) and will report an error.
GetCartesianState	Returns the X, Y, Z, VX, VY and VZ components of the current state of the Spacecraft, referenced from the CentralBody.
GetElementSetToElementSetJacobian	Returns the Jacobian that maps from one element set to another.
GetFromFile	Load the object state from the specified FreeFlyer object file.
GetFromString	Load the object state from the specified string.
GetHeight	Returns the bodydetic height. Objects below the surface will return a negative height.
GetHyperbolicExcessVelocityVector	Returns the X, Y, Z unit vector of the Spacecraft's outbound hyperbolic excess velocity vector (V_inf) referenced to the chosen coordinate frame. This value is only valid for hyperbolic cases (e > 1), it is invalid for elliptical orbits (e < 1) and will report an error.
GetKeplerianState	Returns the A, E, I, RAAN, W, TA components of the state of the Spacecraft, referenced from the CentralBody.
GetLaunchAsymptoteDeclination	Returns the declination of the launch asymptote (DLA). The DLA is the latitude of the outbound hyperbolic asymptote referenced to the chosen coordinate frame. This value is only valid for hyperbolic cases (e > 1), it is invalid for elliptical orbits (e < 1) and will report an error.
GetLaunchAsymptoteRightAscension	Returns the right ascension of the launch asymptote (RLA). The RLA is the equatorial east longitude of the outbound hyperbolic asymptote, referenced to the chosen coordinate frame. This value is only valid for hyperbolic cases (e > 1), it is invalid for elliptical orbits (e < 1) and will report an error.
GetRelativeCartesianPosition	Returns the relative Cartesian position of the calling Spacecraft in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
GetRelativeCartesianState	Returns the relative Cartesian state of the calling Spacecraft in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
GetRelativeCartesianVelocity	Returns the relative Cartesian velocity of the calling Spacecraft in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
GetRelativeCurvilinearPosition	Returns the relative curvilinear position of the calling Spacecraft. The position will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
GetRelativeCurvilinearState	Returns the relative curvilinear state of the calling Spacecraft. The state will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
GetRelativeCurvilinearVelocity	Returns the relative curvilinear velocity of the calling Spacecraft. The velocity will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
IIRV	Improved Inter-Range Vectors give spacecraft position and velocity information in a format used by the Ground Network (GN) for Scheduling. The vectors are in the Geocentric True of Date Rotating reference frame.
InRange	Returns a value of 0 when the specified object is outside the specified ProximityZone(s) on the calling Spacecraft, and returns a value of 1 if the object is within the boundary of one or more ProximityZones. FreeFlyer uses a definitive analytical solution to evaluate this event. Since the evaluation is instantaneous, care must be taken to adjust the propagator step size such that events of interest aren't skipped over.
InRegion	Returns a value of 1 when a Spacecraft's sub satellite point is in a Region and a value of 0 if not in the Region.
InShadow	Returns an integer value indicating whether the Vehicle is in shadow. A value of 1 is returned when the Vehicle is in shadow, and a 0 is returned if it is not.
Intersects	This method provides a sequential listing of events related to the intersection of two objects, where one object moves to at least partially intersect another object. The objects can be a Spacecraft and a ProximityZone or a ProximityZone and a ProximityZone. The start times, end times, ranges at start and end, and duration of the intersection are returned. FreeFlyer uses a definitive analytical solution to evaluate this method.
InTrackSeparation	Returns the in-track separation between the calling Spacecraft and the primarySC in km. This is equivalent to the y-axis separation in the RIC frame.
InView	Returns a value of 1 when the Region, PointGroup, or location specified by an Array is in the view of the specified sensors, and a 0 if not in view. If no sensors are specified, any active sensor on the Spacecraft is used.
LoadRinexNavFile	Loads Broadcast Ephemeris data from a Rinex Navigation file. This method also attaches a sensor to the Spacecraft.
LoadTLE	Loads a state from the specified TLE or TLE file into the Spacecraft and returns the total number of states in the TLE or TLE file. If you are loading a Spacecraft state from a TLE, you should use a SGP4 propagator when running your Mission Plan.
LongAscendingNodeError	Returns the offset from a user-defined global reference longitude grid measured at the Ascending Node, also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. The spacing of the grid is dependent on the number of orbits in the repeat cycle. Note that this property is Earth-specific.
LongAscendingNodeErrorKm	Returns the offset of the actual ascending node on the surface of the central body from a user-defined global reference longitude grid measured at the Ascending Node. The global reference longitude grid is also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. Note that this property is Earth-specific.
LongAscendingNodeErrorRate	Returns the change in the average longitude of ascending node error with respect to time. The longitude of ascending node error is the offset from a user-defined global reference longitude grid measured at the Ascending Node, also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. Note that this property is Earth-specific.
LongDescendingNodeError	Returns the offset from a user-defined global reference longitude grid measured at the Ascending Node, also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. The spacing of the grid is dependent on the number of orbits in the repeat cycle. Note that this property is Earth-specific.
LongDescendingNodeErrorKm	Returns the offset of the actual descending node on the surface of the central body from a user-defined global reference longitude grid measured at the Descending Node. The global reference longitude grid is also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. Note that this property is Earth-specific.
LongDescendingNodeErrorRate	Returns the change in the average longitude of descending node error with respect to time. The longitude of descending node error is the offset from a user-defined global reference longitude grid measured at the Descending Node, also referred to as a repeating groundtrack grid. The repeating groundtrack grid represents the nominal nodal longitudes the Spacecraft would fly over, spaced evenly around the equator. Note that this property is Earth-specific.
MagneticField	Returns the Earth's magnetic field vector at the Spacecraft's current position. The user can specify the coordinate system used to express the magnetic field vector via a Variable or a Sensor. If a Variable or Sensor is not specified, the ICRF is used.
MagRegion	Represents the current magnetic field region the Spacecraft is in, as defined by the National Space Science Data Center's (NSSDC) Satellite Situation Center. It returns an integer value based on the following legend: 0 none of the below 1 Interplanetary Medium 2 Dayside Magnetosheath 3 Nightside Magnetosheath 4 Dayside Magnetosphere 5 Nightside Magnetosphere 6 Plasma Sheet 7 Tail Lobe 8 High Latitude Boundary Layer 9 Low Latitude Boundary Layer 10 Dayside Plasmasphere 11 Nightside Plasmasphere
Maneuver	Executes the specified maneuver and returns the maneuver duration. If the maneuver is an ImpulsiveBurn, then TimeSpan.Zero is returned.
METFromEpoch	This method returns the mission elapsed time (MET) for the calling Spacecraft since the provided referenceEpoch. The output MET is given in DD:HH:MM:SS.SSS format.
MltAscendingNodeError	Returns the offset from a user-defined Ascending Node Mean Local Time (MLT) crossing versus the Spacecraft's MLT crossing at the Ascending Node. This method is primarily used for Sun-Synchronous orbits: where the orientation of the orbit is fixed with respect to the Sun and the Spacecraft has a fixed nodal crossing time requirement. This method requires a single argument, which defines the fixed nodal MLT crossing. This base MLT is used to evaluate the difference between the engineer's requirement and the Spacecraft nodal crossing time.
MltDescendingNodeError	Returns the offset from a user-defined Descending Node Mean Local Time (MLT) crossing versus the Spacecraft's MLT crossing at the Descending Node. This method is primarily used for Sun-Synchronous orbits: where the orientation of the orbit is fixed with respect to the Sun and the Spacecraft has a fixed nodal crossing time requirement. This method requires a single argument, which defines the fixed nodal MLT crossing. This base MLT is used to evaluate the difference between the engineer's requirement and the Spacecraft nodal crossing time.
MoonShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the Moon's shadow. Events are reported for penumbra, umbra, or annular shadow condition geometries. Related Methods: EarthShadowTimes, ShadowTimes
Occult	The Occult method determines if an object is hidden from the view of the calling Spacecraft by the Earth. It returns a '1' when the object is occulted or '0' when it is not. When multiple objects are specified, the method returns a '1' if any of the objects are occulted and '0' if none are occulted.
OrbitApoapsis	The OrbitApoapsis method is a conditional property that can be used with the Step command to advance the Spacecraft to apoapsis, the point in the orbit where the Spacecraft is at its furthest distance from a Planet, Moon, or CelestialObject. FreeFlyer determines this condition by monitoring the dot product between the position and velocity vectors and identifying when this value passes from positive to negative. The celestial body to be used in the apoapsis calculation can be specified using a Variable, String, or CelestialObject. If a celestial body is not specified, the current Spacecraft Central Body will be used.
OrbitApogee	The OrbitApogee method is a conditional property that can be used with the Step command to advance the Spacecraft to apogee, the point in the orbit where the Spacecraft is at its furthest distance from a celestial body. FreeFlyer determines this condition by monitoring the dot product between the position and velocity vectors and identifying when this value passes from positive to negative. The celestial body to be used in the apoapsis calculation can be specified using a Variable, String, or CelestialObject. If a celestial body is not specified, the current Spacecraft Central Body will be used.
OrbitPeriapsis	The OrbitPeriapsis method is a conditional property that can be used with the Step command to advance the Spacecraft to periapsis, the point in the orbit where the Spacecraft is at its closest approach to a Planet, Moon, or CelestialObject. FreeFlyer determines this condition by monitoring the dot product between the position and velocity vectors and identifying when this value passes from negative to positive. The celestial body to be used in the periapsis calculation can be specified using a Variable, String, or CelestialObject. If a celestial body is not specified, the current Spacecraft Central Body will be used.
OrbitPerigee	The OrbitPerigee method is a conditional property that can be used with the Step command to advance the Spacecraft to perigee, the point in the orbit where the Spacecraft is at its closest approach to a celestial body. FreeFlyer determines this condition by monitoring the dot product between the position and velocity vectors and identifying when this value passes from negative to positive. The celestial body to be used in the periapsis calculation can be specified using a Variable, String, or CelestialObject. If a celestial body is not specified, the current Spacecraft Central Body will be used.
OrbitPeriselene	The OrbitPeriselene method is a conditional property that can be used with the Step command to advance the Spacecraft to periselene, the point in the orbit where the Spacecraft is at its closest approach to the Moon. FreeFlyer determines this condition by monitoring the dot product between the position and velocity vectors of the Spacecraft and identifying when this value passes from negative to positive.
OrbitWizardCircularSetup	Initializes the state of a spacecraft with a circular orbit around its central body. The user specifies the desired orbit Altitude or Radius. The Spacecraft's previous Inclination and RAAN are maintained.
OrbitWizardGeoSynchSetup	Initializes the state of a spacecraft with a geosynchronous orbit. The user specifies the desired Longitude at the Equator Crossing, Argument of Perigee, and True Anomaly. The Spacecraft's previous Inclination is maintained.
OrbitWizardInterplanetarySetup	Initializes the state of a spacecraft with an orbit centered around the specified CelestialObject.
OrbitWizardMolniyaSetup	Initializes the state of a spacecraft with a Molniya orbit around the Earth. The user specifies the desired orbit Epoch and Longitude at Apogee and whether to dwell in the Northern or Southern hemisphere. The Spacecraft's previous Inclination, True Anomaly, and RAAN are maintained.
OrbitWizardRelativeMotionCircular	Configures the orbit state of the calling Spacecraft to fly in a circular formation with the specified separation from the primarySC. The circular formation causes the secondarySC to revolve around the primary at a constant distance. The period of the circular relative motion is the same as the orbital period of the primarySC. The circular relative motion of the secondarySC is constrained to a plane that is inclined +/- 30 degrees from primarySC's intrack-crosstrack plane, towards the radial direction. The user specifies the circular formation radius, initial phase angle, and inclination sign. An initial phase angle of zero corresponds to placing the secondarySC with an initial RIC-frame y-axis separation of zero; the z-axis separation would be at its maximum value. This formation orbit wizard is based on the HCW system, so it assumes that the primarySC is in a two-body circular orbit, and that the secondarySC is near to the primary and propagated with the HCW propagator. If this is not the case, the actual propagated formation will not maintain the idealized relative motion.
OrbitWizardRelativeMotionInPlane	Configures the orbit state of the calling Spacecraft to fly in an in-plane formation with the specified in-track separation from the primarySC. The in-plane formation is a simple design, where the two Spacecraft occupy the same orbit with a constant separation in mean anomaly. The specified in-track separation represents the initial RIC-frame y-axis separation between the two Spacecraft. This formation orbit wizard is based on the HCW system, so it assumes that the primarySC is in a two-body circular orbit, and that the secondarySC is near to the primary and propagated with the HCW propagator. If this is not the case, the actual propagated formation will not maintain the idealized relative motion.
OrbitWizardRelativeMotionInTrack	Configures the orbit state of the calling Spacecraft to fly in an in-track formation with the specified in-track separation from the primarySC. The in-track formation is similar to the in-plane formation, but here the Spacecraft will follow each other in the same ground-track. This formation is achieved with a separation in RAAN, which corresponds to a cross-track oscillation of the secondary Spacecraft relative to the primary. The period of the cross-track oscillation is the same as the orbital period of the primary. The specified in-track separation represents the initial RIC-frame y-axis separation between the two Spacecraft. This formation orbit wizard is based on the HCW system, so it assumes that the primarySC is in a two-body circular orbit, and that the secondarySC is near to the primary and propagated with the HCW propagator. If this is not the case, the actual propagated formation will not maintain the idealized relative motion.
OrbitWizardRelativeMotionProjectedCircular	Configures the orbit state of the calling Spacecraft to fly in a projected circular formation with the specified separation from the primarySC. The projected circular formation causes secondarySC to revolve around the primarySC in an ellipse, where the relative trajectory forms a circle projected onto the primarySC's RIC frame Y-Z plane (i.e. the intrack-crosstrack plane). The period of the circular relative motion is the same as the orbital period of primarySC. The circular relative motion of the secondarySC is constrained to a plane that is inclined +/- 26.565 degrees from the primarySC's intrack-crosstrack plane, towards the radial direction. The user specifies the radius of the projected circle, initial phase angle, and inclination sign. An initial phase angle of zero corresponds to placing the secondarySC with an initial RIC-frame y-axis separation of zero; the z-axis separation would be at its maximum value. This formation orbit wizard is based on the HCW system, so it assumes that the primarySC is in a two-body circular orbit, and that the secondarySC is near to the primary and propagated with the HCW propagator. If this is not the case, the actual propagated formation will not maintain the idealized relative motion.
OrbitWizardRelativeMotionSafetyEllipse	Configures the orbit state of the calling Spacecraft to fly in a safety ellipse formation with the specified radial and cross-track offsets, the center of the ellipse in track, the center of the ellipse rate of change in track seen by the primarySC, and the radial phase angle, from the primarySC. This formation orbit wizard is based on the HCW system, so it assumes that the primarySC is in a two-body circular orbit, and that the secondarySC is near to the primary and propagated with the HCW propagator. If this is not the case, the actual propagated formation will not maintain the idealized relative motion.
OrbitWizardSunSyncSetup	Initializes the state of a Spacecraft with a sun-synchronous orbit around the Earth. The user specifies the desired eccentricity and semi-major axis or inclination, and whether these are evaluated at the ascending or descending node. The Spacecraft's previous Argument of Perigee and True Anomaly are maintained.
ParticleFlux	The ParticleFlux method uses the NASA AP-8 and AE-8 particle flux maps to calculate the omnidirectional flux of the specified particle type with the energy level specified by minimumEnergy. The AP/AE maps contain collected data of proton and electron fluxes in the Earth's radiation belt with energy ranges from 0.04 MeV to 7 MeV for electrons and from 0.1 MeV to 400 MeV for protons. See the help file for links to more detailed documentation on the AP/AE maps. The user has the option to specify epoch and longitude offsets to include in the flux calculation.
ParticleFluxTimes	Determines when the spacecraft crosses into or out of a region with a certain level of flux. The ParticleFluxTimes method uses the NASA AP-8 and AE-8 particle flux maps to calculate the omnidirectional flux of the specified particle type with the energy level specified by minimumEnergy. The AP/AE maps contain collected data of proton and electron fluxes in the Earth's radiation belt with energy ranges from 0.04 MeV to 7 MeV for electrons and from 0.1 MeV to 400 MeV for protons. See the help file for links to more detailed documentation on the AP/AE maps. The user has the option to specify epoch and longitude offsets to include in the flux calculation.
PassData	Provides detailed information on the contact between a GroundStation and a Spacecraft. This information includes Acquisition of Signal (AOS), Loss of Signal (LOS), and the maximum elevation of the contact.
PenumbralShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the shadow of a central body. Events are reported for penumbral shadow condition geometries only. If the specified body is a member of the Earth-Moon system, the counterpart not specified will also be considered in shadow evaluation.
PercentShadow	Returns a numerical value representing the percentage-of-shadow the Spacecraft encounters in its orbit with respect to a CelestialObject's shadow.
PlanetPosition	Returns the position vector of the specified planetary body in the ICRF. The planetary body location is defined as the location of the planet's system barycenter and not the actual position of the planetary body.
PlanetVelocity	Returns the velocity vector of the specified planetary body with respect to the CentralBody of the Spacecraft in the ICRF.
ProbabilityOfCollision	Returns the 2D probability of collision (Pc) between two objects. Each object must be propagated to the time of closest approach (TCA). The 2D method assumes a high relative velocity and linear relative motion. Note, FreeFlyer uses Foster's 1992 method, in Cartesian coordinates, to compute the Probability of Collision. For more information on Foster's 1992 method, please refer to: Foster's journal paper "A Parametric Analysis of Orbital Debris Collision Probability and Maneuver Rate for Space Vehicles" from August of 1992.
PutToFile	Convert the object state to FreeFlyer object xml and write to a file.
PutToString	Convert the object state to FreeFlyer object xml and return as a string.
RadialProjSeparation	Returns the distance between the calling Spacecraft and the projection of the argument Spacecraft position on the calling Spacecraft orbit radius vector. The distance returned is measured from the calling Spacecraft and is positive in the direction of the calling Spacecraft radius vector.
RadialSeparation	Returns the difference between the position vector magnitude of the calling Spacecraft and the position vector magnitude of the argument Spacecraft (i.e. radial separation = calling Spacecraft position magnitude - argument Spacecraft position magnitude).
Range	Returns the distance between two Vehicles, a Vehicle and a GroundStation, or a Vehicle and a CelestialObject. The Range method requires a single argument, either a Vehicle, a CelestialObject, a CelestialObject BodyID, or a GroundStation object.
RangeRate	Returns the time rate of change of a Spacecraft's range from either a GroundStation or another Spacecraft.
Rate	Calculates the rate of change with respect to time of a user-defined argument or Variable object. The rate is determined by differencing successive data points and dividing by the difference in time in days, based on the Spacecraft's epoch at each point.
RayPathHeight	The planeto-detic height of the Line-of-Sight vector above the limb of an interfering Celestial Object. The interfering Celestial Object is the central body of the calling Spacecraft unless otherwise specified.
ReferenceEquals	Returns 1 if the argument refers to the calling object and 0 otherwise.
RefreshDailyOrbitNumber	Determines the current daily orbit number for the Spacecraft and automatically fills the Spacecraft.OrbitNumberDaily property with the result.
RelativeAngle	Returns the angle measured between any two celestial objects, Stars, or Vehicles, with the calling Vehicle as the vertex.
ResetMethodStates	Resets any method states associated with the calling object.
Restore	Restore an object's state from a previously saved state.
Save	Save the object's state so that it can be restored later.
SetAttitudeMotionModel	Sets the motion model to use for attitude propagation. With a Kinematic model, the user can configure the Spacecraft attitude using Euler Angles, Quaternions, an Attitude Matrix, or by setting up a Spinner. When AHF is selected, the user can specify an ephemeris containing attitude data to use when propagating the Spacecraft.
SetAttitudeRefFrame	Sets the Spacecraft's attitude reference frame to the desired predefined reference frame or custom coordinate system object. This reinitializes the Spacecraft's attitude representation, which sets the Spacecraft's EulerSequence to the default value of {3,1,2}.
SetCartesianState	Sets the Cartesian state of the Spacecraft, referenced from the CentralBody.
SetCentralBody	Sets the body of the calling object.
SetFixedPosition	Sets the object's position components (X, Y, Z) referenced to the Body Fixed coordinate system of its central body.
SetFixedVelocity	Sets the Spacecraft's velocity components (VX, VY, VZ) referenced to the Body Fixed coordinate system of its central body.
SetKeplerianState	Sets the Keplerian state of the Spacecraft, referenced from the CentralBody.
SetOrientation	Sets the orientation of the reference object using the specified CoordinateSystem object. Note, the AttitudeRefFrame property of the Vehicle must be set to "ICRF".
SetPropagatorType	Sets the propagator type to use to advance the vehicle state in time.
SetRandomState	Populates the Spacecraft state using a random state from a Gaussian distribution described by the provided covariance matrix. The dimension and order of the covariance matrix is assumed to be consistent with the Spacecraft's ProcessAction settings.
SetRelativeCartesianPosition	Sets the calling Spacecraft's Cartesian position in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetRelativeCartesianState	Sets the calling Spacecraft's Cartesian state in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetRelativeCartesianVelocity	Sets the calling Spacecraft's Cartesian velocity in reference to the refSC in the ICRF. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetRelativeCurvilinearPosition	Sets the calling Spacecraft's curvilinear position in reference to the refSC. The position will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetRelativeCurvilinearState	Sets the calling Spacecraft's curvilinear state in reference to the refSC. The state will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetRelativeCurvilinearVelocity	Sets the calling Spacecraft's curvilinear velocity in reference to the refSC. The velocity will be evaluated in the LVC cylindrical curvilinear frame. Note that if the epochs of the calling Spacecraft and the refSC do not match, this will report an error.
SetThreeDModel	Associates a ThreeDModel object to the specified Vehicle.
ShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the shadow of a central body. Events are reported for Penumbra, Umbra, or Annular shadow condition geometries. If the specified body is a member of the Earth-Moon system, the counterpart not specified will also be considered in shadow evaluation.
SimulateTrackingData	Simulate OD tracking data. The measurement usage, noise, bias and error model settings are determined by the settings of the Observation provided as an argument and its associated Observing object.
SpacecraftNoon	Computes the observed noon as seen by the Spacecraft. Noon is defined as the midpoint between the two times when the Spacecraft exits and enters the Earth's umbral cone. This method returns the shadow entry and exits times, as well as the time of Spacecraft Noon.
Step	Steps the Vehicle using the current propagator step size.
StepToApsis	Steps to the specified apsis of the Spacecraft's orbit based on its current central body. The next or previous apsis is selected based on the sign of the current propagator stepsize.
StepToEpoch	Steps the Vehicle to the specified epoch.
StepToNode	Steps to the specified node of the Spacecraft's orbit based on its current central body. The next or previous node is selected based on the sign of the current propagator stepsize. If the spacecraft is already at the node, it will move off the node.
SubSatellitePointInShadow	Returns an integer value indicating whether the sub-satellite point on the Spacecraft's Central Body is in shadow. A value of 1 is returned when the sub-satellite point is in shadow, and a 0 is returned if it is not.
SunTerminatorCrossingTimes	Returns time of entry into darkness, entry into light, and duration in darkness of the spacecraft sub-satellite point. Darkness is defined to be the period when the spacecraft sub-satellite point is within the Earth-Umbra cone.
TCA	Returns the epoch for the time of closest approach (TCA) between two objects. This method returns the local extrema and is more precise if used after narrowing down your closest approaches using the Spacecraft.CloseApproachRange() method.
UmbralShadowTimes	Determines the entry time, exit time, and duration of a Spacecraft's pass through the shadow of a central body. Events are reported for umbral shadow conditions only. If the specified body is a member of the Earth-Moon system, the counterpart not specified will also be considered in shadow evaluation.
VectorToObject	Returns an Array containing the components of the vector originating at the location of the prefix object and terminating at the location of the argument object, in the ICRF.

Name	Description	Attributes
Attitude	Reports the three Euler rotation angles that define the attitude of a Spacecraft with respect to the AttitudeRefFrame. This property has been Deprecated. Please use Spacecraft.EulerAngles instead.	Type: Array Access: Read/Write Units: deg
AttitudeAngles	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. This property has been Deprecated. Please use the EulerAngles vector instead.	Type: Array Access: Read/Write Units: deg
AttitudeLVLH	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. This property has been Deprecated. Please use Spacecraft.EulerAnglesLVLH instead.	Type: Array Access: Read-Only Units: deg
AttitudeMJ2000	Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the ICRF. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence. This property has been Deprecated. Please use Spacecraft.EulerAnglesICRF instead.	Type: Array Access: Read-Only Units: deg
AttitudeUVW	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. This property has been Deprecated. Please use Spacecraft.EulerAnglesUVW instead.	Type: Array Access: Read-Only Units: deg
Bstar	A special "drag" term associated with the Norad SGP4/SDP4 propagators. It measured in units of 1/earth radii and is related to the Spacecraft's Cd. It is obtained by having FreeFlyer import the NORAD Two-Line Element (TLE) state vector. It is also computed by FreeFlyer when a conversion from Cartesian elements to Norad element is invoked. This property has been Deprecated. Please use Spacecraft.SGP4.Bstar instead.	Type: Variable Access: Read/Write Units: er^-1
DragCoeffType	Flag indicating whether Cd or B* is specified This property has been Deprecated. Please use Spacecraft.PhysicalPropertyDefinitionType instead.	Type: Variable Access: Read/Write
EarthShadow	EarthShadow returns an integer value indicating when the Spacecraft is in any portion of the Earth's shadow. A value of 1 is returned when the Spacecraft is in Earth's shadow and a 0 is returned if it is not. The calculation assumes a spherical Earth and spherical Sun. When a Spacecraft is in the penumbra region of the shadow, or partial shadow, EarthShadow returns a 1.	Type: Variable Access: Read-Only
EarthWidth	Returns the angular width of the Earth as observed from the Spacecraft's position.	Type: Variable Access: Read-Only Units: deg
EulerAnglesMJ2000	Returns the three Euler Angles that define the orientation of the Spacecraft Body Coordinate System with respect to the ICRF. This property does not affect the Euler Angle Sequence. Please use Spacecraft.EulerSequence to set the Euler Angle rotation sequence. This property has been Deprecated. Please use Spacecraft.EulerAnglesICRF instead. This property returns values referenced from the system inertial frame rather than the MJ2000 Earth Equator frame.	Type: Array Access: Read-Only Units: deg
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 equinox and the meridian passing through the Greenwich observatory in England. This property has been Deprecated. Please use the CelestialObject.GetPrimeMeridianSiderealTime method. The CelestialObject.GetBodyRotationAngle and CelestialObject.GetBodyRotationAngleInertial methods compute similar values and may also be useful.	Type: Variable Access: Read-Only Units: deg
JRIndex	Represents the current value of the dimensionless geomagnetic index (Kp) (an integer) used by the Jacchia-Roberts atmospheric density model at the Spacecraft's current epoch. These indices are stored in the Jacchia-Roberts File in three-hour increments.	Type: Variable Access: Read-Only
JRPlanetaryIndex	Represents the current value of the dimensionless planetary index (Ap) used by the Jacchia-Roberts and MSIS atmospheric density model at the Spacecraft's current epoch.	Type: Variable Access: Read-Only
JRTemperature	Represents the current value of the exospheric temperature being used by the Jacchia-Roberts atmospheric density model at the Spacecraft's current epoch. This value is stored in the Jacchia-Roberts flux predict file in daily increments.	Type: Variable Access: Read-Only Units: Deg K
L1Position	Returns the instantaneous position of the Earth-Sun L1 libration point in the ICRF. The location of the L1 point is calculated relative to the Sun and the Earth-Moon barycenter.	Type: Array Access: Read-Only Units: km
L2Position	Returns the instantaneous position of the Earth-Sun L2 libration point in the ICRF. The location of the L2 point is calculated relative to the Sun and the Earth-Moon barycenter.	Type: Array Access: Read-Only Units: km
Momentum	Represents the instantaneous specific orbital angular momentum (angular momentum per unit mass) vector referenced in the ICRF.	Type: Array Access: Read-Only Units: km^2/s
MoonDistance	Returns distance measured from the center of the Earth to the center of the Moon, based on the Spacecraft's epoch.	Type: Variable Access: Read-Only Units: km
MoonShadow	Returns an integer value indicating if the Spacecraft is in any portion of the Moon's shadow. Returns a value of 1 when the Spacecraft is in Moon shadow and a 0 if it is not.	Type: Variable Access: Read-Only
MoonVector	Returns the three-dimensional (X, Y, Z) position vector of the Moon with respect to the Earth, referenced to the ICRF.	Type: Array Access: Read-Only Units: km
MoonWidth	Returns the angular width of the Moon as observed from the Spacecraft position.	Type: Variable Access: Read-Only Units: deg
NoradE	The Norad element set Eccentricity. This property has been Deprecated. Please use the SGP4.E property instead.	Type: Variable Access: Read/Write
NoradEquinoctialH	The h component of the Norad based Equinoctial element set, defined by: h = e*sin(RAAN + w), where e, w, and RAAN are the Norad Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialH property instead.	Type: Variable Access: Read/Write
NoradEquinoctialK	The k component of the Norad based Equinoctial element set, defined by: k = e*cos(RAAN + w), where e, w, and RAAN are the Norad Eccentricity, Argument of Perigee and Right Ascension of the Ascending Node respectively. These orbit properties are with respect to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialK property instead.	Type: Variable Access: Read/Write
NoradEquinoctialLongitude	Equinoctial Norad based 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 True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialLongitude property instead.	Type: Variable Access: Read/Write Units: deg
NoradEquinoctialMeanMotion	The Norad element set Mean Motion, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialMeanMotion property instead.	Type: Variable Access: Read/Write Units: orbits/day
NoradEquinoctialP	The p component of the Norad based Equinoctial element set, defined by: p = tan(i/2)sin(RAAN), where i is the orbit inclination and RAAN is the Norad Right Ascension of the Ascending Node. These orbit properties are with respect to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialP property instead.	Type: Variable Access: Read/Write
NoradEquinoctialQ	The q component of the Norad based Equinoctial element set, defined by: q = tan(i/2)cos(RAAN) where i is the orbit inclination and RAAN is the Norad Right Ascension of the Ascending Node. These orbit properties are with respect to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.EquinoctialHK property instead.	Type: Variable Access: Read/Write
NoradI	The Norad element set Inclination, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.I property instead.	Type: Variable Access: Read/Write Units: deg
NoradMA	The Norad element set Mean Anomaly, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.MA property instead.	Type: Variable Access: Read/Write Units: deg
NoradMeanMotion	The Norad element set Mean Motion, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.MeanMotion property instead.	Type: Variable Access: Read/Write Units: orbits/day
NoradRAAN	The Norad element set Right Ascension of the Ascending Node, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.RAAN property instead.	Type: Variable Access: Read/Write Units: deg
NoradW	The Norad element set Argument of Perigee, referenced to the True Equator Mean Equinox coordinate frame. This property has been Deprecated. Please use the SGP4.W property instead.	Type: Variable Access: Read/Write Units: deg
Receivers	The spacecraft's receivers. This property has been Deprecated. GPSReceivers and the Spacecraft.Receivers array have been deprecated. Please use Spacecraft.GNSSReceivers instead. See the Help File Guide on Orbit Determination for more information.	Type: Array of Receiver Access: Read-Only
SCMoonDistance	Returns the distance measured from the Spacecraft to the center of the Moon.	Type: Variable Access: Read-Only Units: km
SCSunDistance	Returns the distance measured from the Spacecraft to the center of the Sun.	Type: Variable Access: Read-Only Units: km
Shadow	Returns either a '1' or a '0' value indicating the Spacecraft is in any portion of either the Earth's or Moon's Shadow based on a penumbral condition.	Type: Variable Access: Read-Only
SunDistance	Returns the magnitude of the position vector from the center of the Sun to the center of the spacecraft's central body.	Type: Variable Access: Read-Only Units: km
SunVector	Returns the three-dimensional (X, Y, Z) position of the Sun with respect to the calling object's central body, referenced to the ICRF.	Type: Array Access: Read-Only Units: km
SunWidth	Returns the angular width of the Sun as observed from the Spacecraft position.	Type: Variable Access: Read-Only Units: deg
SystemTime	The current date and time of the system in the International Atomic Time (TAI) system represented in the Modified Julian Date format.	Type: TimeSpan Access: Read-Only Units: day