Satellite and GNSS Satellite Scope Functions

Example VBS:

1
2
srp = sc.Satellite(0).ForceModel.SolarPressure
bResult = srp.GenerateSPADFile("C:\SPADSphereSRP.txt", "Sphere", 10.0, 5.0)

bool GenerateSPADFileFromNPlate(string outputFile, string shapeType, double gridSpacing)

Use this to create a tabulated area vector file in SPAD format to represent the shape defined by the supplied N-Plate file. You can use the output file as a definition file when using the "Tabulated Area Vector" solar pressure model.

Parameters

Return Value

Example VBS

1
2
srp = sc.Satellite(0).ForceModel.SolarPressure
bResult = srp.GenerateSPADFileFromNPlate("C:\SPADCustomSRP.txt", "C:\MySat.nplate", 5.0)

bool Go()

Use this to generate ephemeris for a specified satellite over the time span given in the satellite's EphemerisGeneration attributes.

Return Value

Example VBS:

1
bResult = sc.Satellite(0).Go

bool InitialCovarianceIsPositiveDefinite()

This checks if the initial orbit error covariance, expressed in the OrbitUncertainty settings, is positive definite.

Return Value

Example VBS:

1
bResult = sc.Satellite(0).OrbitUncertainty.InitialCovarianceIsPositiveDefinite()

bool OEMToSTKEphemeris(string inputFile, string outputFile)

This converts a CCSDS 502 OEM *.oem file to an STK ephemeris *.e file.

Parameters

Return Value

Example VBS:

1
2
bResult = sc.Satellite(0).OEMToSTKEphemeris("C:\od\sat1.oem", _
                                            "C:\od\sat1.e")

bool OutputPropagatorDefn(string outputFile)

This places the propagator definition in a file for the Ansys Systems Tool Kit® (STK®) application to read for initializing a satellite's integrator settings.

Parameters

Return Value

Example VBS:

1
sc.GNSS.GNSSSatellite(0).OutputPropagatorDefn "C:\od\gps1.pg"

bool PropEphem(string filename, string stopTime, bool inclCov, bool updateState)

This is a helper function for InitalStateTool utility. It propagates a satellite state from the current epoch to a new time. You can use it to generate an ephemeris file or move a satellite state to a new epoch. You access this function via the Satellite's EphemerisGeneration attribute scope.

Parameters

Return Value

Example VBS:

1
2
Set sat = sc.Satellite(0)
bResult = sat.EphemerisGeneration.PropEphem("c:\temp\sat.e","1 Jul 2009 05:30:00 UTCG", false, true)

bool ReplaceICRFWithJ2000()

This function looks for references to the ICRF frame and converts them to the J2000 frame. For a satellite, it will transform the orbit from ICRF to J2000. It will also check the ephemeris generation CoordFrame and check each maneuver frame, changing frames from ICRF to J2000. It does not otherwise modify maneuver data. Any changes appear in the Message Viewer.

Calling this function at the scenario level will in turn call this function for each satellite and GNSSSatellite object.

Return Value

Example VBS:

1
bResult = sc.Satellite(0).ReplaceICRFWithJ2000

bool STKEphemerisToOEM(string inputFile, string outputFile)

Use this to convert an STK Ephemeris *.e file to a CCSDS 502 OEM *.oem file. It will convert the file using the Scenario.SatEphemeris.CCSDS and Satellite.CCSDS controls. The output *.oem ephemeris will be in the central body default inertial frame regardless of the input *.e ephemeris coordinate frame.

Parameters

Return Value

Example VBS:

1
bResult = sc.Satellite(0). STKEphemerisToOEM ("C:\od\sat1.e", "C:\od\sat1.oem")

bool Stop()

Use this to stop the generation of satellite ephemeris. The ephemeris file will contain the ephemeris generated up to the time of receiving the Stop command.

Return Value

Example VBS:

1
bResult = sc.Satellite(0).Stop

CAgAttr UpdateCentralBody(string cbName, string propagatorComponentName)

This function changes a satellite's OrbitState central body, ensures that the satellite is using a propagator from the Component Browser, and sets the specific component propagator to use. It converts the satellite's initial state based on the Cartesian elements to an orbit around the new central body. When the choice is Moon or Sun, it also exposes the satellite's DynamicReports attributes and updates the DynamicReports.CentralBody value. You can access this function via the satellite's ReferenceTrajectory attribute scope.

Parameters

Return Value

The function returns a CAgAttr object containing the following properties:

Example JavaScript:

1
2
3
4
5
6
7
8
9
10
11
let cbName = "Moon";
let propagatorComponentName = "Moon HPOP Default v10";
let result = sc.Satellite(0).ReferenceTrajectory.UpdateCentralBody(cbName, propagatorComponentName);
if(result.Succeed.value())
{
    alert("Success: " + result.Messages.value());
}
else
{
    alert("Failed: " + result.Messages.value());
}

bool UpdateInitialConditions(string filename, string newEpoch, string sUnused, bool bPerturb, bool bUpdateCov, bool bUnused)

This computes a new set of initial conditions for a satellite based on the reference ephemeris defined in the given file. You can access this function via the satellite's ReferenceTrajectory attribute scope.

Parameters

Return Value

Example VBS:

1
2
Set sat = sc.GNSS.GNSSSatellite(0)
bResult = sat.ReferenceTrajectory.UpdateInitialConditions( "c:\temp\sat.e","1 Jul 2009 05:30:00 UTCG", "", true, true, false)

The function Satellite.ReferenceTrajectory.UpdateInitialConditions2() is available for ODTK application version 6.5.2 and later. It allows a passing satellite epoch to be in any valid date time format.