Spacecraft Navigation
Spacecraft Navigation — There are a number of Spacecraft fields that are used exclusively to support GMAT's navigation capability.
Description
When using GMAT's navigation capabilities, certain Spacecraft parameters can be "solvedfor." As discussed in the Spacecraft Ballistic/Mass Properties section, the Spacecraft ballistic and mass properties include the coefficient of reflectivity, Cr, and the coefficient of drag, Cd. As discussed in the Spacecraft Orbit State section, you can specify the CartesianState, i.e., the X, Y, Z position (km), and the Vx, Vy, Vz velocity (km/s) of a Spacecraft. As part of GMAT's navigation capability, GMAT can ingest measurements and estimate ("solvefor") values for Cr, Cd, and either the CartesianState, or KeplerianState.
See Also: BatchEstimatorInv
Fields
Field  Description  

AddHardware 
List of Antenna, Transmitter, Receiver, and Transponder objects attached to a Spacecraft


CdSigma 
Standard deviation of the coefficient of
reflectivity, Cd. This field is only used if
the UseInitialCovariance field of the
BatchEstimatorInv resource is set to


CrSigma 
Standard deviation of the coefficient of
reflectivity, Cr. This field is only used if
the UseInitialCovariance field of the
BatchEstimatorInv resource is set to


OrbitErrorCovariance 
State 6x6 error covariance matrix. If CartesianState is estimated, this must be a Cartesian covariance. If KeplerianState is estimated, this must be a Keplerian covariance. Regardless of choice of spacecraft coordinate system, the covariance must be specified in the EarthMJ2000Eq coordinate system. This
field is only used if the
UseInitialCovariance of the
BatchEstimatorInv resource is set to


SolveFors 
List of fields to be solved for. This list must at least include either CartesianState or KeplerianState (but not both). For example, Cr cannot be the only parameter solved for.

Remarks
When estimating CartesianState, the input OrbitErrorCovariance matrix must represent a Cartesian covariance, and when estimating KeplerianState the OrbitErrorCovariance must represent a Keplerian covariance. Note that Keplerian covariance input employs Mean Anomaly (MA) instead of True Anomaly. The current release of GMAT only supports input of Keplerian orbit elements using TA and does not permit explicitly setting an initial MA.
For more details, see the section called “UseInitialCovariance Restrictions” in the Batch Estimator resource.
Examples
Solve for Cr and the spacecraft Cartesian state.
Create Spacecraft Sat
Create BatchEstimatorInv bat
Sat.SolveFors = {CartesianState, Cr}
%User must create a TrackingFileSet
%and set up bat appropriately
BeginMissionSequence
RunEstimator bat
Solve for Cd and the spacecraft Cartesian state assuming that the a priori information is included in the estimation state vector.
Create Spacecraft Sat
Sat.SolveFors = {CartesianState, Cd}
Create BatchEstimatorInv bat
bat.UseInitialCovariance= True
%User must create a TrackingFileSet
%and set up bat appropriately
Create Array Initial_6x6_covariance[6,6]
BeginMissionSequence
Initial_6x6_covariance = ...
diag([1e6 1e70 1e70 1e70 1e70 1e70]) %X pos known very well
Sat.OrbitErrorCovariance = Initial_6x6_covariance
Sat.CrSigma = 1e6 %Cr known very well
RunEstimator bat