Propagator

• For numerical Propagator documentation see Numerical Propagator

• For ForceModel documentation see Force Model

• For SPICE Propagator documentation see SPK-Configured Propagator

• For Code500 ephemeris Propagator documentation see Code500 Ephemeris-Configured Propagator

• For STK ephemeris Propagator documentation see STK Ephemeris-Configured Propagator

Overview

A Propagator object that uses a numerical integrator (as opposed to an ephemeris propagator) is one of a few objects in GMAT that is configured differently in the scripting and in the GUI. In the GUI, you configure the integrator and force model setting on the same dialog box. See the Remarks section below for detailed discussion of GMAT’s numerical integrators as well as performance and accuracy comparisons, and usage recommendations. This resource cannot be modified in the Mission Sequence. However, you can do whole object assignment in the mission,( i.e. myPropagator = yourPropagator ).

When working in the script, you must create a ForceModel object separately from the Propagator and specify the force model using the “FM” field on the propagator object. See the Examples section later in this section for details.

Options

GUI

Settings for the embedded Runge-Kutta integrators. Select the desired integrator from the Type menu.

The Adams-Bashforth-Moulton integrator has additional settings as shown.

Remarks

Best Practices for Using Numerical Integrators

The comparison data presented in a later section suggest that the PrinceDormand78 integrator is the best all purpose integrator in GMAT. When in doubt, use the PrinceDormance78 integrator, and set MinStep to zero so that the integrator’s adaptive step algorithm controls the minimum integration step size. Below are some important comments on GMAT’s step size control algorithms and the dangers of using a non-zero value for the minimum integration step size. The AdamsBashforthMoulton integrator is a low order integrator and we only recommend its use for low precision analysis when a predictor-corrector algorithm is required. We recommend that you study the performance and accuracy analysis documented later in this section to select a numerical integrator for your application. You may need to perform further analysis and comparisons for your application.

Numerical Integrators Overview

The table below describes each numerical integrator in detail.

Performance & Accuracy Comparison of Numerical Integrators

The tables below contain performance comparison data for GMAT's numerical integrators. The first table shows the orbit types, dynamics models, and propagation duration for each test case included in the comparison. Five orbit types were compared: low earth orbit, Molniya, Mars transfer (Type 2), Lunar transfer, and finite burn (case 1 is blow down, and case 2 is pressure regulated). For each test case, the orbit was propagated forward for a duration and then back-propagated to the intial epoch. The error values in the table are the RSS difference of the final position after forward and backward propagation to the initial position. The run time data for each orbit type is normalized on the integrator with the fasted run time for that orbit type. For all test cases the ErrorControl setting was set to RSSStep. Accuracy was set to 1e-12 for all integrators except for AdamsBashfourthMoulton which was set to 1e-11 because of poor performance when Accuracy was set to 1e-11.

Comparing the run time data for each integrator shown in the table below we see that the PrinceDormand78 integrator was the fastest for 4 of the 6 cases and tied with the RungeKutta89 integrator for LEO test case. For the Lunar flyby case, the RungeKutta89 was the fastest integrator, however, in this case the PrinceDormand78 integrator was at least 2 orders of magnitude more accurate given equaivalent Accuracy settings. Notice that the AdamsBashforthMoulton integrator has km level errors for some orbits because it is a low-order integrator.

Fields Unique to the AdamsBashforthMoulton Integrator

The AdamsBashforthMoulton integrator has two additional fields named TargetError and LowerError that are only active when Type is set to AdamsBashforthMoulton. If you are using another integrator type, those fields must be removed from your script file to avoid parsing errors. When working in the GUI, this is performed automatically. See examples below for more details.

Examples

            Create Spacecraft aSat
Create ForceModel aForceModel

Create Propagator aProp
aProp.FM              = aForceModel
aProp.Type            = PrinceDormand78
aProp.InitialStepSize = 60
aProp.Accuracy        = 1e-011
aProp.MinStep         = 0
aProp.MaxStep         = 86400
aProp.MaxStepAttempts = 50
aProp.StopIfAccuracyIsViolated = true

BeginMissionSequence

Propagate aProp(aSat) {aSat.ElapsedDays = .2}
          

            Create Spacecraft aSat
Create ForceModel aForceModel
aForceModel.ErrorControl = None

Create Propagator aProp
aProp.FM              = aForceModel
aProp.Type            = PrinceDormand78
aProp.InitialStepSize = 30

BeginMissionSequence

Propagate aProp(aSat) {aSat.ElapsedDays = .2}
          

            Create Spacecraft aSat
Create ForceModel aForceModel
aForceModel.ErrorControl = RSSStep

Create Propagator aProp
aProp.FM              = aForceModel
aProp.Type            = AdamsBashforthMoulton
aProp.InitialStepSize = 60
aProp.MinStep         = 0
aProp.MaxStep         = 86400
aProp.MaxStepAttempts = 50
%  Note the following fields must be set with decreasing values!
aProp.Accuracy        = 1e-010
aProp.TargetError     = 1e-011
aProp.LowerError      = 1e-013
aProp.StopIfAccuracyIsViolated = true

BeginMissionSequence

Propagate aProp(aSat) {aSat.ElapsedDays = .2}
          

Overview

A ForceModel is a model of the environmental forces and dynamics that affect the motion of a spacecraft. GMAT supports numerous force models such as point mass and spherical harmonic gravity models, atmospheric drag, solar radiation pressure, tide models, and relativistic corrections. A ForceModel is configured and attached to the Propagator object (see the Propagator object for differences between script and GUI configuration when configuring a Propagator). The Propagator, along with the Propagate command, uses a ForceModel to numerically solve the orbital equations of motion (forwards or backwards in time) using the forces configured in the ForceModel object, and may include thrust terms in the case of powered flight. See the discussion below for detailed information on how to configure force models for your application. This resource cannot be modified in the Mission Sequence.

See Also: Propagator

Fields

GUI

Settings for the ForceModel object.

Remarks

Overview of Primary Body/Central Body and Field Interactions

In GMAT, a primary body is a celestial body that is modeled with a complex force model which may include a spherical harmonic gravity model, tides, or drag. A body cannot appear in both the PrimaryBodies and PointMasses fields. GMAT currently requires that there are no more than one primary body per ForceModel, but this behavior will change in future versions and the user interface is designed to naturally support this future development area.

GMAT currently requires that the primary body is either the same as the CentralBody or set to None. If you change the CentralBody in the GUI, GMAT changes the primary body to None, and you can then select between None and the central body. When you select a primary body in the GUI, the Gravity and Drag fields activate and allow you to select models for those forces consistent with the body selected in the PrimaryBodies field. For example, if you select Earth as the primary body, you can only select Earth drag models in the Drag.AtmosphereModel field. See the field list above for available models.

Configuring Gravitational Models

GMAT supports point mass gravity, spherical harmonic, and tide modeling for all celestial bodies. On a Propagator, all celestial bodies are classified into two mutually exclusive categories: PrimaryBodies, and Point Masses. To model a body as a point mass, add it to the PointMasses list. GMAT currently requires that there be only a single body in the PrimaryBodies list. When a primary body is selected, the CentralBody and primary body must be the same.

Bodies modeled as PointMasses use the gravitational parameter defined on the body (i.e. Earth.Mu) in the equations of motion. Bodies defined as PrimaryBodies use the constants defined on the potential file in the equations of motion. GMAT supports four gravity file formats: the .cof format, the STK .grv format, the .gfc format, and the .tab format. You can provide a custom potential file for your application as long as it is one of the supported formats. Potential files defined in the startup file are available in the Model list in the GUI. For example, the following lines in the startup file configure GMAT so that EGM96 is an available option for Model in the GUI when the primary body is Earth:

EARTH_POT_PATH         = DATA_PATH/gravity/earth/
EGM96_FILE             = EARTH_POT_PATH/EGM96.cof 

Below is an example script snippet for configuring a custom gravity model.

Create ForceModel aForceModel
aForceModel.CentralBody = Earth
aForceModel.PrimaryBodies = {Earth}
aForceModel.GravityField.Earth.Degree = 21
aForceModel.GravityField.Earth.Order  = 21
aForceModel.GravityField.Earth.PotentialFile = 'c:\MyData\File.cof'

Overview of Tide Model Field Interactions

By default, the tide data source is set to None and the tide model selector is disabled if no tide model is selected. To use a tide model, first the tide data source must be changed to either Inherited or Tide File, at which point the Tide Model selector becomes enabled to select from the tide models supported by the tide data source. See the field list above for available models. The Inherited option indicates that the data for the tide model is provided either by the gravity potential file or the data is built into GMAT. The tide data contained in a gravity potential file has precedence over any built-in values. The Tide File option enables the file selector to choose a file containing the Love numbers to be used as the data source for the tide model. The tide data contained in a tide file has precedence over all other tide data sources.

Configuring Tide Models

GMAT supports solid tide modeling for all central bodies, and both solid and pole tide modeling for the Earth. Tide models can only be used if a PrimaryBody is set. GMAT contains built-in values for both solid and pole tides for the Earth. External files can also be used to provide the Love numbers to be used in the tide model, either from a gravity file that supports tides, or a separate tide file.

If a gravity file with Love numbers is provided, those Love numbers will be used for the solid tide model calculations. If a tide file is provided, the Love numbers in the tide file will be used. If both a gravity file with Love numbers and a tide file are provided, the Love numbers from both files will be used, with the Love numbers in the tide file having precedence over the gravity file. Only if no tide file is provided and the gravity potential file has no love numbers are GMAT's default Love numbers used for the Earth. GMAT's built-in values are the only data source for pole tides.

Below is an example script snippet for configuring a custom gravity model including Lunar solid tides.

Create ForceModel aForceModel
aForceModel.CentralBody = Luna
aForceModel.PrimaryBodies = {Luna}
aForceModel.GravityField.Luna.Degree = 21
aForceModel.GravityField.Luna.Order  = 21
aForceModel.GravityField.Luna.PotentialFile = 'c:\MyData\File.cof'
aForceModel.GravityField.Luna.TideFile = 'c:\MyData\File.tide'
aForceModel.GravityField.Luna.TideModel = 'Solid'

Tide files use the .tide file extension. You can provide a custom tide file for your application as long as it is in the supported format. Tide files contain the Love numbers to be used to model the solid tides. Tide files can include the k2, k3, and k+ coefficients. The format used by the tide file is 'k {degree} {order} {value}' or 'kplus {order} {value}' for k+.

Below is a sample tide file using the built-in values that GMAT uses for the Earth's Love numbers

k 2 0 0.30190
k 2 1 0.29830
k 2 2 0.30102
k 3 0 0.093
k 3 1 0.093
k 3 2 0.093
k 3 3 0.094
kplus 0 -0.00087
kplus 1 -0.00079
kplus 2 -0.00057

Zero Tide and Tide Free Models

The selection of a tide model is closely linked to the gravitational potential model that is used. Some gravitational potential models incorporate some tidal effects into the gravitational potential model. Two common ways gravitational models handle modeling tidal forces are by being tide-free and zero-tide. Tide free gravitational models contain no effects of tidal forces in the gravitational potential, while zero tide gravitational models contain the permanent (time-independent) effect of tides on the potential. For STK .grv files, the "IncludesPermTides" keyword is recognized to identify if the gravitational potential model includes permanent tide effects, however the coefficients in the "TideFreeValues" and "ZeroTideValues" keyword blocks are currently ignored.

Configuring Drag Models

GMAT supports many density models for Earth including Jacchia-Roberts and various MSISE models. Density models for non-Earth bodies -- the Mars-GRAM model for example -- are included using custom plug-in components and are currently only supported in the script interface.

To configure Earth density models, select Earth as the primary body, In the GUI, this activates the AtmosphereModel list. You can configure the solar flux values using the Setup button next to the AtmosphereModel list after you have selected an atmosphere model. Below is an example script snippet for configuring the NRLMSISE00 density model.

Create ForceModel aForceModel
GMAT aForceModel.PrimaryBodies = {Earth}
GMAT aForceModel.Drag.AtmosphereModel = NRLMSISE00

Note that when you select None for Drag.AtmosphereModel , the fields associated with density configuration, such as Drag.F107, Drag.F107A, and Drag.MagneticIndex and others are inactive and must be removed from your script file to avoid parsing errors. When working in the GUI, this is performed automatically.

The table below describes the limits on altitude for drag models supported by GMAT.

MarsGRAM2005

When PrimaryBody is Mars, you can choose Mars-GRAM 2005 as your atmosphere model. This model is only available when the libMarsGRAM plugin is available and enabled in the GMAT startup file.

When using the MarsGRAM2005 atmosphere model, three new fields are available in the script language (but not the GUI):

See the Fields section for details on these fields.

In addition, the space weather fields are treated as follows:

The Mars-GRAM 2005 input file is a text file in FORTRAN NAMELIST format. Most variables in this file are passed directly to the Mars-GRAM model and are used as intended. However, some are replaced internally by GMAT-supplied values. The following table lists those input variables that are handled specially.

The input file is read by the Mars-GRAM 2005 model code, which has limited error checking. If the input file or data files are incorrect or missing, GMAT may exhibit unintended behavior. Note that local winds returned by the Mars-GRAM 2005 model are not included in GMAT's drag model.

Configuring Space Weather Data for Density Models

GMAT supports several space weather input types for drag modelling including constant flux and Geo-magnetic index values, a historical weather data file, and a predicted weather data file. You can separately configure the data used for historical data and predicted data. For historical data you can choose between constant values and a CSSI space weather file. For predicted data you can choose between constant values and a Schatten predict file. Each of those sources is discussed in detail below.

The precedence for data source is determined by the simulation epoch (i.e. the epoch when density is evaluated), and the epochs contained on the data files

Constant Values

GMAT supports constant flux and Geo-magnetic index values for all Earth density models. You configure GMAT to use those values for historical and predicted data as shown below using NRLMSISE00 for the example.

Create ForceModel aForceModel
GMAT aForceModel.Drag.AtmosphereModel = NRLMSISE00
GMAT aForceModel.Drag.HistoricWeatherSource = 'ConstantFluxAndGeoMag'
GMAT aForceModel.Drag.PredictedWeatherSource = 'ConstantFluxAndGeoMag'
GMAT aForceModel.Drag.F107 = 150
GMAT aForceModel.Drag.F107A = 150
GMAT aForceModel.Drag.MagneticIndex = 3

Historical Space Weather Data

You can provide a Center for Space Standards and Innovation (CSSI) file for historical space weather data. GMAT does not use the predicted portion of the file but does use the historical portion of the data. The CCSI file format is described in detail at the Celestrak website and the files are available for download at that site and here. You configure GMAT to use historical data as shown below.

Create ForceModel aForceModel
GMAT aForceModel.Drag.AtmosphereModel = NRLMSISE00
GMAT aForceModel.Drag.HistoricWeatherSource = 'CSSISpaceWeatherFile'
GMAT aForceModel.Drag.CSSISpaceWeatherFile = 'CSSI_2004To2026.txt'

You can provide a full or relative path to the file, or put the file in GMAT’s data file folders documented in the startup file help.

Predicted Space Weather Data

You configure GMAT to use Schatten predicted data as shown below

Create ForceModel aForceModel
GMAT aForceModel.Drag.AtmosphereModel = NRLMSISE00
GMAT aForceModel.Drag.PredictedWeatherSource = 'SchattenFile'
GMAT aForceModel.Drag.SchattenFile = 'SchattenPredict.txt'
GMAT aForceModel.Drag.SchattenErrorModel = 'Nominal'
GMAT aForceModel.Drag.SchattenTimingModel = 'NominalCycle'

You can provide a full or relative path to the file, or put the file in GMAT’s data file folders documented in the startup file help. Additionally you can choose between Nominal, PlusTwoSigma, and MinusTwoSigma for the SchattenErrorModel, and between NominalCycle, EarlyCycle, and LateCycle for the SchattenTimingModel.

The Schatten file is distributed by the Flight Dynamics Facility (FDF) at Goddard Space Flight Center. You can apply for an account to obtain Schatten file updates at the FDF Forms Interface. Note that GMAT reads the raw file containing all permutation of mean, +2 sigma, and -2 sigma, and nominal, early and late solar cycles. The files from the FDF must be modified to include keywords that indicate when data starts and ends as shown below:

           NOMINAL TIMING      EARLY TIMING        LATE TIMING      
 mo. yr.  mean +2sig -2sig ap mean +2sig -2sig ap mean +2sig -2sig ap
BEGIN_DATA 
  2 2011    92  107   76    9  105  125   85   10   77   87   66    8
  3 2011    93  110   77    9  106  128   86   10   79   89   67    8
  4 2011    95  112   78    9  108  129   87   10   80   92   69    8
END_DATA

Data must be formatted according to FORMAT(I3,I5,I6,11I5), and no comments or blank lines can occur between the BEGIN_DATA and END_DATA keywords.

Configuring SRP Models

GMAT supports a spherical SRP model, and an SRP file for high fidelity SRP modelling. Both models use a dual cone model for central body shadowing of the spacecraft. See the Spacecraft Ballistic/Mass Properties documentation for configuring a SPAD file for a spacecraft. The script snippet below shows how to configure two ForceModels, one that use Spherical and on that uses a SPADFile.

% A spherical SRP model
Create ForceModel aForceModel_1
aForceModel_1.PrimaryBodies = {Earth}
aForceModel_1.SRP = On
aForceModel_1.SRP.Model = Spherical

% A SPAD SRP model
Create ForceModel aForceModel_2
aForceModel_2.PrimaryBodies = {Earth}
aForceModel_2.SRP = On
aForceModel_2.SRP.Model = SPADFile

You can define the solar flux using two approaches which are currently only supported in the script interface. One approach is to define the flux value using the SRP.Flux field and the value of an astronomical unit (in km) using the Nominal_Sun field as shown in the following example.

Create ForceModel aForceModel
aForceModel.PrimaryBodies = {Earth}
aForceModel.SRP = On
aForceModel.SRP.Flux = 1367
aForceModel.SRP.Nominal_Sun = 149597870.691

An alternative approach is to define the flux pressure at 1 astronomical unit using the Flux_Pressure field as shown below..

Create ForceModel aForceModel
aForceModel.PrimaryBodies = {Earth}
aForceModel.SRP = On
aForceModel.SRP.Flux_Pressure = 4.53443218374393e-006
aForceModel.SRP.Nominal_Sun = 149597870.691

If you mix flux settings, as shown in the example below, GMAT will use the last approach in the script. Here, GMAT will use the Flux_Pressure setting.

Create ForceModel aForceModel
aForceModel.PrimaryBodies = {Earth}
aForceModel.SRP = On
aForceModel.SRP.Flux = 1370
aForceModel.SRP.Nominal_Sun = 149597870
aForceModel.SRP.Flux_Pressure = 4.53443218374393e-006

Variational Equations and the STM

GMAT can optionally propagate the orbit State Transition Matrix (STM). For more information on how to configure GMAT to compute the STM, see the Propagate command documentation.

Examples

            Create Spacecraft aSat

Create ForceModel aForceModel
aForceModel.CentralBody = Earth
aForceModel.PointMasses = {Earth}

Create Propagator aProp
aProp.FM = aForceModel

BeginMissionSequence

Propagate aProp(aSat) {aSat.ElapsedDays = .2}
          

            Create Spacecraft aSat

Create ForceModel aForceModel
aForceModel.CentralBody = Earth
aForceModel.PrimaryBodies = {Earth}
aForceModel.PointMasses = {Sun, Luna}
aForceModel.SRP = On
aForceModel.RelativisticCorrection = On
aForceModel.ErrorControl = RSSStep
aForceModel.GravityField.Earth.Degree = 20
aForceModel.GravityField.Earth.Order = 20
aForceModel.GravityField.Earth.PotentialFile = 'EGM96.cof'
aForceModel.GravityField.Earth.TideModel = 'None'
aForceModel.Drag.AtmosphereModel = MSISE90
aForceModel.Drag.F107 = 150
aForceModel.Drag.F107A = 150
aForceModel.Drag.MagneticIndex = 3
aForceModel.SRP.Flux = 1359.388569998901
aForceModel.SRP.SRPModel = Spherical;
aForceModel.SRP.Nominal_Sun = 149597870.691

Create Propagator aProp
aProp.FM = aForceModel

BeginMissionSequence

Propagate aProp(aSat){aSat.ElapsedDays = .2}
          

            Create Spacecraft aSpacecraft;
aSpacecraft.DryMass   = 2000
aSpacecraft.SPADSRPFile = '..\data\vehicle\spad\SphericalModel.spo'
aSpacecraft.SPADSRPScaleFactor = 1;

Create ForceModel aFM;
aFM.SRP = On;
aFM.SRP.SRPModel = SPADFile

Create Propagator aProp;
aProp.FM = aFM;

BeginMissionSequence

Propagate aProp(aSpacecraft) {aSpacecraft.ElapsedDays = 0.2}
          

            Create Spacecraft moonSat
GMAT moonSat.DateFormat = UTCGregorian
GMAT moonSat.Epoch.UTCGregorian = 01 Jun 2004 12:00:00.000
GMAT moonSat.CoordinateSystem = MoonMJ2000Eq
GMAT moonSat.DisplayStateType = Cartesian
GMAT moonSat.X = -1486.792117191545200
GMAT moonSat.Y = 0.0
GMAT moonSat.Z = 1486.792117191543000
GMAT moonSat.VX = -0.142927729144255
GMAT moonSat.VY = -1.631407624437537
GMAT moonSat.VZ = 0.142927729144255

Create CoordinateSystem MoonMJ2000Eq
MoonMJ2000Eq.Origin = Luna
MoonMJ2000Eq.Axes   = MJ2000Eq

Create ForceModel MoonLP165P
GMAT MoonLP165P.CentralBody = Luna
GMAT MoonLP165P.PrimaryBodies = {Luna}
GMAT MoonLP165P.SRP = On
GMAT MoonLP165P.SRP.Flux = 1367
GMAT MoonLP165P.SRP.Nominal_Sun = 149597870.691
GMAT MoonLP165P.Gravity.Luna.PotentialFile = ../data/gravity/luna/LP165P.cof
GMAT MoonLP165P.Gravity.Luna.Degree = 20
GMAT MoonLP165P.Gravity.Luna.Order = 20

Create Propagator RKV89
GMAT RKV89.FM = MoonLP165P

BeginMissionSequence

Propagate RKV89(moonSat) {moonSat.ElapsedSecs = 300}
          

Description

An SPK-configured Propagator propagates a spacecraft by interpolating user-provided SPICE kernels. You configure a Propagator to use an SPK kernel by setting the Type field to SPK. SPK kernels and the NAIFId are defined on the Spacecraft Resource. You control propagation, including stopping conditions, using the Propagate command. This resource cannot be modified in the Mission Sequence. However, you can do whole object assignment in the mission,( i.e. myPropagator = yourPropagator ).

See Also: Spacecraft, Propagate

Fields

GUI

To configure a Propagator to use SPK files, on the Propagator dialog box, select SPK in the Type menu. There are four fields you can configure for an SPK propagator including StepSize, CentralBody, EpochFormat, and StartEpoch. Note that changing the EpochFormat setting converts the input epoch to the selected format. You can also type FromSpacecraft into the StartEpoch field and the Propagator will use the epoch of the Spacecraft as the initial propagation epoch.

Remarks

To use an SPK-configured Propagator, you must specify the SPK kernels and NAIFId on the Spacecraft, configure a Propagator to use SPK files as opposed to numerical methods, and configure the Propagate command to use the configured SPK propagator. The subsections and examples below discuss each of these items in detail.

Configuring Spacecraft SPK Kernels

To use an SPK-configured Propagator, you must add the SPK kernels to the Spacecraft and define the spacecraft's NAIFId. SPK Kernels for selected spacecraft are available here. Two sample vehicle spk kernels, (GEOSat.bsp and MoonTransfer.bsp) are distributed with GMAT for example purposes. An example of how to add spacecraft kernels via the script interface is shown below.

            Create Spacecraft aSpacecraft
GMAT aSpacecraft.NAIFId = -123456789
GMAT aSpacecraft.OrbitSpiceKernelName = {...
                                    '..\data\vehicle\ephem\spk\GEOSat.bsp'}
          

To add Spacecraft SPK kernels via the GUI:

1. On the Spacecraft dialog box, click the SPICE tab.

2. Under the SPK Files list, click Add.

3. Browse to locate and select the desired SPK file

4. Repeat to add all necessary SPK kernels

5. In the NAIF ID field, enter the spacecraft integer NAIF id number. Note: For a given mission, each spacecraft should have a unique NAIF ID if the spacecraft are propagated with an SPK propagator.

You can add more than one kernel to a spacecraft as shown via scripting below, where the files GEOSat1.bsp and GEOSat2.bsp are dummy file names used for example purposes only and are not distributed with GMAT. In the script, you can use relative path or absolute path to define the location of an SPK file. Relative paths are defined with respect to the GMAT bin directory of your local installation.

            Create Spacecraft aSpacecraft
aSpacecraft.OrbitSpiceKernelName ={'C:\MyDataFiles\GEOSat1.bsp',...
                                   'C:\MyDataFiles\GEOSat2.bsp'}
          

            Create Propagator spkProp
spkProp.EpochFormat = 'UTCGregorian'
spkProp.StartEpoch = '22 Jul 2014 11:29:10.811'

Create Propagator spkProp2
spkProp2.EpochFormat = 'TAIModJulian'
spkProp2.StartEpoch = '23466.5'
          

            Create Propagator spkProp
spkProp.Type = SPK
spkProp.StepSize = 300

          

            Propagate spkProp(aSpacecraft) {aSpacecraft.Earth.Periapsis}
          

            Propagate BackProp spkProp(aSpacecraft) {aSpacecraft.ElapsedDays = -1.5}
          

Examples

            Create Spacecraft aSpacecraft;
aSpacecraft.Epoch.UTCGregorian = '02 Jun 2004 12:00:00.000'
aSpacecraft.NAIFId = -123456789
aSpacecraft.OrbitSpiceKernelName = {'..\data\vehicle\ephem\spk\GEOSat.bsp'}

Create Propagator spkProp
spkProp.Type = SPK
spkProp.StepSize = 300
spkProp.CentralBody = Earth
spkProp.StartEpoch = FromSpacecraft

Create OrbitView EarthView
EarthView.Add = {aSpacecraft, Earth, Luna}
EarthView.ViewPointVector = [ 30000 -20000 10000 ]
EarthView.ViewScaleFactor = 2.5

BeginMissionSequence
Propagate spkProp(aSpacecraft) {aSpacecraft.TA = 90}
Propagate spkProp(aSpacecraft) {aSpacecraft.ElapsedDays = 2.4}
          

            Create Spacecraft aSpacecraft
aSpacecraft.NAIFId = -123456789
aSpacecraft.OrbitSpiceKernelName = {...
                          '..\data\vehicle\ephem\spk\MoonTransfer.bsp'}

Create Propagator spkProp
spkProp.Type = SPK
spkProp.StepSize = 300
spkProp.CentralBody = Earth
spkProp.EpochFormat = 'UTCGregorian'
spkProp.StartEpoch = '22 Jul 2014 11:29:10.811'

Create OrbitView EarthView
EarthView.Add = {aSpacecraft, Earth, Luna}
EarthView.ViewPointVector = [ 30000 -20000 10000 ]
EarthView.ViewScaleFactor = 30

BeginMissionSequence
Propagate spkProp(aSpacecraft) {aSpacecraft.ElapsedDays = 12}
          

Description

A Code500 ephemeris-configured Propagator propagates a spacecraft by interpolating or stepping along a user-provided Code500-format binary ephemeris file. You configure a Propagator to use a Code500 ephemeris by setting the Type field to Code500. The Code500 ephemeris file is specified on the Spacecraft.EphemerisName resource. The user controls propagation, including stopping conditions, using the Propagate command. This resource cannot be modified in the Mission Sequence. However, you can do whole object assignment in the mission sequence, (i.e. myPropagator = yourPropagator ).

The Propagator CentralBody option is not applicable to the Code500 propagator and should not be used with the Code500 propagator type. GMAT will automatically detect and use the central body of the ephemeris file. The Propagate command should be used to traverse the ephemeris file. GMAT will throw an error message and terminate when attempting to propagate outside the bounds of the ephemeris file.

Code500 ephemeris files are binary-format files. As discussed in the EphemerisFile help, GMAT can generate Code500 ephemeris files in both little-endian and big-endian binary format (via EphemerisFile.OutputFormat). The ephemeris propagator can read Code500 ephemeris files in either endian format. The endian format of the ephemeris file will be automatically detected by GMAT.

See Also: Spacecraft, Propagate, EphemerisFile

Fields

The only Propagator fields applicable to the Code500 ephemeris propagator are EpochFormat, StartEpoch, StepSize and Type.

GUI

To configure a Propagator from the GMAT GUI to use Code500 ephemeris files, select and open a Propagator from the Resources tree. In the Integrator category select Code500 from the Type drop-down box. This will display the Code500 propagator options dialog. There are four fields displayed for a Code500 propagator - StepSize, CentralBody, EpochFormat, and StartEpoch. Note that changing the EpochFormat setting converts the input epoch to the selected format. You can also type FromSpacecraft into the StartEpoch field and the Propagator will use the epoch of the Spacecraft as the initial propagation epoch. The CentralBody field is displayed to the user, but is unused when the integrator type is Code500.

Remarks

There is currently no GUI option to assign the Code500 ephemeris file to the Spacecraft resource. You must specify the Code500 ephemeris file on the Spacecraft.EphemerisName parameter via script. The subsections below provide examples of how to do this.

            Create Spacecraft aSpacecraft

aSpacecraft.EphemerisName = '../data/vehicle/ephem/code500/sat_leo.ephem'

BeginMissionSequence
          

            Create Propagator Code500Prop

Code500Prop.Type     = 'Code500'
Code500Prop.StepSize = 60.

BeginMissionSequence
          

Examples

            Create Spacecraft aSpacecraft

% Ephem file span is 4/20/2015 - 4/30/2015

aSpacecraft.EphemerisName = '../data/vehicle/ephem/code500/sat_leo.ephem'
aSpacecraft.DateFormat    = UTCGregorian
aSpacecraft.Epoch         = '22 Apr 2015 00:00:00.000'

Create Propagator Code500Prop

Code500Prop.Type       = 'Code500'
Code500Prop.StepSize   = 60.
Code500Prop.StartEpoch = 'FromSpacecraft'

Create ReportFile PropReport

PropReport.Filename     = 'EphemPropagator_Code500_ForwardProp.txt'
PropReport.WriteHeaders = True

BeginMissionSequence

While aSpacecraft.ElapsedDays <= 1

    Propagate Code500Prop(aSpacecraft)

    Report PropReport aSpacecraft.UTCGregorian aSpacecraft.TAIModJulian ...
        aSpacecraft.X aSpacecraft.Y aSpacecraft.Z ...
        aSpacecraft.VX aSpacecraft.VY aSpacecraft.VZ

EndWhile
          

An additional, more detailed, example of use of the Code500 ephemeris propagator is shown in the Ex_Code500_EphemerisCompare.script file provided in the samples\Navigation directory. This script generates a report showing the difference, in RIC coordinates, between the orbits in two different Earth-centered Code500 ephemeris files.

Description

An STK ephemeris-configured Propagator propagates a spacecraft by interpolating or stepping along a user-provided STK-format binary ephemeris file. You configure a Propagator to use an STK ephemeris by setting the Type field to STK. The STK ephemeris file is specified on a Spacecraft resource using the Spacecraft.EphemerisName field. The user controls propagation, including stopping conditions, using the Propagate command. This resource cannot be modified in the Mission Sequence. However, you can do whole object assignment in the mission sequence, (i.e. myPropagator = yourPropagator ).

The Propagator CentralBody option is not applicable to the STK propagator and should not be used with the STK propagator type. GMAT will automatically detect and use the central body of the ephemeris file. The Propagate command should be used to traverse the ephemeris file. GMAT will throw an error message and terminate when attempting to propagate outside the bounds of the ephemeris file. The STK propagator includes code that steps the spacecraft to the ephem boundary before stepping out of the span of the ephem

STK ephemeris files are ASCII files conforming to the Satellite Tool Kit TimePosVel specifications. As discussed in the EphemerisFile help, GMAT can generate STK ephemeris files using the EphemerisFile.OutputFormat field. The STK propagator works with STK formatted files, starting with STK 4.0, or GMAT STK ephemerides.

See Also: Spacecraft, Propagate, EphemerisFile

Fields

The only Propagator fields applicable to the STK ephemeris propagator are EpochFormat, StartEpoch, StepSize and Type.

GUI

To configure a Propagator from the GMAT GUI to use STK ephemeris files, select and open a Propagator from the Resources tree. In the Integrator category select STK from the Type drop-down box. This will display the STK propagator options dialog. There are four fields displayed for an STK propagator that affect propagation - StepSize, CentralBody, EpochFormat, and StartEpoch. Note that changing the EpochFormat setting converts the input epoch to the selected format. You can also type FromSpacecraft into the StartEpoch field and the Propagator will use the epoch of the Spacecraft as the initial propagation epoch. The CentralBody field is displayed to the user, but is unused when the integrator type is STK.

Implementation Notes

Position interpolation by default is performed using a 7th order Hermite-Newton divided difference interpolator using function value only data (i.e. position data for the position interpolation). Velocity interpolation uses the derivative of the position polynomial to produce interpolated values.

For segmented ephemerides, the interpolator restarts at segment boundaries. If an ephemeris segment has fewer than 8 points, the interpolator activates derivative information for the position interpolation by including the velocity data for the interpolation. The order of the interpolator changes to match the number of points in the segment (for n data points, order = 2n - 1 for position and n - 1 for velocity). The first time this happens, a warning notice is posted to the message window indicating the order of the velocity interpolation. Subsequent changes are not reported, but the interpolation order will adapt to the number of points in subsequent segments.

Propagating with stopping conditions can show sub-millisecond related differences.

STK ephemeris files can set an ephemeris epoch that is very different from the data in the file by setting a distant in time Scenario Epoch, and compensating using the time offset for each ephemeris point in the file. This can lead to round-off issues in propagation, particularly when propagating to the end of an ephemeris (or back propagating to the start).

Remarks

There is currently no GUI option to assign the STK ephemeris file to the Spacecraft resource. You must specify the STK ephemeris file on the Spacecraft.EphemerisName parameter via script. The subsections below provide examples of how to do this.

            Create Spacecraft aSpacecraft;

aSpacecraft.EphemerisName = '../data/vehicle/ephem/stk/SampleSTKEphem.e';

BeginMissionSequence;
          

            Create Propagator STKProp;

STKProp.Type     = 'STK';
STKProp.StepSize = 60;

BeginMissionSequence;
          

Examples

            
%
%   Spacecraft
% 
Create Spacecraft STKSat;

GMAT STKSat.DateFormat = UTCGregorian;
GMAT STKSat.Epoch = '01 Jan 2000 12:00:00.000';
GMAT STKSat.EphemerisName = '../data/vehicle/ephem/stk/SampleSTKEphem.e';

%
%   Propagator
%

Create Propagator STKProp;

GMAT STKProp.Type = STK;
GMAT STKProp.StepSize = 60;
GMAT STKProp.CentralBody = Earth;
GMAT STKProp.EpochFormat = 'A1ModJulian';
GMAT STKProp.StartEpoch = 'FromSpacecraft';

%
%   Output 
%

Create OrbitView OrbitView1;
GMAT OrbitView1.SolverIterations = Current;
GMAT OrbitView1.UpperLeft = [ 0 0 ];
GMAT OrbitView1.Size = [ 0 0 ];
GMAT OrbitView1.RelativeZOrder = 0;
GMAT OrbitView1.Maximized = false;
GMAT OrbitView1.Add = {STKSat, Earth};


%----------------------------------------
%---------- Arrays, Variables, Strings
%----------------------------------------

Create Array initialState[6,1] finalState[6,1];

%
% Miscellaneous variables.
%

Create String initialEpoch finalEpoch;

%
%   Mission Sequence
%

BeginMissionSequence;

GMAT [initialEpoch, initialState, finalEpoch, finalState] = ...
   GetEphemStates('STK', STKSat, 'UTCGregorian', EarthMJ2000Eq);

GMAT STKSat.Epoch = initialEpoch;

While STKSat.ElapsedDays <= 1
      
   Propagate STKProp(STKSat);

EndWhile;
          

This example is provided in the samples directory, in the Ex_2017a_STKEphemPropagation script.