Description

Classical (Keplerian) coordinate type.

Public Properties

CoordinateSystem	The coordinate system and coordinate epoch.
CoordinateSystemType	Coordinate system being used.
Location	The location of the spacecraft in the orbit at epoch.
LocationType	The element used for specifying spacecraft location in the orbit at epoch.
Orientation	The orbit orientation.
SizeShape	The size and shape of the orbit.
SizeShapeType	The pair of elements used for specifying orbit size and shape.
StateEpoch	Smart epoch component allows the user to configure the state epoch explicitly or implicitly (using a pre-defined or custom time instant component).
SupportedCoordinateSystemTypes	Returns an array of supported coordinate system types

Interfaces

Implemented Interface
IAgOrbitState

CoClasses that Implement IAgOrbitStateClassical

Name
AgOrbitStateClassical

Example

Set initial state of satellite and propagate

[Python - STK API]

# IAgSatellite satellite: Satellite object
keplerian = satellite.Propagator.InitialState.Representation.ConvertTo(AgEOrbitStateType.eOrbitStateClassical)
keplerian.SizeShapeType = AgEClassicalSizeShape.eSizeShapeAltitude
keplerian.LocationType = AgEClassicalLocation.eLocationTrueAnomaly
keplerian.Orientation.AscNodeType = AgEOrientationAscNode.eAscNodeLAN

# Assign the perigee and apogee altitude values:
keplerian.SizeShape.PerigeeAltitude = 500      # km
keplerian.SizeShape.ApogeeAltitude = 600       # km

# Assign the other desired orbital parameters:
keplerian.Orientation.Inclination = 90         # deg
keplerian.Orientation.ArgOfPerigee = 12        # deg
keplerian.Orientation.AscNode.Value = 24       # deg
keplerian.Location.Value = 180                 # deg

# Apply the changes made to the satellite's state and propagate:
satellite.Propagator.InitialState.Representation.Assign(keplerian)
satellite.Propagator.Propagate()

Set initial state of satellite and propagate

[MATLAB]

% IAgSatellite satellite: Satellite object
keplerian = satellite.Propagator.InitialState.Representation.ConvertTo('eOrbitStateClassical'); % Use the Classical Element interface
keplerian.SizeShapeType = 'eSizeShapeAltitude';  % Changes from Ecc/Inc to Perigee/Apogee Altitude
keplerian.LocationType = 'eLocationTrueAnomaly'; % Makes sure True Anomaly is being used
keplerian.Orientation.AscNodeType = 'eAscNodeLAN'; % Use LAN instead of RAAN for data entry

% Assign the perigee and apogee altitude values:
keplerian.SizeShape.PerigeeAltitude = 500;      % km
keplerian.SizeShape.ApogeeAltitude = 600;       % km

% Assign the other desired orbital parameters:
keplerian.Orientation.Inclination = 90;         % deg
keplerian.Orientation.ArgOfPerigee = 12;        % deg
keplerian.Orientation.AscNode.Value = 24;       % deg
keplerian.Location.Value = 180;                 % deg

% Apply the changes made to the satellite's state and propagate:
satellite.Propagator.InitialState.Representation.Assign(keplerian);
satellite.Propagator.Propagate();

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IAgOrbitStateClassical Interface