GNSSOrbitalElements.java
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*
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* Unless required by applicable law or agreed to in writing, software
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package org.orekit.propagation.analytical.gnss.data;
import org.orekit.annotation.DefaultDataContext;
import org.orekit.attitudes.AttitudeProvider;
import org.orekit.data.DataContext;
import org.orekit.frames.Frame;
import org.orekit.frames.Frames;
import org.orekit.propagation.analytical.gnss.GNSSPropagator;
import org.orekit.propagation.analytical.gnss.GNSSPropagatorBuilder;
import org.orekit.time.TimeStamped;
/** This interface provides the minimal set of orbital elements needed by the {@link GNSSPropagator}.
*
* @author Pascal Parraud
*
*/
public interface GNSSOrbitalElements extends TimeStamped {
/**
* Gets the PRN number of the GNSS satellite.
*
* @return the PRN number of the GNSS satellite
*/
int getPRN();
/**
* Gets the Reference Week of the GNSS orbit.
*
* @return the Reference Week of the GNSS orbit within [0, 1024[
*/
int getWeek();
/**
* Gets the Reference Time of the GNSS orbit as a duration from week start.
*
* @return the Reference Time of the GNSS orbit (s)
*/
double getTime();
/**
* Gets the Semi-Major Axis.
*
* @return the Semi-Major Axis (m)
*/
double getSma();
/**
* Gets the Mean Motion.
*
* @return the Mean Motion (rad/s)
*/
double getMeanMotion();
/**
* Gets the Eccentricity.
*
* @return the Eccentricity
*/
double getE();
/**
* Gets the Inclination Angle at Reference Time.
*
* @return the Inclination Angle at Reference Time (rad)
*/
double getI0();
/**
* Gets the Rate of Inclination Angle.
*
* @return the Rate of Inclination Angle (rad/s)
*/
double getIDot();
/**
* Gets the Longitude of Ascending Node of Orbit Plane at Weekly Epoch.
*
* @return the Longitude of Ascending Node of Orbit Plane at Weekly Epoch (rad)
*/
double getOmega0();
/**
* Gets the Rate of Right Ascension.
*
* @return the Rate of Right Ascension (rad/s)
*/
double getOmegaDot();
/**
* Gets the Argument of Perigee.
*
* @return the Argument of Perigee (rad)
*/
double getPa();
/**
* Gets the Mean Anomaly at Reference Time.
*
* @return the Mean Anomaly at Reference Time (rad)
*/
double getM0();
/**
* Gets the Amplitude of the Cosine Harmonic Correction Term to the Argument of Latitude.
*
* @return the Amplitude of the Cosine Harmonic Correction Term to the Argument of Latitude (rad)
*/
double getCuc();
/**
* Gets the Amplitude of the Sine Harmonic Correction Term to the Argument of Latitude.
*
* @return the Amplitude of the Sine Harmonic Correction Term to the Argument of Latitude (rad)
*/
double getCus();
/**
* Gets the Amplitude of the Cosine Harmonic Correction Term to the Orbit Radius.
*
* @return the Amplitude of the Cosine Harmonic Correction Term to the Orbit Radius (m)
*/
double getCrc();
/**
* Gets the Amplitude of the Sine Harmonic Correction Term to the Orbit Radius.
*
* @return the Amplitude of the Sine Harmonic Correction Term to the Orbit Radius (m)
*/
double getCrs();
/**
* Gets the Amplitude of the Cosine Harmonic Correction Term to the Angle of Inclination.
*
* @return the Amplitude of the Cosine Harmonic Correction Term to the Angle of Inclination (rad)
*/
double getCic();
/**
* Gets the Amplitude of the Sine Harmonic Correction Term to the Angle of Inclination.
*
* @return the Amplitude of the Sine Harmonic Correction Term to the Angle of Inclination (rad)
*/
double getCis();
/**
* Gets the Earth's universal gravitational parameter.
*
* @return the Earth's universal gravitational parameter
*/
double getMu();
/**
* Gets the mean angular velocity of the Earth of the GNSS model.
*
* @return the mean angular velocity of the Earth of the GNSS model
*/
double getAngularVelocity();
/**
* Gets the duration of the GNSS cycle in seconds.
*
* @return the duration of the GNSS cycle in seconds
*/
double getCycleDuration();
/**
* Get the propagator corresponding to the navigation message.
* <p>
* The attitude provider is set by default to be aligned with the EME2000 frame.<br>
* The mass is set by default to the
* {@link org.orekit.propagation.Propagator#DEFAULT_MASS DEFAULT_MASS}.<br>
* The ECI frame is set by default to the
* {@link org.orekit.frames.Predefined#EME2000 EME2000 frame} in the default data
* context.<br>
* The ECEF frame is set by default to the
* {@link org.orekit.frames.Predefined#ITRF_CIO_CONV_2010_SIMPLE_EOP
* CIO/2010-based ITRF simple EOP} in the default data context.
* </p><p>
* This constructor uses the {@link DataContext#getDefault() default data context}
* </p>
* @return the propagator corresponding to the navigation message
* @see #getPropagator(Frames)
* @see #getPropagator(Frames, AttitudeProvider, Frame, Frame, double)
* @since 12.0
*/
@DefaultDataContext
default GNSSPropagator getPropagator() {
return new GNSSPropagatorBuilder(this).build();
}
/**
* Get the propagator corresponding to the navigation message.
* <p>
* The attitude provider is set by default to be aligned with the EME2000 frame.<br>
* The mass is set by default to the
* {@link org.orekit.propagation.Propagator#DEFAULT_MASS DEFAULT_MASS}.<br>
* The ECI frame is set by default to the
* {@link org.orekit.frames.Predefined#EME2000 EME2000 frame} in the default data
* context.<br>
* The ECEF frame is set by default to the
* {@link org.orekit.frames.Predefined#ITRF_CIO_CONV_2010_SIMPLE_EOP
* CIO/2010-based ITRF simple EOP} in the default data context.
* </p>
* @param frames set of frames to use
* @return the propagator corresponding to the navigation message
* @see #getPropagator()
* @see #getPropagator(Frames, AttitudeProvider, Frame, Frame, double)
* @since 12.0
*/
default GNSSPropagator getPropagator(final Frames frames) {
return new GNSSPropagatorBuilder(this, frames).build();
}
/**
* Get the propagator corresponding to the navigation message.
* @param frames set of frames to use
* @param provider attitude provider
* @param inertial inertial frame, use to provide the propagated orbit
* @param bodyFixed body fixed frame, corresponding to the navigation message
* @param mass spacecraft mass in kg
* @return the propagator corresponding to the navigation message
* @see #getPropagator()
* @see #getPropagator(Frames)
* @since 12.0
*/
default GNSSPropagator getPropagator(final Frames frames, final AttitudeProvider provider,
final Frame inertial, final Frame bodyFixed, final double mass) {
return new GNSSPropagatorBuilder(this, frames).attitudeProvider(provider)
.eci(inertial)
.ecef(bodyFixed)
.mass(mass)
.build();
}
}