KeplerianElements.java
/* Copyright 2002-2022 CS GROUP
* Licensed to CS GROUP (CS) under one or more
* contributor license agreements. See the NOTICE file distributed with
* this work for additional information regarding copyright ownership.
* CS licenses this file to You under the Apache License, Version 2.0
* (the "License"); you may not use this file except in compliance with
* the License. You may obtain a copy of the License at
*
* http://www.apache.org/licenses/LICENSE-2.0
*
* Unless required by applicable law or agreed to in writing, software
* distributed under the License is distributed on an "AS IS" BASIS,
* WITHOUT WARRANTIES OR CONDITIONS OF ANY KIND, either express or implied.
* See the License for the specific language governing permissions and
* limitations under the License.
*/
package org.orekit.files.ccsds.ndm.odm;
import org.orekit.files.ccsds.section.CommentsContainer;
import org.orekit.files.ccsds.section.Data;
import org.orekit.frames.Frame;
import org.orekit.orbits.KeplerianOrbit;
import org.orekit.orbits.PositionAngle;
import org.orekit.time.AbsoluteDate;
/** Container for Keplerian elements.
* @author sports
* @since 6.1
*/
public class KeplerianElements extends CommentsContainer implements Data {
/** Epoch of state vector and optional Keplerian elements. */
private AbsoluteDate epoch;
/** Orbit semi-major axis (m). */
private double a;
/** Mean motion (the Keplerian Mean motion in rad/s).
* <p>
* Used in OMM instead of semi-major axis if MEAN_ELEMENT_THEORY = SGP/SGP4.
* </p>
*/
private double meanMotion;
/** Orbit eccentricity. */
private double e;
/** Orbit inclination (rad). */
private double i;
/** Orbit right ascension of ascending node (rad). */
private double raan;
/** Orbit argument of pericenter (rad). */
private double pa;
/** Orbit anomaly (rad). */
private double anomaly;
/** Orbit anomaly type (mean or true). */
private PositionAngle anomalyType;
/** Gravitational coefficient. */
private double mu;
/** Simple constructor.
*/
public KeplerianElements() {
a = Double.NaN;
meanMotion = Double.NaN;
e = Double.NaN;
i = Double.NaN;
raan = Double.NaN;
pa = Double.NaN;
anomaly = Double.NaN;
mu = Double.NaN;
}
/** {@inheritDoc}
* <p>
* We check neither semi-major axis nor mean motion here,
* they must be checked separately in OPM and OMM parsers
* </p>
*/
@Override
public void validate(final double version) {
super.validate(version);
checkNotNull(epoch, StateVectorKey.EPOCH);
checkNotNaN(e, KeplerianElementsKey.ECCENTRICITY);
checkNotNaN(i, KeplerianElementsKey.INCLINATION);
checkNotNaN(raan, KeplerianElementsKey.RA_OF_ASC_NODE);
checkNotNaN(pa, KeplerianElementsKey.ARG_OF_PERICENTER);
checkNotNaN(anomaly, KeplerianElementsKey.MEAN_ANOMALY);
}
/** Get epoch of state vector, Keplerian elements and covariance matrix data.
* @return epoch the epoch
*/
public AbsoluteDate getEpoch() {
return epoch;
}
/** Set epoch of state vector, Keplerian elements and covariance matrix data.
* @param epoch the epoch to be set
*/
public void setEpoch(final AbsoluteDate epoch) {
refuseFurtherComments();
this.epoch = epoch;
}
/** Get the orbit semi-major axis.
* @return the orbit semi-major axis
*/
public double getA() {
return a;
}
/** Set the orbit semi-major axis.
* @param a the semi-major axis to be set
*/
public void setA(final double a) {
refuseFurtherComments();
this.a = a;
}
/** Get the orbit mean motion.
* @return the orbit mean motion
*/
public double getMeanMotion() {
return meanMotion;
}
/** Set the orbit mean motion.
* @param motion the mean motion to be set
*/
public void setMeanMotion(final double motion) {
this.meanMotion = motion;
}
/** Get the orbit eccentricity.
* @return the orbit eccentricity
*/
public double getE() {
return e;
}
/** Set the orbit eccentricity.
* @param e the eccentricity to be set
*/
public void setE(final double e) {
refuseFurtherComments();
this.e = e;
}
/** Get the orbit inclination.
* @return the orbit inclination
*/
public double getI() {
return i;
}
/**Set the orbit inclination.
* @param i the inclination to be set
*/
public void setI(final double i) {
refuseFurtherComments();
this.i = i;
}
/** Get the orbit right ascension of ascending node.
* @return the orbit right ascension of ascending node
*/
public double getRaan() {
return raan;
}
/** Set the orbit right ascension of ascending node.
* @param raan the right ascension of ascending node to be set
*/
public void setRaan(final double raan) {
refuseFurtherComments();
this.raan = raan;
}
/** Get the orbit argument of pericenter.
* @return the orbit argument of pericenter
*/
public double getPa() {
return pa;
}
/** Set the orbit argument of pericenter.
* @param pa the argument of pericenter to be set
*/
public void setPa(final double pa) {
refuseFurtherComments();
this.pa = pa;
}
/** Get the orbit anomaly.
* @return the orbit anomaly
*/
public double getAnomaly() {
return anomaly;
}
/** Set the orbit anomaly.
* @param anomaly the anomaly to be set
*/
public void setAnomaly(final double anomaly) {
refuseFurtherComments();
this.anomaly = anomaly;
}
/** Get the type of anomaly (true or mean).
* @return the type of anomaly
*/
public PositionAngle getAnomalyType() {
return anomalyType;
}
/** Set the type of anomaly.
* @param anomalyType the type of anomaly to be set
*/
public void setAnomalyType(final PositionAngle anomalyType) {
refuseFurtherComments();
this.anomalyType = anomalyType;
}
/**
* Set the gravitational coefficient.
* @param mu the coefficient to be set
*/
public void setMu(final double mu) {
refuseFurtherComments();
this.mu = mu;
}
/**
* Get the gravitational coefficient.
* @return gravitational coefficient
*/
public double getMu() {
return mu;
}
/** Generate a keplerian orbit.
* @param frame inertial frame for orbit
* @return generated orbit
*/
public KeplerianOrbit generateKeplerianOrbit(final Frame frame) {
return new KeplerianOrbit(a, e, i, pa, raan, anomaly, anomalyType, frame, epoch, mu);
}
}