User Manual 3.3 Analytical propagation - Historique des versions

Admin : /* Scope */

2018-03-01T13:28:43Z

Scope

← Version précédente		Version du 1 mars 2018 à 13:28
Ligne 10 :		Ligne 10 :
	* Propagation mode selection		* Propagation mode selection
	* Propagation frame selection		* Propagation frame selection
	More information can be found [~~ORB_ANA_Home~~ here].		More information can be found [[User Manual 3.3 Analytical propagation\|here]].

	=== Javadoc ===		=== Javadoc ===

Admin : /* Precomputed ephemeris */

2018-03-01T13:06:42Z

Precomputed ephemeris

← Version précédente		Version du 1 mars 2018 à 13:06
Ligne 295 :		Ligne 295 :
	*'''Integrated Ephemeris''' : With the ephemeris generation mode, a bounded propagator is available at the end of the numerical propagation and gives access to the orbit state at any date inside the propagation time interval. These intermediate states are computed thanks to the object IntegratedEphemeris. This object stores for each step the interpolator used, during the integration.		*'''Integrated Ephemeris''' : With the ephemeris generation mode, a bounded propagator is available at the end of the numerical propagation and gives access to the orbit state at any date inside the propagation time interval. These intermediate states are computed thanks to the object IntegratedEphemeris. This object stores for each step the interpolator used, during the integration.
	*'''Ephemeris''' : Another class is available for the same purpose (access to the orbit state at any given date): the Ephemeris class. An instance is built from a list of spacecraft states - that is to say, discrete ephemeris. Here, the intermediate states (keplerian, cartesian, ... parameters, depending on the kind of Orbit used) are computed with Hermite interpolators, which provides less accuracy than with the IntegratedEphemeris object. But for most uses, notably event detection, the Ephemeris class appears reliable enough.		*'''Ephemeris''' : Another class is available for the same purpose (access to the orbit state at any given date): the Ephemeris class. An instance is built from a list of spacecraft states - that is to say, discrete ephemeris. Here, the intermediate states (keplerian, cartesian, ... parameters, depending on the kind of Orbit used) are computed with Hermite interpolators, which provides less accuracy than with the IntegratedEphemeris object. But for most uses, notably event detection, the Ephemeris class appears reliable enough.
	*'''Lagrange Ephemeris''' and '''Hermite Ephemeris''' : These class are deprecated since 3.1. They were extending AbstractPropagator and they were computing propagation using Lagrange or Hermite interpolation between spacecraftstates. These possibilities has been replaced by the use of a simple propagator based on a PVCoordinateProvider which compute position, velocity with a Lagrange or Hermite interpolation between position, velocity ephemeris. This way of propagation is described in [~~ORB_Ephemeris_Home~~ Ephemeris] page.		*'''Lagrange Ephemeris''' and '''Hermite Ephemeris''' : These class are deprecated since 3.1. They were extending AbstractPropagator and they were computing propagation using Lagrange or Hermite interpolation between spacecraftstates. These possibilities has been replaced by the use of a simple propagator based on a PVCoordinateProvider which compute position, velocity with a Lagrange or Hermite interpolation between position, velocity ephemeris. This way of propagation is described in [[User Manual 3.3 Ephemeris\|Ephemeris]] page.

	=== Covariance matrix propagation ===		=== Covariance matrix propagation ===

Admin : /* Scope */

2018-03-01T13:05:48Z

Scope

← Version précédente		Version du 1 mars 2018 à 13:05
Ligne 4 :		Ligne 4 :
	This section describes the analytical propagators (Keplerian, Eckstein-Hechler, 2D...) provided by the Orekit library.		This section describes the analytical propagators (Keplerian, Eckstein-Hechler, 2D...) provided by the Orekit library.
	It also explains some interpolation algorithms, like the Hermite or Lagrange interpolation for ephemeris and the covariance matrix interpolation.		It also explains some interpolation algorithms, like the Hermite or Lagrange interpolation for ephemeris and the covariance matrix interpolation.
	Numerical propagators are détails in [~~ORB_ANA_Home~~ Numerical propagation].		Numerical propagators are détails in [[User Manual 3.3 Analytical propagation\|Numerical propagation]].

	This section is specific to analytical propagators. However all propagators inherit the same interface <code>Propagator</code> and provide the same services:		This section is specific to analytical propagators. However all propagators inherit the same interface <code>Propagator</code> and provide the same services:

Admin : Admin a déplacé la page Analytical propagation vers User Manual 3.3 Analytical propagation sans laisser de redirection

2018-02-28T10:45:56Z

Admin a déplacé la page Analytical propagation vers User Manual 3.3 Analytical propagation sans laisser de redirection

← Version précédente	Version du 28 février 2018 à 10:45
(Aucune différence)

Admin le 28 février 2018 à 08:39

2018-02-28T08:39:53Z

← Version précédente		Version du 28 février 2018 à 08:39
Ligne 425 :		Ligne 425 :
	\|[{{JavaDoc3.3}}/fr/cnes/sirius/patrius/propagation/PVCoordinatePropagator.html ...]		\|[{{JavaDoc3.3}}/fr/cnes/sirius/patrius/propagation/PVCoordinatePropagator.html ...]
	\|}		\|}

			[[Category:User_Manual_3.3_Orbit_Propagation]]

Admin le 28 février 2018 à 08:39

2018-02-28T08:39:06Z

← Version précédente		Version du 28 février 2018 à 08:39
Ligne 1 :		Ligne 1 :
			__NOTOC__

	== Introduction ==		== Introduction ==
	=== Scope ===		=== Scope ===

Admin : /* Contents */

2018-02-28T08:37:10Z

Contents

Voir les modifications

Admin : /* The trigonometric polynomial part */

2018-02-28T08:27:48Z

The trigonometric polynomial part

← Version précédente		Version du 28 février 2018 à 08:27
Ligne 257 :		Ligne 257 :
	The right ascension of ascending node and its mean longitude are connected as per the following equation :		The right ascension of ascending node and its mean longitude are connected as per the following equation :

	<center>		<center><math>l_{NAC} = \Omega- \theta</math></center>
	<math>\l_{NAC} = \Omega- \theta</math></center>

	where <math>\theta</math> represent the Earth rotation angle and <math>\Omega</math> the polynomial part of the right ascension of ascending node.		where <math>\theta</math> represent the Earth rotation angle and <math>\Omega</math> the polynomial part of the right ascension of ascending node.

Admin : /* The trigonometric polynomial part */

2018-02-28T08:27:17Z

The trigonometric polynomial part

← Version précédente		Version du 28 février 2018 à 08:27
Ligne 248 :		Ligne 248 :
	For each circular parameter (a, ex, ey, i, lna, <math>\alpha</math>), the formula of the trigonometric polynomial part used is:		For each circular parameter (a, ex, ey, i, lna, <math>\alpha</math>), the formula of the trigonometric polynomial part used is:

	~~(% style="text-align:~~ center~~;" %)~~		<center>
	<math>\sum_{p=0}^{p_{max}}A(p)\cos\left(n_p.\alpha_C+k_p.l_{NAC} + \Phi_p\right)</math>		<math>\sum_{p=0}^{p_{max}}A(p)\cos\left(n_p.\alpha_C+k_p.l_{NAC} + \Phi_p\right)</math></center>

	where <math>\alpha_C</math> and <math>l_{NAC}</math> represent the secular evolution of the latitude argument and longitude of the ascending node. This secular evolution corresponds to the "centered part" of the respective parameter models.		where <math>\alpha_C</math> and <math>l_{NAC}</math> represent the secular evolution of the latitude argument and longitude of the ascending node. This secular evolution corresponds to the "centered part" of the respective parameter models.
Ligne 257 :		Ligne 257 :
	The right ascension of ascending node and its mean longitude are connected as per the following equation :		The right ascension of ascending node and its mean longitude are connected as per the following equation :

	~~(% style="text-align:~~ center~~;" %)~~		<center>
	<math>\l_{NAC} = \Omega- \theta</math>		<math>\l_{NAC} = \Omega- \theta</math></center>

	where <math>\theta</math> represent the Earth rotation angle and <math>\Omega</math> the polynomial part of the right ascension of ascending node.		where <math>\theta</math> represent the Earth rotation angle and <math>\Omega</math> the polynomial part of the right ascension of ascending node.

Admin : /* Features Description */

2018-02-28T08:26:30Z

Features Description

← Version précédente		Version du 28 février 2018 à 08:26
Ligne 47 :		Ligne 47 :
	== Features Description ==		== Features Description ==

	Analytical propagators can be defined without any [{{JavaDoc3.3}}//org/orekit/attitudes/AttitudeProvider.html AttitudeProvider].		Analytical propagators can be defined without any [{{JavaDoc3.3}}/org/orekit/attitudes/AttitudeProvider.html AttitudeProvider].
	<syntaxhighlight lang="java">		<syntaxhighlight lang="java">
	KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);		KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);
	</syntaxhighlight>		</syntaxhighlight>

	In that case, the methods'''''getAttitudeProvider''''', '''''getAttitudeProviderForces''''' or '''''getAttitudeProviderEvents''''' will return a null-value. If propagation is performed using a propagator defined without any AttitudeProvider, the final SpacecraftState will have a null attitude value.		In that case, the methods '''''getAttitudeProvider''''', '''''getAttitudeProviderForces''''' or '''''getAttitudeProviderEvents''''' will return a null-value. If propagation is performed using a propagator defined without any AttitudeProvider, the final SpacecraftState will have a null attitude value.

	A single AttitudeProvider could be added to the propagator:		A single AttitudeProvider could be added to the propagator:
Ligne 85 :		Ligne 85 :

	// Extrapolator definition		// Extrapolator definition
	//~~<del></del><del></del><del></del><del></del><del></del>~~---		//-----------------------
	KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);		KeplerianPropagator extrapolator = new KeplerianPropagator(initialOrbit);

	// Extrapolation at a final date different from initial date		// Extrapolation at a final date different from initial date
	//~~<del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del>~~-		//---------------------------------------------------------
	double delta_t = 100000.0; // extrapolation duration in seconds		double delta_t = 100000.0; // extrapolation duration in seconds
	AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);		AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);
Ligne 120 :		Ligne 120 :

	// Extrapolator definition		// Extrapolator definition
	//~~<del></del><del></del><del></del><del></del><del></del>~~---		//-----------------------
	EcksteinHechlerPropagator extrapolator = new EcksteinHechlerPropagator(initialOrbit,		EcksteinHechlerPropagator extrapolator = new EcksteinHechlerPropagator(initialOrbit,
	ae, muIntegration, FramesFactory.getCIRF(), c20, c30, c40, c50, c60, ParametersType.OSCULATING);		ae, muIntegration, FramesFactory.getCIRF(), c20, c30, c40, c50, c60, ParametersType.OSCULATING);

	// Extrapolation at a final date different from initial date		// Extrapolation at a final date different from initial date
	//~~<del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del><del></del>~~-		//--------------------------------------------
	double delta_t = 100000.0; // extrapolation duration in seconds		double delta_t = 100000.0; // extrapolation duration in seconds
	AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);		AbsoluteDate extrapDate = initDate.shiftedBy(delta_t);
Ligne 165 :		Ligne 165 :

	Lyddane propagators <code>LyddaneSecularPropagator</code> and <code>LyddaneLongPeriodPropagator</code> are analytical propagators providing different mean elements: only secular effects for the former, secular plus long period ones for the latter. They only consider the effects of J2, J3, J4, J5 zonal harmonics. The propagation frame has to be nearly inertial, and an equatorial one is recommended (e.g. CIRF). When osculating elements are provided as input to <code>LyddaneSecularPropagator</code>and <code>LyddaneLongPeriodPropagator</code>, the osculating orbits that results from propagation are identical.		Lyddane propagators <code>LyddaneSecularPropagator</code> and <code>LyddaneLongPeriodPropagator</code> are analytical propagators providing different mean elements: only secular effects for the former, secular plus long period ones for the latter. They only consider the effects of J2, J3, J4, J5 zonal harmonics. The propagation frame has to be nearly inertial, and an equatorial one is recommended (e.g. CIRF). When osculating elements are provided as input to <code>LyddaneSecularPropagator</code>and <code>LyddaneLongPeriodPropagator</code>, the osculating orbits that results from propagation are identical.
	Lyddane propagators are'''not to be used''':		Lyddane propagators are '''not to be used''':
	* for orbits whose eccentricity is larger than 0.9.		* for orbits whose eccentricity is larger than 0.9.
	* for orbits whose inclination is close to the critical one (~63.43 deg and ~116.57 deg with a recommended margin of at least 0.5 deg).		* for orbits whose inclination is close to the critical one (~63.43 deg and ~116.57 deg with a recommended margin of at least 0.5 deg).
Ligne 237 :		Ligne 237 :

	==== The "centered" part ====		==== The "centered" part ====
	The "centered" part is any function having one variable : the date. It is represented by an object of type [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/UnivariateDateFunction.html UnivariateDateFunction].		The "centered" part is any function having one variable : the date. It is represented by an object of type [{{JavaDoc3.3}}/org/orekit/propagation/analytical/twod/UnivariateDateFunction.html UnivariateDateFunction].
	At this time, the following univariate functions are available :		At this time, the following univariate functions are available :
	* piecewise functions of date which are reprensented by the extensions of [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/AbstractDateIntervalFunction.html AbstractDateIntervalFunction] :		* piecewise functions of date which are reprensented by the extensions of [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/AbstractDateIntervalFunction.html AbstractDateIntervalFunction] :
	** [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/DateIntervalLinearFunction.html piecewise linear functions of date]		** [{{JavaDoc3.3}}/org/orekit/propagation/analytical/twod/DateIntervalLinearFunction.html piecewise linear functions of date]
	** [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/DateIntervalParabolicFunction.html piecewise 2nd order polynomial functions of date]		** [{{JavaDoc3.3}}/org/orekit/propagation/analytical/twod/DateIntervalParabolicFunction.html piecewise 2nd order polynomial functions of date]
	* [{{JavaDoc3.3}}//org/orekit/propagation/analytical/twod/DatePolynomialFunction.html polynomial functions of date]		* [{{JavaDoc3.3}}/org/orekit/propagation/analytical/twod/DatePolynomialFunction.html polynomial functions of date]
	[[File:Analytical2DPropagatorCenteredFunction.png\|center]]		[[File:Analytical2DPropagatorCenteredFunction.png\|center]]

Ligne 295 :		Ligne 295 :
	These pseudopropagators are bounded and based on a set of ephemerides :		These pseudopropagators are bounded and based on a set of ephemerides :

	'''*Integrated Ephemeris''' : With the ephemeris generation mode, a bounded propagator is available at the end of the numerical propagation and gives access to the orbit state at any date inside the propagation time interval. These intermediate states are computed thanks to the object IntegratedEphemeris. This object stores for each step the interpolator used, during the integration.		*'''Integrated Ephemeris''' : With the ephemeris generation mode, a bounded propagator is available at the end of the numerical propagation and gives access to the orbit state at any date inside the propagation time interval. These intermediate states are computed thanks to the object IntegratedEphemeris. This object stores for each step the interpolator used, during the integration.
	'''*Ephemeris''' : Another class is available for the same purpose (access to the orbit state at any given date): the Ephemeris class. An instance is built from a list of spacecraft states - that is to say, discrete ephemeris. Here, the intermediate states (keplerian, cartesian, ... parameters, depending on the kind of Orbit used) are computed with Hermite interpolators, which provides less accuracy than with the IntegratedEphemeris object. But for most uses, notably event detection, the Ephemeris class appears reliable enough.		*'''Ephemeris''' : Another class is available for the same purpose (access to the orbit state at any given date): the Ephemeris class. An instance is built from a list of spacecraft states - that is to say, discrete ephemeris. Here, the intermediate states (keplerian, cartesian, ... parameters, depending on the kind of Orbit used) are computed with Hermite interpolators, which provides less accuracy than with the IntegratedEphemeris object. But for most uses, notably event detection, the Ephemeris class appears reliable enough.
	'''*Lagrange Ephemeris''' and '''Hermite Ephemeris''' : These class are deprecated since 3.1. They were extending AbstractPropagator and they were computing propagation using Lagrange or Hermite interpolation between spacecraftstates. These possibilities has been replaced by the use of a simple propagator based on a PVCoordinateProvider which compute position, velocity with a Lagrange or Hermite interpolation between position, velocity ephemeris. This way of propagation is described in [ORB_Ephemeris_Home Ephemeris] page.		*'''Lagrange Ephemeris''' and '''Hermite Ephemeris''' : These class are deprecated since 3.1. They were extending AbstractPropagator and they were computing propagation using Lagrange or Hermite interpolation between spacecraftstates. These possibilities has been replaced by the use of a simple propagator based on a PVCoordinateProvider which compute position, velocity with a Lagrange or Hermite interpolation between position, velocity ephemeris. This way of propagation is described in [ORB_Ephemeris_Home Ephemeris] page.

	=== Covariance matrix propagation ===		=== Covariance matrix propagation ===

User Manual 3.3 Analytical propagation - Historique des versions

Admin : /* Scope */

Admin : /* Precomputed ephemeris */

Admin : /* Scope */

Admin : Admin a déplacé la page Analytical propagation vers User Manual 3.3 Analytical propagation sans laisser de redirection

Admin le 28 février 2018 à 08:39

Admin le 28 février 2018 à 08:39

Admin : /* Contents */

Admin : /* The trigonometric polynomial part */

Admin : /* The trigonometric polynomial part */

Admin : /* Features Description */

Admin : /* Scope */

Admin : /* Precomputed ephemeris */

Admin : /* Scope */

Admin : Admin a déplacé la page Analytical propagation vers User Manual 3.3 Analytical propagation sans laisser de redirection

Admin : /* Contents */

Admin : /* The trigonometric polynomial part */

Admin : /* The trigonometric polynomial part */

Admin : /* Features Description */