This package provides tools to propagate orbital states with different methods.
Propagation is the prediction of the evolution of a system from an initial state.
In Orekit, this initial state is represented by a
SpacecraftState, which is a
simple container for all needed information : orbit, mass, kinematics, attitude, date,
frame, and which can also hold any number user-defined additional data like battery
status or operating mode for example.
Propagation is based on a top level
Propagation interface (which in facts extends
PVCoordinatesProvider interface) and several implementations.
Depending on the needs of the calling application, all propagators can be used in
slave mode: This mode is used when the user wants to completely drive evolution of
time with his own loop. The (slave) propagator is passive: it computes this result and
returns it to the calling (master) application, without any intermediate feedback.
Users often use this mode in loops, each target propagation time representing the next
small time step. In that case the events detection is made but the step handler does
nothing, actions are managed directly by the calling application.
master mode: This mode is used when the user needs to have some custom function
called at the end of each finalized step during integration. The (master) propagator
is active: the integration loop calls the (slave) application callback methods at each
finalized step, through the step handler. Users often use this mode with only a single
call to propagation with the target propagation time representing the end final date.
ephemeris generation mode: This mode is used when the user needs random access
to the orbit state at any time between the initial and target times, and in no
sequential order. A typical example is the implementation of search and iterative
algorithms that may navigate forward and backward inside the propagation range before
finding their result.
CAVEATS: Be aware that this mode cannot support events that modify spacecraft initial state.
Be aware that since this mode stores all intermediate results, it may be memory-intensive
for long integration ranges and high precision/short time steps.
The recommended mode is master mode. It is very simple to use and allow user to get
rid of concerns about synchronizing force models, file output, discrete events. All
these parts are handled separately in different user code parts, and Orekit takes
care of all management. The following class diagram shows the main interfaces used
for master mode.
This mode also lets the propagator choose its own step size,
but still let user choose a different step size for its output, which greatly increases
performances. The following sequence diagram shows how a user-provided step handler is
called when propagation is done in master mode.
The following sequence diagram shows how user can handle themselves time when
propagation is done in slave mode, with a loop managed at user level.
The following sequence diagram shows how user can use ephemeris generation mode to
do a two phases propagation, one first phase to generate the epehemeris, and a second
phase using the generated ephemeris.
All propagators, including analytical ones, support discrete events handling during
propagation. This feature is activated by registering events detectors as defined by
EventDetector interface to the propagator, each
associated with an
EventHandler that will be triggered automatically at event
occurrence. The following class diagram shows the main interfaces used for event handling.
At each propagation step, the propagator checks the registered events detectors for
the occurrence of some event, as shown in the following sequence diagram. If an event
occurs, then the corresponding action is triggered, which can notify the propagator to
resume propagation (possibly with an updated state) or to stop propagation.
Resuming propagation with changed state is used for example in the
class. When the maneuver is triggered, the orbit is changed according to the velocity increment
associated with the maneuver and the mass is reduced according to the consumption. This
allows the handling simple maneuvers transparently even inside basic propagators like
Kepler or Eckstein-Heschler.
Stopping propagation is useful when some specific state is desired but its real occurrence
time is not known in advance. A typical example would be to find the next ascending node
or the next apogee. In this case, we can register a
NodeDetector or an
object and launch a propagation with a target time far away in the future. When the
event is triggered, it notifies the propagator to stop and the returned value is the state
exactly at the event time.
Users can define their own events, typically by extending the
AbstractDetector abstract class.
There are also several predefined events detectors already available, amongst which :
- a simple
DateDetector, which is simply triggered at a predefined date, and can be reset to add new dates on the run (which is useful to set up delays starting when a previous event is been detected),
ElevationDetector, which is triggered at raising or setting time of a satellite with respect to a ground point, taking atmospheric refraction into account and either constant elevation or ground mask when threshold elevation is azimuth-dependent,
ElevationExtremumDetector, which is triggered at maximum (or minimum) satellite elevation with respect to a ground point,
AltitudeDetectorwhich is triggered when satellite crosses a predefined altitude limit and can be used to compute easily operational forecasts,
DihedralFieldOfViewDetector, which are triggered when some target enters or exits a satellite sensor field of view,
GeographicZoneDetector, which are triggered the spacecraft enters or leave a zone, which can be non-convex, split in different sub-zones, have holes, contain the pole,
EclipseDetector, which is triggered when some body enters or exits the umbra or the penumbra of another occulting body,
ApsideDetector, which is triggered at apogee and perigee,
NodeDetector, which is triggered at ascending and descending nodes,
PositionAngleDetector, which is triggered when satellite angle on orbit crosses some value (works with either anomaly, latitude argument or longitude argument and with either true, eccentric or mean angles),
LongitudeExtremumDetector, which are triggered when satellite position with respect to central body reaches some predefined values,
AlignmentDetector, which is triggered when satellite and some body are aligned in the orbital plane...
EventShifter is also provided in order to slightly shift the events occurrences times.
A typical use case is for handling operational delays before or after some physical event
EventSlopeFilter is provided when user is only interested in one kind of events that
occurs in pairs like raising in the raising/setting pair for elevation detector, or
eclipse entry in the entry/exit pair for eclipse detector. The filter does not simply
ignore events after they have been detected, it filters them before they are located
and hence save some computation time by not doing an accurate search for events that
will ultimately be ignored.
EventEnablingPredicateFilter is provided when user wants to filter out some
events based on an external condition set up by a user-provided enabling predicate
function. This allow for example to dynamically turn some events on and off during
propagation or to set up some elaborate logic like triggering on elevation first
time derivative (i.e. one elevation maximum) but only when elevation itself is above
some threshold. The filter does not simply ignore events after they have been detected,
it filters them before they are located and hence save some computation time by not doing
an accurate search for events that will ultimately be ignored.
Event occurring can be automatically logged using the
All propagators can be used to propagate user additional states on top of regular
orbit attitude and mass state. These additional states will be available throughout
propagation, i.e. they can be used in the step handlers, in the event detectors and
events handlers and they will also be available in the final propagated state.
There are three main cases:
- if a deterministic way to compute the additional state from the spacecraft state
is known, then the user can put this computation in an implementation of the
AdditionalStateProviderinterface and register it to the propagator, which will call it each time it builds a
- if a differential equation drives the additional state evolution, then the user
can put this equation in an implementation of the
AdditionalEquationsinterface and register it to an integrator-based propagator, which will integrated it and provide the integrated state.
- if no evolution laws are provided to the propagator, but the additional state is available in the initial state, then the propagator will simply copy the initial value throughout propagation, without evolving it.
The first two cases correspond to additional states managed by the propagator, the
last case not being considered as managed. The list of states managed by the propagator
is available using the
The following class diagram shows the available propagators
KeplerianPropagator is based on Keplerian-only motion. It depends only on µ.
This analytical model is suited for near-circular orbits and inclination neither
equatorial nor critical. It considers J2 to J6 potential zonal coefficients,
and uses mean parameters to compute the new position.
Note that before version 7.0, there was a large inconsistency in the generated
orbits. It was fixed as of version 7.0 of Orekit, with a visible side effect.
The problem is that if the circular parameters produced by the Eckstein-Hechler
model are used to build an orbit considered to be osculating, the velocity deduced
from this orbit was inconsistent with the position evolution! The reason is
that the model includes non-Keplerian effects but it does not include a corresponding
circular/Cartesian conversion. As a consequence, all subsequent computation involving
velocity were wrong. This includes attitude modes like yaw compensation and Doppler
effect. As this effect was considered serious enough and as accurate velocities were
considered important, the propagator now generates Cartesian orbits which are built
in a special way to ensure consistency throughout propagation.
A side effect is that if circular parameters are rebuilt by user from these propagated
Cartesian orbit, the circular parameters will generally not match the initial
orbit (differences in semi-major axis can exceed 120 m). The position however will
match to sub-micrometer level, and this position will be identical to the positions
that were generated by previous versions (in other words, the internals of the models
have not been changed, only the output parameters have been changed). The correctness
of the initialization has been assessed and is good, as it allows the subsequent orbit
to remain close to a numerical reference orbit.
If users need a more definitive initialization of an Eckstein-Hechler propagator, they
should consider using a propagator converter to initialize their Eckstein-Hechler
propagator using a complete sample instead of just a single initial orbit.
This analytical model is dedicated to Two-Line Elements (TLE) propagation.
Differential effects adapter¶
This model is used to add to an underlying propagator some effects it does not
take into account. A typical example is to add small station-keeping maneuvers
to a pre-computed ephemeris or reference orbit which does not take these maneuvers
into account. The additive maneuvers can take both the direct effect (Keplerian part)
and the induced effect due for example to J2 which changes ascending node rate when
a maneuver changed inclination or semi-major axis of a Sun-Synchronous satellite.
Numerical propagation is one of the most important parts of the Orekit project.
Based on Apache Commons Math ordinary differential equations integrators, the
NumericalPropagator class realizes the interface between space mechanics and
mathematical resolutions. Despite its utilization seems daunting on first sight,
it is in fact quite straigthforward to use.
Simple propagation of equations of motion¶
The mathematical problem to integrate is a dimension-seven time-derivative equations
system. The six first elements of the state vector are the orbital parameters, which
may be any orbit type (
CartesianOrbit) in meters and radians, and the last element is the mass in
kilograms. It is possible to have more elements in the state vector if
have been added (typically
PartialDerivativesEquations which is an implementation of
AdditionalEquations devoted to integration of Jacobian matrices). The time derivatives are
computed automatically by the Orekit using the Gauss equations for the first parameters
corresponding to the selected orbit type and the flow rate for mass evolution
during maneuvers. The user only needs to register the various force models needed for
the simulation. Various force models are already available in the library and specialized
ones can be added by users easily for specific needs.
The integrators (first order integrators) provided by Apache Commons Math need
the state vector at t0, the state vector first time derivative at t0,
and then calculates the next step state vector, and asks for the next first
time derivative, etc. until it reaches the final asked date. These underlying numerical
integrators can also be configured. Typical tuning parameters for adaptive stepsize
integrators are the min, max and perhaps start step size as well as the absolute and/or
relative errors thresholds. The following code snippet shows a typical setting for Low
Earth Orbit propagation:
// steps limits final double minStep = 0.001; final double maxStep = 1000; final double initStep = 60; // error control parameters (absolute and relative) final double positionError = 10.0; final double tolerances = NumericalPropagator.tolerances(positionError, orbit, orbit.getType()); // set up mathematical integrator AdaptiveStepsizeIntegrator integrator = new DormandPrince853Integrator(minStep, maxStep, tolerances, tolerances); integrator.setInitialStepSize(initStep); // set up space dynamics propagator propagator = new NumericalPropagator(integrator);
The following sequence diagram show the internal calls triggered during a numerical propagation.
The important thing to note is that the force model are decoupled from the integration process,
they only need to compute an acceleration.
Propagating both equations of motions and additional equations¶
Sometimes, simple equations of motion are not enough. Some additional parameters can be useful
alongside with the trajectory, like dual parameters in optimal control or Jacobians (also
called state-transition matrices). This second case especially useful for computing sensitivity
of a trajectory with respect to initial state changes or with respect to force models parameters
Orekit provides a common way to handle both cases: additional equations. Users can register sets
of additional equations alongside with additional initial states. These equations will be propagated
by the numerical integrator. They will not be used for step control, though, so integrating with
or without these equations should not change the trajectory and no tolerance setting is needed for
One specific implementation of additional equations is the partial derivatives equations which
propagate Jacobian matrices, both with respect to initial state and with respect to force model
The above class diagram shows the design of the partial derivatives equations. As can be seen,
the PartialDerivativesEquations class implements the AdditionalEquations interface and as such
can be registered by user in a numerical propagator. the propagator will propagate both the
main set of equations corresponding to the equations of motion and the additional set corresponding
to the Jacobians of the main set. This additional set is therefore tightly linked to the main set
and in particular depends on the selected force models. The various force models add their
direct contribution directly to the main set, just as in simple propagation. They also add a
contribution to the Jacobians thanks to the AccelerationJacobiansProvider interface, each force
model being associated with an acceleration Jacobians provider. Some force models like solar
radiation pressure implement this interface by themselves. Some more complex force model do not
implement the interface and will be automatically wrapped inside a Jacobianizer class which will
use finite differences to compute the local Jacobians.
Semianalytical propagation is an intermediate between analytical and numerical propagation.
It retains the best of both worlds, speed from analytical models and accuracy from numerical models.
Semianalytical propagation is implemented using Draper Semianalytical Satellite Theory (DSST).
Since version 7.0, both mean elements equations of motion models and short periodic terms
have been implemented and validated.
Since 9.0, most of the Orekit propagators (in fact all of them except DSST) have both a regular
version the propagates states based on classical real numbers (i.e. double precision numbers)
and a more general version that propagates states based on any class that implements the
RealFieldElement interface from Hipparchus. Such classes mimic real numbers in the way they
support all operations from the real field (addition, subtraction, multiplication, division,
but also direct and inverse trigonometric functions, direct and inverse hyperbolic functions,
logarithms, powers, roots...).
A very important implementation of the
RealFieldElement interface is the
class, which in addition to compute the result of the canonical operation (add, multiply, sin,
atanh...) also computes its derivatives, with respect to any number of variables and to any
derivation order. If for example a user starts a computation with 6 canonical variables px,
py, pz, vx, vy, vz to represent an initial state and then performs a propagation. At the
end for all produced results (final position, final velocity but also geodetic altitude
with respect to an ellipsoid body or anything that Orekit computes), then for these results
one can retrieve its partial derivatives up to the computed order with respect to the 6
canonical variables. So if for example in a step handler you compute a geodetic altitude h,
you also have ∂³h/∂px²∂vz or any of the 84 components computed at order 3 for each value
(1 value, 6 first order derivatives, 14 second order derivatives, 56 third order derivatives).
DerivativeStructure class also provides Taylor expansion, which allow to extrapolate
the result accurately to close values. This is an implementation of Taylor Algebra. Its two
main uses in space flight dynamics are
- accurate propagation of uncertainties (to much higher accuracy than simple covariance matrices)
- very fast Monte-Carlo analyses
Orekit implementations of field propagators support all features from classical propagators:
propagation modes, events (all events detectors), frames transforms, geodetic points. The
propagators available are Keplerian propagator, Eckstein-Heschler propagator, SGP4/SDP4
propagator, and numerical propagator with all Hipparchus integrators (fixed steps or adaptive
stepsizes) and all force models (including all atmosphere models). All attitude modes are
One must be aware however of the combinatorial explosion of computation size. For p derivation
parameters and o order, the number of components computed for each value is given by the
binomial coefficient (o+p¦p). As an example 6 parameters and order 6 implies every single
double in a regular propagation will be replaced by 924 numbers in field propagation. These
numbers are all combined together linearly in addition and subtraction, but quadratically
in multiplication and divisions. The
DerivativeStructure class is highly optimized, but having
both high order and high number of parameters remains inherently costly.
Once the propagation has been performed, however, evaluating a Taylor expansion, for example in
a Monte-Carlo application is very fast. So even if the propagation ends up to be for example a
hundred of times slower than regular propagation, depending on the number of derivatives, the
payoff is still very important as soon as we evaluate a few hundreds of points. As Monte-Carlo
analyses more often use several thousands of evaluations, the payoff is really interesting.
Another important implementation of the
RealFieldElement interface is the
class, which computes the same operation on a number of components of a tuple, hence
allowing to perform parallel orbit propagation in one run. Each spacecraft will correspond
to one component of the tuple. The first spacecraft (component at index 0) is the reference.
There is a catch, however. In many places in orbit propagations, there are conditional
statements the depend on the current state (for example is the spacecraft in eclipse
or still in Sun light). As a single choice is allowed, the outcome of the check is based
on the reference spacecraft only (i.e. fist component of the tuple) and the conditional
branch is selected according to this reference spacecraft. The spacecrafts represented by
the other components of the tuple will follow the same branch in the algorithm, even despite
they may not be in the same conditions. This means that using
Tuple for orbit propagation
works only for close enough spacecrafts. This is well suited for finite differences,
formation flying or co-positioning, but this is not suited for constellations.