Orbit Propagator
Propagate orbit of one or more spacecraft
- Library:
Aerospace Blockset / Aerospace Blockset CubeSat Simulation Library
Description
The Orbit Propagator block propagates the orbit of one or more spacecraft by a propagation method:
Kepler universal variable formulation (quicker)
Numerical integration (more accurate)
You can define initial orbital states in the Orbit tab as:
A set of orbital elements
Position and velocity state vectors in International Celestial Reference Frame (ICRF) or fixed-frame coordinate systems.
The block uses quaternions, which are defined using the scalar-first convention.
The Orbit Propagator block is available only in the Aerospace Blockset CubeSat Simulation Library, available through the Add-On Explorer.
To access the Aerospace Blockset CubeSat Simulation Library, type
asbCubeSatBlockLib in the MATLAB® Command Window.
For more information on the coordinate systems the Orbit Propagator block uses, see Algorithms.
Limitations
The Orbit Propagator block is available only through the Add-On Explorer.
Ports
Input
Aicrf — Applied acceleration
3-element vector | m-3 array
Acceleration applied to the spacecraft with respect to the port coordinate system (ICRF or fixed-frame), specified as a 3-element vector or m-by-3 array, at the current time step.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision).Select the Input external accelerations check box.
Data Types: double
φθψ — Moon libration angles
3-element vector
Moon libration angles for transformation between the ICRF and Moon-centric fixed-frame using the Moon-centric Principal Axis (PA) system, specified as a 3-element vector. To get these values, use the Moon Libration block.
Note
The fixed-frame used by this block when Central body is set
to Moon is the Mean Earth/pole axis (ME) system. For more
information, see Algorithms.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision).Set Central body to
Moon.Select the Input Moon libration angles check box.
Data Types: double
αδW — Right ascension, declination, and rotation angle
3-element vector
Central body spin axis instantaneous right ascension, declination, and rotation angle, specified as a 3-element vector. This port is available only for custom central bodies.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision).Set Central body to
Custom.Set Central body spin axis source to
Port.
Data Types: double
Output
Xicrf — Position of spacecraft
3-element vector | m-by-3, where m is number of spacecraft
array
Position of the spacecraft with respect to (ICRF or fixed-frame), returned as a 3-element vector or m-by-3 array, where m is number of spacecraft, at the current time step. The size of the initial conditions provided in the Orbit tab control the port dimension.
Data Types: double
Vicrf — Velocity
3-element vector | m-by-3 array, where m is number of
spacecraft array
Velocity of the spacecraft with respect to ICRF or fixed-frame, returned as a 3-element vector or m-by-3 array, where m is number of spacecraft array, at the current time step. The size of the initial conditions provided in the Orbit tab control the port dimension.
Data Types: double
qicrf2ff — Transformation
4-element vector (scalar first)
Transformation between the ICRF coordinate system and fixed-frame, returned as a 4-element vector (scalar first), at the current time step.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision).Select the Output quaternion (ICRF to Fixed-frame) check box.
Data Types: double
tutc — Time at current time step
scalar | 6-element vector
Time at current time step, returned as a:
scalar — If you specify the Start data/time parameter as a Julian date.
6-element vector — If you specify the Start data/time parameter as a Gregorian date with six elements (year, month, day, hours, minutes, seconds).
This value is equal to the Start date/time parameter value + the elapsed simulation time.
Dependencies
To enable this parameter, select the Output current date/time (UTC Julian date) check box.
Data Types: double
Parameters
Main
Propagation method — Orbit propagation method
Kepler (unperturbed) (default) | Numerical (high precision)
Orbit propagation method, specified as:
Kepler (unperturbed)— Uses a universal variable formulation of the Kepler problem to determine the spacecraft position and velocity at each time step. This method is faster thanNumerical (high precision).Numerical (high precision)— Determine the spacecraft position and velocity at each time step using numerical integration. This option models central body gravity based on the settings in the Central body tab. This method is more accurate thanKepler (unperturbed), but slower.
Programmatic Use
Block Parameter:
propagator |
| Type: character vector |
Values:
'Kepler (unperturbed)' | 'Numerical (high
precision)' |
Default:
'Kepler (unperturbed)' |
Input external accelerations — Input additional accelerations
off (default) | on
To enable additional external accelerations to be included in the integration of the spacecraft equations of motion, select this check box. Otherwise, clear this check box.
Dependencies
To enable this check box, set Propagation method to
High precision.
Programmatic Use
Block Parameter:
accelIn |
| Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
External acceleration coordinate frame — Frame for additional accelerations
ICRF (default) | Fixed-frame
Input additional accelerations, specified as ICRF or
Fixed-frame. These accelerations are included in
integration of the spacecraft equations of motion.
Dependencies
To enable this parameter:
Set Propagation method to
Numerical (high precision)Select the Input external accelerations check box
Programmatic Use
Block Parameter:
accelFrame |
| Type: character vector |
Values:
'ICRF' | 'Fixed-frame' |
Default:
'ICRF' |
State output coordinate frame — Port coordinate frame
ICRF (default) | Fixed-frame
Coordinate frame for output ports, specified as ICRF or
Fixed-frame. These port labels are affected:
Output port X
Output port V
Dependencies
To enable this parameter, set Propagation method to
Numerical (high precision).
Programmatic Use
Block Parameter:
outportFrame |
| Type: character vector |
Values:
'ICRF' | 'Fixed-frame' |
Default:
'ICRF' |
Start date/time (UTC Julian date) — Initial start time for simulation
juliandate (202 0, 1, 1, 12, 0, 0) (default) | valid scalar Julian date | valid Gregorian date including year, month, day, hours, minutes, seconds as
6-element vector for Gregorian dates
Initial start date and time of simulation, specified as a Julian or Gregorian date. The block defines initial conditions using this value.
Tip
To calculate the Julian date, use the juliandate function.
Programmatic Use
Block Parameter:
startDate |
| Type: character vector |
Values: 'juliandate(2020, 1, 1,
12, 0, 0)' | valid scalar Julian date | valid Gregorian date including
year, month, day, hours, minutes, seconds as 6-element vector |
Default:
'juliandate(2020, 1, 1, 12, 0, 0)' |
Output current date/time (UTC Julian date) — Add output port tutc
on (default) | off
To output the current date or time, select this check box. Otherwise, clear this check box.
Programmatic Use
Block Parameter:
dateOut |
| Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
Action for out-of-range input — Out-of-range block behavior
Warning (default) | Error | None
Out-of-range block behavior, specified as follows:
| Action | Description |
|---|---|
None
| No action. |
Warning
| Warning displays in the MATLAB Command Window. Model simulation continues. |
Error (default) | MATLAB returns an exception. Model simulation stops. |
Programmatic Use
Block Parameter:
action |
| Type: character vector |
Values: 'None' |
'Warning' | 'Error' |
Default:
'Warning' |
Orbit
Define the initial states of the space craft.
Initial state format — Input method for initial states of orbit
Orbital elements (default) | ICRF state vector | Fixed-frame state vector
Input method for initial states of orbit, specified as Orbital
elements, ICRF state vector, or
Fixed-frame state vector.
Dependencies
Available options are based on Propagation method settings:
| Kepler (unperturbed) | Numerical (high precision) |
|---|---|
| Orbital elements | Orbital elements |
| ICRF state vector | ICRF state vector |
| — | Fixed-frame state vector |
Programmatic Use
Block Parameter
stateFormatKep when propagator is set to
Kepler (unperturbed), stateFormatNum when
propagator is set to Numerical (high
precision) |
| Type: character vector |
Values:
'Orbital elements' | 'ICRF state vector'
when propagator is set to 'Kepler
(unperturbed)' | 'Orbital elements' |
'ICRF state vector' | 'Fixed-frame state'
when propagator is set to 'Numerical (high
precision)' |
Default:
'Orbital elements' |
Orbit type — Orbit classification
Keplerian (default) | Elliptical equatorial | Circular | Circular equatorial
Orbit classification, specified as:
Keplerian— Model elliptical, parabolic, and hyperbolic orbits using six standard Keplerian orbital elements.Elliptical equatorial— Fully define an equatorial orbit, where inclination is 0 or 180 degrees and the right ascension of the ascending node is undefined.Circular— Define a circular orbit, where eccentricity is 0 and the argument of periapsis is undefined. To fully define a circular orbit, selectCircular equatorial.Circular equatorial— Fully define a circular orbit, where eccentricity is 0 and the argument of periapsis is undefined.
Dependencies
To enable this parameter, set Initial state format to
Orbital elements.
Programmatic Use
Block Parameter:
orbitType |
| Type: character vector |
Values:
'Keplerian' | 'Elliptical equatorial' |
'Circular inclined' | 'Circular
equatorial' |
Default:
'Keplerian' |
Semi-major axis — Half of major axis of ellipse
6786000 (default) | 1-D array of size m, number of spacecraft
Half of ellipsis major axis, specified as a 1-D array whose size is the number of spacecraft.
For parabolic orbits, this block interprets this parameter as the periapsis radius (distance from periapsis to the focus point of orbit).
For hyperbolic orbits, this block interprets this parameter as the distance from periapsis to the hyperbola center.
Dependencies
To enable this parameter, set Initial state format to
Orbital elements.
Programmatic Use
Block Parameter:
semiMajorAxis |
| Type: character vector |
| Values: scalar | 1- or 2-D array of size that is number of spacecraft |
Default:
'6786000' |
Eccentricity — Deviation of orbit
0.01 (default) | scalar | value between 0 and 1, or greater than
1 for Keplerian orbit type | 1-D array of size m, number of spacecraft
Deviation of the orbit from a perfect circle, specified as a scalar or 1-D array of size that is number of spacecraft, where eccentricity is the shape of the ellipse.
If Orbit type is set to Keplerian,
value can be:
1for parabolic orbitGreater than
1for hyperbolic orbit
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit type to
KeplerianorElliptical equatorial.
Programmatic Use
Block Parameter:
eccentricity |
| Type: character vector |
Values:
0.01 | scalar | value between 0 and
1, or greater than 1 for Keplerian orbit
type | 1-D array of size m, number of spacecraft |
Default:
'0.01' |
Inclination (deg) — Tilt angle of CubeSat orbital plane
50 (default) | scalar | 1-D array of size m number of spacecraft | degrees between 0 and 180 | radians between 0 and pi
Vertical tilt of the ellipse with respect to the reference plane measured at the ascending node, specified as a scalar or 1-D array of size m number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elementsOrbit type to
KeplerianorCircular inclined
Programmatic Use
Block Parameter:
inclination |
| Type: character vector |
| Values: 50 | scalar | 1-D array of size m number of spacecraft | degrees between 0 and 180 | radians between 0 and pi |
Default:
'50' |
RAAN (deg) — Angular distance in equatorial plane
95 (default) | scalar value between 0 and 360 | 1-D array of size m number of spacecraft
Right ascension of ascending node (RAAN), specified as a value between
0 and 360, specified as a scalar or 1-D array
of size that is m number of spacecraft, in specified units. RAAN is
the angular distance along the reference plane from the ICRF x-axis
to the location of the ascending node (the point at which the spacecraft crosses the
reference plane from south to north).
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit type to
KeplerianorCircular inclined.
Programmatic Use
Block Parameter:
raan |
| Type: character vector |
Values:
'95' | scalar value between 0 and
360 | 1-D array of size m number of
spacecraft |
Default:
'95' |
Argument of periapsis (deg) — Angle from spacecraft ascending node to periapsis
93 (default) | value between 0 and 360 | 1-D array of size m number of spacecraft
Angle from the spacecraft ascending node to periapsis (closest point of orbit to the central body), specified as a 1-D array of size m that is number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elementsOrbit type to
Keplerian
Programmatic Use
Block Parameter:
argPeriapsis |
| Type: character vector |
Values: 93 | scalar value between
0 and 360 | 1-D array of size
m number of spacecraft |
Default:
'93' |
True anomaly — Angle between periapsis and initial position of spacecraft
203 (default) | scalar value between 0 and 360 | 1-D array of size m number of spacecraft
Angle between periapsis (closest point of orbit to the central body) and the initial position of spacecraft along its orbit at Start date/time, specified as a scalar or 1-D array of size that is number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit type to
KeplerianorElliptical inclined.
Programmatic Use
Block Parameter:
trueAnomaly |
| Type: character vector |
Values:
'203' | scalar value between 0 and
360 | 1-D array of size m number of
spacecraft |
Default:
'203' |
Argument of latitude (deg) — Angle between ascending node and initial position of spacecraft
200 (default) | scalar | value between 0 and 360 | 1-D array of size m number of spacecraft
Angle between the ascending node and the initial position of spacecraft along its orbit at Start date/time, specified as a scalar or 3-element vector or 1-D array of size number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit Type to
Circular inclined.
Programmatic Use
Block Parameter:
argLat |
| Type: character vector |
Values:
'200' | scalar value between 0 and
360 | 3-element vector or 1-D array of size number of
spacecraft |
Default:
'200' |
Longitude of periapsis (deg) — Angle between ICRF x-axis and eccentricity vector
100 (default) | scalar | value between 0 and 360 | 1-D array of size m number of spacecraft
Angle between the ICRF x-axis and the eccentricity vector, specified as a scalar or 3-element vector or 1-D array of size number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit type to
Circular equatorial.
Programmatic Use
Block Parameter:
lonPeriapsis |
| Type: character vector |
Values: 100 | scalar value between
0 and 360 | 3-element vector or 1-D array
of size number of spacecraft |
Default:
'100' |
True longitude (deg) — Angle between ICRF x-axis and initial position of spacecraft
150 (default) | scalar | value between 0 and 360 | 1-D array of size m, number of spacecraft
Angle between the ICRF x-axis and the initial position of spacecraft along its orbit at Start date/time, specified as a scalar or 1-D array of size m, number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements.Orbit type to
Elliptical equatorial.
Programmatic Use
Block Parameter:
trueLon |
| Type: character vector |
Values:
'150' | scalar value between 0 and
360 | 3-element vector or 2-D array of size
m-by-3 array of spacecraft |
Default:
'150' |
ICRF position — Cartesian position vector of spacecraft
[3649700.0 3308200.0 -4676600.0] (default) | 3-element vector for single spacecraft or 2-D array of size
m-by-3 array of multiple spacecraft
Cartesian position vector of spacecraft in ICRF coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or 2-D array of size m-by-3 array of multiple spacecraft.
Dependencies
To enable this parameter, set Initial state format to
ICRF state vector.
Programmatic Use
Block Parameter:
inertialPosition |
| Type: character vector |
Values:
[3649700.0 3308200.0 -4676600.0] | 3-element vector for single
spacecraft or 2-D array of size m-by-3 array of multiple
spacecraft |
Default:
'[3649700.0 3308200.0 -4676600.0]' |
ICRF velocity — Cartesian velocity vector of spacecraft
[-2750.8 6666.4 2573.4] (default) | 3-element vector for single spacecraft or 2-D array of size
m-by-3 array of multiple spacecraft
Cartesian velocity vector of spacecraft in ICRF coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or 2-D array of size m-by-3 array of multiple spacecraft.
Dependencies
To enable this parameter, set Initial state format to
ICRF state vector.
Programmatic Use
Block Parameter:
inertialVelocity |
| Type: character vector |
Values:
[-2750.8 6666.4 2573.4] | 3-element vector for single
spacecraft or 2-D array of size m-by-3 array of multiple
spacecraft |
Default:
'[-2750.8 6666.4 2573.4]' |
Fixed-frame position — Position vector of spacecraft
[-4142689.0 -2676864.7 -4669861.6] (default) | 3-element vector for single spacecraft or 2-D array of size
m-by-3 array of multiple spacecraft
Cartesian position vector of spacecraft in fixed-frame coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or 2-D array of size m-by-3 array of multiple spacecraft.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).set Initial state format to
Fixed-frame state vector.
Programmatic Use
Block Parameter:
fixedPosition |
| Type: character vector |
Values:
'[-4142689.0 -2676864.7 -4669861.6]' | 3-element vector for
single spacecraft or 2-D array of size m-by-3 array of multiple
spacecraft |
Default:
'[-2750.8 6666.4 2573.4]' |
Fixed-frame velocity — Velocity vector of spacecraft
[1452.7 -6720.7 2568.1] (default) | 3-element vector for single spacecraft or 2-D array of size
m-by-3 array of multiple spacecraft
Cartesian velocity vector of spacecraft in fixed-frame coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or 2-D array of size m-by-3 array of multiple spacecraft.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Initial state format to
Fixed-frame state vector.
Programmatic Use
Block Parameter:
fixedVelocity |
| Type: character vector |
Values:
'[1452.7 -6720.7 2568.1]' | 3-element vector for single
spacecraft or 2-D array of size m-by-3 array of multiple
spacecraft |
Default:
'[1452.7 -6720.7 2568.1]' |
Central Body
Central body — Celestial body around which spacecraft orbits
Earth (default) | Moon | Mercury | Venus | Mars | Jupiter | Saturn | Uranus | Neptune | Custom
Celestial body, specified as Earth,
Moon, Mercury,
Venus, Mars,
Jupiter, Saturn,
Uranus, Neptune, or
Custom, around which the spacecraft defined in the
Orbit tab orbits.
Programmatic Use
Block Parameter:
centralBody |
| Type: character vector |
Values:
'Earth' | 'Moon'
|'Mercury' | 'Venus' |
'Mars' | 'Jupiter' |
'Saturn' | 'Uranus' |
'Neptune' | 'Custom' | |
Default:
'Earth' |
Gravitational potential model — Control gravity model for central body
Spherical harmonics when
Central body set to Earth,
Moon, Mars, or
Custom, Oblate ellipsoid when Central
body set to Mercury,
Venus, Jupiter,
Saturn, Uranus, or
Neptune (default) | None | Point-mass | Oblate ellipsoid (J2)
Control the gravity model for the central body, specified as
Spherical harmonics,
Point-mass, or Oblate ellipsoid
(J2).
Dependencies
To enable this parameter, set Propagation method to
Numerical (high precision). Available options are based
on Central body settings:
| Earth, Moon, Mars, or Custom | Mercury, Venus, Jupiter, Saturn, Uranus, or Neptune |
|---|---|
None | None |
Spherical harmonics | Oblate ellipsoid (J2) |
Point-mass | Point-mass |
Oblate ellipsoid (J2) | — |
Programmatic Use
Block Parameter:
gravityModel when centralBody set to
'Earth', 'Moon',
'Mars', or 'Custom' |
gravityModelnoSH when centralBody set to
Mercury, Venus,
Jupiter, Saturn, Uranus,
or Neptune |
| Type: character vector |
Values:
'Spherical harmonics' | 'None' |
'Point-mass' | 'Oblate ellipsoid (J2)'
when centralBody set to 'Earth',
'Moon', 'Mars', or
'Custom'; 'Point-mass' | 'Oblate
ellipsoid (J2)' when centralBody set to
Mercury, Venus,
Jupiter, Saturn, Uranus,
or Neptune |
Default:
'Spherical harmonics' when centralBody set
to 'Earth', 'Moon',
'Mars', or 'Custom'; 'Oblate
ellipsoid (J2)' when centralBody set to
Mercury, Venus,
Jupiter, Saturn, Uranus,
or Neptune |
Spherical harmonic model — Spherical harmonic model
EGM2008 for Central
body set to Earth,
LP-100K for Central body set to
Moon, GMM2B for
Central body set to Mars, (default) | EGM96 | EIGEN-GL04C | LP-165P
Spherical harmonic gravitational potential model, specified according to the specified Central body.
Dependencies
To enable this parameter, set Propagation method to
Numerical (high precision). Available options are based
on Central body settings:
| Central body | Spherical Harmonic Model Option |
|---|---|
| Earth | EGM2008, EGM96, or EIGEN-GL04C |
| Moon | LP-100K or LP-165P |
| Mars | GMM2B |
Programmatic Use
Block Parameter:
'earthSH' when centralBody set to
'Earth' | 'moonSH' when
centralBody set to 'Moon' |
'marsSH' when centralBody set to
'Mars' |
| Type: character vector |
Values:
'EGM2008' | 'EGM96' |
'EIGEN-GL04C' when centralBody set to
'earthSH'; 'LP-100K' |
'LP-165P' when centralBody set to
'moonSH'; 'GMM2B' when
centralBody set to 'marsSH' |
Default:
'Spherical harmonics' |
Spherical harmonic coefficient file — Harmonic coefficient MAT-file
aerogmm2b.mat (default) | harmonic coefficient MAT-file
Harmonic coefficient MAT-file that contains definitions for a custom planetary model, specified as a character vector or string.
This file must contain:
| Variable | Description |
|---|---|
| Re | Scalar of planet equatorial radius in meters (m). |
| GM | Scalar of planetary gravitational parameter in meters cubed per second squared (m3/s2) . |
| degree | Scalar of maximum degree. |
| C | (degree+1)-by-(degree+1) matrix containing normalized spherical harmonic coefficients matrix, C. |
| S | (degree+1)-by-(degree+1) matrix containing normalized spherical harmonic coefficients matrix, S. |
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Gravitational potential model to
Spherical harmonics.
Programmatic Use
Block Parameter:
shFile |
| Type: character vector |
Values:
'aerogmm2b.mat' | harmonic coefficient MAT-file |
Default:
'aerogmm2b.mat' |
Degree — Degree of harmonic model
120 (default) | scalar | maximum of 2159
Degree of harmonic model, specified as a double scalar:
| Planet Model | Recommended Degree | Maximum Degree |
|---|---|---|
| 120 | 2159 |
| 70 | 360 |
| 60 | 100 |
| 60 | 165 |
| 60 | 80 |
| 70 | 360 |
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Earth,Moon,Mars, orCustom.Gravitational potential model to
Spherical harmonics.
Programmatic Use
Block Parameter:
shDegree |
| Type: character vector |
Values:
'80' | scalar |
Default:
'80' |
Use Earth orientation parameters (EOPs) — Use Earth orientation parameters
on (default) | off
Select this check box to use Earth orientation parameters for the transformation between the ICRF and fixed-frame coordinate systems. Otherwise, clear this check box.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Earth.
Programmatic Use
Block Parameter:
useEOPs |
| Type: character vector |
Values:
'on' | 'off' |
Default:
'on' |
IERS EOP data file — Earth orientation data
aeroiersdata.mat (default) | MAT-file
Custom list of Earth orientation data, specified in a MAT-file.
Dependencies
To enable this parameter:
Select the Use Earth orientation parameters (EOPs) to check box.
Set Propagation method to
Numerical (high precision).Set Central body to
Earth.
Programmatic Use
Block Parameter:
eopFile |
| Type: character vector |
Values:
'aeroiersdata.mat' | MAT-file |
Default:
'aeroiersdata.mat' |
Input Moon libration angles — Moon libration angle rate
off (default) | on
To specify libration angles (φ θ ψ) for Moon orientation, select this check box.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Moon.
Programmatic Use
Block Parameter:
useMoonLib |
| Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
Output quaternion (ICRF to Fixed-frame) — Add output transformation quaternion port
off (default) | on
To add output transformation quaternion port for the quaternion transformation from the ICRF to the Fixed-frame coordinate system, select this check box.
Dependencies
To enable this check box, set Propagation method to
Numerical (high precision).
Programmatic Use
Block Parameter:
outputTransform |
| Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
Central body spin axis source — Central body spin source
Port (default) | Dialog
Central body spin axis, specified as Port or
Dialog. The block uses the spin axis to calculate the
transformation from the ICRF to the fixed-frame coordinate system for the custom
central body.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.
Programmatic Use
Block Parameter:
cbPoleSrc |
| Type: character vector |
Values:
'Port' | 'Dialog' |
Default:
'Port' |
Spin axis right ascension (RA) at J2000 (deg) — Right ascension of central body spin axis at J2000
317.68143 (default) | double scalar
Right ascension of central body spin axis at J2000 (2451545.0 JD, 2000 Jan 1 12:00:00 TT), specified as a double scalar.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbRA |
| Type: character vector |
Values:
'317.68143' | double scalar |
Default:
'317.68143' |
Spin axis RA rate (deg/century) — Right ascension rate of central body spin axis
-0.1061 (default) | double scalar
Right ascension rate of the central body spin axis, specified as a double scalar, in specified angle units/century.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbRARate |
| Type: character vector |
Values:
'-0.1061' | double scalar |
Default:
'-0.1061' |
Spin axis declination (Dec) at J2000 (deg) — Declination of central body spin axis at J2000
52.88650 (default) | double scalar
Declination of the central body spin axis at J2000 (2451545.0 JD, 2000 Jan 1 12:00:00 TT), specified as a double scalar.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbDec |
| Type: character vector |
Values:
'52.88650' | double scalar |
Default:
'52.88650' |
Spin axis Dec rate (deg/century) — Declination rate of central body spin axis
-0.0609 (default) | double scalar
Declination rate of the central body spin axis, specified as a double scalar, in specified angle units/century.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbDecRate |
| Type: character vector |
Values:
'-0.0609' | double scalar |
Default:
'-0.0609' |
Initial rotation angle at J2000 (deg) — Rotation angle of central body x-axis
176.630 (default) | double scalar
Rotation angle of the central body x axis with respect to the ICRF x-axis at J2000 (2451545.0 JD, 2000 Jan 1 12:00:00 TT), specified as a double scalar, in specified angle units.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbRotAngle |
| Type: character vector |
Values:
'176.630' | double scalar |
Default:
'176.630' |
Rotation rate (deg/day) — Rotation rate of central body x-axis
350.89198226 (default) | double scalar
Rotation rate of the central body x axis with respect to the ICRF x-axis (2451545.0 JD, 2000 Jan 1 12:00:00 UTC), specified as a double scalar, specified angle units/day.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Central body spin axis source to
Dialog.
Programmatic Use
Block Parameter:
cbRotRate |
| Type: character vector |
Values:
'350.89198226' | double scalar |
Default:
'350.89198226' |
Equatorial radius — Equatorial radius
3396200 (default) | double scalar
Equatorial radius for a custom central body, specified as a double scalar.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Gravitational potential model to
Point-massorOblate ellipsoid (J2).
Programmatic Use
Block Parameter:
customR |
| Type: character vector |
Values:
'3396200' | double scalar |
Default:
'3396200' |
Flattening — Flattening ratio
0.00589 (default) | double scalar
Flattening ratio for custom central body, specified as a double scalar.
Dependencies
To enable this parameter, set:
Central body to
Custom.Gravitational potential model to
Point-massorOblate ellipsoid (J2).
Programmatic Use
Block Parameter:
customF |
| Type: character vector |
Values:
'0.00589' | double scalar |
Default:
'0.00589' |
Gravitational parameter — Gravitational parameter
4.305e13 (default) | double scalar
Gravitational parameter for a custom central body, specified as a double scalar.
Dependencies
To enable this parameter, set:
Central body to
Custom.Gravitational potential model to
Point-massorOblate ellipsoid (J2).
Programmatic Use
Block Parameter:
customMu |
| Type: character vector |
Values:
'4.305e13' | double scalar |
Default:
'4.305e13' |
Second degree zonal harmonic (J2) — Most significant or largest spherical harmonic term
1.0826269e-03 (default) | double scalar
Most significant or largest spherical harmonic term, which accounts for oblateness of a celestial body, specified as a double scalar.
Dependencies
To enable this parameter, set:
Propagation method to
Numerical (high precision).Central body to
Custom.Gravitational potential model to
Oblate ellipsoid (J2).
Programmatic Use
Block Parameter:
customJ2 |
| Type: character vector |
Values:
'1.0826269e-03' | double scalar |
Default:
'1.0826269e-03' |
Units
Units — Parameter and port units
Metric (m/s) (default) | Metric (km/s) | Metric (km/h) | English (ft/s) | English (kts)
Parameter and port units, specified as:
| Units | Distance Units | Velocity Units | Acceleration Units |
|---|---|---|---|
Metric (m/s) | meters | meters/sec | meters/sec2 |
Metric (km/s) | kilometers | kilometers/sec | kilometers/sec2 |
Metric (km/h) | kilometers | kilometers/hour | kilometers/hour2 |
English (ft/s) | feet | feet/sec | feet/sec2 |
English (kts) | nautical mile | knots | knots/sec |
Programmatic Use
Block Parameter:
units |
| Type: character vector |
Values:
'Metric (m/s)' | 'Metric (km/s)' |
'Metric (km/h)' | 'English (ft/s)' |
'English (kts)' |
Default:
'Metric (m/s)' |
Angle units — Angle units
Degrees (default) | Radians
Parameter and port units for angles, specified as
Degrees or Radians.
Programmatic Use
Block Parameter:
angleUnits |
| Type: character vector |
Values:
'Degrees' | 'Radians' |
Default:
'Degrees' |
Time format — Time format for start date and time output
Julian date (default) | Gregorian
Time format for Start date/time (UTC Julian date) and output
port tutc, specified as
Julian date or Gregorian.
Programmatic Use
Block Parameter:
timeFormat |
| Type: character vector |
Values:
'Julian date' | 'Gregorian' |
Default:
'Julian date' |
Algorithms
The Orbit Propagator block works in the ICRF and fixed-frame coordinate systems:
ICRF — International Celestial Reference Frame. This frame can be treated as equal to the ECI coordinate system realized at J2000 (Jan 1 2000 12:00:00 TT. For more information, see ECI Coordinates.
Fixed-frame — Fixed-frame is a generic term for the coordinate system that is fixed to the central body (its axes rotate with the central body and are not fixed in inertial space).
When Propagation method is
Numerical (high precision), Central Body isEarth, and the Use Earth orientation parameters (EOPs) check box is selected, the Fixed-frame for Earth is the International Terrestial Reference Frame (ITRF). This reference frame is realized by the IAU2000/2006 reduction from the ICRF coordinate system using the earth orientation parameter file provided. If the Use Earth orientation parameters (EOPs) check box is cleared, the block still uses the IAU2000/2006 reduction, but with Earth orientation parameters set to0.When Propagation method is
High precision (numerical), Central Body isMoon, and the Input Moon libration angles check box is selected, the fixed-frame coordinate system for the Moon is the Mean Earth/pole axis frame (ME). This frame is realized by two transformations. First, the values in the ICRF frame are transformed into the Principal Axis system (PA), the axis defined by the libration angles provided as inputs to the block. For more information, see Moon Libration. The states are then transformed into the ME system using a fixed rotation from the Report of the IAU/IAG Working Group on cartographic coordinates and rotational elements: 2006. If Input Moon libration angles check box is cleared, the fixed frame is defined by the directions of the poles of rotation and prime meridians defined in the Report of the IAU/IAG Working Group on cartographic coordinates and rotational elements: 2006.When Propagation method is
Numerical (high precision)and Central Body isCustom, the fixed-frame coordinate system is defined by the poles of rotation and prime meridian defined by the block input α, δ, W, or the spin axis properties.
In all other cases, the fixed frame for each central body is defined by the directions of the poles of rotation and prime meridians defined in the Report of the IAU/IAG Working Group on cartographic coordinates and rotational elements: 2006.
References
[1] Vallado, David. Fundamentals of Astrodynamics and Applications, 4th ed. Hawthorne, CA: Microcosm Press, 2013.
[2] Gottlieb, R. G., "Fast Gravity, Gravity Partials, Normalized Gravity, Gravity Gradient Torque and Magnetic Field: Derivation, Code and Data," Technical Report NASA Contractor Report 188243, NASA Lyndon B. Johnson Space Center, Houston, Texas, February 1993.
[3] Konopliv, A. S., S. W. Asmar, E. Carranza, W. L. Sjogen, D. N. Yuan., "Recent Gravity Models as a Result of the Lunar Prospector Mission, Icarus", Vol. 150, no. 1, pp 1–18, 2001.
[4] Lemoine, F. G., D. E. Smith, D.D. Rowlands, M.T. Zuber, G. A. Neumann, and D. S. Chinn, "An improved solution of the gravity field of Mars (GMM-2B) from Mars Global Surveyor", Journal Of Geophysical Research, Vol. 106, No. E10, pp 23359-23376, October 25, 2001.
[5] Seidelmann, P.K., Archinal, B.A., A’hearn, M.F. et al. Report of the IAU/IAG Working Group on cartographic coordinates and rotational elements: 2006. Celestial Mech Dyn Astr 98, 155–180 (2007).