Orbit Propagator Numerical (high precision)
Propagate orbit of one or more spacecraft using Kepler universal variable formulation
Since R2020b
Libraries:
Aerospace Blockset /
Spacecraft /
Spacecraft Dynamics
Alternative Configurations of Orbit Propagator Numerical (high precision) Block:
Orbit Propagator Kepler (unperturbed)
Description
The Orbit Propagator blocks propagate the orbit of one or more spacecraft by a propagation method. The Orbit Propagator Numerical (high precision) and Orbit Propagator Kepler (unperturbed) blocks are alternative configurations of the same block.
Orbit Propagator Kepler (unperturbed) — Kepler universal variable formulation (quicker)
Orbit Propagator Kepler Numerical (high precision) — More accurate
The size of the provided initial conditions determines the number of spacecraft being modeled. If you supply more than one value for a parameter in the Orbit tab, the block outputs a constellation of satellites. Any parameter with a single provided value is expanded and applied to all the satellites in the constellation. For example, if you provide a single value for all the parameters on the block except True anomaly, which contains six values, the block creates a constellation of six satellites, varying true anomaly only.
The block applies the same expansion behavior to input port A_icrf (applied acceleration). This port accepts either a single value expanded to all spacecraft being modeled, or individual values to apply to each spacecraft.
For more information on the propagation methods the Orbit Propagator blocks use, see Orbit Propagation Methods.
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.
For more information on the coordinate systems the Orbit Propagator blocks use, see Coordinate Systems.
Atmospheric Drag
To help model the drag on spacecraft for high precision orbit propagation, the Orbit Propagator Numerical (high precision) block supports drag. Atmospheric drag affects spacecraft flying at low Earth orbit (LEO); it is less relevant further away from Earth. For the atmospheric drag equation, see Atmospheric Drag.
Third Body Gravity Effects
To include the effects of point-mass Third Body gravity, select which third bodies to include. These calculations require an Ephemeris model and data. For the third body contribution equation, see Third Body Gravity.
Solar Radiation Pressure
To include the effects of solar radiation pressure (SRP) on the spacecraft, provide an eclipse fraction or specify a shadow model to compute the fraction. Consider taking solar radiation pressure into account when atmosphere drag is negligible and pointing accuracy requirements are strict. For the solar radiation pressure equation, see Solar Radiation Pressure.
Examples
Constellation Modeling with the Orbit Propagator Block
Propagate the orbits of a constellation of satellites and compute and visualize access intervals between the individual satellites and several ground stations.
Mission Analysis with the Orbit Propagator Block
Compute and visualize line-of-sight access intervals between satellites and a ground station.
Lunar Mission Analysis with the Orbit Propagator Block
Compute and visualize access intervals between the Apollo Command and Service module and a rover on the lunar surface. The module's orbit is modeled using Reference Trajectory #2 from the NASA report Variations of the Lunar Orbital Parameters of the Apollo CSM-Module [2]. This is a lunar orbit studied by NASA for the Apollo program. The example uses:
High Precision Orbit Propagation of the International Space Station
Propagate the order of the International Space Station (ISS) using high precision numerical orbit propagation.
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
m — Spacecraft mass used by atmospheric drag calculation
scalar | vector of size numSat
Spacecraft mass used by atmospheric drag calculation, specified as scalar or vector of size numSat. numSat is the number of spacecraft.
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Mass source parameter to
Dialog
.
Data Types: double
ρ — Atmospheric density
scalar
Atmospheric density to calculate acceleration due to atmospheric drag.
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Port
.
Data Types: double
F107a — 81-day average Ottawa F10.7 cm solar flux
scalar
81-day average Ottawa F10.7 cm solar flux, centered on the current day specified in Start date/time. These F107 Average values correspond to the 10.7 cm radio flux at the actual distance of the Earth from the Sun. This site provides both classes of values:
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.
Data Types: double
F107 — Daily Ottawa F10.7 cm solar flux
scalar
Daily Ottawa F10.7 cm solar flux, centered on the current day specified in Start date/time. The f107Daily values do not correspond to the radio flux at 1 AU. This site provides both classes of values:
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.
Data Types: double
aph — Daily magnetic index information
N-by-7 array
Daily magnetic index information (aph), specified as an N-by-7 array. The magnetic index information consists of:
Daily magnetic index (AP) |
3 hour AP for current time |
3 hour AP for 3 hours before current time |
3 hour AP for 6 hours before current time |
3 hour AP for 9 hours before current time |
Average of eight 3 hour AP indices from 12 to 33 hours before current time |
Average of eight 3 hour AP indices from 36 to 57 hours before current time |
The effects of daily magnetic index are not large or established below 80,000 m. For more information, see Limitations on NRLMSISE-00 Atmosphere Model.
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.
Data Types: double
Flags — Variation flags
array of 23
Variation flags, specified as an array of 23, to enable or disable particular
variations for the outputs. You can specify one of the following values for a field.
The default value for each field is 1
.
0.0
Removes the value effect on the output.
1.0
Applies the main and the cross-term effects of that value on the output.
2.0
Applies only the cross-term effect of that value on the output.
Field | Description |
---|---|
Flags(1)
| F10.7 effect on mean |
Flags(2)
| Independent of time |
Flags(3)
| Symmetrical annual |
Flags(4)
| Symmetrical semiannual |
Flags(5)
| Asymmetrical annual |
Flags(6)
| Asymmetrical semiannual |
Flags(7)
| Diurnal |
Flags(8)
| Semidiurnal |
Flags(9)
| Daily AP. If you set this field to -1, the block uses the entire
matrix of magnetic index information (APH) instead of
|
Flags(10)
| All UT, longitudinal effects |
Flags(11)
| Longitudinal |
Flags(12)
| UT and mixed UT, longitudinal |
Flags(13)
| Mixed AP, UT, longitudinal |
Flags(14)
| Terdiurnal |
Flags(15)
| Departures from diffusive equilibrium |
Flags(16)
| All exospheric temperature variations |
Flags(17)
| All variations from 120,000 meter temperature (TLB) |
Flags(18)
| All lower thermosphere (TN1) temperature variations |
Flags(19)
| All 120,000 meter gradient (S) variations |
Flags(20)
| All upper stratosphere (TN2) temperature variations |
Flags(21)
| All variations from 120,000 meter values (ZLB) |
Flags(22)
| All lower mesosphere temperature (TN3) variations |
Flags(23)
| Turbopause scale height variations |
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set the Flags source parameter to
Port
.
Data Types: double
Cd — Atmospheric drag coefficient
scalar | vector of size numSat
Atmospheric drag coefficient, specified as a scalar or vector of size numSat.
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Drag coefficient source parameter to
Port
.
Data Types: double
Ad — Atmospheric drag area
scalar | vector of size numSat
Atmospheric drag area, specified as a scalar or vector of size numSat.
Dependencies
To enable this port:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Drag area source parameter to
Port
.
Data Types: double
Rcb, 3 — Position of one or more custom bodies
3-element vector | numCustom3rdBodiesx3-element vector
Position of one or more custom bodies with respect to the current central body, in the ICRF coordinate frame. numCustom3rdBodies is the number of custom bodies. For more information, see Custom gravitational parameter (m^3/s^2).
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision)
.Select the Include third body point-mass gravity check box.
Select the Custom check box.
Data Types: double
Fraction — Fraction of solar disk
between 0
(full eclipse) and 1
(full sunlight) | scalar | numSat
Fraction of solar disk visible from the spacecraft position. numSat is the number of spacecraft.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision)
.Select the Include solar radiation pressure (SRP) check box.
Set Eclipse fraction source to
Port
.
Data Types: double
RC — Reflectivity coefficient
scalar or numSat of values in the range
1
≤RC≤2
Reflectivity coefficient, specified as a scalar or numSat of values in the
range
1
≤RC≤2
.
numSat is the number of spacecraft.
Dependencies
To enable this port:
Set Propagation method to
Numerical (high precision)
.Select the Include solar radiation pressure (SRP) check box.
Set the Reflectivity coefficient source parameter to
Port
.
Data Types: double
Asrp — Cross section area of the spacecraft
scalar | numSat
Cross section area of the spacecraft seen by the Sun, specified as a scalar or numSat. The block uses this value to calculate the acceleration due to solar radiation pressure.
Dependencies
To enable this port:
Set Propagation method parameter to
Numerical (high precision)
.Select the Include solar radiation pressure (SRP) check box.
Set SRP area source parameter to
Port
.
Data Types: double
Output
Xicrf — Position of spacecraft
3-element vector | numSat-by-3
Position of the spacecraft with respect to (ICRF or fixed-frame), returned as a 3-element vector or numSat-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. numSat is the number of spacecraft.
Data Types: double
Vicrf — Velocity
3-element vector | numSat-by-3 array
Velocity of the spacecraft with respect to ICRF or fixed-frame, returned as a 3-element vector or numSat-by-3 array, at the current time step. numSat-by-3 array. The size of the initial conditions provided in the Orbit tab control the port dimension.
Data Types: double
qicrf2ff — Transformation
4-element quaternion (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)
| 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
Numerical (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 vector 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 (2020, 1, 1, 12, 0, 0)
(default) | valid scalar Julian date | valid Gregorian date including year, month, day, hours, minutes, seconds as
6-element vector
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' |
Mass source — Mass source
Dialog
(default) | Port
Mass source, specified as Dialog
or
Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Programmatic Use
Block Parameter:
massSrc |
Type: character vector |
Values:
'Dialog' | 'Port' |
Default:
'Dialog' |
Mass — Spacecraft mass used by atmospheric drag calculation
4.0
(default) | scalar | vector of size numSat
Spacecraft mass used by atmospheric drag calculation, specified as scalar or vector of size numSat. numSat is the number of spacecraft.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Mass source to
Dialog
.
Programmatic Use
Block Parameter:
mass |
Type: character vector |
Values: scalar | vector of size numSat |
Default:
'4.0' |
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 in the Diagnostic Viewer, model simulation continues. |
Error (default) | Error in the Diagnostic Viewer, 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
.
Programmatic Use
Block Parameter
stateFormatNum |
Type: character vector |
Values:
'Orbital elements' | 'ICRF state vector' |
'Fixed-frame state vector' |
Default:
'Orbital elements' |
Orbit Type — Orbit classification
Keplerian
(default) | Elliptical equatorial
| Circular
| Circular equatorial
Orbit classification, specified as:
Keplerian
— Model elliptical orbits using six standard Keplerian orbital elements.Elliptical equatorial
— 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.Circular equatorial
— Define a circular orbit, where eccentricity is 0 or 10 degrees. Argument of periapsis and the right ascension of the ascending node are 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) | 1D array of size numSat
Half of ellipsis major axis, specified as a 1D 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 | 1D array of size m, 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 | 1D array of size numSat
Deviation of the orbit from a perfect circle, specified as a scalar or 1D array of size that is number of spacecraft.
If Orbit type is set to Keplerian
,
value can be:
0
for circular orbitBetween
0
and1
for elliptical orbit1
for parabolic orbitGreater than
1
for hyperbolic orbit
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements
.Orbit type to
Keplerian
orElliptical 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 | 1D array of size numSat |
Default:
'0.01' |
Inclination (deg) — Tilt angle of orbital plane
50 (default) | scalar | 1D array of size numSat | 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 1D array of size numSat, in specified units. numSat is the number of spacecraft.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements
Orbit type to
Keplerian
orCircular inclined
Programmatic Use
Block Parameter:
inclination |
Type: character vector |
Values: 50 | scalar | 1D array of size numSat | 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
| 1D array of size numSat
Right ascension of ascending node (RAAN), specified as a scalar value between
0
and 360
or 1D array of size
numSat, in specified units. numSat is the
number of spacecraft. 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
Keplerian
orCircular inclined
.
Programmatic Use
Block Parameter:
raan |
Type: character vector |
Values:
95 | scalar value between 0 and
360 | 1D array of size m number of
spacecraft |
Default:
'95' |
Argument of periapsis (deg) — Angle from spacecraft ascending node to periapsis
93 (default) | degrees between 0
and 360
| radians between 0
and 2*pi
| 1D 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 1D array of size m that is number of spacecraft, in specified units.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements
Orbit type to
Keplerian
Programmatic Use
Block Parameter:
argPeriapsis |
Type: character vector |
Values:
'95' | scalar value between 0 and
360 | 1D array of size numSat |
Default:
'93' |
True anomaly — Angle between periapsis and initial position of spacecraft
203 (default) | scalar | degrees between 0
and 360
| radians between 0
and 2*pi
| 1D array of size numSat
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 1D array of size numSat, in specified units. numSat is the number of spacecraft.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements
.Orbit type to
Keplerian
orElliptical inclined
.
Programmatic Use
Block Parameter:
trueAnomaly |
Type: character vector |
Values:
'203' | scalar | degrees between 0 and
360 | radians between 0 and
2*pi | 1D array of size numSat |
Default:
'203' |
Argument of latitude (deg) — Angle between ascending node and initial position of spacecraft
200 (default) | scalar | degrees between 0
and 360
| radians between 0
and 2*pi
| 1D array of size numSat
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 1D 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 | degrees between 0 and
360 | radians between 0 and
2*pi | 1D array of size numSat |
Default:
'200' |
Longitude of periapsis (deg) — Angle between ICRF x-axis and eccentricity vector
100 (default) | scalar | degrees between 0
and 360
| radians between 0
and 2*pi
| 1D array of size numSat
Angle between the ICRF x-axis and the eccentricity vector, specified as a scalar or 3-element vector or 1D array of size 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:
lonPeriapsis |
Type: character vector |
Values: 100 | scalar | degrees between
0 and 360 | radians between
0 and 2*pi | 1D array of size
m, number of spacecraft |
Default:
'100' |
True longitude (deg) — Angle between ICRF x-axis and initial position of spacecraft
150 (default) | scalar | degrees between 0
and 360
| radians between 0
and 2*pi
| 1D array of size numSat
Angle between the ICRF x-axis and the initial position of spacecraft along its orbit at Start date/time, specified as a scalar or 1D array of size numSat, in specified units. numSat is the number of spacecraft.
Dependencies
To enable this parameter, set:
Initial state format to
Orbital elements
.Orbit type to
Circular equatorial
.
Programmatic Use
Block Parameter:
trueLon |
Type: character vector |
Values:
'150' | scalar | degrees between 0 and
360 | radians between 0 and
2*pi | 1D array of size numSat |
Default:
'150' |
ICRF position — Cartesian position vector of spacecraft
[3649700.0 3308200.0 -4676600.0]
(default) | 3-element vector | | numSat-by-3 array
Cartesian position vector of spacecraft in ICRF coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or numSat-by-3 array for multiple spacecraft. numSat is the number of 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
Configure the central body environment around which the spacecraft orbits.
Central body — Celestial body around which spacecraft orbits
Earth
(default) | Moon
| Mercury
| Venus
| Mars
| Jupiter
| Saturn
| Uranus
| Neptune
| Sun
| Custom
Celestial body, specified as Earth
,
Moon
, Mercury
,
Venus
, Mars
,
Jupiter
, Saturn
,
Uranus
, Neptune
,
Sun
, 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' | 'Sun' |
'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 |
| — | Maximum degree is extracted from the spherical harmonic coefficient file. |
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
.
Additionally, it must satisfy one of these criteria:
Gravitational potential model is set to either
Spherical harmonics
orOblate ellipsoid (J2)
.External acceleration coordinate frame is set to
Fixed-frame
.State vector output coordinate frame is set to
Fixed-frame
.
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
None
,Point-mass
, orOblate 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-mass
,Oblate ellipsoid (J2)
, orSpherical harmonics
.
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
None
,Point-mass
, orOblate 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' |
Drag
Configure the atmospheric drag environment.
Include atmospheric drag — Option to include atmospheric drag
off (default) | on
To include atmospheric drag, select this check box.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.
Programmatic Use
Block Parameter:
useDrag |
Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
Atmospheric density source — Source of atmospheric density value
Dialog
(default) | Port
Source of atmospheric density value, specified as
Dialog
or Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Programmatic Use
Block Parameter:
atmosSrc |
Type: character vector |
Values:
'Dialog' | 'Port' |
Default:
'Dialog' |
Atmospheric model — Atmospheric model
NRLMSISE-00
(default)
Atmospheric model for atmospheric drag calculation, specified as
NRLMSISE-00
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.
Programmatic Use
Block Parameter:
atmosModel |
Type: character vector |
Values:
'NRLMSISE-00' |
Default:
'NRLMSISE-00' |
Flux source — Source of space weather data
Dialog
(default) | Port
Source for historical and predicted flux and geomagnetic indices used by
atmospheric density calculation, specified as Dialog
or Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.
Programmatic Use
Block Parameter:
fluxSrc |
Type: character vector |
Values:
'Dialog' | 'Port' |
Default:
'Dialog' |
Space weather data file — MAT file of space weather data
aeroSpaceWeatherData.mat
(default) | MAT file name
MAT file of space weather data. This file is the output from the aeroReadSpaceWeatherData
function.
Aerospace Blockset™ includes a default space weather data file,
aeroSpaceWeatherData.mat
. To use the most recent data available,
use aeroReadSpaceWeatherData
to generate a new MAT file, and
specify the file name for this parameter. For more information, see aeroReadSpaceWeatherData
.
If the file is not on the MATLAB® path, specify the full pathname. The MAT file must contain these variables from the space weather data file:
YEAR
MONTH
DAY
AP1
AP2
AP3
AP4
AP5
AP6
AP7
AP8
AP_AVG
F107_OBS
F107_DATA_TYPE
F107_OBS_CENTER81
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set Atmospheric model to
NRLMSISE-00
.
Programmatic Use
Block Parameter:
SpaceWeatherDataFile |
Type: character vector |
Values:
'aeroSpaceWeatherData.mat' | MAT file name |
Default:
'aeroSpaceWeatherData.mat' |
F10.7 extrapolation method — Extrapolation method for f107Average
and f107Daily
None - clip
(default) | Constant
| Least squares fit
Extrapolation method for f107Average
and f107Daily
for times outside the range of the MAT file data, specified as one of these
values.
Method | Description |
---|---|
None - clip | Set |
Constant | Set |
Least squares fit | Approximate
where:
|
Note
For f107Average
values,if part of the 81-day period falls
inside the MAT file range of space weather data, the Solar Flux and
Geomagnetic Index block uses the actual daily values from the
overlapping portion. To calculate the average for the nonoverlapping portion, the
block uses the clipped, constant, or least squares daily value.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set Atmospheric model to
NRLMSISE-00
.
Programmatic Use
Block Parameter: F107ExtrapMethod |
Type: character vector |
Values: 'None - clip' | 'Constant' | 'Least squares fit' |
Default: 'None - clip' |
F10.7 extrapolation value — Extrapolation value to assign to f107Average
and to calculate f107Daily
150.0
(default) | scalar
Extrapolation value to assign to f107Average
and to calculate f107Daily
, specified as a scalar.
Tunable: Yes
Dependencies
This value is assigned when:
F10.7 extrapolation method is set to
Constant
.Time specified by the Year, DOY, and UT (sec) ports is outside the range of the data in the MAT file.
Programmatic Use
Block Parameter: F107ExtrapValue |
Type: character vector |
Values: '150.0' | scalar |
Default: '150.0' |
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set Atmospheric model to
NRLMSISE-00
.Set F10.7 extrapolation method to
Constant
.
Magnetic index extrapolation method — Extrapolation method for magnetic index information
None - clip
(default) | Constant
| IGRF
Extrapolation method for magnetic index information values for times outside the range of the MAT file data, specified as one of these values.
Method | Description |
---|---|
None - clip | Set elements of magnetic index information to the nearest data point available in the MAT file. |
Constant | Set elements of magnetic index information to a constant value specified by the Magnetic index extrapolation method parameter. The elements of magnetic index for times outside this range are based on clipped values of the horizontal magnetic field strength at these time limits. |
IGRF | Calculate the elements of magnetic index information using the International Geomagnetic Reference Field. Because this model is defined for times between January 1, 1900, 12:00 AM UTC and January 1, 2025, 12:00 AM UTC, the predictions for times outside this range are clipped to values at these time limits. |
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set Atmospheric model to
NRLMSISE-00
.
Programmatic Use
Block Parameter:
MagneticIndexExtrapMethod |
Type: character vector |
Values: 'None - clip' |
'Constant' | 'IGRF' |
Default: 'None -
clip' |
Magnetic index extrapolation value — Extrapolation value used to calculate magnetic index elements
4.0
(default) | scalar
Extrapolation value used to calculate magnetic index elements, specified as a scalar.
Tunable: Yes
Dependencies
This parameter is enabled when Magnetic index extrapolation method is set to Constant
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set Atmospheric model to
NRLMSISE-00
.
Programmatic Use
Block Parameter: MagneticIndexExtrapValue |
Type: character vector |
Values: '4.0' | scalar |
Default: '4.0' |
Flags source — Variation flag source
Dialog
(default) | Port
Variation flag source, specified as Dialog
or
Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set the Atmospheric model parameter to
NRLMSISE-00
.
Programmatic Use
Block Parameter:
fluxFlagsSrc |
Type: character vector |
Values:
'Dialog' | 'Port' |
Default:
'Dialog' |
Flags — Variation flags
ones(1,23)
(default)
Variation flags, specified as an array of 23 (ones(1,23)
).
You can specify one of the following values for a field. The default value for
each field is 1
.
0.0 — Removes the effect on the output.
1.0 — Applies the main and the cross-term effects of that value on the output.
2.0 — Applies only the cross-term effect of that value on the output.
The array has these fields.
Field | Description |
---|---|
Flags(1)
| F10.7 effect on mean |
Flags(2)
| Independent of time |
Flags(3)
| Symmetrical annual |
Flags(4)
| Symmetrical semiannual |
Flags(5)
| Asymmetrical annual |
Flags(6)
| Asymmetrical semiannual |
Flags(7)
| Diurnal |
Flags(8)
| Semidiurnal |
Flags(9)
| Daily AP. If you set this field to -1, the block uses
the entire matrix of magnetic index information (APH)
instead of |
Flags(10)
| All UT, longitudinal effects |
Flags(11)
| Longitudinal |
Flags(12)
| UT and mixed UT, longitudinal |
Flags(13)
| Mixed AP, UT, longitudinal |
Flags(14)
| Terdiurnal |
Flags(15)
| Departures from diffusive equilibrium |
Flags(16)
| All exospheric temperature variations |
Flags(17)
| All variations from 120,000 meter temperature (TLB) |
Flags(18)
| All lower thermosphere (TN1) temperature variations |
Flags(19)
| All 120,000 meter gradient (S) variations |
Flags(20)
| All upper stratosphere (TN2) temperature variations |
Flags(21)
| All variations from 120,000 meter values (ZLB) |
Flags(22)
| All lower mesosphere temperature (TN3) variations |
Flags(23)
| Turbopause scale height variations |
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Flags source to
Dialog
.Set the Atmospheric density source parameter to
Dialog
.Set the Atmospheric model parameter to
NRLMSISE-00
.
Programmatic Use
Block Parameter:
fluxFlags |
Type: character vector |
Values:
'ones(1,23)' |
Default:
'ones(1,23)' |
Include anomalous oxygen in density calculation — Option to include anomalous oxygen in density calculation
off (default) | on
To include anomalous oxygen in density calculations, select this check box.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Atmospheric density source parameter to
Dialog
.Set the Atmospheric model parameter to
NRLMSISE-00
.
Programmatic Use
Block Parameter:
useOxygen |
Type: character vector |
Values:
'off' | 'on' |
Default:
'off' |
Drag coefficient source — Source of drag coefficient
Dialog
(default) | Port
Source of drag coefficient, specified as Dialog
or
Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Programmatic Use
Block Parameter:
dragCoeffSrc |
Type: character vector |
Values:
'Dialog' | 'Source' |
Default:
'Dialog' |
Drag coefficient — Spacecraft coefficient of drag
2.179 (default) | scalar | vector of size numSat
Spacecraft coefficient of drag used by atmospheric drag calculation, specified as a scalar or as a vector of size numSat.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Drag coefficient source parameter to
Dialog
.
Programmatic Use
Block Parameter:
dragCoeff |
Type: character vector |
Values: scalar | vector of size numSat |
Default:
'2.179' |
Drag area source — Source of drag area
Dialog
(default) | Port
Source of drag area, specified as Dialog
or
Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Programmatic Use
Block Parameter:
dragAreaSrc |
Type: character vector |
Values:
'Dialog' | 'Source' |
Default:
'Dialog' |
Drag area — Area to compute acceleration due to atmospheric drag
1.0 (default) | scalar | vector of size numSat
Area to compute acceleration due to atmospheric drag, specified as a scalar or as a vector of size numSat. This area of the spacecraft is perpendicular to the spacecraft relative velocity.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to
Earth
.Select the Include atmospheric drag check box.
Set the Drag area source parameter to
Dialog
.
Programmatic Use
Block Parameter:
dragArea |
Type: character vector |
Values: scalar | vector of size numSat |
Default:
'1.0' |
Third Body
Configure third body point mass for gravitational acceleration.
Ephemeris model — Ephemeris model
DE405
(default) | DE421
| DE423
| DE430
| DE432t
Select one of these ephemerides models defined by the Jet Propulsion Laboratory. The block uses ephemeris data to calculate relative celestial positions required for third body point mass gravity and solar radiation pressure.
Ephemeris Model | Description |
---|---|
| Released in 1998. This ephemeris takes into account the Julian date range 2305424.50 (December 9, 1599) to 2525008.50 (February 20, 2201). This block implements these ephemerides with respect to the International Celestial Reference Frame version 1.0, adopted in 1998. |
| Released in 2008. This ephemeris takes into account the Julian date range 2414992.5 (December 4, 1899) to 2469808.5 (January 2, 2050). This block implements these ephemerides with respect to the International Celestial Reference Frame version 1.0, adopted in 1998. |
| Released in 2010. This ephemeris takes into account the Julian date range 2378480.5 (December 16, 1799) to 2524624.5 (February 1, 2200). This block implements these ephemerides with respect to the International Celestial Reference Frame version 2.0, adopted in 2010. |
| Released in 2013. This ephemeris takes into account the Julian date range 2287184.5 (December 21, 1549) to 2688976.5 (January 25, 2650). This block implements these ephemerides with respect to the International Celestial Reference Frame version 2.0, adopted in 2010. |
| Released in April 2014. This ephemeris takes into account the Julian date range 2287184.5, (December 21, 1549 ) to 2688976.5, (January 25, 2650). This block implements these ephemerides with respect to the International Celestial Reference Frame version 2.0, adopted in 2010. |
Note
This block requires that you download ephemeris data using the Add-On Explorer. To
start the Add-On Explorer, in the MATLAB Command Window, type aeroDataPackage
. in the MATLAB desktop toolstrip, click Add-Ons .
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.
Programmatic Use
Block Parameter: ephemerisModel |
Type: character vector |
Values: DE405 | DE421 | DE423 | DE430 |
Default: 'DE405' |
Limit ephemerides date range — Option to enable start and end of range of ephemeris data
off
(default) | on
Control how much data is loaded into memory during simulation and how much data is included in generated code for the block:
Select this check box to limit the loading of ephemeris data to a specified date range.
Clear this check box to include data for the complete date range defined in the Ephemeris model table.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.
Programmatic Use
Block Parameter:
useDateRange |
Type: character vector |
Values: 'off' |
'on'
|
Default: 'off' |
Start date (JD) — Start date of ephemerides date range
juliandate(2020, 1, 1)
(default) | Julian date
Start date of ephemerides date range, specified as a Julian date.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Select the Limit ephemerides date range parameter.
Programmatic Use
Block Parameter: startDate |
Type: character vector |
Values: 'juliandate(2020, 1, 1)' | Julian date |
Default: 'juliandate(2020, 1, 1)' |
End date (JD) — End date of ephemerides date range
juliandate(2050, 1, 1)
(default) | Julian date
End date of ephemerides date range, specified as a Julian date.
Dependencies
To enable this parameter, select the Limit ephemerides date range check box.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Limit ephemerides date range parameter to
on
.
Programmatic Use
Block Parameter: endDate |
Type: character vector |
Values: 'juliandate(2050, 1, 1)' | Julian format date |
Default: 'juliandate(2050, 1, 1)' |
Include third body point-mass gravity — Acceleration due to third body gravity
off
(default) | on
To include acceleration due to third body gravity in orbit propagation calculation, select this check box. Otherwise, clear this check box.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.
Programmatic Use
Block Parameter: useThirdBodyGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Sun — Gravitational acceleration due to Sun
on
(default) | off
To include gravitational acceleration due to the Sun (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Sun
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeSunGravity |
Type: character vector |
Values: 'off' | 'on'
|
Default: 'on' |
Mercury — Gravitational acceleration due to Mercury
off
(default) | on
To include gravitational acceleration due to the Mercury (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Mercury
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeMercuryGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Venus — Gravitational acceleration due to Venus
off
(default) | on
To include gravitational acceleration due to the Venus (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Venus
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeVenusGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Earth — Gravitational acceleration due to Venus
on
(default) | off
To include gravitational acceleration due to the Earth (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Earth
.Set the Include third body point-mass gravity to
on
.
Programmatic Use
Block Parameter: includeVenusGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Moon — Gravitational acceleration due to Moon
on
(default) | off
To include gravitational acceleration due to the Moon (point mass), select this check box. Otherwise, clear it.
Programmatic Use
Block Parameter: includeMoonGravity |
Type: character vector |
Values: 'off' | 'on' | |
Default: 'on' |
Mars — Gravitational acceleration due to Mars
off
(default) | on
To include gravitational acceleration due to the Mars (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Mars
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeMarsGravity |
Type: character vector |
Values: 'off' | 'on' | |
Default:
'off' |
Jupiter — Gravitational acceleration due to Jupiter
off
(default) | on
To include gravitational acceleration due to the Jupiter (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Jupiter
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeJupiterGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Saturn — Gravitational acceleration due to Saturn
off
(default) | on
To include gravitational acceleration due to the Saturn (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Saturn
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeSaturnGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Uranus — Gravitational acceleration due to Uranus
off
(default) | on
To include gravitational acceleration due to the Uranus (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Uranus
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeUranusGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Neptune — Gravitational acceleration due to Neptune
off
(default) | on
To include gravitational acceleration due to the Neptune (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Neptune
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeNeptuneGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Pluto — Gravitational acceleration due to Pluto
off
(default) | on
To include gravitational acceleration due to the Pluto (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Central Body parameter to any value other than
Pluto
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includePlutoGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Custom — Gravitational acceleration due to custom planet
off
(default) | on
To include gravitational acceleration due to a custom planet (point mass), select this check box. Otherwise, clear it.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include third body point-mass gravity parameter to
on
.
Programmatic Use
Block Parameter: includeCustomGravity |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Custom gravitational parameter (m^3/s^2) — Custom gravitational acceleration
42.828314258067e12
(default) | scalar | vector of length numCustom3rdBodies
Custom gravitational acceleration for custom third body, specified as a scalar or as a vector of length numCustom3rdBodies. numCustom3rdBodies is the number of rows provided to the R_cb input port. Provide more than one value to include multiple custom bodies.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include third body point-mass gravity parameter to
on
.Set the Custom parameter to
on
.
Programmatic Use
Block Parameter: customThirdBodyMu |
Type: character vector |
Values: '42.828314258067e12' | scalar | vector of length numCustom3rdBodies |
Default: '42.828314258067e12' |
SRP
Include solar radiation pressure (SRP) — Acceleration due to solar radiation pressure
off
(default) | on
To include acceleration due to solar radiation pressure in orbit propagation calculation, select this check box. Otherwise, clear the check box.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.
Programmatic Use
Block Parameter: useSRP |
Type: character vector |
Values: 'off' |
'on' |
Default: 'off' |
Eclipse fraction source — Source of eclipse fraction
Dialog
(default) | Port
Source of the eclipse fraction (fraction of solar disk visible from the spacecraft location), specified as Dialog
or Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include solar radiation pressure (SRP) parameter to
on
.
Programmatic Use
Block Parameter: useSRP |
Type: character vector |
Values: 'Dialog' |
'Port' |
Default: 'Dialog' |
Shadow model — Shadow model
Dual cone
(default) | Cylindrical
Shadow model for eclipse calculations, specified as one of these values.
Cylindrical
— Fraction can be 0.0 (Umbra) or 1.0 (Sunlight).Dual cone
— Fraction can be 0.0 (Umbra), between 0.0 and 1.0 (Penumbra or Antumbra), or 1.0 (Sunlight).
Programmatic Use
Block Parameter: shadowModel |
Type: character vector |
Values: 'Dual cone' | 'Cylindrical' |
Default:
'Dual cone' |
Include Earth — Option to include Earth
on
(default) | off
Option to include Earth as a secondary occulting body in eclipse calculations when central body is Moon.
Dependencies
Set the Central body parameter to
Moon
.
Programmatic Use
Block Parameter:
includeEarth |
Type: character vector |
Values:
'off' | 'on' |
Default:
'on' |
Include Moon — Option to include Moon
on
(default) | off
Option to include Moon as a secondary occulting body in eclipse calculations when central body is Earth.
Dependencies
Set the Central body parameter to
Earth
.
Programmatic Use
Block Parameter:
includeMoon |
Type: character vector |
Values:
'off' | 'on' |
Default:
'on' |
Reflectivity coefficient source — Source for spacecraft coefficient of reflectivity
'Dialog'
(default) | 'Port'
Source for the spacecraft coefficient of reflectivity, specified as
Dialog
or Port
.
Dependencies
To enable this check box:
Set the Propagation method parameter to
Numerical (high precision)
.Select the Include solar radiation pressure (SRP) check box.
Programmatic Use
Block Parameter:
reflectivityCoeffSrc |
Type: character vector |
Values:
'Dialog' | 'Port' |
Default:
'on' |
Reflectivity coefficient — Spacecraft coefficient of reflectivity
1.8
(default) | scalar | 2D array of size numSat | between [1,2]
Spacecraft coefficient of reflectivity used by solar radiation pressure calculation, specified as a scalar, 2D array of size numSat. numSat is the number of spacecraft.
Tunable: Yes
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include solar radiation pressure (SRP) parameter to
on
.Set the Reflectivity coefficient source parameter to
Dialog
.
Programmatic Use
Block Parameter:
reflectivityCoeff |
Type: character vector |
Values:
'1.8' | scalar | 2D array of size numSat |
between [1,2] |
Default:
'1.8' |
SRP area source — Source for spacecraft solar radiation pressure area
Dialog
(default) | Port
Source for the spacecraft solar radiation pressure (SRP) area, specified as
Dialog
or Port
.
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include solar radiation pressure (SRP) parameter to
on
.
Programmatic Use
Block Parameter: srpAreaSrc |
Type: character vector |
Values: 'Dialog' | 'Port' |
Default:
'Dialog' |
SRP area — Cross section area of the spacecraft seen by the Sun
1.0
(default) | scalar | 2D array of size numSat
Cross section area of the spacecraft seen by the Sun, specified as a scalar or as a 2D array of size numSat. numSat is the number of spacecraft.
Tunable: Yes
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include solar radiation pressure (SRP) parameter to
on
.Set the SRP area source parameter to
Dialog
.
Programmatic Use
Block Parameter:
srpArea |
Type: character vector |
Values:
'1.0' | scalar | 2D array of
size numSat |
Default:
'1.0' |
Solar flux pressure (W*s/m^3) — Spacecraft solar flux pressure
4.5344321e-6
(default) | scalar
Solar flux pressure, specified as a scalar.
Tunable: Yes
Dependencies
To enable this parameter:
Set the Propagation method parameter to
Numerical (high precision)
.Set the Include solar radiation pressure (SRP) parameter to
on
.
Programmatic Use
Block Parameter: fluxPressure |
Type: character vector |
Values: 4.5344321e-6 | scalar |
Default: '4.5344321e-6' |
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 | Velocity | Acceleration | Mass | Area | Density |
---|---|---|---|---|---|---|
Metric (m/s) | meters | meters/sec | meters/sec2 | Kilograms | m2 | kg/m3, some density outputs 1/m3 |
Metric (km/s) | kilometers | kilometers/sec | kilometers/sec2 | Kilograms | m2 | kg/m3, some density outputs 1/m3 |
Metric (km/h) | kilometers | kilometers/hour | kilometers/hour2 | Kilograms | m2 | kg/m3, some density outputs 1/m3 |
English (ft/s) | feet | feet/sec | feet/sec2 | Slugs | feet2 | lbm/ft3, some density outputs 1/ft3 |
English (kts) | nautical mile | knots | knots/sec | Slugs | feet2 | lbm/ft3, some density outputs 1/ft3 |
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' |
Alternative Configurations
Orbit Propagator Kepler (unperturbed) — Propagate orbit of one or more spacecraft Kepler universal variable formulation
The Orbit Propagator
Kepler (unperturbed) block propagates the orbit of one or more spacecraft by
using the Kepler universal variable formulation. To enable this block, set
Propagation method to Kepler
(unperturbed)
.
Libraries:
Aerospace Blockset /
Spacecraft /
Spacecraft Dynamics
Algorithms
Coordinate Systems
The Orbit Propagator Numerical (high precision) block works in the ICRF and fixed-frame coordinate systems:
ICRF — International Celestial Reference Frame (ICRF). 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).
To describe a point in space, you need a frame of reference that does not rotate with respect to the stars. The ICRF, with the origin at the center of the Earth and orthogonal vectors I, J, and K, is used as the frame of reference. The fundamental plane is the IJ-plane, which is closely aligned with the equator with a small offset that changes over time because of precession and nutation of the rotation axis of the Earth.
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). For high precision orbit propagation methods
When the central body is earth, and the Earth orientation parameters (EOPs) are used, the Fixed-frame for Earth is the International Terrestrial 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 Earth orientation parameters are not used, the block still uses the IAU2000/2006 reduction, but with Earth orientation parameters set to
0
.When the central body is moon, and the moon libration angles are provided as input, 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" [5]. If the Moon libration angles input is not provided, 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" [5].
When the Central Body is custom, 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" [5].
Orbit Propagation Methods
The Aerospace Blockset supports two top-level orbit propagation methods:
Kepler (unperturbed)
and Numerical (high
precision)
.
For more information, see Orbit Propagation Methods.
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).
[6] Montenbruck, Oliver, and Gill Eberhard. Satellite Orbits: Models, Methods, and Applications. Springer, 2000.
Extended Capabilities
C/C++ Code Generation
Generate C and C++ code using Simulink® Coder™.
Version History
Introduced in R2020bR2023b: Orbit Propagator Block Changes
The Orbit Propagator block has been updated to take into account:
The effects of third body gravity on orbit propagation.
Space weather data from a data file generated by the
aeroReadSpaceWeatherData
function.Solar radiation pressure (SRP).
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