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Spacecraft Dynamics

Model dynamics of one or more spacecraft

Since R2021b

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  • Spacecraft Dynamics block icon

Libraries:
Aerospace Blockset / Spacecraft / Spacecraft Dynamics

Description

The Spacecraft Dynamics block models translational and rotational dynamics of spacecraft using numerical integration. It computes the position, velocity, attitude, and angular velocity of one or more spacecraft over time. For the most accurate results, use a variable step solver with low tolerance settings (less than 1e-8). To trade off accuracy for speed, use larger tolerances, depending on your mission requirements.

You can define initial orbital states as

  • A set of orbital elements.

  • Position and velocity state vectors.

To propagate orbital states, the block uses the gravity model selected for the current central body. It also includes external accelerations and forces that you provide as inputs to the block. To define initial attitude states, use quaternions, direction cosine matrices (DCMs), or Euler angles.

To propagate attitude states, the block uses moments provided as inputs to the block and mass properties defined on the block.

Aerospace Blockset™ uses quaternions that are defined using the scalar-first convention.

The Spacecraft Dynamics block supports scalar and vector expansion. The block parameter and input port dimensions determine the number of the output signals and the number of spacecraft. After scalar and vector expansion, all parameters in the Orbit, Mass, and Attitude tabs and all input ports except for φθψ (Moon libration angles) and αδW (right ascension, declination, and rotation angle) input ports are defined for each spacecraft.

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, Attitude, or Mass tabs, 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 the block input ports. All input ports support expansion except Moon libration angles (when Central body is Moon) and Spin axis right ascension (RA) at J2000, Spin axis declination (Dec) at J2000, and Initial rotation angle at J2000 (when Central body is Custom). All other ports accept either a single value expanded to all spacecraft being modeled, or individual values applied to each spacecraft.

For more information on the coordinate systems and rotational and translational dynamics the Spacecraft Dynamics block uses, see Algorithms.

Atmospheric Drag

To help model the drag on spacecraft for high precision orbit propagation, the Spacecraft Dynamics 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 Translational Dynamics.

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 Translational Dynamics.

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 Translational Dynamics.

Examples

Getting Started with the Spacecraft Dynamics Block

Getting Started with the Spacecraft Dynamics Block

Model six degree-of-freedom rigid-body dynamics of a spacecraft or constellation of spacecraft with the Spacecraft Dynamics block from Aerospace Blockset™.

Developing the Apollo Lunar Module Digital Autopilot

Developing the Apollo Lunar Module Digital Autopilot

Develop Apollo Lunar Module digital autopilot using Simulink® and Aerospace Blockset.

Analyzing Spacecraft Attitude Profiles with Satellite Scenario

Analyzing Spacecraft Attitude Profiles with Satellite Scenario

Propagate orbit and attitude states in Simulink and visualize computed trajectory and attitude profile in a satellite scenario.

Hohmann Transfer with the Spacecraft Dynamics Block

Hohmann Transfer with the Spacecraft Dynamics Block

Model a Hohmann transfer of a spacecraft between two circular coplanar orbits.

Ports

Input

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Force applied to the spacecraft center of mass in the body frame, specified as a 3-element vector or numSat-by-3 array at the current time step. numSat is the number of spacecraft.

Dependencies

To enable this port, select the Input body forces parameter.

Data Types: double

Moment applied to the spacecraft with respect to mass in the body frame, specified as a 3-element vector or numSat-by-3 array at the current time step. numSat is the number of spacecraft.

Dependencies

To enable this port, select the Input body moments parameter.

Data Types: double

External acceleration to apply to the spacecraft with respect to the ICRF or fixed-frame at the current timestep, specified as a 3-element vector or m-by-3 array.

Dependencies

To enable this port, select the Input external accelerations parameter.

To specify the acceleration coordinate frame, set the External acceleration coordinate frame parameter.

Data Types: double

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 Central body to Moon.

  • Select the Input Moon libration angles parameter.

Data Types: double

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 Central body to Custom.

  • Set Central body spin axis source to Port.

Data Types: double

Spacecraft mass at the current timestep. numSat is the number of spacecraft.

Dependencies

To enable this port, set Mass type to Custom Variable.

Data Types: double

Rate of change of mass (positive if accreted, negative if ablated) at the current timestep, specified as a scalar or 1D array of size numSat. numSat is the number of spacecraft.

Dependencies

To enable this port, set Mass type to Simple Variable.

Data Types: double

Spacecraft inertia tensor, specified as a 3-by-3 array or 3-by-3-by-numSat array at the current timestep. numSat is the number of spacecraft.

Dependencies

To enable this port, set Mass type to Custom Variable.

Data Types: double

Rate of change of inertia tensor matrix, specified as a 3-by-3 array or 3-by-3-by-numSat array at the current time step. numSat is the number of spacecraft.

Dependencies

To enable this port, set Mass type to Custom Variable.

Data Types: double

Relative velocity at which the mass is accreted to or ablated from the body in body-fixed axes, specified as a 3-element vector or numSat-by-3 array. numSat is the number of spacecraft.

Dependencies

To enable this port:

  • Set Mass type to Custom Variable or Simple Variable.

  • Select the Include mass flow relative velocity parameter.

Data Types: double

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 parameter.

  • Set the Atmospheric density source parameter to Port.

Data Types: double

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:

https://www.ngdc.noaa.gov/stp/space-weather/solar-data/solar-features/solar-radio/noontime-flux/penticton/

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 parameter.

  • Set the Atmospheric density source parameter to Dialog.

Data Types: double

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:

https://www.ngdc.noaa.gov/stp/space-weather/solar-data/solar-features/solar-radio/noontime-flux/penticton/

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 parameter.

  • Set the Atmospheric density source parameter to Dialog.

Data Types: double

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 parameter.

  • Set the Atmospheric density source parameter to Dialog.

Data Types: double

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.

FieldDescription
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 APH(:,1)

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 parameter.

  • Set the Atmospheric density source parameter to Dialog.

  • Set the Flags source parameter to Port.

Data Types: double

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 parameter.

  • Set the Drag coefficient source parameter to Port.

Data Types: double

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 parameter.

  • Set the Drag area source parameter to Port.

Data Types: double

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 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

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

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

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Position of the spacecraft with respect to the ICRF or fixed-frame outport coordinate frame, returned as a 3-element vector or numSat-by-3 array at the current time step. numSat is the number of spacecraft.

Dependencies

  • To change the output coordinate frame for this port, set the State vector output coordinate frame parameter.

  • The size of the initial conditions provided in the Mass, Orbit, or Attitude tab control the port dimension.

Data Types: double

Velocity of the spacecraft with respect to the ICRF or fixed-frame outport coordinate frame, returned as a 3-element vector or numSat-by-3 array at the current time step. numSat is the number of spacecraft.

Dependencies

  • To change the output coordinate frame for this port, set the State vector output coordinate frame parameter.

  • The size of the initial conditions provided in the Mass, Orbit, or Attitude tab control the port dimension.

Data Types: double

Total inertial acceleration of the spacecraft with respect to the ICRF, returned as a 3-element vector or numSat-by-3 array at the current timestep. numSat is the number of spacecraft.

Dependencies

  • To enable this port, select the Output total inertial acceleration parameter

  • The size of the initial conditions provided in the Orbit tab control the port dimension.

Data Types: double

Spacecraft attitude quaternion, returned as a (scalar first) quaternion rotation from the body axis to the outport frame, as a 4-element quaternion, or numSat-by-4 array (scalar first) at the current time step. numSat is the number of spacecraft.

Dependencies

The coordinate frame and attitude format of this port depends on these settings:

  • To specify the attitude reference coordinate frame, set the Attitude reference coordinate frame parameter.

  • Set Attitude representation to Quaternion.

Data Types: double

Spacecraft attitude direction cosine matrix (DCM), returned as a3-by-3 array or numSat-by-3-by-3 array. numSat is the number of spacecraft.

Dependencies

The coordinate frame and attitude format of this port depends on these settings:

  • To specify the attitude reference coordinate frame, set the Attitude reference coordinate frame parameter.

  • Set Attitude representation to DCM.

Data Types: double

Spacecraft attitude Euler angles, returned as a 3-element vector or numSat-by-3 array. numSat is the number of spacecraft.

Dependencies

The coordinate frame and attitude format of this port depend on these settings:

  • To specify the attitude reference coordinate frame, set the Attitude reference coordinate frame parameter.

  • Set Attitude representation to Euler angles.

Data Types: double

Angular rate of the spacecraft relative to the attitude reference coordinate frame, returned as a 3-element vector or numSat-by-3 array, expressed as body axis angular rates PQR. numSat is the number of spacecraft.

Dependencies

The attitude reference coordinate frame depends on the Attitude reference coordinate frame parameter.

Data Types: double

Body angular acceleration relative to the ICRF frame, returned as a 3-element array or numSat-by-3 array. numSat is the number of spacecraft.

Dependencies

To enable this port, select the Output total inertial angular acceleration parameter.

The attitude reference coordinate frame depends on the Attitude reference coordinate frame parameter.

Data Types: double

Coordinate system transformation between the ICRF and fixed-frame coordinate system at the current timestep, returned as a 4-element array.

Dependencies

To enable this port, select the Output quaternion (ICRF to Fixed-frame) parameter.

Data Types: double

Time at current time step, returned as a:

  • scalar — If you specify the Start data/time parameter as a Julian date.

  • 6-element array — If you specify the Start data/time parameter as a Gregorian date with six elements (year, month, day, hours, minutes, seconds).

This value equals the Start date/time parameter value plus the elapsed simulation time.

Dependencies

To enable this parameter, select the Output current date/time (UTC Julian date) parameter.

Data Types: double

Fuel tank status at the current timestep, returned as a scalar or numSat-element array, returned as:

  • 1 — Tank is full.

  • 0 — Tank is not full or empty.

  • -1 — Tank is empty.

numSat is the number of spacecraft.

Dependencies

To enable this port:,

  • Set Mass type to Simple Variable.

  • Select the Output fuel tank status parameter.

Data Types: double

Parameters

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Main

To enable external forces to be included in the integration of the spacecraft equations of motion in the body frame, select this parameter. Otherwise, clear this parameter.

Programmatic Use

Block Parameter: forcesin
Type: character vector
Values: 'off' | 'on'
Default: 'off'

To enable external moments to be included in the integration of the spacecraft equations of motion in the body frame, select this parameter. Otherwise, clear this parameter.

Programmatic Use

Block Parameter: momentsIn
Type: character vector
Values: 'off' | 'on'
Default: 'off'

To enable additional external accelerations to be included in the integration of the spacecraft equations of motion, select this parameter. Otherwise, clear this parameter.

Programmatic Use

Block Parameter: accelIn
Type: character vector
Values: 'off' | 'on'
Default: 'off'

Frame for acceleration input port A, specified as ICRF or Fixed-frame.

Dependencies

To enable this parameter, select the Input external accelerations parameter.

Programmatic Use

Block Parameter: accelFrame
Type: character vector
Values: 'ICRF' | 'Fixed-frame'
Default: 'ICRF'

Position and velocity state output port coordinate frame setup, specified as ICRF or Fixed-frame.

Programmatic Use

Block Parameter: outportFrame
Type: character vector
Values: 'ICRF' | 'Fixed-frame'
Default: 'ICRF'

Enable the total acceleration output computed by the block with respect to the ICRF or fixed-frame outport coordinate frame. This acceleration includes all external accelerations, forces, and internal environmental accelerations that act on the spacecraft.

Note

Do not use this port as part of a simulation loop (in other words, do not feed this output back into the block).

Tunable: Yes

Dependencies

To change the output coordinate frame for this port, set the State vector output coordinate frame parameter.

Programmatic Use

Block Parameter: AccelOut
Type: character vector
Values: 'on' | 'off'
Default: 'off'

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.

Tunable: Yes

Dependencies

The data format for this parameter is controlled by the Time format parameter.

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 1D or 6-element array
Default: 'juliandate(2020, 1, 1, 12, 0, 0)'

To output the current date or time, select this parameter. Otherwise, clear this parameter.

Dependencies

The data format for this parameter is controlled by the Time format parameter.

Programmatic Use

Block Parameter: dateOut
Type: character vector
Values: 'off' | 'on'
Default: 'off'

Out-of-range block behavior action. Specify one of these options.

ActionDescription
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'

Mass

Spacecraft mass type, specified as:

  • Fixed — Mass and inertia are constant throughout the simulation.

  • Simple Variable — Mass and inertia vary linearly as a function of mass rate.

  • Custom Variable — Instantaneous mass, inertia, and inertia rate are inputs to the block.

Programmatic Use

Block Parameter: massType
Type: character vector
Values: 'Fixed' | 'Simple Variable' | 'Custom Variable'
Default: 'Fixed'

Initial mass of rigid body spacecraft, specified as scalar or vector of size numSat. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set the Mass type parameter to either Fixed or Simple variable.

Programmatic Use

Block Parameter: mass
Type: character vector
Values: scalar | vector of size numSat
Default: '4.0'

Spacecraft empty (dry) mass, specified as a scalar or vector of size numSat. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: emptyMass
Type: character vector
Values: 1D array of size numSat | 1D array of size numSat
Default: '3.5'

Spacecraft full (wet) mass, specified as a scalar or vector of size numSat. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: fullMass
Type: character vector
Values: scalar | vector of size numSat
Default: '4.0'

Initial inertia tensor matrix of the spacecraft, specified, as a 3-by-3 array for a single spacecraft or a 3-by-3-by-numSat array for multiple spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set Mass type to Fixed.

Programmatic Use

Block Parameter: inertia
Type: character vector
Values: '[0.2273, 0, 0; 0 0.2273 0; 0 0 .0040]' | 3-by-3 array | 3-by-3-by-numSat array
Default: '[0.2273, 0, 0; 0 0.2273 0; 0 0 .0040]'

Empty (dry) inertia tensor matrix, specified as a 3-by-3 array for a single spacecraft or a 3-by-3-by-numSat array for multiple spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: emptyInertia
Type: character vector
Values: 3-by-3 array | 3-by-3-by-numSat array
Default: [0.1989, 0, 0; 0 0.1989 0; 0 0 .0035]

Full (wet) inertia tensor matrix, specified as a 3-by-3 array for a single spacecraft or a 3-by-3-by-numSat array for multiple spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: fullInertia
Type: character vector
Values: 3-by-3 array | 3-by-3-by-numSat array
Default: [0.2273, 0, 0; 0, 0.2273, 0; 0, 0, .0040]

To enable mass flow velocity to the block, select this parameter. The mass flow velocity is the relative velocity in the body frame at which the mass is accreted or ablated. To disable mass flow velocity to the block, clear this parameter.

Dependencies

To enable this parameter, set Mass type to Simple variable or Custom variable.

Programmatic Use

Block Parameter: useMassFlowRelativeVelocity
Type: character vector
Values: 'on' | 'off'
Default: 'off'

To limit the mass flow when the spacecraft mass is full or empty, select this parameter. Otherwise, clear this parameter.

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: limitMassFlow
Type: character vector
Values: 'on' | 'off'
Default: 'on'

To enable fuel tank status, select this parameter. Otherwise, clear this parameter.

Dependencies

To enable this parameter, set Mass type to Simple variable.

Programmatic Use

Block Parameter: outputFuelStatus
Type: character vector
Values: 'on' | 'off'
Default: 'on'

Orbit

Define the initial states of the spacecraft.

Input method for initial states of orbit, specified as Orbital elements, ICRF state vector, or Fixed-frame state vector.

Programmatic Use

Block Parameter stateFormatNum when propagator is set to High precision (numerical)
Type: character vector
Values: 'Orbital elements' | 'Orbital elements' | 'ICRF state vector' | 'Fixed-frame state' when propagator is set to 'High precision (numerical)'
Default: 'Orbital elements'

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, select Circular 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'

Half of ellipse major axis, specified as a 1D array of size numSat. numSat 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.

Tunable: Yes

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 numSat
Default: '6786000'

Deviation of the orbit from a perfect circle, specified as a scalar or 1D array of size numSat. numSat is the number of spacecraft.

If Orbit type is set to Keplerian, this value can be:

  • 1 for parabolic orbit

  • Greater than 1 for hyperbolic orbit

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit type to Keplerian or Elliptical 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'

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.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements

  • Set Orbit type to Keplerian or Circular 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'

Right ascension of ascending node (RAAN), specified as a value between 0 and 360, specified as a scalar 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 International Celestial Reference Frame (ICRF) x-axis to the location of the ascending node — the point at which the spacecraft crosses the reference plane from south to north.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit type to Keplerian or Circular inclined.

Programmatic Use

Block Parameter: raan
Type: character vector
Values: '95' | scalar value between 0 and 360 | 1D array of size numSat
Default: '95'

Angle from the spacecraft ascending node to periapsis (closest point of orbit to the central body), specified as a 1D array of size numSat, in specified units. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements

  • Set Orbit type to Keplerian

Programmatic Use

Block Parameter: argPeriapsis
Type: character vector
Values: 93 | scalar value between 0 and 360 | 1D array of size numSat
Default: '93'

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.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit type to Keplerian or Elliptical inclined.

Programmatic Use

Block Parameter: trueAnomaly
Type: character vector
Values: '203' | scalar value between 0 and 360 | 1D array of size numSat
Default: '203'

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 numSat, in specified units. numSat is number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit Type to Circular inclined.

Programmatic Use

Block Parameter: argLat
Type: character vector
Values: '200' | scalar value between 0 and 360 | 1D array of size numSat
Default: '200'

Angle between the ICRF x-axis and the eccentricity vector, specified as a scalar or 3-element vector or 1D array of size numSat, in specified units. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit type to Elliptical equatorial.

Programmatic Use

Block Parameter: lonPeriapsis
Type: character vector
Values: 100 | scalar value between 0 and 360 | 1D array of size numSat
Default: '100'

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 or a numSat-by-3 vector, in specified units. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Initial state format to Orbital elements.

  • Set Orbit type to Circular equatorial.

Programmatic Use

Block Parameter: trueLon
Type: character vector
Values: '150' | scalar value between 0 and 360 | 1D array of size numSat | numSat-by-3 vector
Default: '150'

Cartesian position vector of spacecraft in ICRF coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or a numSat-by-3 array for multiple spacecraft. numSat is the number of spacecraft.

Tunable: Yes

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 | numSat-by-3 array
Default: '[3649700.0 3308200.0 -4676600.0]'

Cartesian velocity vector of spacecraft in ICRF coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or a numSat-by-3 array for multiple spacecraft. numSat is the number of spacecraft.

Tunable: Yes

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 | 2-D array of size numSat-by-3 array
Default: '[-2750.8 6666.4 2573.4]'

Cartesian position vector of spacecraft in fixed-frame coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or a numSat-by-3 array for multiple spacecraft. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, 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 | numSat-by-3 array
Default: '[-2750.8 6666.4 2573.4]'

Cartesian velocity vector of spacecraft in fixed-frame coordinate system at Start date/time, specified as a 3-element vector for single spacecraft or a numSat-by-3 array for multiple spacecraft. numSat is the number of spacecraft.

Tunable: Yes

Dependencies

To enable this parameter, set 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 | numSat-by-3 array
Default: '[1452.7 -6720.7 2568.1]'

Attitude

Attitude and angular rate coordinate frame with respect to the attitude and angular rate initial conditions, specified as:

  • ICRF

  • Fixed-frame

  • NED

  • LVLH

Programmatic Use

Block Parameter: attitudeFrame
Type: character vector
Values: 'ICRF' | 'Fixed-frame' | 'NED' | 'LVLH'
Default: 'ICRF'

Orientation format for spacecraft attitude (initial condition and output port), specified as Quaternion, DCM, or Euler angles.

Programmatic Use

Block Parameter: attitudeFrame
Type: character vector
Values: 'Quaternion' | 'DCM' | 'Euler angles'
Default: 'Quaternion'

Spacecraft initial attitude (orientation) of the spacecraft provided as either a quaternion, DCM, or Euler angle set with respect to Attitude representation.

Tunable: Yes

Dependencies

This parameter name and value format changes depending on the Attitude representation parameter.

Parameter NameAttitude Representation SettingValue Format

Initial quaternion

Quaternion

  • 4-element vector

  • numSat-by-4 array

Initial DCM

DCM

  • 3-by-3 array

  • numSat-by-3-by-3 array

Initial Euler angles

Euler angles

  • 3-element vector

  • numSat-by-3 array

Programmatic Use

Block Parameter: attitude
Type: character vector
Values: 4-element vector | numSat-by-4 array | 3-by-3 array | numSat-by-3-by-3 array | 3-element array | numSat-by-3 array
Default: '[1, 0, 0, 0]'

Rotation angle sequence for Euler angle attitude representation.

Tunable: Yes

Dependencies

To enable this parameter, set Attitude representation to Euler angles.

Programmatic Use

Block Parameter: rotationOrder
Type: character vector
Values: 'ZYX' | 'ZYZ' |'ZXY' | 'ZXZ' | 'YXZ' | 'YXY' | 'YZX' | 'YZY' | 'XYZ' | 'XYX' | 'XZY' | 'XZX'
Default: 'ZYX'

Initial body-fixed angular rates (PQR) with respect to Attitude reference coordinate frame.

Tunable: Yes

Programmatic Use

Block Parameter: attitudeRate
Type: character vector
Values: | 3-element vector | numSat-by-3 array
Default: [0, 0, 0]

Enable output total vehicle acceleration computed by the block with respect to the ICRF attitude reference coordinate frame. This acceleration includes all moments that act on the spacecraft.

Tunable: Yes

Programmatic Use

Block Parameter: angAccelOut
Type: character vector
Values: 'on' | 'off'
Default: 'off'

Select this parameter to enable the use of the gravity gradient torque in the block rotational dynamics equations. Otherwise, clear this parameter.

Tunable: Yes

Programmatic Use

Block Parameter: angAccelOut
Type: character vector
Values: 'on' | 'off'
Default: 'on'

Central Body

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'

Control the gravity model for the central body by specifying as Spherical harmonics, Point-mass, or Oblate ellipsoid (J2).

Dependencies

Available options are based on Central body settings.

Earth, Moon, Mars, or CustomMercury, Venus, Jupiter, Saturn, Uranus, or Neptune
Spherical harmonicsOblate ellipsoid (J2)
Point-massPoint-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' | '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 gravitational potential model, specified according to the specified Central body.

Dependencies

Available options are based on Central body settings:

Central bodySpherical Harmonic Model Option
EarthEGM2008, EGM96, or EIGEN-GL04C
MoonLP-100K or LP-165P
MarsGMM2B

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'

Rotational rate of a custom central body, specified as a scalar.

Dependencies

To enable this parameter, set Central body to Custom.

Programmatic Use

Block Parameter: 'customOmega'
Type: character vector
Values: '4.06124975e-3' | scalar
Default: '4.06124975e-3'

Harmonic coefficient MAT file that contains definitions for a custom planetary model, specified as a character vector or string.

This file must contain these variables:

VariableDescription
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 Central body to Custom.

  • Set 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 of harmonic model, specified as a scalar.

Planet ModelRecommended DegreeMaximum Degree

EGM2008

120

2159

EGM96

70

360

LP100K

60

100

LP165P

60

165

GMM2B

60

80

EIGENGL04C

70

360

Dependencies

To enable this parameter:

  • Set Central body to Earth, Moon, Mars, or Custom.

  • Set Gravitational potential model to Spherical harmonics.

Programmatic Use

Block Parameter: shDegree
Type: character vector
Values: '80' | scalar
Default: '80'

Select this parameter to use Earth orientation parameters for the transformation between the ICRF and fixed-frame coordinate systems. Otherwise, clear this parameter.

Dependencies

To enable this parameter, set Central body to Earth.

Additionally, it must satisfy one of these criteria:

  • Gravitational potential model is set to either Spherical harmonics or Oblate ellipsoid (J2).

  • External acceleration coordinate frame is set to Fixed-frame.

  • State vector output coordinate frame is set to Fixed-frame.

  • Attitude reference coordinate frame is set to Fixed-frame or NED.

Programmatic Use

Block Parameter: useEOPs
Type: character vector
Values: 'on' | 'off'
Default: 'on'

Custom list of Earth orientation data, specified in a MAT file.

Dependencies

To enable this parameter:

  • Select the Use Earth orientation parameters (EOPs) parameter.

  • Set Central body to Earth.

Programmatic Use

Block Parameter: eopFile
Type: character vector
Values: 'aeroiersdata.mat' | MAT-file
Default: 'aeroiersdata.mat'

To specify Euler libration angles (φ θ ψ) for Moon orientation, select this parameter.

Dependencies

To enable this parameter, set Central body to Moon.

Programmatic Use

Block Parameter: useMoonLib
Type: character vector
Values: 'off' | 'on'
Default: 'off'

To add output transformation quaternion port for the quaternion transformation from the ICRF to the fixed-frame coordinate system, select this parameter. Otherwise, clear this parameter.

Programmatic Use

Block Parameter: outputTransform
Type: character vector
Values: 'off' | 'on'
Default: 'off'

Central body spin axis source, 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 Central body to Custom.

Programmatic Use

Block Parameter: cbPoleSrc
Type: character vector
Values: 'Port' | 'Dialog'
Default: 'Port'

Right ascension of central body spin axis at J2000 (2451545.0 JD, 2000 Jan 1 12:00:00 TT), specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Central body spin axis source to Dialog.

Programmatic Use

Block Parameter: cbRA
Type: character vector
Values: '317.68143' | double scalar
Default: '317.68143'

Right ascension rate of the central body spin axis, specified as a double scalar, in specified angle units/century.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Central body spin axis source to Dialog.

Programmatic Use

Block Parameter: cbRARate
Type: character vector
Values: '-0.1061' | double scalar
Default: '-0.1061'

Declination of the central body spin axis at J2000 (2451545.0 JD, 2000 Jan 1 12:00:00 TT), specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Central body spin axis source to Dialog.

Programmatic Use

Block Parameter: cbDec
Type: character vector
Values: '52.88650' | double scalar
Default: '52.88650'

Declination rate of the central body spin axis, specified as a double scalar, in specified angle units/century.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Central body spin axis source to Dialog.

Programmatic Use

Block Parameter: cbDecRate
Type: character vector
Values: '-0.0609' | double scalar
Default: '-0.0609'

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.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set 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 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, in angle units/day.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set 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 for a custom central body, specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter, set Gravitational potential model to None, Point-mass, or Oblate ellipsoid (J2).

Programmatic Use

Block Parameter: customR
Type: character vector
Values: '3.3962e6' | double scalar
Default: '3.3962e6'

Flattening ratio for custom central body, specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Gravitational potential model to Point-mass, Oblate ellipsoid (J2), or Spherical harmonics.

Programmatic Use

Block Parameter: customF
Type: character vector
Values: '0.00589' | double scalar
Default: '0.00589'

Gravitational parameter for a custom central body, specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set Gravitational potential model to None, Point-mass, or Oblate ellipsoid (J2).

Programmatic Use

Block Parameter: customMu
Type: character vector
Values: '4.305e13' | double scalar
Default: '4.305e13'

Most significant or largest spherical harmonic term, which accounts for oblateness of a celestial body, specified as a double scalar.

Tunable: Yes

Dependencies

To enable this parameter:

  • Set Central body to Custom.

  • Set 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.

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'

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 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'

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'

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'

Extrapolation method for f107Average and f107Daily for times outside the range of the MAT file data, specified as one of these values.

MethodDescription
None - clip

Set f107Average and f107Daily to the nearest data point available in the MAT file.

Constant

Set f107Average and f107Daily to a constant value specified by the Magnetic index extrapolation method parameter value.

Least squares fit

Approximate f107Average and f107Daily using a least-squares fit of the space weather data from October 1, 1957, to December 1, 2040. This method uses a trigonometric function of the form:

a + b*cos(c*t + d*sin(e*t)),

where:

  • a is 128.2351780622538

  • b is 54.3285872213434

  • c is 0.0015460708364

  • d is 0.2429462096495

  • e is 0.0015563188188

  • t is jdCurrent - jdReftime

  • jdCurrent is the current Julian date

  • jdRef is 2436112.5

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'

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.

Extrapolation method for magnetic index information values for times outside the range of the MAT file data, specified as one of these values.

MethodDescription
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'

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'

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'

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.

FieldDescription
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 APH(:,1).

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)'

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'

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'

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'

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'

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.

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

DE405

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.

DE421

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.

DE423

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.

DE430

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.

DE432t

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'

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 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 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)'

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'

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'

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'

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'

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'

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'

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'

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'

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'

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'

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'

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'

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 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

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'

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 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'

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'

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'

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'

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'

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'

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, 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

Parameter and port units, specified as shown here.

UnitsForcesMomentMassInertiaDistanceVelocityAccelerationAreaDensity
Metric (m/s)NewtonNewton-meterKilogramsKilogram m2metersmeters/secmeters/sec2m2kg/m3, some density outputs 1/m3
Metric (km/s)NewtonNewton-meterKilogramsKilogram m2kilometerskilometers/seckilometers/sec2m2kg/m3, some density outputs 1/m3
Metric (km/h)NewtonNewton-meterKilogramsKilogram m2kilometerskilometers/hourkilometers/hour2m2kg/m3, some density outputs 1/m3
English (ft/s)Pound-forceFoot-poundSlugsSlug ft2feetfeet/secfeet/sec2feet2lbm/ft3, some density outputs 1/ft3
English (kts)Pound-forceFoot-poundSlugsSlug ft2nautical mileknotsknots/secfeet2lbm/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)'

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 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

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References

[1] Vallado, David. Fundamentals of Astrodynamics and Applications. 4th ed. Hawthorne, CA: Microcosm Press, 2013.

[2] Vepa, Ranjan. Dynamics and Control of Autonomous Space Vehicles and Robotics. New York: Cambridge University Press, 2019.

[3] Stevens, Frank L., and Brian L. Stevens. Aircraft Control and Simulation. 2nd ed. Hoboken, NJ: John Wiley & Sons, 2003.

[4] Gottlieb, R. G. Fast Gravity, Gravity Partials, Normalized Gravity, Gravity Gradient Torque and Magnetic Field: Derivation, Code and Data. NASA Contractor Report 188243. Houston: NASA, February 1993.

[5] 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 150, no. 1 (2001): 1–18.

[6] Lemoine, F. G. et al. "An Improved Solution of the Gravity Field of Mars (GMM-2B) from Mars Global Surveyor." Journal of Geophysical Research 106, no. E10 (2001): 23359–23376.

[7] Seidelmann, P. Kenneth et al. "Report of the IAU/IAG Working Group on Cartographic Coordinates and Rotational Elements: 2006." Celestial Mech Dyn Astr 98 (20017): 155–180 (2007).

[8] Standish, E. M. "JPL Planetary and Lunar Ephemerides." DE405/LE405. Interoffice memorandum. JPL IOM 312.F-98-048. August 26, 1998.

Extended Capabilities

C/C++ Code Generation
Generate C and C++ code using Simulink® Coder™.

Version History

Introduced in R2021b

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See Also

Blocks

Functions
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