Simple Variable Mass 6DOF ECEF (Quaternion)

Implement quaternion representation of six-degrees-of-freedom equations of motion of simple variable mass in Earth-centered Earth-fixed (ECEF) coordinates

Library

Equations of Motion/6DOF

Description

The Simple Variable Mass 6DOF ECEF (Quaternion) block considers the rotation of a Earth-centered Earth-fixed (ECEF) coordinate frame (X_ECEF, Y_ECEF, Z_ECEF) about an Earth-centered inertial (ECI) reference frame (X_ECI, Y_ECI, Z_ECI). The origin of the ECEF coordinate frame is the center of the Earth, additionally the body of interest is assumed to be rigid, an assumption that eliminates the need to consider the forces acting between individual elements of mass. The representation of the rotation of ECEF frame from ECI frame is simplified to consider only the constant rotation of the ellipsoid Earth (ω_e) including an initial celestial longitude (L_G(0)). This excellent approximation allows the forces due to the Earth's complex motion relative to the “fixed stars” to be neglected.

The translational motion of the ECEF coordinate frame is given below, where the applied forces [F_x F_y F_z]^T are in the body frame. Vre_b is the relative velocity in the wind axes at which the mass flow ( $\dot{m}$ ) is ejected or added to the body axes.

$\begin{matrix} {\bar{F}}_{b} = [\begin{matrix} F_{x} \\ F_{y} \\ F_{z} \end{matrix}] = m ({\dot{\bar{V}}}_{b} + {\bar{ω}}_{b} \times {\bar{V}}_{b} + D C M_{b f} {\bar{ω}}_{e} \times {\bar{V}}_{b} + D C M_{b f} ({\bar{ω}}_{e} \times ({\bar{ω}}_{e} \times {\bar{X}}_{f}))) \\ + \dot{m} (\bar{V} r e_{b} + D C M_{b f} ({\bar{ω}}_{e} \times {\bar{X}}_{f})) \\ A_{b b} = [\begin{matrix} {\dot{u}}_{b} \\ {\dot{v}}_{b} \\ {\dot{w}}_{b} \end{matrix}] = \frac{{\bar{F}}_{b} - \dot{m} ({\bar{V}}_{r e_{b}} + D C M_{b f} (w_{e} \times X_{f}))}{m} \\ - [{\bar{ω}}_{_{b}} \times {\bar{V}}_{b} + D C M {\bar{ω}}_{e} \times {\bar{V}}_{b} + D C M_{b f} ({\bar{ω}}_{e} ({\bar{ω}}_{e} \times X_{f}))] \\ A_{b} _{e c e f} = \frac{{\bar{F}}_{b} - \dot{m} ({\bar{V}}_{r e_{b}} + D C M_{b f} (ω_{e} \times X_{f}))}{m} \end{matrix}$

where the change of position in ECEF ${\dot{\bar{x}}}_{f} ({\dot{\bar{x}}}_{i})$ is calculated by

${\dot{\bar{x}}}_{f} = D C M_{f b} {\bar{V}}_{b}$

and the velocity of the body with respect to ECEF frame, expressed in body frame $({\bar{V}}_{b})$ , angular rates of the body with respect to ECI frame, expressed in body frame $({\bar{ω}}_{b})$ . Earth rotation rate $({\bar{ω}}_{e})$ , and relative angular rates of the body with respect to north-east-down (NED) frame, expressed in body frame $({\bar{ω}}_{r e l})$ are defined as

$\begin{array}{l} {\bar{V}}_{b} = [\begin{matrix} u \\ v \\ w \end{matrix}] {\bar{ω}}_{r e l} = [\begin{matrix} p \\ q \\ r \end{matrix}] {\bar{ω}}_{e} = [\begin{matrix} 0 \\ 0 \\ ω_{e} \end{matrix}] \\ {\bar{ω}}_{b} = {\bar{ω}}_{r e l} + D C M_{b f} {\bar{ω}}_{e} + D C M_{b e} {\bar{ω}}_{n e d} \\ {\bar{ω}}_{n e d} = [\begin{matrix} \dot{l} \cos μ \\ - \dot{μ} \\ - \dot{l} \sin μ \end{matrix}] = [\begin{matrix} V_{E} / (N + h) \\ - V_{N} / (M + h) \\ V_{E} \tan μ / (N + h) \end{matrix}] \end{array}$

The rotational dynamics of the body defined in body-fixed frame are given below, where the applied moments are [L M N]^T, and the inertia tensor I is with respect to the origin O.

$\begin{array}{l} {\bar{M}}_{b} = [\begin{matrix} L \\ M \\ N \end{matrix}] = \bar{I} {\dot{\bar{ω}}}_{b} + {\bar{ω}}_{b} \times (\bar{I} {\bar{ω}}_{b}) + \dot{I} {\bar{ω}}_{b} \\ I = [\begin{matrix} I_{x x} & - I_{x y} & - I_{x z} \\ - I_{y x} & I_{y y} & - I_{y z} \\ - I_{z x} & - I_{z y} & I_{z z} \end{matrix}] \end{array}$

The inertia tensor is determined using a table lookup which linearly interpolates between I_full and I_empty based on mass (m). The rate of change of the inertia tensor is estimated by the following equation.

$\dot{I} = \frac{I_{f u l l} - I_{e m p t y}}{m_{f u l l} - m_{e m p t y}} \dot{m}$

The integration of the rate of change of the quaternion vector is given below.

$[\begin{matrix} {\dot{q}}_{0} \\ {\dot{q}}_{1} \\ {\dot{q}}_{2} \\ {\dot{q}}_{3} \end{matrix}] = - \frac{1}{2} [\begin{matrix} 0 & ω_{b} (1) & ω_{b} (2) & ω_{b} (3) \\ - ω_{b} (1) & 0 & - ω_{b} (3) & ω_{b} (2) \\ - ω_{b} (2) & ω_{b} (3) & 0 & - ω_{b} (1) \\ - ω_{b} (3) & - ω_{b} (2) & ω_{b} (1) & 0 \end{matrix}] [\begin{matrix} q_{0} \\ q_{1} \\ q_{2} \\ q_{3} \end{matrix}]$

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

Parameters

Main

Units

Specifies the input and output units:

Units	Forces	Moment	Acceleration	Velocity	Position	Mass	Inertia
`Metric (MKS)`	Newton	Newton meter	Meters per second squared	Meters per second	Meters	Kilogram	Kilogram meter squared
`English (Velocity in ft/s)`	Pound	Foot pound	Feet per second squared	Feet per second	Feet	Slug	Slug foot squared
`English (Velocity in kts)`	Pound	Foot pound	Feet per second squared	Knots	Feet	Slug	Slug foot squared

Mass type

Select the type of mass to use:

`Fixed`	Mass is constant throughout the simulation (see 6DOF ECEF (Quaternion)).
`Simple Variable`	Mass and inertia vary linearly as a function of mass rate.
`Custom Variable`	Mass and inertia variations are customizable (see Custom Variable Mass 6DOF ECEF (Quaternion)).

The Simple Variable selection conforms to the previously described equations of motion.

Initial position in geodetic latitude, longitude and altitude

The three-element vector for the initial location of the body in the geodetic reference frame. Latitude and longitude values can be any value. However, latitude values of +90 and -90 may return unexpected values because of singularity at the poles.

Initial velocity in body axes

The three-element vector containing the initial velocity of the body with respect to the ECEF frame, expressed in the body frame.

Initial Euler orientation

The three-element vector containing the initial Euler rotation angles [roll, pitch, yaw], in radians. Euler rotation angles are those between the body and NED coordinate systems.

Initial body rotation rates

The three-element vector for the initial angular rates of the body with respect to the NED frame, expressed the body frame, in radians per second.

Initial mass

The mass of the rigid body.

Inertia

A scalar value for the inertia of the body.

Empty mass

A scalar value for the empty mass of the body.

Full mass

A scalar value for the full mass of the body.

Empty inertia matrix

A 3-by-3 inertia tensor matrix for the empty inertia of the body.

Full inertia matrix

A 3-by-3 inertia tensor matrix for the full inertia of the body.

Include mass flow relative velocity

Select this check box to add a mass flow relative velocity port. This is the relative velocity at which the mass is accreted or ablated.

Include inertial acceleration

Select this check box to enable an additional output port for the accelerations in body-fixed axes with respect to the inertial frame. You typically connect this signal to the accelerometer.

Planet model

Specifies the planet model to use: Custom or Earth (WGS84).

Equatorial radius of planet

Specifies the radius of the planet at its equator. The units of the equatorial radius parameter should be the same as the units for ECEF position. This option is only available when Planet model is set to Custom.

Flattening

Specifies the flattening of the planet. This option is only available when Planet model is set to Custom.

Rotational rate

Specifies the scalar rotational rate of the planet in rad/s. This option is only available when Planet model is set to Custom.

Celestial longitude of Greenwich source

Specifies the source of Greenwich meridian's initial celestial longitude:

`Internal`	Use celestial longitude value from Celestial longitude of Greenwich.
`External`	Use external input for celestial longitude value.

Celestial longitude of Greenwich

The initial angle between Greenwich meridian and the x-axis of the ECI frame. This parameter appears if you set Celestial longitude of Greenwich source to Internal.

Include inertial acceleration

Select this check box to enable an additional output port for the accelerations in body-fixed axes with respect to the inertial frame. You typically connect this signal to the accelerometer.

State Attributes

Assign unique name to each state. You can use state names instead of block paths during linearization.

To assign a name to a single state, enter a unique name between quotes, for example, 'velocity'.
To assign names to multiple states, enter a comma-delimited list surrounded by braces, for example, {'a', 'b', 'c'}. Each name must be unique.
If a parameter is empty (' '), no name assignment occurs.
The state names apply only to the selected block with the name parameter.
The number of states must divide evenly among the number of state names.
You can specify fewer names than states, but you cannot specify more names than states.
For example, you can specify two names in a system with four states. The first name applies to the first two states and the second name to the last two states.
To assign state names with a variable in the MATLAB^® workspace, enter the variable without quotes. A variable can be a character vector, cell array, or structure.

Quaternion vector: e.g., {'qr', 'qi', 'qj', 'qk'}

Specify quaternion vector state names.

Default value is ''.

Body rotation rates: e.g., {'p', 'q', 'r'}

Specify body rotation rate state names.

Default value is ''.

Velocity: e.g., {'U', 'v', 'w'}

Specify velocity state names.

Default value is ''.

ECEF position: e.g., {'Xecef', 'Yecef', 'Zecef'}

Specify the ECEF position state names.

Default value is ''.

Inertial position: e.g., {'Xeci', 'Yeci', 'Zeci'}

Specify the inertial position state names.

Default value is ''.

Celestial longitude of Greenwich: e.g., 'LG'

Specify the Celestial longitude of Greenwich state name.

Default value is ''.

Mass: e.g., 'mass'

Specify mass state name.

Default value is ''.

Inputs and Outputs

Input	Dimension Type	Description
First	Vector	Contains the three applied forces in body-fixed axes.
Second	Vector	Contains the three applied moments in body-fixed axes.
Third	Scalar	Contains one or more rates of change of mass (positive if accreted, negative if ablated).
Fourth (Optional)	Three-element vector	Contains one or more relative velocities at which the mass is accreted to or ablated from the body-fixed axes.

Output	Dimension Type	Description
First	Three-element vector	Contains the velocity of body respect to ECEF frame, expressed in ECEF frame.
Second	Three-element vector	Contains the position in the ECEF reference frame.
Third	Three-element vector	Contains the position in geodetic latitude, longitude and altitude, in degrees, degrees and selected units of length respectively.
Fourth	Three-element vector	Contains the body rotation angles [roll, pitch, yaw], in radians. Euler rotation angles are those between body and NED coordinate systems.
Fifth	3-by-3 matrix	Applies to the coordinate transformation from ECI axes to body-fixed axes.
Sixth	3-by-3 matrix	Applies to the coordinate transformation from NED axes to body-fixed axes.
Seventh	3-by-3 matrix	Applies to the coordinate transformation from ECEF axes to NED axes.
Eighth	Three-element vector	Contains the velocity of body with respect to ECEF frame, expressed in body frame.
Ninth	Three-element vector	Contains the relative angular rates of body with respect to NED frame, expressed in body frame, in radians per second.
Tenth	Three-element vector	Contains the angular rates of the body with respect to ECI frame, expressed in body frame, in radians per second.
Eleventh	Three-element vector	Contains the angular accelerations of the body with respect to ECI frame, expressed in body frame, in radians per second squared.
Twelfth	Three-element vector	Contains the accelerations in body-fixed axes with respect to body frame.
Thirteenth	Scalar	Is an element containing a flag for fuel tank status: 1 indicates that the tank is full. 0 indicates that the integral is neither full nor empty. -1 indicates that the tank is empty.
Fourteenth (Optional)	Three-element vector	Contains the accelerations in body-fixed axes with respect to ECEF frame.

Assumptions and Limitations

This implementation assumes that the applied forces are acting at the center of gravity of the body.

This implementation generates a geodetic latitude that lies between ±90 degrees, and longitude that lies between ±180 degrees. Additionally, the MSL altitude is approximate.

The Earth is assumed to be ellipsoidal. By setting flattening to 0.0, a spherical planet can be achieved. The Earth's precession, nutation, and polar motion are neglected. The celestial longitude of Greenwich is Greenwich Mean Sidereal Time (GMST) and provides a rough approximation to the sidereal time.

The implementation of the ECEF coordinate system assumes that the origin is at the center of the planet, the x-axis intersects the Greenwich meridian and the equator, the z-axis is the mean spin axis of the planet, positive to the north, and the y-axis completes the right-hand system.

The implementation of the ECI coordinate system assumes that the origin is at the center of the planet, the x-axis is the continuation of the line from the center of the Earth toward the vernal equinox, the z-axis points in the direction of the mean equatorial plane's north pole, positive to the north, and the y-axis completes the right-hand system.

References

Stevens, B. L., and F. L. Lewis, Aircraft Control and Simulation, Second Edition, John Wiley & Sons, New York, 2003.

McFarland, Richard E., A Standard Kinematic Model for Flight simulation at NASA-Ames, NASA CR-2497.

“Supplement to Department of Defense World Geodetic System 1984 Technical Report: Part I - Methods, Techniques and Data Used in WGS84 Development,” DMA TR8350.2-A.

Documentation

Simple Variable Mass 6DOF ECEF (Quaternion)

Library

Description

Parameters

Main

State Attributes

Inputs and Outputs

Assumptions and Limitations

References

See Also

Extended Capabilities

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

Introduced in R2006a

Aerospace Blockset Documentation

Support

Documentation

Simple Variable Mass 6DOF ECEF (Quaternion)

Library

Description

Parameters

Main

State Attributes

Inputs and Outputs

Assumptions and Limitations

References

See Also

Extended Capabilities

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

Introduced in R2006a

Aerospace Blockset Documentation

Support

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