Help: Orbit generator

Table of Contents	ECSS	Model Page
Background information	Orbit generator
	Orbit generator (Earth)

Overview

The SPENVIS orbit generator can be used to compute low altitude orbits, geostationary orbits, highly eccentric orbits and flybys. It takes into account the oblateness of the Earth, the gravitational attraction of Sun and Moon, air drag (by means of the CIRA atmospheric model) and solar radiation pressure.

The spacecraft trajectory is defined in terms of a mission, i.e. in addition to defining the spacecraft orbital elements, the time period the spacecraft is in orbit has to be specified as well. A mission can be subdivided in segments, with different orbital characteristics.

Once the mission has been defined, orbital parameters have to be specified for each mission segment. Once this is done, a page with a summary table of the mission segments is presented and the trajectory generation can be started.

Mission definition

A mission is defined by two quantities:

the number of mission segments, each with its own set of orbital parameters;
(advanced users can define up to ten segments while non-advanced users are limited to five)
the time of the end of the mission (the start of the mission is set to the starting point of the orbit for the first segment). The mission end can be specified by the total mission duration, by an end date, or by the duration of the last segment (in this case, the mission length is defined as the sum of the segment lengths).

The start time of each mission segment is set to the start time of the orbit defined for each segment. When more than one segment is defined, the end time of each segment (except for the last segment) is set to the start time of the next segment by default. Alternatively, the start of a segment can also be defined as the end of a previous segment.

The segment and mission lengths and epochs defined in this fashion are used by the environment and effects models that run on a spacecraft trajectory. Hence, these parameters need not be specified when running the models in question. In particular, the radiation sources and effects models use the segment and mission lengths to scale orbit averaged fluxes to segment and mission fluences, or to add dose contributions from trapped particles to those from solar protons (which are defined for the total mission length only).

Trajectory uploads

Advanced users have also the option to upload a trajectory file.

Attitude vectors

Attitude vectors of the satellite are generated along its orbit. These parameters are required for the directional trapped proton flux model, the ATOMOX tool that integrates atmospheric densities over a trajectory, and the illumination tool that calculates incident electromagnetic radiation on an oriented plate.

Attitude definition Four sets of pre-defined directions can be selected by advanced users (for the other users, the default setting parallel to the velocity vector is used):

One satellite axis is parallel to the velocity vector (Z);
X-axis: lies in the orbital plane pointing away from the Earth;
Y-axis: perpendicular to the orbital plane and points to the South.
One satellite axis points to the zenith (Z);
X-axis: lies in the horizontal plane and points to the geographic North;
Y-axis: lies in the horizontal plane and points to the geographic West.
Satellite axes are parallel to the Geocentric Inertial axes (X, Y, Z).
One satellite axis is parallel to the Sun direction (Z);
Y-axis: perpendicular to Z and the Earth rotation axis;
X-axis: perpendicular to Y and Z.

These Z-directions are illustrated in the figure above, with N pointing to the geographic north (inertial system Z-axis), V the velocity vector, Z pointing to the zenith, L_n the local axis parallel to the GEI Z-axis, and S pointing to the Sun.

Solar radiation pressure and atmospheric drag

Advanced users have the option to input parameters for Solar radiation pressure and atmospheric drag.

The Solar radiation pressure parameter is defined as 0.451x10^-8 K A/M, where:

K is the material parameter:

K = 1 - gamma + rho	plate
K = 1 - gamma	sphere
K = 2	plate with perfect specular reflection
K = 1	sphere with perfect specular reflection or perfect absorption
K = 1.44	sphere with perfect diffuse reflection

rho and gamma are the reflectivity and transmittivity of the satellite.

A is the total reflective area (m²) of satellites and plates; for spherical satellites it is the cross sectional area.
M (kg) is the mass of the satellite.

The atmospheric drag parameter (or ballistic coefficient) is defined as 0.5x10⁶ C_D A/M, where C_D is a dimensionless drag coefficient, A is the cross sectional area (m²) of the satellite, and M (kg) is the mass of the satellite. The atmospheric drag is evaluated using the NRLMSISE-00 model.

Orbital parameters

A spececraft orbit is described by means of a number of parameters, a start date and a duration. The start date corresponds to the date and time of the first point written on the output file. The duration can be specified as a number or orbits, or directly as a duration in days (non-integer numbers of orbits or days are allowed). It is not expedient to generate a trajectory file for the whole of a typical mission duration, instead a duration should be selected that guarantees coverage of all geographic locations. For LEO and GEO trajectories, this is typically of the order of one day, for intermediate trajectories a longer duration may be needed. For very high altitude trajectories, a short duration is sufficient, as the models that can be run on a trajectory do not depend on spacecraft location outside the magnetosphere.

It is recommended to produce graphical representations of the trajectory before proceeding with the environment models.

Five different orbit types can be selected:

general (elliptical) orbit;
hyperbolic (flyby) orbit;
circular heliosynchronous orbit;
geostationary orbit;
near Earth interplanetary orbit.

For the general orbit, all parameters have to be entered, while the last three cases are provided as special cases that only require a restricted number of parameters. When an orbit type is selected, the appropriate input fields are displayed on the page. The segment title is used for annotating the report and output files.

Advanced users have the option to specify the orbit using two-line elements. These are general perturbation element sets generated by NORAD to predict position and velocity of Earth orbiting objects. A complete description of TLEs and current and archived sets can be found at T.S. Kelso's CelesTrak site.

General orbit parameters

Altitude

The orbit altitude can be specified in three ways, by providing the following sets of parameters:

the altitude of the perigee and apogee of the trajectory, respectively, above the mean radius of the Earth;
the length of the semi-major axis and the eccentricity;
the altitude of a circular orbit.

Note that for general (elliptical) orbits the semi-major axis is positive and the eccentricity must be greater or equal (circular orbit) to zero and less than one.

Finally, for a circular orbit the semi-major axis and the eccentricity of the osculating ellipse are calculated in order to minimise the altitude variations due to the oblate Earth (J2 approximation).

Inclination

The orbit inclination is the angle between the orbital plane and the equatorial plane, measured at the ascending node in the direction of orbital motion. The orbit is called direct when the inclination is smaller than 90° and retrograde when the inclination is larger than 90°.

Right ascension of the ascending node

The right ascension of the ascending node is the angle in the equatorial plane between the line of nodes and the direction to the vernal equinox, measured from the vernal equinox (the direction of the intersection of the ecliptic and equatorial planes) towards the ascending node. Alternatively, the longitude of perigee or apogee can be specified.

Argument of perigee

The argument of perigee is the angle measured in the orbital plane from the ascending node to the perigee.

True anomaly

The true anomaly is the angle from the perigee direction to the satellite direction.

Hyperbolic orbit parameters

The orbit altitude can be specified by providing the length of the semi-major axis and the eccentricity. Note that for hyperbolic orbits the semi-major axis has a negative value and the eccentricity must be greater than one.

The remaining orbital elements (inclination, right ascension of the ascending note, argument of perigee and true anomaly) are defined in the same way as in the general orbit case.

Heliosynchronous orbit parameters

A circular heliosynchronous orbit is defined by specifying its altitude and the local solar time. The orbit inclination is derived from the altitude.

Geostationary orbit parameters

A geostationary orbit is defined by specifying its (east) longitude.

Near Earth interplanetary orbit parameters

This option is available to generate a coordinate set outside of the Earth's magnetosphere. It is defined as a circular orbit at 90,000 km altitude and inclination 0°. Hence, it is not a true interplanetary trajectory, but it can be used as an approximation for the interplanetary environment. Some models use the distance from the Sun as an input parameter, which can be entered here.

Advanced users have the option to set the output resolution of the orbit generator. Up to three time steps (s) can be set for three different regions defined by a limiting altitude (km).

Mission summary

When all mission segment orbits have been defined, the mission summary page is presented. This page provides a table with the orbit types and durations (or number of orbits) and the start and end times for each mission segment. This table can be used for a final check of the mission definition. The orbit generator will check the segment definitions for overlapping segments and other inconsistencies. If any discrepancies are found, an error message will be displayed.

The mission summary page is skipped when an uploaded trajectory is used.

Pressing the button will start the calculation and bring up the "Results" page.

The button calls up the model selection page for consecutive runs of multiple models. This feature is available for advanced users only.

Warning: using these buttons deletes all existing output from the orbit generator and from any model that uses this output, in order to ensure consistency in the outputs.

Last update: Mon, 12 Mar 2018