focussuite focusleo: the ultimate flight dynamics solution for Low Earth Orbit satellites operation
focusleo is a member of GMV's focussuite and brings to satellite operators a solution to the domain of flight dynamics operation of low (LEO) and medium (MEO) Earth orbiting satellites. focusleo is a completely integrated application that provides full life cycle support of a wide variety of satellites through an unsurpassed collection of mission independent and mission specific functionality that brings to satellite operators a perfect combination of professional off-the-shelf solutions with the flexibility and customer focus of a world leading supplier of flight dynamics custom solutions. The computation layer of focusleo is based in part on ESA´s NAPEOS technology. It provides scheduling capabilities for the routine operations. It supports currently the Galileo, Globalstar and O3B platforms.
Flight Dynamics Functions Supported by focusleo computational Layer
Mission Specific Functions
Providing full lifecycle support of the corresponding satellite platform with a degree of accuracy 100% compatible with native systems focusleo supports all mission specific needs, such as:
- Maneuver Implementation
- Fuel Mass Evolution
- Station Keeping Maneuver Preparation
- De-orbiting maneuvers computation
- AOCS support functions
- Contingency support functions
Mission Independent Functions
Ingestion of Solar Flux Data and Earth Orientation
Modules in charge of ingesting in the system the Solar flux data, Earth Orientation parameters and leap seconds information generated by External Agencies.
Pre-Processing of Tracking Data
This module ingests and pre-process input tracking data. It ingests the raw tracking data, soothes, reduces it, and applies calibration. It supports the following formats:
- Satellite Laser Ranging
- Precise Range and Range Rate equipment
- GPS (RINEX and Double differences in phase and pseudo-range)
- Multipurpose Tracking System (MPTS) (range, doppler, azimuth and elevation)
- Altimetry (PO-DACC, ESOC)
Orbit Determination and Maneuver Estimation
This module provides estimations of the spacecraft orbit that are more accurate than the predictions generated from previous orbit reconstructions.
The resulting orbit is uniquely defined by:
- An initial state vector
- An initial epoch
- The list of force models applied
- Parameters associated with these models (extended state vector)
It is able to estimate all arc-dependent parameters required including:
- Scale factors for S/C maneuvers
- Aerodynamic force coefficients for each S/C and per ar
- Radiation pressure coefficients (solar radiation) for each S/C and per arc
- Bias per different altimeter on the altimetry measurements
- Station biases per arc and per pass
- Tracking station positions
- Station parameters per pass or per time interval
- Pass and arc parameters for the same station in the same run
The Orbit Determination component is also responsible of calibrating the maneuvers from the tracking data. For several revolutions after the maneuver(s) have been performed, a calibration run is executed. This activity will continue until sufficient measurement data have been accumulated before and after the maneuver to provide accurate estimations of both the spacecraft motion and the maneuver calibration factors.
The Orbit Determination component shall take advantage of the Orbit Propagation component to propagate the orbit within the orbit reconstruction period. This ensures that the orbit modeling is consistent between reconstruction and prediction.
This module computes the evolution of an orbit forward and backward in time. It supports the propagation of impulsive and long maneuvers. The force model used by this module is fully configurable by the operator. The algorithms and models have been implemented to make the system compliant with the latest recommendations of the International Earth Rotation Service (IERS) standards.
This module generates the geometrical events:
- Ascending and descending node crossing times
- Apogee and Perigee crossing times.
- Umbra and penumbra crossing times of Earth and Moon eclipses.
- Sun and Moon AOS/LOS for a Sun sensor
- Sun and Moon AOS/LOS for an Earth sensor
- Sun-Satellite-Earth Colinearity
- Sun-Satellite-Station Colinearity
- Station visibilities
- Satellite local time
- Day/night terminator crossing
Target Orbit Definition
This module is responsible for computing/defining the target orbit for the different phases of an Earth-observation mission or just defining it in case of other kind of mission. This reference orbit can include:
- Sun-synchronous orbit
- Ground track repetition
- Constant local time at node crossing (ascending or descending).
- Frozen eccentricity vector.
The propagation model used for the definition of the reference orbit can be customized by the operator. It computes a reference orbit to be used for maneuver planning. The characteristics of the reference orbit supported are:
- Repeat ground track
- Frozen eccentricity
- Constant local time at ascending or descending node
- Overflight or a particular point on the Earth
The reference orbit file generated contains the same number of days as the repeat cycle. However, since it is designated internally as a reference orbit, it can be used for analysis over multiple cycles and not just for the cycle that it contains.
Orbit Analysis and Orbit Control
The orbit analysis module is responsible for comparing the current orbit with the reference one in terms of:
- Ground-track deviation: separation of the sub-satellite point from the reference ground track
- Ascending node status: separation in inclination, altitude, phase and local time of ascending node
Moreover, in case an exit from the dead-band (defined as maximum acceptable separation from the reference ground track) is detected, then the exit time is computed and one of the following actions is executed, depending on the type of exit detected:
- Eastward exit at ascending node -> maintenance maneuver is required
- Westward exit at ascending node -> braking maneuver could be required -> warning provided to the user
- Exit at northernmost latitude -> inclination maintenance maneuver is required -> warning provided to the user
The Orbit Control module is responsible for evaluating and optimizing the maneuvers to maintain the current orbit within the limits, defined by the user, of the reference ground track. Different algorithms are available, depending on mission requirements, and tandem operations are supported.
This module is responsible for manipulating a reference orbit in order to:
- Shift it in time
- Displace the longitude of the initial node
- Synchronize it with defined ascending node condition
It creates the ground-track and nodes files for the new generated reference orbit.
Antenna Pointing Elements
A set of modules are available to generate antenna pointing elements in different formats (i.e. NASA IIRV, ESA STDM,…)
Two Line Elements (TLE) Management
A set of modules are available providing the following functionalities:
- Ingestion of NORAD TLE database and generation of internal orbit file for the requested satellite.
- Generation of satellite TLE set
Databases supported by focusleo
- Central Body: It contains the definition of different geophysical models used by the system, i.e. nutation model, Gravity Field Model, Dynamic Solid Tide Model,…
- Tracking Stations and Tracking Sites: It contains data related to ground-stations used for tracking purposes and corresponding sites
- Satellites: It contains satellite-specific data, such as its identifiers, mass and radiation and aerodynamical coefficients.
- Transponders: It contains data relative to on-board tracking instruments and transponders
- Thrusters: It contains the data relative to the thruster equipment of the satellites, including, among others, thruster alignment, specific impulse and characteristic durations
- Instruments: It contains basically the location and data that define the different parameters of the sensors modeled by the FDS (generic sensors with pyramidal and conical field of view), wheels, gyroscopes.