Table of Contents

  The Background 
  The Partners
  The Objectives
  The Satellites
  The Formation Flying Sensors
  Other Instruments
  The Propulsion Systems
  The Mission
  The GNC Experiments
  Autonomous Formation Flying (SSC)
  Autonomous Rendezvous (SSC)
  Proximity Operations, Final Approach and Recede (SSC)
  Autonomous Formation Control (DLR)
  Autonomous Orbit Keeping (DLR)
  Autonomous Formation Flying (CNES)
  Forced RF-Based Motion; Collision Avoidance (CNES)
  The Project Phases
  The Verification Process

THE BACKGROUND

The PRISMA project was established in early 2005, when SSC formulated a mission concept consisting of two satellites and a series of experiments in order to test formation flying and rendezvous technology in a real space mission.

The formation flying and rendezvous technology was chosen because it represents techniques which is crucial in any upcoming multi-satellite mission.

Formation Flying, the discipline to manoeuvre several spacecraft as one entity with a high level of autonomy, has several promising applications, such as building large antennas and telescopes using interferometer techniques.

Equally promising, but in a different field, is the Rendezvous technology, which can pave way for autonomous in-orbit inspection, servicing or assembly missions.

A common feature for both areas is that the relative position control loop, and its sensors, is actually implemented on-board rather than in the ground control. It is consequently very much about autonomy and about Guidance, Navigation and Control (GNC) technology.

The PRISMA project was established in early 2005, when SSC formulated a mission concept consisting of two satellites and a series of experiments in order to test formation flying and rendezvous technology in a real space mission.The formation flying and rendezvous technology was chosen because it represents techniques which is crucial in any upcoming multi-satellite mission. Formation Flying, the discipline to manoeuvre several spacecraft as one entity with a high level of autonomy, has several promising applications, such as building large antennas and telescopes using interferometer techniques. Equally promising, but in a different field, is the Rendezvous technology, which can pave way for autonomous in-orbit inspection, servicing or assembly missions.A common feature for both areas is that the relative position control loop, and its sensors, is actually implemented on-board rather than in the ground control. It is consequently very much about autonomy and about Guidance, Navigation and Control (GNC) technology.

The mission concept attracted DLR to contribute with GPS sensors and a navigation system, Alcatel (later replaced by CNES) to join with the Formation Flying RF sensor (FFRF) under development, and DTU to contribute with a development of the DTU star camera technology into a Vision Based Sensor (VBS).

Together with the SSC and Swedish National Space Board’s (SNSB) ambition to test in flight two new motor technologies developed in Sweden, PRISMA was established as a test bed for GNC technology and sensor/motor technology. The basic funding from the Swedish National Space Board was agreed in early 2005. The Mission Requirements Document could be agreed in April 2005.

A key driver for the cooperation was that all participants were invited by SSC to participate with their own experiments and thus sharing mission time and resources as compensation for the “investment” of development time and hardware to the project. Especially, DLR and CNES have developed their own Guidance and Control software in order to run closed loop experiments during their experiment slots. This software resides in the SSC OBSW.

The SSC Prisma Team picture

THE PARTNERS

The main partners and their contributions are the following: 

• German Aerospace Center (DLR), contributing with a GPS navigation system, comprising both the GPS hardware and navigation software. DLR also perform autonomous formation flying experiments with dedicated software embedded in the SSC overall GNC software. 
  
• CNES, contributing with the formation flying RF (FFRF) sensor and the associated manoeuvring experiment and software Also the Spanish agncy CDTI is involved in the development of the FFRF instrument. CNES also perform autonomous formation flying experiments with dedicated software embedded in the SSC overall GNC software. The complete package is called FFIORD experiment, Formation Flying In-Orbit Demonstration. 
  
• Technical University of Denmark (DTU), contributing with a Vision Based Sensor (VBS).

Other partners are: 

• ECAPS (a subsidiary company to SSC), exploring the new HPGP motor system (a replacement system for hydrazine)
• Nanospace (a subsidiary company to SSC), exploring the new Micropropulsion motor system based on MEMS tehnology 
• Techno Systems (TSD), contributing with a versatile Digital Video System
• Institute of Space Physics in Kiruna, Sweden, exploring a newly developed mass spectrometer with a MEMS velocity filter.

The distribution of task and responsibilities in the project can be depicted as in the figure below:

 

 

 

 

THE OBJECTIVES

Primary objectives

The primary objective of PRISMA is to demonstrate new technology in the field of Formation Flying and Rendezvous technique. This shall be done by a series of experiments divided in Guidance, Navigation and Control (GNC) experiments and sensor/actuator experiments. The GNC experiment sets consist of closed loop orbit control experiments conducted by SSC, DLR and CNES as in the following table:

 

Table 1

GNC Experiment Sets
Passive formation flying
Autonomous formation flying	SSC
Autonomous formation control	DLR
RF-based formation flying	CNES
Forced motion
Proximity Operations Final Approach and Recede	SSC
Forced RF-based Motion Collision Avoidance	CNES
Autonomous Rendezvous	SSC

Related to these GNC tests, hardware related tests will be conducted according to the following:

 

Table 2

Hardware Related Tests
Flight demo of HPGP Motor	SSC
Flight demo of Microthruster Motor	Nanospace
Validation of RF Sensor (FFRF)	CNES
Validation of Vision Based Sensor (VBS)	DTU

 

 

Secondary Objectives

In addition to the primary objectives, the mission has the following secondary objectives:

 

• Provide test flight for newly developed Data Handling System and The Power Conditioning and Distribution Unit with battery management
  electronics (SSC).
  
• Act as model project for new model based development of on-board software (SSC). 
  
• Demonstrate Autonomous Orbit Keeping of a single spacecraft (DLR).
  
• Demonstratea a newly developed Ground Support and Operational Support Equipment for multi-vehicle missions, the RAMSES system (SSC). 
  
• Provide a test flight for a Digital Video System developed by Techno System in Italy.
  
• Provide a test flight for a MEMS-based particle mass spectrometer from the Institute of Space Physics in Kiruna, Sweden. 

Link to PRISMA presentation 

THE SATELLITES

 

The PRISMA mission consists of two satellites: Mango and Tango, see picture below.

 

 

 

Mango is a 3-axis stabilized and has full 3D delta-V manoeuvrability independent of the spacecraft’s attitude. Mango is equipped with three propulsion systems, where the main system, a hydrazine propulsion system with 6 thrusters, has approximately 120 m/s delta-V capability. The central body of MANGO has exterior dimensions 750×750×820 mm. When deployed, the distance between the tips of the solar panels is 2600 mm.

The Tango satellite has a simplified, yet 3-axis stabilizing, magnetic attitude control system and no orbit manoeuvre capability. The Tango body is 570×740×295 mm.
The wet mass of the two spacecraft is approximately 190 kg. Mango is 150 kg and Tango is 40 kg.


The Mango and Tango architecture design can be seen in the following two pictures.

 

Block diagram of Mango

 

Block diagram of Tango

 

 

 

As can be seen in the previous block diagrams, most units and functions are redundant. In fact, the system has been built in order to be single failure tolerant in all essential functions. A Failure, Detection, Isolation and Recovery system (FDIR) has been implemented which enables autonomous switching to redundant units in most failure situations.

A specific complexity of flying a multi-satellite mission is of course the formation safety, meaning risk of collision or formation evaporation. Therefore, the FDIR system also can detect potential collision or evaporation risk, and go to “safe orbit”, an inherently collision-free orbit, with a minimum of autonomous maneuvers.

 

 

THE FORMATION FLYING SENSORS

GPS H/W

The GPS system and the navigation functions are a contribution of the German Space Center, DLR. The physical GPS system is presented in the picture below.

 

The MAIN and the TARGET spacecraft will, for redundancy purposes, each carry two independent GPS receivers that are operated in a cold configuration. Increased flexibility for handling non-zenith pointing attitudes is provided by two GPS antennas on each spacecraft, which are selected by an onboard algorithm for optimum GPS coverage or may, alternatively, be set by ground command.

In accord with the envisaged application, the miniature Phoenix receiver has been adopted for the PRISMA mission. Phoenix is a twelve channel single-frequency GPS receiver based on a commercial-off-the-shelf hardware platform and qualified by DLR for use in low Earth orbit (LEO). It offers single-frequency Coarse/Acquisition (C/A) code and carrier tracking and can be aided with a priori trajectory information to safely acquire GPS signals under high dynamic conditions. Upon tracking, Phoenix outputs a One-Pulse-Per-Second signal which is applied for the onboard time synchronization. Particularly important for formation flying missions is Phoenix’s feature to align the message time tags to integer GPS seconds which facilitates the differential processing of raw GPS data in the subsequent navigation filter.

The receiver is built around the GP4020 baseband processor of Zarlink, which combines the correlator, a microcontroller core with a 32 bit ARM7TDMI microprocessor and several peripheral functions in a single package. Phoenix provides a code tracking accuracy of better than 0.5 m and a carrier-phase accuracy of better than 1 mm at 45 dB-Hz. With a mass of the receiver board of 20 gr, a size of 70 x 47 x 15 mm and a power consumption of 0.85 W at begin of life, the receiver is particularly suited for small satellite missions like PRISMA.


GPS S/W and Navigation functions

The GPS-based GNC software on MAIN contributes both to the Basic Software (BSW) in terms of GPS sensor message processing as well as to the Application Software (ASW). The latter encapsulates all top-level guidance, navigation and control applications. The GPS interface (GIF) handles GPS raw data formats and ephemerides, and performs data sampling as well as coarse editing of MAIN and TARGET GPS data prior to the GPS-based orbit determination (GOD).

For an efficient software implementation, the architecture comprises two cores: the ORB core which is executed every 30 s and the GNC core, executed once per second to clearly separate the computational intensive GPS-based orbit determination from functions with low computational burden. The GPS-based orbit prediction (GOP) function evaluates the orbit, provided by the orbit determination function, at a 1 Hz rate and accounts for orbit manoeuvres which have not been taken into account already by GOD. It outputs MAIN and TARGET orbit states as well as associated quality indicators which are used by other onboard GNC functions as well as by the autonomous formation control (AFC) function implementing the guidance and control algorithms for formation acquisition and control.

GOD implements a reduced-dynamic orbit determination which accounts for the complex gravity field of the Earth (complete to order and degree 20), third body accelerations from the Sun and the Moon, as well as atmospheric drag and solar radiation pressure. Any deficiency in the assumed force model is absorbed in an empirical acceleration vector defined in a co-moving frame. A symmetric filter design has been chosen which adjusts the absolute states of both spacecraft. The relative spacecraft state is simply computed by differencing the absolute states. In this way, no explicit model for the relative motion is required and the interdependency of absolute and relative dynamics is fully exploited. For absolute state estimation, an ionosphere-free combination of pseudorange and L1 carrier-phase measurements is applied while a precise relative state is implicitly determined from single-difference carrier phase measurements.

FFRF system (CNES)

As partner on the PRISMA mission, CNES participates with the experiment FFIORD, Formation Flying In-orbit Ranging Demonstration, consisting of the delivery of the RF metrology instrument (FFRF), and GNC software implemented on the Mango satellite. CNES delivers the Formation Flying Radio Frequency (FFRF) subsystem, an RF metrology package which is baselined as first stage metrology sensor for many future European FF missions. The FFRF subsystem development has been performed with Thalse Alenia Space (TAS) as prime contractor, (contracted by both CNES and the Spanish agency CDTI).


The FFRF subsystem is in charge of the coarse relative positioning of 2 to 4 satellites. It produces relative position, velocity and line-of-sight (LOS) as inputs to GNC subsystem for which it provides coarse metrology measurements. As the first element in the future Formation Flying metrology system chain, the FFRF sensor ensures initial good relative navigation accuracy for the subsequent optical metrology subsystems (coarse optical lateral metrology, fine optical metrology, and fine longitudinal metrology).

As partner on the PRISMA mission, CNES participates with the experiment FFIORD, Formation Flying In-orbit Ranging Demonstration, consisting of the delivery of the RF metrology instrument (FFRF), and GNC software implemented on the Mango satellite. CNES delivers the Formation Flying Radio Frequency (FFRF) subsystem, an RF metrology package which is baselined as first stage metrology sensor for many future European FF missions. The FFRF subsystem development has been performed with Thalse Alenia Space (TAS) as prime contractor, (contracted by both CNES and the Spanish agency CDTI). The FFRF subsystem is in charge of the coarse relative positioning of 2 to 4 satellites. It produces relative position, velocity and line-of-sight (LOS) as inputs to GNC subsystem for which it provides coarse metrology measurements. As the first element in the future Formation Flying metrology system chain, the FFRF sensor ensures initial good relative navigation accuracy for the subsequent optical metrology subsystems (coarse optical lateral metrology, fine optical metrology, and fine longitudinal metrology).

 

 

The FFRF subsystem consists of one FFRF terminal and up to 4 sets of antennas on each satellite in a constellation. On PRISMA, MAIN is equipped with a triplet (1 Rx/Tx master and 2 Rx slaves) and TARGET is fitted with 3 single Rx/Tx antennas for omni-directional coverage. The antennas are manufactured by SAAB Ericsson Space.

The terminal operates by dual frequency in S-band, and multi-satellite signal exchange is managed by TDMA. Ranging and angular measurements are extracted from received signals and are used for computing relative position (1 cm accuracy), velocity and LOS (1° accuracy). In addition to providing relative navigation measurements, the FFRF subsystem also provides time bias synchronization between the satellite clocks as well as an inter satellite link (ISL) as auxiliary functionality (12 kbit/s or 4 kbit/s bitrate).

For further information, see http://smsc.cnes.fr/PRISMA/index.htm


Vision Based Sensor (DTU, Denmark)

DTU has developed the microASC, a fully autonomous miniature stellar reference instrument, for use onboard spacecrafts as an attitude reference sensor. The microASC is designed for highly flexible configurations and can host from 1 to 4 camera head units (CHU), located at suitable places and directions on a spacecraft, such that a fully redundant blinding free attitude sensor configuration can be achieved.
Onboard PRISMA two CHUs are used as standard attitude reference sensors, with their pointing directions such that simultaneous blinding by Sun, Earth and the Target SC is avoided during the complex fly-around maneuvers.

The third port on the microASC is also equipped with a standard CHU. This CHU is however pointed in the forward direction, such that the target SC can be in its field of view for most mission phases.

The fourth port is equipped with a CHU with a modified focal length, iris and electronic shutter, so as to enable operations at high light conditions at close range.

The VBS data processing is performed in the microASC, which is made possible by its huge spare processing power for its standard operation. The basic SW has four different modes of operation: 

• The Far Range mode is entered whenever both stars and the target SC is detectable in the FOV,
  
• The intermediate range mode is entered when the target is to bright to allow for star detection
  
• The short range mode when features of the target in the FOV is discernible.

The cooperative mode operates in a similar fashion, however instead of relying on natural illumination, five LEDs on each face of the target, arranged in specific patterns, are used to generate feature points in the image. To ensure a sufficient contrast ratio to the natural illumination background, the VBS cameras and LEDs are operated in synchronous pulse mode.

The non-cooperative short range mode operation principle is based on a 3D model database of the target stored in the microASC. This model contains information on target features, their relative distances and location. Features of the target image obtained using natural illumination from Sun and Earth albedo, is matched to the database model, whereby both relative position and pose information extraction is possible.

 

 

OTHER INSTRUMENTS

PRISMA also acts as a test flight for two other instruments, one Digital Video System, and one Particla Analyzer, the PRIMA unit.

Digital Video System (Techno Systems, Italy)

The DVS system is developed by Techno Systems in Italy. It is a camera system with the following features:

• High-resolution color CCD,
• Video or still frames
• High framerate (up to 20 fps)
• Compression factor 1 to 60

For further info, see http://www.tsdev.it/digitalvideosystems_for_PRISMA.htm

PRIMA Mass Analyzer (IRF)

The Swedish Institute of Space Physics (IRF) participates in the project with a completely new type of instrument, PRIMA (PRIsma Mass Analyzer). PRIMA is a low energy (<100eV) ion mass analyzer based on the Solar WInd Monitor (SWIM) sensor developed for the Indian Chandrayaan-1 mission. 

The Swedish Institute of Space Physics (IRF) participates in the project with a completely new type of instrument, PRIMA (PRIsma Mass Analyzer). PRIMA is a low energy (<100eV) ion mass analyzer based on the Solar WInd Monitor (SWIM) sensor developed for the Indian Chandrayaan-1 mission. 

For more info see http://www.irf.se/Topical/?dbfile=Prisma%20and%20PRIMA&dbsec=Administration

THE PROPULSION SYSTEMS

PRISMA Mango satellite is equipped with three propulsion systems,

• The nominal hydrazine system (red in the figure)
• The HPGP system (green in the figure)
• The Micropropulsion system (cyan in the figure).

The hydrazine system was chosen as the nominal system since both the other systems, the HPGP and the Microprop system, is new develop-ments, for which the PRISMA satellite is the flight demonstration mission.

 

The Hydrazine propulsion system consists of six 1-N thrusters directed towards the MAIN S/C Centre of Gravity (COG), giving torque-free translational capability. The propellant tank contains 11 kg of usable fuel and gives approximately 110 m/s delta-V over the mission. Firing times will range from 0.1s (requested typically at autonomous formation flying and proximity operations) up to steady state burns up to 2 minutes.

The Hydrazine system has been designed by Swedish Space Corporation, based on procured components from mainly US suppliers.

The High Performance Green Propellant (HPGP) motor experiment is a new propulsion system introducing environmentally friendly, non-toxic ADN-based fuel which theoretically gives 10% better impulse and 30% higher density than hydrazine. The actual development tests have been ongoing for several years, and is now qualified for flight.

The performance has proved to be better than hydrazine in all modes of operation.

The HPGP propulsion system has two 1 N thruster, also directed towards COG. The system provides redundancy to the main hydrazine system if any nominal thruster should fail. The propellant tank contains 5 kg of usable fuel and gives approximately 60 m/s delta-V over the mission.

The development, involving the propellant itself and a compatible 1-N thruster and catalyst bed, is driven by ECAPS, a subsidiary company to SSC, and has been supported by ESA for several years.

The Micropropulsion system is based on MEMS technology (Micro-ElectroMechanical Systems) and is under development by NanoSpace, a subsidiary of SSC, on contracts from ESA and the Swedish national space board.

The micropropulsion system is developed by the SSC subsidiary company Nanospace. It shall be capable of delivering accurate thrust ranging from tenths of micro- Newtons up to a milli-Newtons. Such a system would be a potential candidate for future missions where extremely low and accurate thrust is requested, such as Darwin, Gaia, Xeus, Proba-3, Simbol-X and others. The key component is the golf-ball sized thruster module (see figure) containing a silicon wafer stack with four complete rocket engines with integrated flow control valves, filters, and heaters.

Extremely small internal heaters inside the thrust chamber increase the performance of the system in terms of specific impulse. The propellant is Nitrogen. The four thrusters are orthogonally distributed in the equator plane of the golf ball sized thruster module. The thrust will be too low for PRISMA to actually utilize in the fairly disturbed LEO orbit it is in. However, the functionality and performance of the thruster system will be flight demonstrated during the mission, mainly via analyzing the command torque to the reaction wheel, but also by analyzing the velocity increments generated via the Precise Orbit Determination functionality.

THE MISSION

PRISMA will be launched by a Dnepr rocket together with the French satellite Picard. The orbit is a sun synchronous orbit with 725 km altitude and 06.00h ascending node. The two satellites Mango and Tango are clamped together as depicted in Figure. 


Tango is separated from Mango after the first commissioning in orbit after an initial commissioning campaign during which all on-board equipment is checked out. The separation of TANGO is observed with MANGO’s on-board Digital Video System (DVS).

During the whole mission, ground only communicates with MANGO. Mango in turn communicates with TANGO via an intersatellite Link on 400 MHz. This link is specified to range at least 10 km
After separation, a series of manoeuvring, sensor and motor experiments starts. This sequence of experiments
is planned in increasing order of difficulty. Each experiment group is divided into an early harvest part and a part that completes the experiment. The different experiment sets will also be ordered in a way that is considered increasing in difficulty or complexity.
The initial experiments consist of passive formation flying checking out the GPS navigation functionality, mainly provided in the DLR contribution.. These functionalities will be fundamental to the continuation of the mission and to the correct functionality of Safe Orbit Control.
These experiments are followed by DLR’s passive formation flying, and a checkout of the VBS, and RF instruments during a campaign of GPS based forced motion.
The most complex functionality is considered to be the VBS autonomous rendezvous during which the far range detection capabilities of the VBS is examined and during which a fully autonomous rendezvous with TANGO is performed down to the closest possible approach distance.
The total mission duration for all nominal operations is planned to be around 300 days.

Link to Mission Timeline

The orbit is a Dawn-Dusk Sun Synchrounous orbit, meaning that the orbit plane is close to perpendicular to the sun direction all around the year. This simplifies the power generation since the solar panels can be directed to the sun in a natural way during most of the time, see figure below.

The orbit is a Dawn-Dusk Sun Synchrounous orbit, meaning that the orbit plane is close to perpendicular to the sun direction all around the year. This simplifies the power generation since the solar panels can be directed to the sun in a natural way during most of the time, see figure below.

   

The GNC experiments are further described in the following sections.

THE GNC EXPERIMENTS

The GNC manoeuvring experiments consists of SSC, DLR, CNES, and DTU experiments. These include closed loop orbit control of the MAIN spacecraft.

The different experiments will be distributed over the mission length in a sequence with increasing level of complexity, ensuring early harvest results for all parties in the beginning of the mission.

SSC is responsible for the overall experiment planning and design, and has designed experiment sequences in all experiment sets. DLR and CNES each has dedicated autonomous formation flying experiments based on their respective sensor system contributions; the GPS and FFRF systems. Both have dedicated software embedded in the GNC core software created by SSC. DTU supports all VBS-based experiments with the highly sophisticated functionality of the VBS camera system.

The figure below depicts the overall GNC mode diagram.

GNC experiments under SSC responsibility

The SSC GNC experiments to be carried out can be sorted into four different groups:

    1. Autonomous Formation Flying (AFF) – Passive formation flying based on GPS. 
    2. Autonomous rendezvous based on VBS 
    3. Proximity Operations – 3D proximity precision forced motion based on GPS and/or VBS. 
    4. Final Approach and Recede – Very close VBS based operation.

The different experiments are summarized in
Table 1 where also typical relative distances are given.


Table 1

Experiment	Distance range (m)	Sensor
Autonomous Formation Flying	20 – 5000	GPS
Homing and Rendezvous	10 – 100 000	VBS
Proximity Operations	5 – 100	VBS and/or GPS
Final Approach and Recede		VBS

These four experiments are selected because they highlight different important aspects of formation flying and each of the experiments is based on one or several model scenarios.


AUTONOMOUS FORMATION FLYING (SSC)

Link to film section

Autonomous Formation Flying (AFF) experiment concerns passive relative orbits. Such orbits make use of the slightly different orbits of the two spacecraft to create a natural periodic motion. The resulting orbit is in this way passive and the only control required is to maintain the orbit counter-acting external disturbances, such as solar pressure, aerodynamic drag and gravitational irregularities.

The orbit control is based on relative GPS with navigation data provided by the navigation filter supplied by DLR.

The model missions for the AFF are situations occurring in On-Orbit Assembly and Inspection, passive apertures, and loose formations.
Several different in-plane as well as out-of-plane orbits will be part of this experiment. The different orbits will also be applied at different distances.

As shown in the mode diagram, the AFF Mode is the central mode in the mode architecture. The reason is that the AFF orbit control is considered to be the most robust of the different experiments. It is also considered to be the simplest mode in terms of operations. For these reasons, the mode will be used as a parking mode while preparing the other experiments.

AUTONOMOUS RENDEZVOUS (SSC)

Link to film section


The Autonomous Rendezvous (ARV) experiment consists of complete autonomous approach and rendezvous based on optical information only, provided by the VBS system.

The different phases are:

• the autonomous location of the TARGET spacecraft,
• orbit phasing and aligning to the Targt orbit
• intermediate transfer (several steps)
• final approach to the TARGET spacecraft through a predetermined approach corridor. 

All of these steps are designed to be automatic but with possibilities for commanded escape.

The typical model mission for this experiment is a Mars Sample Return Mission but servicing in geosynchronous orbit as well as assembly in escape orbit are also relevant model scenarios.

PROXIMITY OPERATIONS, FINAL APPROACH AND RECEDE (SSC)

Link to film section


The 3D Proximity Operations experiment consists of forced trajectories in the close proximity of the TARGET spacecraft.
The navigation takes place in a network of flight-paths defined about a virtual structure defined around the TARGET spacecraft as illustrated in figure below.

Typical model missions are On-Orbit Servicing, Inspection, and Assembly.
The proximity operation experiment has two different branches. The first is based on relative GPS where the navigation is retrieved from the DLR navigation filter. The second is VBS-based where the navigation is made using the VBS. The VBS can operate with a cooperative TARGET equipped with an LED pattern but also with a non-cooperative TARGET where there are no aids for the VBS other than the geometry of the TARGET itself. In these operational phases, the VBS delivers both the relative distance and the relative attitude to the TARGET spacecraft.

The controller used in the proximity operations experiment is based on a non-linear model predictive control law.

The 3D Proximity Operations experiment consists of forced trajectories in the close proximity of the TARGET spacecraft. The navigation takes place in a network of flight-paths defined about a virtual structure defined around the TARGET spacecraft as illustrated in figure below.Typical model missions are On-Orbit Servicing, Inspection, and Assembly. The proximity operation experiment has two different branches. The first is based on relative GPS where the navigation is retrieved from the DLR navigation filter. The second is VBS-based where the navigation is made using the VBS. The VBS can operate with a cooperative TARGET equipped with an LED pattern but also with a non-cooperative TARGET where there are no aids for the VBS other than the geometry of the TARGET itself. In these operational phases, the VBS delivers both the relative distance and the relative attitude to the TARGET spacecraft. The controller used in the proximity operations experiment is based on a non-linear model predictive control law.

AUTONOMOUS FORMATION CONTROL (DLR)

Based on the sole use of GPS as navigation sensor, the Spaceborne Autonomous Formation Flying Experiment (SAFE) will be conducted to demonstrate a fully autonomous, robust and precise formation flying of spacecraft. To this end, a fuel-optimized guidance and control algorithm will be employed at typical spacecraft distances of 100 to 2000 m, which is representative of future bi-static radar satellite formation flying missions.

The guidance concept applies the eccentricity¬/inclination vector separation (D’Amico et al. 2006) which avoids collision hazard from along-track position uncertainties through the proper separation of the two spacecraft in radial and cross-track direction. A deterministic impulsive feedback control is applied for formation keeping which is, through its low computational burden, ideally suited for an onboard implementation.

SAFE will exploit the versatility and generality of the relative eccentricity/inclination vector control in order to acquire and maintain safe close formation flying configurations in complete autonomy over long time intervals (weeks). For control window sizes of 5 m, 7 m and 20 m in the relative inclination, eccentricity and mean argument of latitude, respectively, and in the presence of realistic sensor and actuator performances typically six orbit maneuvers are executed per day requiring an overall daily delta-v budget of about 0.04 m/s.


AUTONOMOUS ORBIT KEEPING (DLR)

To be written.


AUTONOMOUS FORMATION FLYING (CNES)

The CNES Formation Flying mode is invoked on command from the central AFF mode (see mode diagram). This enables CNES embedded S/W to take over navigation and guidance functions. Navigation is now based on the FFRF instrument.
The CNES experiments in closed loop is intended to demonstrate e.g.

• Proximity operations:
  • Station-keeping at different distances and offset positions from the orbit track (VBAR)
  •  Low speed translations in-plane and out of plane
• Rendezvous 
• Collision avoidance (autonomous transfer on Football orbit in 1 or 2 manoeuvres)
• Stand-by on a relative orbit
• Recovery after system anomaly.

For more info on the CNES FF, see http://smsc.cnes.fr/PRISMA/GP_missionF.htm


FORCED RF-BASED MOTION; COLLISION AVOIDANCE (CNES)


THE PROJECT PHASES

THE VERIFICATION PROCESS
The verification and test process has to a large extent been similar to a normal satellite project. On system level, there are however some new elements specific for the multisatellite project.

The verification in a very brief overview is depited in the figure below. The main test verification elements are the following:

• The STM tests. This was a structure test model with the main purpose of qualifying the new structure of both Mango and Tango, the separation system between the two satellites, and certain propulsion equipment.
   
• The BTM environment. This was a test environment where a mixture of EM and FM units was connected to the EM data handling system and power systems, in order to interatively debug I/F problems and check basic functionality. Certain rudimentay S/W was delivered from the OBSW team. 
• The Satlab S/W development and test environment.
• The Integratied test environment (also called the SFPT tests (the System Functional and Performance Tests).
• The Environmental test campaign.

 

 

The Satlab environment
The Software System Tests take place in a real-time software validation facility developed at SSC, called SatLab. The environment is illustrated below and consists of one engineering model (EM) of the computer board for each of the MAIN and TARGET spacecraft. The complete onboard software runs on a single computer for each of the satellites. In the real case, the boards are connected to the sensor and actuator interface electronics via a CAN bus.

SATLAB ENVIRONMENT

The SatLab environment includes a simulator, SatSim, which simulates the interface electronics, sensors, actuators, mechanical dynamics, and space environment for both of the satellites.

The System Functional and Performance Tests environment
The System Functional and Performance Tests (SFPT) tests are a series of test on spacecraft level aiming at verifying as much functionality as possible with the real spacecraft hardware included. The series of tests comprises both Open Loop Tests and Closed Loop tests(where the correct response to a predefined sensor stimuli is verified) and closed loop tests (where actualthe following (not exhaustive) test list:

The Open Loop Tests
The purpose of the open loop testing is to verify the correct response of the system to a well-known stimuli through-out the whole chain of functions that is from the sensor, via harness and interface electronics to the onboard data bus, to the onboard computer and the different layers of S/W, back through the system to a specific actuator. The test comprises e.g. illumination of the sun sensors and sun presence sensors and the corresponding response on magnetic torques or reaction wheels. This is valid for both S/C individually.

GPS Closed Loop MAIN/TARGET