Challenges in Control and Autonomy of Communications Satellites

Copyright © IF AC Automatic Control in Aerospace, IFAC Seoul, Korea, 1998

CHALLENGES IN CONTROL AND AUTONOMY OF COMMUNICATIONS SATELLITES

E. Gottzein, W. Fichter, A. Jablonski, O. Juckenhofel, M. Mittnacht, C. Miiller, Muller, M. Surauer

Daimler-Benz Aerospace, Dornier Domier Satellitensysteme GmbH, PO Box 80 11 69, 81663 Munich, Germany

The rapid growth, predicted in space communications stimulated ideas beyond present solutions. In addition to geosynchronous satellites, which presently carry nearly all space transponders in use, new low and intermediate orbit constellations of satellites and even constellations in eccentric orbits are in design and planning. The paper gives an overview of this scenario to derive the new and challenging requirements on the satellites attitude and orbit control system (AOeS) (AOCS) and presents solutions, based on examples from current projects. For the first time, commercial AGCS AOCS have to be produced in series and the technologies employed are driven equally by performance, quality and effort (cost and schedule). With the high number of satellites, to be operated simultaneously in constellations, on-board autonomy with minimum ground intervention during operation becomes a "must". This concerns nominal operation as well as reconfiguration and repair in case of failures. New methods for autonomous failure management, and autonomous orbit detennination determination and station keeping are therefore a central subject of the paper. Copyright © 1998 IFAC

Keywords: Attitude Control, Orbit Determination, Station Keeping, GPS, On-Board Autonomy, Failure Management, Reconfiguration, Satellite Constellations.

launched August 19, 1964, mass 37.5 kg, power 29 Watts. Like most of the early satellites, SYNCOM-3 was spinstabilized. Positioned above the Pacific Ocean, it enabled European viewers to watch the Tokyo Olympic Games live.

1. SPACE COMMUNICATIONS MARKET DEVELOPMENT In 1945 the British scientist and author Arthur C. Clark suggested to place three satellites in an equatorial Earth orbit with a radius of 42,164.5 km, corresponding to an orbital period of one sidereal day (23h56min). Stationed 120 deg apart, each would be able to cover one third of the populated area of the Earth, and the constellation of all three together would enable global communication.

The idea of geosynchronous orbital period was adapted to the specific geographic requirements in the former Soviet Union. The first MOLNIYA communications satellite, launched April 23, 1965, operated from a quasisynchronous elliptical orbit, altitude at perigee 538 km and at apogee 39,300 km, inclination with respect to the equator 65.5 deg. One satellite of this series guaranteed communication for 8 hours and three properly synchronized

The first geosynchronous communications satellite (GEO satellite) realizing Clarks idea was SYNCOM-3, which was

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in relation to each other could provide 24 hour coverage for areas in the northern hemisphere. The first MOLNIYA was already three-axis stabilized, power 700 Watts. MOLNIYA satellites of the ORBITA network were used for the hotline between the White House and the Krernlin, which was established for security reasons in 1971.

Satellite constellations of many satellites in low Earth orbit (LEO satellites) are in preparation to meet the projected demand for high and low data rate satellite services, many of them using intersatellite links. The low data rate (LDR) network constellations Globalstar and Iridium, expected to be operational in the near future, will allow mobile telephone communications between any point on Earth. The data rate of the Ka-band ISL of Iridium is 25 Mbps.

In the following period, international organizations were founded by groups of nations to procure and launch telecommunications satellites and to provide communications satellite services all over the world. The global operators INTELSAT (founded in 1964) and INTERSPUTNIK were supplemented by regional operators like EUTELSAT, ARABSAT etc .. A period of rapid growth in satellite communications services followed, stimulated by a large demand for telecommunications and broadcasting services. A few years ago new applications like Mobile and High Data Rate (HDR) services to any place around the world and ultimately the "INTERNET IN THE SKY" accelerated this trend even more.

The advanced HDR constellation Teledesic is targeted towards high speed data transmission between centers and uses Ka-band ISL with 1.5 Gbps.

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Fig.1.2 shows the increase in satellite market value to be expected in the next years to fulfill the demand for increasing payload capacity. For GEO satellites, the market demand shows a cyclic trend with peaks around 1998 and 2010. A specific market segment for servICIng the densely populated high latitude northern areas will be covered by satellite systems in highly eccentric orbits (HEOs). The quasi-synchronous active arcs over latitudes between 45 and 70 deg provide improved visibility for Ka-band links and minimize atmospheric damping.

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Overall satellite industry growth by sector: Launches, manufacturing & operations.

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The Global SPACEWA yrM Network scheduled to be operational between 2002 to 2005 proposes to station 4 . 2 regional geosynchronous systems at geosynchronous altitude over North America, the Pacific Rim, Europe/Africa and Central/South America and to interconnect the satellites by Ka-band intersatellite link.

GEO-Satellites (28 GHz, 15 GHz, 6 GHz) LEO-Satellites (28 GHz, 1.6 GHz)

57% 43%

What are the messages from Fig.1.1 and Fig.1.2 to the satellite industry and particular the satellite designers? 1. The international market, international customer situation and the amount of capital needed makes global market and

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partner strategy a "MUST" for space companies. This means international competitiveness of the product in cost, quality and performance, but also a basically open attitude for cooperation and worksharing, in particular for the designers. satellite designers.

0.1 deg in North/South and EastIW EastIWest a box of ± 0.1 est direction, on collocated positions < ± 0.05 deg.

There are various types of communications satellites in the geostationary orbit. The main applications are summarized below:

2. For the first time in space history, commercial satellites are produced in series. This requires re-engineering the product and design and production procedures. 3. Comparing the 30 % sales growth of the satellite procurement market with the much bigger revenue generated by the satellite services market, challenges the satellite designers to improve operational availability and autonomy and operational life time of their systems. The attitude and orbit control (AOCS) subsystem plays a key role in all three of the above challenges. •

Operational life time can be increased by prudent management of on-board resources, such as propellant and the use of electrical propulsion systems

•

Operational availability can be increased by minimizing "outage" time, through an on-board autonomous failure detection, isolation, and recovery (FDIR) system.

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Operational cost and reliability can be reduced by autonomous orbit determination and position keeping.

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CC-, Ku-Band) Television Distribution CC-. Satellites on fixed or movable orbit position, operated by national or international entities or companies (e. g. Asiasat, Hispasat, Eutelsat, Intelsat). In areas with high rain attenuation they are equipped mainly with C-band payloads or they have mixed Ku-band and C-band transponders. Antenna coverages are reconfigurable or can be repointed.

•

Direct TV CKu-Band) Satellites on fixed orbit positions, operated by national or international companies (e.g. Nahuelsat, Eutelsat, USDBS). Ku-band payload transponders with high RF power up to 200 watts provide high EIRP for direct TV or HDTV. Antenna beams can be reconfigurable or 0.1 deg). In repointable with very accurate pointing areas with high payload channel demand satellites are collocated on the same position (e.g. ASTRA with 4 satellites).

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CC-, Ku-Band) Telecommunication Bundle CC-. Satellites on fixed or movable orbit position, operated by national or international entities or companies (e.g. Kopernikus, Hispasat, Eutelsat, Intelsat). In areas with high rain attenuation, they are equipped mainly with Cband payloads or they have mixed Ku-band and C-band transponders. Antenna coverages are reconfigurable or can be repointed (complex antenna farms). Trunk service between distribution centers with large ground antennas allow a high number of channels at low power level.

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Connections,. Data Transfer CC-. CC-, Telecommunications Connections Ku-Band) Satellites on fixed or movable orbit position, operated by national or international entities or companies (e.g. Romantis, Sinosat, Galaxy, Spaceway, Eutelsat). In areas with high rain attenuation they are equipped mainly with C-band payloads or they have mixed Kuband and C-band transponders. Antenna coverages are reconfigurable or can be repointed (complex antenna farms). Service between individual V-Sat terminals at medium power levels.

•

Mobile Communications CL-Band) Satellites on fixed or movable orbit position, operated 11). Payload by national companies (e.g. M-Sat, Aussat II). in L-band for mobile communications (truck services),

On-board autonomy is increasingly important for space constellations, using intersatellite links, because the ephemeris of the individual satellites have to be known precisely on-board, while their visibility (time and occurrence) to ground stations is limited. The following chapters deal with all three of the above AOCS related tasks, presenting solutions by examples from GEO, HEO and LEO projects. 2. SCENARIOS AND REQUIREMENTS

2.1 Scenarios GEO Satellites. Most of the satellites in the geosynchronous orbit are three axis stabilized communications satellites. The large power demand is provided by flat solar array panels oriented towards the Sun. Most satellites are injected into elliptical transfer orbit (apogee at 36,000 km, perigee at :::::200 km, inclination depending on launch site between 5 ::::200 and 28 deg). In order to realize the high amount of velocity increments for injection into the geosynchronous orbit and for station keeping, generally bi-propellant propulsion systems are used. Satellite lifetime is now extended beyond 15 years, where electric propulsion for station keeping becomes very attractive. Orbit positions shall be kept within

143

large largereflector-antenna reflector-antennaororphased phasedarray arrayantenna antennawith with nationwide nationwidecoverage. coverage. LEO/MED LEO-constellations are LEO/MEDConstellations. Constellations. LEO-constellations are composed composedofoflarge largenumbers numbersofofsatellites satellitesononvarious variousLEOLEOorbits orbitscovering coveringalmost almostthe thewhole wholeEarth Earthsurface. surface.The Theorbit orbit altitude altitude and and availability availability requirements requirements are are defining defining the the number numberofofsatellites satellites(between (between2020and and288). 288).They Theyform forma a complete completenet netfor forcommunications communicationsservices. services.The Thetraffic trafficisis controlled either by the satellites or gateway controlled either by the satellites or gateway ground ground stations. stations. Systems Systems with with permanent permanent visibility visibility offer offer a a continuous controlling continuoustraffic trafficlink, link,which whichisisguaranteed guaranteedbybycontrolling the constellation to form a closed the constellation to form a closednet netand andbybyhandover handover procedures procedures when when satellites satellites are are disappearing disappearing from from users users visibility. visibility.The Thesatellites satellitesare arenormally normallyinjected injectedinto intoorbit orbitinin multiple multiplelaunch launchconfiguration, configuration,and andfinal final positioning positioningand and orbit maintenance is performed by own propulsion orbit maintenance is performed by own propulsionsystems. systems.

Table Table2.1 2.1a)a)LEOIMEO LEOIMEOConstellations Constellations System System Orbit OrbitPlanes Planes

Ellipso Ellipso (ellipt.) (elli.et.) 22

Ellipso Ellipso ECCO ECCO Teledesic Teledesic (circ.) (circ.) 1 1212 equat. equat.

Satellites Satellites per 4+1 6+1 perPlane Plane 4+1 6+1 Satellites Satellites TotalOp. 66 88 TotalOp. Orbit 0 OrbitInclin. Inclin.[0][0] 116 116 0 Altitude 7500/670 Altitude 7500/670 8060 8060 Apo./Peri. [km] Apo./Peri. [km] Band Band [kg] Mass Mass [kg] 700-800 700-800 700-800 700-800 OperationalNear OperationallYear 2000 2000

• • Paging, Paging,Message MessageTransfer Transfer(Very (VeryLow LowData DataRates RatesininL-LBand UHF} Bandandlor and/orUHF) Mobile Mobile communications communications networks networks forfor low low data data rate rate transmission. ...)... ) transmission.The Thesatellites satellites(e.g. (e.g.Orbcom, Orbcom,LEO LEOOne, One, are arevery verysmall smalland andserve serveasasrelays relaysusing usingalso alsoUHF UHFwith with good goodpropagation. propagation.Low Lowpointing pointingaccuracy accuracyallows allowsfor for simple simpleattitude attitudecontrol control(e.g. (e.g.gravity gravitygradient). gradient). Big BigLEOslMEOs LEOslMEOs (Iridium, (Iridium,Globalstar, Globalstar,ICO-P, ICO-P,Ellipso) Ellipso) • • Iridium Iridium- Independent - IndependentMobile MobileCommunications CommunicationsNetwork Network The Iridium network consists of The Iridium network consists of6666satellites satellitesplus plusspares spares onon6 6nearly nearlypolar polarorbit orbitplanes planes(i=86 (i=86deg) deg)with with780 780km km altitude. altitude.ItItisisananindependent independentsystem systemwith withononboard board switching communications switchingand androuting routingofof communicationslinks linksfor forvoice voice SIC link in L-band service. The user (cell-phone) service. The user (cell-phone) SIC link in L-bandand and the Ka-band. 6 6 phased phased array array the inter-satellite inter-satellite link link inin Ka-band. antenna antennaconnects connectsthe theadjacent adjacentsatellites satellitesininplane planeand and neighboring planes to form a worldwide traffic net. neighboring planes to form a worldwide traffic net. • • Globalstar Mobile Mobile Communications Communications Network Network Globalstar Controlled by Gateways Controlled by Gateways The TheGlobalstar Globalstarsystem systemconsists consistsofof4848 satellites satellitesplus plus spares sparesonon8 8orbit orbitplanes planeswith with5252deg deginclination inclinationinin1,400 1,400 km km height. height. The The system system switching switching and and routing routing ofof communications communicationslinks linksfor formainly mainlyvoice voiceservice serviceare aredone done bybythe thegateway gatewayground groundstations. stations.The Theuser user(cell-phone) (cell-phone) SIC SIClink linkininL-band L-bandand andthe thegateway gatewaylink linkininC-band. C-band.The The user userantenna antennalink linkfor forreceive receiveand andtransmit transmitisisformed formedbyby8 8 beams. beams.Power Powerlevel levelisistraffic trafficdependent. dependent.The Thesatellites satellites are nadir pointed with yaw steering. are nadir pointed with yaw steering.

1111 288 288 00 9090 2000 2000 1350 1350

280 280

Orbcom Orbcom GlobalGlobal- ICO ICO star star 88 33 22

Orbit OrbitPlanes Planes Satellites Satellites per 88 perPlane Plane Satellites Satellites TotalOp. 2424 TotalOp. Orbit OrbitInclin. Inclin.[0][0] 4545 Altitude 770 Altitude 770 Apo./Peri. [km] Apo./Peri. [km] Band Band [kg] Mass [kg] 4040 Mass OperationallYear OperationallYear 1997 1997

Little LittleLEOs LEOs

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KaKa 1500 1500 2002 2002

Table Table2.1 2.1b)b}LEOIMEO LEOIMEOConstellations Constellations System System

To Toavoid avoiddebris debrisininorbit, orbit,atatend endofoflife lifeororinincase caseofoffailure failure the thesatellites satellitesmust mustbebetransferred transferredinto intoa agraveyard graveyardorbit orbitoror de-orbited de-orbitedinto intothe theatmosphere. atmosphere.

• • ICO ICO- Telecom - TelecomService Service AAsatellite satellitesplus plusspares spares satellitesystem systemconsisting consistingofof2020satellites operating on 2 planes with 45 deg inclination operating on 2 planes with 45 deg inclinationinin10,300 10,300 km km altitude. altitude. The The satellites satellites provide provide global global teletelecommunications services for individual user communications services for individual userterminals. terminals. Satellite Satellitecontrol controlsimilar similartotoLEOs. LEOs.

Iridium Iridium 66

6+1 6+1

1010+ + 2 2 1111+ +1 1

4848 5252 1400 1400

2020 6666 4545 8686 10300 10300 780 780

LL 450 450 1998 1998

LL 2450 2450 689 689 2000 2000 1998 1998

Next NextGeneration GenerationLEOs LEOs There are There aremany manyconstellation constellationprograms programsplanned plannedtotoreplace replace and improve the existing LEOs. The predicted and improve the existing LEOs. The predictedgrowth growthinin traffic trafficcapacity capacityrequires requiresmore morepowerful powerfulsystems. systems.Satellites Satellites with longer lifetime and more available with longer lifetime and more availableelectrical electricalpower power will willuse useelectric electricpropulsion propulsionfor forstation stationkeeping keepingand andalso alsofor for orbit orbitraising raisingtotosave savelaunch launchcosts. costs.This Thisleads leadstotomaneuvers maneuvers with withvery verylong longduration durationatatvery verylow lowthrust thrustlevels. levels.

144 144

MegaLEOs (Teledesic, Sky Bridge) •

orbital planes. Satellites in constellations have to be synchronized properly relative to each other, in particular when intersatellite links are used to form an orbital data way. way.

Teledesic Very big satellite system consisting of 288 satellites in 12 orbital planes in 1,350 km height. The system will provide high data rate broadcast services, video, worldwide. The teleconferencing to fixed locations worldwide. satellites provide electrical power in the range of 6,000 Watts and electric propulsion is envisaged. Satellite control, station keeping and orbit maintenance are fully autonomous.

The performance requirements of the attitude control satellites, system are mission dependent. For GEO satellites, ±O.l °0 in generally, the beam pointing shall be kept within ±0.1 pitch and roll and ±0.2° in yaw. yaw. High power satellites for respectively. direct TV request ±0.05° and ±O.125°, respectively. Satellites with large antenna farms will need individual loop. Intelligent beam pointing control in closed or open loop. control algorithms shall compensate for long term and periodical distortions. LEOs normally allow for pointing 0.5-100 and 3-5 3_5 00 for simple satellites, respectively. errors of 0.5-1 HEO satellites have similar requirements to GEOs, but pointing shall be directed towards the center of coverage area instead of the nadir direction. This requires attitude maneuvers to compensate for motion along the active arc. For both LEOs and HEOs yaw steering is required to avoid double axis solar array drives, and to provide defined sun orientation for proper power and thermal conditions.

HEO Constellations. High data rate service (Ka-band), rings. augmenting and connecting fiber optic rings.

•

Pentriad Pentriad system of 9 high power satellites (plus spares) will be operated from 3 linked highly elliptical orbits type). The system can provide broadband (Molniya type). satellite service, including multicasting and direct to home service in the northern hemisphere. From the quasi stationary active arc the coverage areas in North quasistationary America, Europe and North Asia can be served from high elevation angles. Primary broadband channels of 155 Mbps can be grouped to virtually channel up to 3,8 Gbps. The satellites will point to the center of the coverage areas with yaw steering. Station and orbit keeping maneuvers will be performed in the non active part of the orbit.

Satellite communications service providers are demanding on-board autonomy to reduce operational costs, to minimize outages of communication channels, and to increase overall reliability of the network. Autonomous procedures for failure identification, reconfiguration and recovery are necessary to meet the above demands. The autonomy requirements are different for single satellites and for LEO constellations. constellations.

2.2 Requirements

Due to the large number of satellites to be controlled and the short visibility over the ground station all functions for keeping, health attitude control, station keeping, status, etc. shall be housekeeping, configuration management, etc. addition, for performed without ground involvement. In addition, (e.g. Iridium or Teledesic) the orbit some constellations (e.g. parameters of the constellation shall be controlled as well as the satellite interlinks.

The requirements to the satellite control systems can be divided into the categories performance and autonomy. The performance criteria for orbit control are optimal use of propellant to achieve the desired orbit and maintain station. accuracy of orbital elements, while on station. GEOs are injected from transfer orbit into circular orbit by quasi-impulsive apogee boost maneuver (circularization and inclination).On orbit, station keeping has to be performed to maintain latitude and longitude with a typical accuracy of ±O.1 and ±O.05 deg respectively. Co-positioning of three ±D.1 and more satellites in "one box" requires coordinated relative station keeping to avoid radio signal interference or even collision.

3. CONTROL SYSTEMS, CHALLENGES AND SOLUTIONS 3.1 3.1 GEO Satellites

The improvement of AOeS performance for AOCS geosynchronous communications satellites is illustrated in Fig.3.1 Fig.3.l for the representative example of the Spacebus AOeS. AOCS.

During launch, LEOs and HEOs are first injected into intermediate parking orbits. Transfer to final orbits is performed by a number of low thrust burns, of several hours duration, using chemical or electrical thrusters. In addition to altitude, eccentricity and inclination, the nodes of the orbits have to be adjusted and maintained by shifting of the

Key features of the AOCS 3000B of e.g. EUTELSAT ill III are: are:

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In the following paragraph, problem areas in the control system design of big GEO satellites are described using the representative example of the Spacebus 3000B platform.

Fig. 3.1. Performance of the Spacebus AOCS •

Attitude control by momentum wheels in all three axis in Normal Mode Advantage: Extension of satellite life time by three years

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Station keeping in Normal Mode under Normal Mode attitude control Advantages: Double failure tolerance with 14 thrusters, easy adaptation to electrical thrusters

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Momentum Management during station keeping Advantage: Extension of satellite life time

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On-board Failure Management Advantage: Autonomous operation up to two weeks Advantage: Double failure tolerance for gyros, sun sensors and thrusters Survival of Earth sensor double failure

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Fig. The block diagram of the AOCS 3000B is shown in Fig. 4.1, the autonomous failure management is described in chapter 4.1.

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The following list gives an overview of the AOCS 3000 components. The wheel and sensor arrangement is al., presented in Fig. 3.2. For more details see (Surauer, et aI., 1993). Sensors • Two-axis Earth sensors (two) xlz -plane • Two-axis Sun sensors, distributed in the SIC X/Z (6 heads) • Three-axis gyro units (two)

Flexible Structures The critical parts are the large flexible solar generators and antenna reflectors, which lead to elastic modes within the control system bandwidth. The moments of inertia (MOl) of solar generators about the SIC center of mass is typically 3 to 4 times the MOl of the SIC body around roll and yaw axes for this type of satellite. The elastic modes are very lightly damped. Characteristic values for elastic eigenfrequencies (fixed base) are: Normal bending:

146

first mode second mode

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0.20-0.30 Hz

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•

The free-free natural frequencies increase by a factor of about 2 to 2.5 for the first normal and in-plane bending modes, they are only slightly higher for the second modes. normal bending and torsional modes.

De-orbiting Every SIC has to be transferred to graveyard orbits with 99.3 % probability of success at the end of life (12 to 15 years) or in case of SIC failure.

In section 4, methods and procedures to improve on-board autonomy in particular for GEO satellites will be discussed. 3.2 LEO Constellations

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Low acceleration propellant sloshing At launch, about 50 to 60% of total satellite mass are made up by propellants. After the final orbit position has been reached, the remaining fuel for station keeping is still considerable (around 30% of total satellite mass). Depending on the tank structure the dynamics of the liquids, the fuel sloshing, exhibits nearly undamped oscillations. The lower modes of the highly non-linear fluid dynamic equations are approximated by springlmass or pendulum systems. The sloshing masses spring/mass of these models may reach up to 900 kg for apogee boost and up to 600 kg for station keeping maneuvers. The eigenvalues change with the level of propellants remaining in the tanks and increase with the SIC acceleration, which also determines the orientation of bums. During transients, the propellant during thruster burns. sloshing amplitudes as high as 0.5 m may be encountered during station keeping. The first propellant 0.1 Hz for apogee boost and eigenvalues are around 0.1 about 0.02 Hz for station keeping maneuvers.

Fig.3.3 illustrates the Globalstar constellation, a typical LEO constellation for communication purposes. Table 3.1 gives the corresponding technical data.

Propellant sloshing and flexible modes require robust and advanced non-linear control system design to avoid instabilities and limit cycles and to maintain high attitude pointing accuracy even under severe sensor noise and transient conditions, e.g. during station keeping. •

•

Fig. 3.3. Illustration of the Globalstar system.

Pointing Accuracy Attitude pointing accuracy has to be maintained under severe sensor noise conditions during on-orbit normal modes. Critical are transients at beginning and end of station keeping maneuvers and yaw performance in the SunlEarth co-linearity region. region. SunJEarth

Table 3.1 Technical data of the Globalstar constellation. No of orbital planes No. of satellites per plane Orbital altitude Orbital period Orbit inclination

Co-positioning III The increasing demand for transponder capacity In distinct orbital positions has led to co-positioning of three and more satellites in one orbital box of <± 0.1 deg in North/South and <± 0.05 deg in EastlWest directions. The box size of 70 by 35 km demands, that relative position of satellites is controlled to avoid collision. To do this with minimum ground intervention, requires onboard autonomous orbit position determination and control.

8 6 1414 km 113 min 52 deg

The attitude accuracy requirements (3cr) of the Globalstar 3.2. satellites in the roll and pitch axis is shown in Table 3.2. The required attitude accuracy in the yaw axis is 1.2 deg, which includes also sensor errors, when Sun sensors or magnetometer magneto meter are used as yaw sensors. sensors. The following list gives an overview of the control system hardware. More detailed information is given in (Alexander, et al., 1997).

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Table 3.2 Attitude accuracy requirements.

Control error [deg] Sensor error [deg] [de~]

Roll 0.2

Pitch 0.2

0.25

0.25

Sensors 2-axis Earth sensor • In the spacecraft x/z•• 2-axis Sun sensors, distributed In plane • Magnetometer • GPS receiver for navigation purposes, time reference, and for backup attitude sensing.

Fig. 3.4. Schematic of GIobalstar Globalstar yaw-steering.

When yaw-steering is applied, the satellite body has to periodically perform large-angle yaw slews, according to a reference yaw profile that depends on the Earth-satelliteSun configuration (see Fig.3.4). This is done with internal control torques generated by 4 redundant wheels. The control task becomes difficult, when the 3-axis attitude tracking has to be performed only with 2-axis attitude measurements. Solutions to this problem are described in measurements. al.,, 1996; Briiderle, et aI., al., 1996). (Fichter, et aI.

Actuators 4 momentum wheels • 2 magnetic torquers for total momentum control • • 5 mono-propellant thrusters, IN, for orbit correction and acquisition modes The on-board processing electronics (OBPE) is based on a I750A CPU, 64kRAM, 128kPROM. The OBPE is fully 1750A redundant.

Minimum Equipment Configuration. In order to reduce the overall control system costs of a satellite constellation, it is important to minimize the equipment and recurring costs. The Globalstar satellites are equipped with Earth sensors, Sun sensors, and magnetometers, no gyros are implemented. All attitude control modes, including acquisition modes, operate completely gyroless. gyro less. Appropriate estimation and control algorithms were developed to compensate the lack of direct rate measurements. They are based on directional measurements of the Sun sensor and its derivatives. The principles of these algorithms are described in (Surauer, et aI., 1996). al.,

Usin oao Globalstar as a representative example, the following . paragraphs para;aphs describe two distinct features of constellatIon cons.tellation satellites, their impact on the control system desIgn, design, and future challenges.

Available Power. The available power on-board of each satellite is a very important parameter, because it limits the communication capacity. In order to maximize the available on-board power, the solar arrays have to be oriented perpendicular to the sunline, i.e. two mechanical degrees of freedom are required for the solar array orientation. This can either be realized by a 2-axis solar array drive, i.e. the rotated,, but also tilted. An solar arrays are not only rotated alternative approach is to use a conventional solar array drive with one degree of freedom (rotation), in conjunction with the yaw motion of the complete satellite as a second degree of freedom. This is implemented in the Globalstar satellites and is commonly known as yaw-steering . Both methods described above have severe implications on the attitude control system design: In the case of solar array tilting, large products of inertia result in large gravity gradient disturbance torques, which are difficult to reject when magnetic torquers, i.e. restricted, time-varying torque capability, are used as attitude control actuators (in addition to one momentum wheel in the orbit normal axis). axis). Control approaches and optimization techniques to this control problem are described in (Bals, et al., aI., 1996).

Future Challenges The attitude control system of current constellation satellites operates fairly autonomously. However, there is a necessity for further development and a further potential for cost reduction in the following areas:

• ••

•• ••

148

Implementation of autonomous position control and constellation keeping. Low cost spaceborne GPS receiver for navigation purposes. Exploitation of intersatellite links for control purposes. Implementation of functional redundancies for reliable de-orbiting.

3.3 HEO Satellites Ground Control

&

Payload

Fig. 3.5 shows the orbits of the PENTRIAD HEO satellite apogee/perigee: 43,000/3,000 system (inclination: 64.3 deg, apogee/perigee: km). To avoid two degrees of freedom solar generators, yaw steering, similar to Globalstar will be performed to point solar panels into the direction of the Sun. Different from the nadir pointing of the GEO and LEO systems described before, HEO satellites have to perform accurate Center of Coverage Area (CCA) pointing to distinct centers during passage of their active arc. To calculate the pointing bias from Sun and Earth reference, accurate knowledge of satellite orbit and position is needed on-board. Pointing to coverage area centers requires, that pitch and roll maneuvers are performed in addition to the yaw steering. steering.

---

m· e=.. e=. s.. .. SUI..

Actuator Equipl'Mnt

.

.. ~::,~~~i Fig. Fig. 3.6. On-board AOCS Software Modularity.

3.4 Trends in Implementation Performance, quality and cost are design drivers for future communications satellites. satellites. To achieve this requires simplicity of design, multiple (re)-use of hardware and software and potential for growth and change by modularity. The hierarchical, modular architecture underlying AOCS software design at DSS is illustrated in Fig.3.6. Fig.3 .6. Hierarchy and modularity of software design simplify the handling and adaptation of the large number of AOCS modes (typically more than 20).

Sensors

Fig. 3.5.

PENTRIAD (9 satellites in 3 planes, 5 active arcs): Observer on Earth fixed system.

The control system hardware is similar to pointing accuracy on the active arc is requirements, while the relaxed pointing non-active arc is determined by the maneuvers.

Actuators. HK&Control

serial Serial links

TMlTC

Dala Bus

PIl

P/l

T'-A

Oata

El ~ ,~~4~=4~~~

Perigee Passage

Globalstar. CCA similar to GEO accuracy on the station keeping

Senal Link Star conhgllallon

AODS RM AM SGM

= Att~ude Orbit Determination Determination Anitude and Ortlit =Reconfiguration Module :: = Safeguard Memory

Fig. 3.7. Central processing node - architecture. For this type of orbit, disturbance torque magnitudes vary significantly, depending on orbital altitude and SIC orientation. This effect can be used to optimize momentum management and momentum/reaction wheel sizing.

Concerning hardware and on-board processing, similar considerations led to the decomposition in physical and functional entities of Fig. 3.7. By consequently optimizing the distribution of functions and interfaces of the central processing node, the costs, mass, and power can be further reduced, while computing performance is increased to accommodate the new functions, which are required for the AOCS projects in planning.

The attitude measurement system has to be adapted to cover the range from perigee to apogee altitude. altitude. Experience gained with the three axis stabilized transfer of GEOs can be used in the design of this system.

149

4. ON-BOARD AUTONOMY

concept of (usually dual) equipment redundancy, functional redundancy is used, e.g. for back-up modes.

Autonomy of Earth oriented communications and navigation satellites is mainly motivated by operational and economical reasons. The activities are targeted towards increasing operational availability and safety and reducing operational cost.

HPC1

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Operational safety is increased by an improved failure detection, isolation, and recovery (FDIR) system, implemented in the on-board computer. The FDIR system of Spacebus3000B is explained in more detail in 4.1.

Sun Senior

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Reduction of nominal operational cost can be achieved by autonomous or semi-autonomous orbit determination and orbit control (station keeping). This is true not only for geostationary satellites, but also in particular for constellation satellites in low Earth orbits. In case of geostationary satellites, autonomous orbit determination has a larger impact on the overall control system design, because it requires additional hardware or at least hardware modifications. When orbit determination is implemented, autonomous orbit control can then be realized with existing on-board equipment and software adaptation only.

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Fig.4.l. Schematic of Spacebus 3000B AOCS Fig.4.1 shows an example of AOCS hardware for present Communication Satellites. Safeguard Memory (SGM) with failure history register (spacecraft and equipment), Reconfiguration Module (RM) with processor watchdog and status relays, and independent high priority control lines (HPC1I2) are the hardware on which the FDIR is based, independently from the AOCS On-Board Computer Unit (OBCU).

For constellation satellites in low Earth orbits the use of GPS receivers for navigation purposes is state-of-the-art. In addition, GPS receivers serve as a time reference. In the geostationary orbit, GPS receivers or even GPS/GLONASS receivers can also be used for orbit determination. The next section briefly describes the corresponding navigation techniques. Alternatively, a second approach for (semi) autonomous orbit determination is described for geostationary applications, based on the existing sensor configuration of Spacebus 3000B. This leads to a limited autonomy period. The advantage: Only equipment, which is already available on board is used.

4.1

c---_

The FDIR S/W functions are decentralized to the maximum possible extend and hierarchically structured in four levels:

Level Zero failures. Identification and repair handled locally on equipment level (BITE), no interruption of Normal Mode operation, status message to ground.

Autonomous Failure Management

Level One failures. Detection of equipment failures by software criteria in OBCU and recovery by switching to redundant unit.

For Geosynchronous Communication Satellites, constraints on outage time are very severe. Typically, less than a total of 60 minutes accumulated outage time over an operational life of 15 years is permitted under the threat of penalty. In addition, the system has to operate without access to the ground for periods of at least 48 hours at any time. To meet these requirements failure identification and repair have to be performed on board with minimal interruption of communication links.

Level Two failures. Watchdog monitoring of OBCU and initiation of reconfiguration sequences by RM in three steps: Step 1: Warmstart (Software Reset) Step 2: Coldstart (Power Supply OFF/ON) Step 3: Reconfiguration to predefined new configuration

The FDIR System of the Spacebus 3000B is based on the hierarchical classification of failures, which facilitates identification and allows for repair with minimum intervention of AOCS functions. In addition to the classical

Level Three failures. Hardware monitoring of critical equipment, such as propulsion or power systems and reconfiguration of AOCS by pre-established sequences

150

stored instored the RM. ToRM. support the failure on ground, about 26500 km which corresponds to an orbital periodperiod of in the To support the analysis failure analysis on ground, about 26500 km which corresponds to an orbital of about half a day. geostationary orbit (semi-major the failure storedis in the in the the occurrence and the and status about half aSince day. the Since the geostationary orbit (semi-major the occurrence theofstatus of the is failure stored axis 42200 km) is km) above GPS GPS are are failure history which iswhich part ofis the axis 42200 is the above theorbit, GPS and orbit, andsignals GPS signals failure register, history register, partSGM. of the SGM. radiatedradiated towardstowards the Earth, only GPS signals of satellites the Earth, only GPS signals of satellites beyondbeyond the Earth be can received. This situation is shown in all of above fail is fail the issystem Only, ifOnly, be received. This situation is shown in thecan Earth if the all of the measures above measures the system Fig.4.3. switchedswitched to the ultimate sun-pointing Safe Mode to secure FigA.3. to the ultimate sun-pointing Safe Mode to secure of spacecraft and provide time fortime analysis and survival survival of spacecraft and provide for analysis and visibility: GEO-GPS satellites visibility: GEO-GPS sat.llites 5 5(0): 65.1 %5(1): ".5(2): ".5(3): ".5(4): ". , repair. To restart Normal Mode operation, the satellite has , 1.2 0"45(4) 5(0): 65.123.6% 5(1 ), 23.6 10.1 %5(2), 10.1 1.2"45(3) 5 0 "4 repair. To restart Normal Mode operation, the satellite has to go through the full acquisition sequence, which may take to go through the full acquisition sequence, which may take up several in the worst up minutes several minutes in thecase. worst case. 4

By applying this minimum intervention strategystrategy for for By applying this minimum intervention handlinghandling satellite satellite anomalies, most failures can be can repaired anomalies, most failures be repaired without without losing pointing and therefore withoutwithout outage outage and losing pointing and therefore and interruption of service. interruption of service.

3

2

4.2.

Orbit Determination of Geostationary Satellites Orbit Determination of Geostationary Satellites 4.2. Using GPS Signals Using GPS Signals

1

ii II

1

2

3

4 1 5 2 6 3 7 4 8 5 9 6 107 118 129 13 10 14 11 15 12 16 13 17 14 18 15 19 16 20 17 21 18 22 19 23 20 24 21 22 23 24 time(hours)

time(hours)

Fig.4.4.FigAA. Visibility, longitude 70 deg,70receiver antenna gain gain Visibility, longitude deg, receiver antenna 10dB 10 dB The geometrical configuration of the of GPS-satellites, GEO- GEOThe geometrical configuration the GPS-satellites, satellite,satellite, and theand Earth veryto poor the leads Earth to leads very GPS poor visibility GPS visibility conditions, seen from geostationary orbit. Fig.4.4 showsshows conditions, seenthe from the geostationary orbit. FigAA the number of visible GPS satellites over over one orbit the number of visible GPS satellites one orbit revolution, based on the following assumptions: on the following assumptions: revolution, based • • • conditions. The satellite GPS satellite constellation VisibilityVisibility conditions. The GPS constellation nominally of 24 satellites. The semi-major of 24 satellites. The semi-major axis is axis is nominally consists consists

GPS constellation Fig.4.2. FigA.2. GPS constellation

of IGPS OdB;satellite GPS satellite half beam • Receiver 10dB; half beam Receiver antennaantenna gain ofgain to 21deg, GPS satellite antenna width width up to 21updeg, GPS satellite antenna pattern.pattern. • threshold CIN threshold of 37dBHz for signal acquisition. CIN of 37dBHz for signal acquisition. in order • Artificially increased Earth radius Artificially increased Earth radius (1000 (1000 km), inkm), order avoid tropospheric and ionospheric to avoidtotropospheric and ionospheric effects.effects. Geometric vilibility conditiona GEO - GPS-..r.llites 10000r----...,.--~--~-__,_-____,--_r__-__r_I

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Visibility from GEO Fig.4.5.FigA.5. Visibility from GEO

Geometry GPS - GEO Fig.4.3. FigA.3. Geometry GPS - GEO

151

151

Fig.4.5 shows the accumulated time interval, interval, when more Fig.4.5 simultaneously, as aa than 22 GPS satellites are visible simultaneously, function of the geostationary longitude. It can be seen, that the visibility conditions vary significantly. Even with a constant satellite longitude over a full GEO satellite lifetime, the worst case visibility conditions will be encountered, because of the inertial drift of the GPS constellation.

the state of the underlying dynamics. This can be expressed as

lei) +[J[J +1(i)

=

,;p(i) = C(i)cI>(i,O)x(O) C(i)(i,O)i(O) + !¥J(i)

c· b(i) +

(4)

where the measurement matrix is given by From the visibility plots it is clear that there are less than 4 i.e. a direct kinematic satellites available at one time instant, i.e. possible. Therefore Kalman navigation solution is not possible. filtering has to be applied. Most publications in the past proposed Kalman filtering based on a system model that includes not only the orbit dynamics, but also the receiver clock dynamics. This approach leads to the conclusion, that the receiver clock precision determines the accuracy of the navigation solution. With other words: words: a relatively precise receiver clock is required for reasonable navigation accuracy. However, it would be more favorable to achieve accuracy. good navigation results without stringent requirements on the receiver clock.

[~~ ~ ~ ~j m1T

III]

C(i) = = : [

m~

(5)

0 0 0

~p(i) Sp(i)

is a (nxl) vector, when n pseudoranges are measured time. For nn;::: ~ 2 (practically n=2 or n=3 in the GEO at a time. orbit) the clock bias b can be eliminated by forming differences of the pseudorange measurements. measurements. This can formally expressed by a multiplication mUltiplication of eq. (4) with T from the left, where

T -_[~o.

o

1 0

(A verin, et al., 1996), an alternative approach is In (Averin, described based on single differences of pseudorange measurements. It eliminates the receiver clock bias and requires only the modeling of the well-known orbit dynamics. This leads to relatively accurate navigation results without stringent requirements on the receiver clock.

T=

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0

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-1

, dim [(n-l) [en-I) x n]. nJ . (6)

-1

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Method. The pseudorange measurement Single Difference Method. between the geostationary satellite and the GPS satellite is following: given by the following:

o(i) t5(i)

T(i )t,.p(i) =T(i)C(i)(i,O)x(O) T(i)C(i)(i,O)x(O) + T(i)!1j(i) = T(i)DP(i) T(i)~l (i) .

(7)

(1)

The Kalman filter is fed by the pre-processed (differenced) measurements (ti) /Xi)..

A linearization with respect to a reference value (best estimation) yields

Performance Analysis. The following natural disturbance forces are taken into account for the performance analysis:

!1p;

= p;- p ; = ~;

• • • •

(2)

where

Solar pressure Earth oblateness Triaxiality of the Earth Gravitation of Sun and Moon

(3)

The uncertainty of the solar pressure is assumed to be inertially constant with a magnitude of 10% of its nominal value, all other disturbance uncertainties are modeled as (1(j) of its nominal random noise, with a magnitude of 5% (la) value.

is the unit vector (line-of-sight direction) from the GEO satellite to the GPS satellite. In order to apply Kalman filtering, the pseudorange measurement has to be related to

The measurement noise is modeled as a second order Gauss-Markov process with a time constant of 120 s and a standard deviation of 23 m, representing selective availability (SA) effects. Receiver noise and other

152

measurement errors are modeled as white noise. All simulation runs are based on worst case visibility conditions, i.e. not more than 2 GPS satellites are visible at a time.

Table 4.2 shows the navigation accuracy where GPS signals are used for navigation. Results are given for each axis (orbit normal, along-track, radial), and for different disturbance forces. The maximum error occurs in Xdirection, which is caused by SA and solar pressure. The position deviations in [%] are computed with respect to the reference case.

Fig.4.6 shows the time histories of the navigation errors in the reference case, where all disturbance forces are assumed to be known exactly, i.e. only the measurement noise (SA) degrades the navigation accuracy.

Table 4.2 Results of the performance analysis. Table 4.1 Statistics of the reference case. Reference case

Orbit normal Y Along track X Radial Z

Standard deviation [m]

Mean value [m] 0.1 -36.8 -7.4

51.1 182.1 99.8

Position error

Max. Deviation [m] 152.0 878.6 505.9

Reference case +5% (random) Earth oblateness +5% (random) -4.6 145 Earth triaxiality + 10% (constant) 12.5 171 +10% solar pressure + 5% (random) Sun and 6.6 162 grave Moon SEav.

Table 4.1 summarizes the corresponding statistics for the reference case. A deviation of 1000 m in X-direction 10.33 corresponds to an angular EastIW est deviation of 1.5 10EastIWest deg in the geostationary orbit. Note, that a common station keeping window has a width of ± 0.05 deg in EfW EIW direction, which corresponds approximately to ± 32 km.

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Along Radial Z track X [%] [m] [%] [m] 879 506 0.5 883 0 506

In Table 4.3 the performance of GPS navigation is compared with the navigation based on GLONASS satellites only, as well as the combination of both (GNSS). Only position errors in the X-axis (along track, worst case axis) are shown shown..

y-axis estimation error (,.lerenoe)

1000

Orbit normal Y [%] [m] 152 1.3 154

Table 4.3 Position error in X - direction. direction .

1500000

time[s] 1irnlI[s)

GPS Position error X [%] [m] 879 Reference case 0.5 883 +5% Earth oblateness 879 0 +5% Earth triaxiality ty triaxiali 4 913 +10% solar pressure +5% Sun and 0.7 885 grave Moon SEav.

ll-lIXis estimation error (,.lerenoe)

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153

GLONASS Position error X [%] [m] 787 0.4 790

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4.3.

Orbital Motion. Typically, deviations of a geostationary satellite from its nominal position must not exceed ±O.05° in EastlWest EastIWest and ±O.lo ±O.l ° in North/South direction. Because of these small deviations, the orbital motion can be linearized with good precision with respect to the ideal geostationary orbit, orbit, represented by the reference coordinate system (x (x"n Yn y" Zr) in Fig.4.8. Figo4.8.

Semi-Autonomous Orbit Determination Based on Earth and Sun Sensor Measurements

An alternative approach to orbit determination is to exploit 2-axis Earth and Sun sensor measurements. These sensors are already part of the attitude determination system on many geostationary satellites. The implementation of the orbit determination is therefore less expensive and the system complexity is not increased. As a disadvantage, the sensor measurements do not contain enough information to fully autonomously determine the orbit. Besides that, satellite flight data (telemetry) show, that the sensor configuration is sensitive to thermal deformation of the satellite structure.

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A semi-autonomous orbit determination will be presented, exhibiting the following key features: features:

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• Propagation of the satellite North/South motion based on exact mathematical models. EastIWest motion based on • Estimation of the satellite EastIWest Earth and Sun sensor measurements. • Calibration of the Earth/Sun sensor configuration regarding thermal deformation of the satellite structure and sensor misalignment.

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Figo4.8. Orbit reference, attitude reference and body Fig.4.8. frame (subscript r, r', b). The resulting equations of relative motion are well known as Hill's equation. For a representation with respect to small y/ageo deviations A = x/ageo (longitudinal deviation), ~ = y/'ageo (latitudinal deviation), r = z/ageo (radial deviation), the following set of equations is valid:

Sensor Configuration. The sensor configuration consists of a 2-axis Earth sensor and a 2-axis Sun sensor. The Earth sensor is aligned along the satellite Earth pointing +zb-axis and furnishes information on roll and pitch attitude (4),8) (<1>,9) of the satellite. The Sun sensor heads (SSH) are distributed in Figo4.7) such that the direction the satellite xJzb-plane (see Fig.4.7) of the Sun can always be measured in the form of a unit vector Sb. Sb'

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---

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Measurement Principle. Principle. To outline the measurement principle, a satellite body fixed (and therefore attitude dependent) system (Xb, Yb, Zb) with its origin in the center of mass of the satellite is defined. defined.

" " -"-

.....

Fig.4.7. Figo4.7. Typical distribution of Earth sensor (IRES) and Sun sensor heads (SSH) in the xJzb-plane of a geostationary satellite (SIRIUS2).

154

4A Dominant perturbing forces and their effect on the Table 4.4 orbital motion of geostationary satellites.

The effect of longitudinal and latitudinal deviations of the satellite from its nominal position on the measurement FigA.9 and Fig.4.1 FigA.l 0, respectively. quantity is presented in Fig.4.9 Summing up, measurements of Earth and Sun sensors can be processed to a scalar quantity featuring:

Effect on GEO

Source Earth potential: Zonal effects Tesseral effects Luni-solar gravitation Solar pressure Station keeping maneuvers

EfW oscillation, drift EfW drift rate N/S oscillation EfW oscillation, drift EfW and N/S accelerations

• • •

FigA.l1 shows the Orbit Determination Principle. Fig.4.11 combined influence of orbit deviations on the angle between Earth and Sun directions. There is some ambiguity in the effects of longitudinal and latitudinal deviations on the scalar measurement, i.e. a deviation angle A may have the same effect on yet) as latitudinal deviations.

The measured Sun vector Sb is related to its calculated reference Sr by a transformation about small orbit and s. is known bv ephemeris data. attitude deviation angles, Sr

Yo

y

Fig.4.11.Combined FigA.ll.Combined influence of orbit deviation angles on the angle between Sun and Earth direction

FigA.9. Influence of longitudinal deviations on the angle Fig.4.9. between Sun and Earth direction.

This can also be shown analytically: The EastfWest (A) and the North/South (~) motion of the satellite are not fully observable with measurements yet), see (Fichter and luckenhbfel, 1998). Juckenhofel,

The resulting measurement quantity yet) corresponds to the difference of the nominal (Yo) and the actual angle (y) between the Sun and Earth directions:

Measurements of Earth and Sun sensors can only be used to estimate either one of the two subsystems: the North/South motion or the EastfWest motion. The solution to the subsystem that is not supported by the measurements has to be made available by other means, e.g. by propagation

(9)

,e~S"

For practical autonomous orbit determination purposes it is favorable to estimate the EastfWest motion (using the Earth and Sun sensor data) rather than the North/South motion because the former contains integral behavior (drift) and uncertain disturbances (solar pressure) which would lead to relatively fast divergence with pure propagation. In contrast the North/South motion can be accurately propagated over long periods of time.

y

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155

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As an advantage of the Kalman filter, the estimator can be i.e. measurements are taken run in ,Redundancy Mode', i.e. with a single Sun sensor head in a limited section of the orbit only. The accuracy of the estimation in ,Redundancy Mode' compared to ,Normal Mode' (360 deg Sun sensor significantly. measurements) is not decreased significantly.

I

Fig.4.13 compares SIRIUS2 SIRlUS2 latitudinal ground station Fig.4.l3 results. The tracking data (triangles) to propagation results. propagator is initialized with tracking data, no station performed. The propagation error is keeping maneuver is performed. significantly smaller than one thousands of a degree after more than a week of propagation. Its magnitude is in the accuracy. range of tracking accuracy.

Fig.4.14 represents estimation results in ,Redundancy Mode '. During the propagation sequence (240 deg) the Mode'. AII ::; 0.004 deg. error increases to IA FigA.15 shows the filter response to propagation errors. Fig.4.l5 North/South maneuver modeling errors (six months station keeping) are assumed to be of identical orientation (very conservative) and represented as error in the initial condition of the North/South propagation (Ll~ (~~ ::; 0.01 deg). Note that the resulting estimation error vanishes for Sry = O.

Errors in the propagation of the North/South motion contribute to the measurements Yet) yet) and therefore directly East/West motion. motion. effect the observation of the EastIWest

System B: B: EastlWest EastIWest Filter. The estimation is based on a discrete Kalman filter algorithm (Hill's equations supported by Earth and Sun sensor measurements). The estimation accuracy is limited (worst case) by

All simulation results are based on a ,cold start' operation (zero initial filter condition) and calibrated sensors.

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The magnitude of k(t) is too big to neglect these sensor errors in the course of autonomous orbit determination. Calibration of the sensor configuration, i.e. determination of the relevant k(t) is the key to successful orbit determination. determination.

Fig.4.15. FigA.1S. Long term influence of N/S propagation errors ~~:: latitudinal errors (top) related to resulting ll~ longitudinal estimation error and Sun vector ycomponent Sry. Sry .

157

At Daimler-Benz Aerospace AG a long term investigation is currently conducted to analyze flight data and to identify a suitable model for the error functions .

4.5 shows an error budget on this assumption. The resulting orbit determination accuracy of ±O.02 deg in satellite longitude is believed to be sufficient to control the satellite in the limits of ±O.05 deg position tolerance.

Additionally, modifications to future bus design are suited to reduce the influences of thermal deformation: mounting the Sun sensor head that is used for orbit determination on a common bracket with the Earth sensor reduces the relative motion of the sensors, i.e. the magnitude of k(t).

5. STANDARD AOCS PROCESS 5.1 Introduction

The growing demand and worldwide competItlOn in communications satellites leads for the first time to the necessity to streamline the product and manufacturing process for series production. To reduce costs and delivery time while maintaining high standards in quality and performance requires re-engineering of the process by which the complex AOCS subsystems are designed, developed, and manufactured.

Conclusions. A simple approach to orbit determination is based exclusively on Earth and Sun sensor measurements. These sensors are already part of the attitude determination system of many geostationary satellites. Therefore, introduction of autonomous orbit determination neither increases system complexity nor hardware expenses.

As a disadvantage, the available measurements do not contain sufficient orbit information to estimate North/South and EastIWest motion. A reasonable solution to this observability problem is the propagation of the stable and well modeled North/South motion. A filter can then be designed to estimate the EastlWest motion, especially the satellite drift.

In this chapter, as an example, the approach developed at Daimler-Benz Aerospace to fulfill all of the above needs is described. The main steps of the re-engineering process are:

Flight data analysis has shown, that time varying sensor misalignment reduces the accuracy of the estimation in an unacceptable way. As this is not a sensor but a satellite specific problem, this holds true for the whole class of direction measuring sensors. Orbit determination with direction measuring devices therefore reduces to a calibration problem. A solution to the problem for the presented Earth and Sun sensor configuration can therefore easily be adopted for the application of star sensors for example.

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streamlining the organization, by bringing all AOCS related activities together in a product center,

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introducing standardized tools and standardized Development, Verification and Test Facilities,

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developing a modular on-board software and hardware architecture,

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centralizing procurement components and equipment.

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Fig.5.1 shows the key elements of the improved AOCS process, see (Lange, et aI., 1997).

Table 4.5 Error budget: Satellite longitude Constant

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Fig.5.l. Key elements of improved AOCS process

158

In the following the interest is focus sed on the technical aspects of standardization of the development and test facilities.

loop. For performance tests of the first flight model of both GEO satellites and constellation satellites, the full set of tests is performed at the DBT facility, by using the TMlTC interface as realized in the ground station, to verify the AOCS operation procedures.

5.2 StandardizedJacilities

The three main facilities shown in Fig.S.2, see (Lange, et al., 1997), are •

Software Development Facility (SDF),

•

Software Verification Facility (SVF) and

•

Dynamic Bench Test Facility (DBT). Software Development Facility

Software Verification Facility

For the following flight models of the GEO satellites a subset of these tests is selected for representative performance tests on the DBT facility. For serial production of Constellation Satellites, parallel hardware in the loop test benches are used for functional tests to comply with the high delivery rates. Fig.5.3 shows the DBT facility and Table 5.1 the DBT specific data.

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Following the design, the control algorithms are transferred to the Software Development Facility, where they are directly coded to Flight Software in ADA by the control engineers. This allows the functional tests of AOCS algorithms in the closed loop and simultaneously the software at a very early stage of the process. In addition, sensitivity analysis and module tests are performed. For easy use the simulation environment is built with Matrix/M in a modular architecture, with defined interfaces to the SIC-model and the equipment models (written in C). Adaptations, extensions or an exchange of single models can be realized easily.

4 Fig.5.3. Dynamic Bench Test Facility

Table 5.1 Dynamic Bench Test Facility data

• • •

In the Software Verification Facility closed loop tests under real time conditions using the on-board computer, the data handling software, the operating system and the AOCS software from the SDF are performed.

•

In the third step performance and acceptance tests are done for qualification with not only the computer but also sensors and actuators (as much as possible) in the closed control

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Accuracy Min. velocity Max velocity Outer axes Inner axes 2 Sun simulators 2 Earth simulators

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Workshop on Spacecraft Attitude and Orbit Control, Control, ESTEC. Noordwijk, The Netherlands. ESTEC Averin, S., V. Vinogradov, N. Ivanov and V. Salischev (1996). Application of Differential Method for Relative Positioning of Geostationary Satellites with Use of GLONASS and GPS Navigation signals. Proceedings of the 5thth International Conference on Differential Navigation DSNS-96, Volume 2. Bals, J., W. Fichter and M. Surauer (1996). Optimization of Magnetic Attitude and Angular Momentum Control for Low Earth Orbit Satellites. Proceedings of 3. 3. International Conference on Spacecraft Guidance, Guidance, Navigation, and Control Systems, ESTEC ESTEC. Noordwijk, Noordwijk, The Netherlands Briiderle, E., W. Fichter, B. Lange, N. Furumoto and J. Rodden (1996). Dynamic Momentum Bias for Yaw Steering. Proceedings of 13. /3. IFAC World Congress. San Francisco. Eichhorn, E., G. Farnetani, H.D. Fischer and J. Seidl (1997). The Development, Test and Validation of the ADCS for the Spacebus Family. Proceedings of ESAlIFAC International Workshop on Spacecraft ESTEC. Noordwijk, The Attitude and Orbit Control, ESTEC Netherlands. Netherlands. Fichter, W., M. Surauer and P. Zentgraf (1996). Control Design for Generalized Normal Mode Operation of Bias Momentum Satellites. Control Engineering Practice, Volume 4, Number 10. Juckenhofel (1998). Observability of a Fichter, W. and O. JuckenhOfel Geostationary Orbit with Earth and Sun Sensor Measurements. The Journal of Guidance, Control, and Dynamics. Submitted for publication. Lange, G., W.. Oesterlin, and H. H. Widmann G. , M. Mollenhoff, W (1997). Approach for a Standard AOCS Development Process. Proceedings of ESAlIFAC International Workshop on Spacecraft Attitude and Orbit Control, ESTEC. Noordwijk, The Netherlands. ESTEC Stinshoff, K., H.D. Fischer, M. Surauer and G. Lange (1997). The Role and Significance of Dynamic Bench Testing in AOCS Development and Verification. Proceedings of ESAlIFAC International Workshop on Spacecraft Attitude and Orbit Control, ESTEC ESTEC. Noordwijk, The Netherlands. H. Bittner, W. W. Fichter and H.D. Fischer Surauer, M., H. (1993). Advanced Attitude and Orbit Control Concepts for Three-Axis-Stabilized Communication and Application Satellites. IFAC Symposia Series. Series. Number 12. M., P. Zentgraf and W. Fichter (1996). Attitude Surauer, M., Control of Geostationary Satellites with Minimal Use of Gyroscopes. Gyroscopes. Proceedings of 3. International Conference on Spacecraft Guidance, Navigation, and Control ESTEC. Noordwijk, The Netherlands. Systems, ESTEC

In all three facilities the same qualified SIC-model and the same qualified equipment models are used. By this provision adaptation and qualification only have to be realized once and no differences in reference runs due to different models have to be explained. All three facilities are based on UNIX operating systems.

5.3 Conclusions By streamlining the AOCS process for all application satellites, cost and delivery time have been reduced considerably. By the same procedures, flexibility, performance and quality have been improved. New software and hardware modules can be added easily, and modifications and adaptations to the AOCS as well as to the test benches can be realized in an easy and controlled way. Modifications, in particular to the AOCS AOCS,, can be introduced and demonstrated systematically and quickly by use of the three test benches: This is a step towards linking the AOCS to the SIC rapid prototyping process.

6. SUMMARY Following a brief introduction into the development of the space communication market, current and future scenarios are described and corresponding requirements on the attitude and orbit control system are derived. State-of-the art attitude control systems are outlined, and future challenges for the enhancements of these systems are presented. A major part of future AOCS requirements is spacecraft autonomy, which can be roughly divided into two areas: areas: autonomous failure management and autonomous orbit determination and control. Both areas were described in detail. competItIon and the new situation of series Worldwide competItIOn production has not only an impact on the design and the product itself, but also on the development and production process. A suitable approach ("standard AOCS process") is outlined.

ACKNOLEDGEMENT The authors would like to thank Mr. Colin Rogers from Eutelsat, Paris for providing flight data of the Eutelsat2 FM2 satellite. REFERENCES Alexander, R., E. Briiderle, E. Groegor, W. Schrempp and H. Widmann (1997). Attitude and Orbit Control for Globalstar. Proceedings of ESAlIFAC International

160