Challenges in the control and autonomy of communications satellites

Control Engineering Practice 8 (2000) 409}427

Challenges in the control and autonomy of communications satellites E. Gottzein, W. Fichter*, A. Jablonski, O. JuckenhoK fel, M. Mittnacht, C. MuK ller, M. Surauer DaimlerChrysler Aerospace, Dornier Satellitensysteme GmbH, PO Box 80 11 69, 81663 Munich, Germany Received 13 November 1998; accepted 18 August 1999

Abstract The predicted rapid growth in space communications has stimulated many ideas beyond today's solutions. In addition to geosynchronous satellites, new low and intermediate orbit constellations of satellites and even constellations in eccentric orbits are being designed and planned. The paper gives an overview of the new and challenging requirements for satellite attitude and orbit control systems (AOCS), and presents solutions based on examples from current projects. New methods for autonomous failure management, autonomous orbit determination and station keeping are discussed.  2000 Elsevier Science ¸td. All rights reserved. Keywords: Attitude control; Orbit determination; Station keeping; GPS; On-board autonomy; Failure management; Recon"guration; Satellite constellations

1. Market development of space communications In 1945 the British scientist and author Arthur C. Clark suggested placing three satellites in an equatorial Earth orbit with a radius of 42,164.5 km, corresponding to an orbital period of one sidereal day (23 h 56 min). Stationed 1203 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 "rst geosynchronous communications satellite (GEO satellite) realizing Clarks idea was SYNCOM-3, which was launched August 19, 1964 (mass 37.5 kg, power 29 W). Like most of the early satellites, SYNCOM-3 was spin stabilized. Positioned above the Paci"c Ocean, it enabled European viewers to watch the Tokyo Olympic Games live. The idea of the geosynchronous orbital period was adapted to the speci"c geographic requirements of the former Soviet Union. The "rst MOLNIYA communications satellite, launched April 23, 1965, operated from a quasi-synchronous elliptical orbit with an altitude at perigee of 538 km and at apogee of 39,300 km, the inclination with respect to the equator was 65.53. One satellite of this series guaranteed communication for eight hours

* Corresponding author. Tel.: #49-89-607-26144; fax: #49-89-60723583. E-mail address: walter."[email protected] (W. Fichter).

and the three properly synchronized to each other could provide 24 h coverage for areas in the northern hemisphere. The "rst MOLNIYA was already three-axis stabilized with a power of 700 W. MOLNIYA satellites of the ORBITA network were used for the hotline between the White House and the Kremlin, which was established for security reasons in 1971. In the following period, international organizations were founded by groups of nations to procure and launch communications satellites and to provide tele-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 telecommunication 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. Fig. 1 shows the expected growth of the market for satellite services by sectors, Fig. 2 shows an overview of the market value of the satellites which are needed to meet these demands. At present, the revenue share of GEO systems in the commercial market is 90%. The GEO system market will continue to grow due to the expansion in direct broadcast systems and regional GEO mobile applications. It will grow further once the Ka-band intersatellite link (ISL) systems are completed.

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market demand shows a cyclical trend with peaks around 1998 and 2010. A speci"c 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 453 and 703 provide improved visibility for Ka-band links and minimize atmospheric damping. For the year 2003 the following market shares are predicted: E GEO-Satellites (28 GHz, 15 GHz, 6 GHz): 57%. E LEO-Satellites (28 GHz, 1.6 GHz): 43%. Fig. 1. Overall satellite industry growth by sector: Launches, manufacturing and operations.

Fig. 2. Satellite market value overview.

The Global SPACEWAY2+ Network scheduled to be operational between 2002 to 2005 proposes to station 4 regional geosynchronous systems over North America, the Paci"c Rim, Europe/Africa and Central/South America and to interconnect the satellites by Ka-band intersatellite link. Satellite constellations made up of many satellites in low Earth orbits (LEO satellites) are being prepared to meet the projected demand for high and low data rate satellite services, many of which will use 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 Kaband ISL of Iridium is 25 Mbps. The advanced HDR constellation Teledesic is targeted towards high speed data transmission between centers and uses Ka-band ISL with 1.5 Gbps. Fig. 2 shows the increase in satellite market value expected to ful"ll the demand for increasing payload capacity in the next few years. For GEO satellites, the

What are the messages of Figs. 1 and 2 to the satellite industry and in particular satellite designers? 1. The international market and customer situation as well as the amount of capital needed makes a global market and partnership strategy a must for space companies. This means international competitiveness of the product in terms of cost, quality and performance, but also an open attitude to cooperation and worksharing, in particular for the satellite designers. 2. For the "rst time in space history, commercial satellites are being produced in series. This requires reengineering of the product, the design and production procedures. 3. The 30% sales growth in the satellite procurement market, compared with the much bigger revenue generated by the satellite services market, challenges the satellite designers to improve operational availability as well as the autonomy and operational lifetime of their systems. The attitude and orbit control (AOCS) subsystem plays a key role in all three of the above challenges. E Operational lifetime can be increased by a prudent management of on-board resources, such as the propellant and the use of electrical propulsion systems. E Operational availability can be increased by minimizing the &outage' time, through an on-board autonomous failure detection, isolation, and recovery (FDIR) system. E Operational cost and reliability can be reduced by autonomous orbit determination and position keeping. On-board autonomy is increasingly important for space constellations using intersatellite links, because the ephemeris of the individual satellites have to be known on-board precisely, even though their visibility (time and occurrence) to ground stations is limited. The following sections deal with all three of the above AOCS related tasks, presenting solutions by examples from GEO, HEO and LEO projects.

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

2. Scenarios and requirements 2.1. Scenarios 2.1.1. GEO satellites Most of the satellites in geosynchronous orbit are three-axis stabilized communications satellites. The large power demand is provided by #at solar array panels oriented towards the Sun. Most satellites are injected into an elliptical transfer orbit (apogee at 36,000 km, perigee at &200 km, inclination depending on launch site between 53 and 283). In order to realize the large velocity increments needed for injection into the geosynchronous orbit and for station keeping, bi-propellant propulsion systems are generally used. Satellite lifetimes now extend beyond 15 yr, so electric propulsion for station keeping becomes very attractive. Orbit positions should be kept within a box of $0.13 in North/South and East/West direction, on collocated positions ($0.053. There are various types of communications satellite in geostationary orbit. The main applications are summarized below: E Television distribution (C-, Ku-Band): Satellites on "xed 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 recon"gurable or can be repointed. E Direct TV (Ku-Band): Satellites on "xed orbit positions, operated by national or international companies (e.g. Nahuelsat, Eutelsat, USDBS). Ku-band payload transponders with high RF power of up to 200 W provide high EIRP for direct TV or HDTV. Antenna beams can be recon"gured or repointed with very high accuracy (0.13). In areas with high payload channel demand, satellites are collocated in the same position (e.g. ASTRA with four satellites).

411

E Telecommunication bundle (C-, Ku-Band): Satellites on "xed or movable orbit positions, 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 C-band payloads or they have mixed Ku-band and C-band transponders. Antenna coverages are recon"gurable 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. E Telecommunications connections, data transfer (C-, KuBand): Satellites on "xed or movable orbit positions, 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 Ku-band and C-band transponders. Antenna coverages are recon"gurable or can be repointed (complex antenna farms). Service between individual V-Sat terminals is at medium power levels. E Mobile communications (L-Band): Satellites on "xed or movable orbit positions, operated by national companies (e.g. M-Sat, Aussat II). Payload is in the L-band for mobile communications (truck services), with large re#ector-antenna or phased array antenna for nationwide coverage. 2.1.2. LEO/MEO constellations LEO constellations are composed of large numbers of satellites on various LEO orbits covering almost the whole of the Earth's surface (Tables 1 and 2). The orbit altitude and availability requirements de"ne the number of satellites (between 20 and 288). They form a complete network for communications services. The tra$c is controlled either by the satellites themselves or gateway ground stations. Systems with permanent visibility o!er a continuous tra$c link, which is guaranteed by controlling the constellation to form a closed net and by handover procedures when satellites disappear from users visibility. The satellites are normally injected into orbit in

Table 1 LEO/MEO constellations System

Ellipso (ellipt.)

Ellipso (circ.)

ECCO

Teledesic

Orbit planes Satellites per plane satellites Total Op. Orbit Inclin. (3) Altitude Apo./Peri. (km) Band Mass (kg) Operational year

2 4#1 8 116 7500/670

6#1 6 0 8060

1 equat. 11#1 11 0 2000

12 24#n 288 90 1350

700}800

280

Ka 1500 2002

700}800 2000

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Table 2 LEO/MEO constellations System Orbit planes Satellites per plane satellites Total Op. Orbit Inclin. (3) Altitude Apo./Peri. (km) Band Mass (kg) Operational year

Orbcom

Global-star

ICO

Iridium

3 8 24 45 770

8 6#1 48 52 1400

2 10#2 20 45 10,300

6 11#1 66 86 780

40 1997

L 450 1998

2450 2000

L 689 1998

a multiple launch con"guration, and the "nal positioning and orbit maintenance is performed by their own propulsion systems. To avoid debris in orbit, at the end of a satellite's life or in case of failure they must be transferred into a graveyard orbit or de-orbited into the atmosphere.

E ICO } telecom service: A satellite system that consists of 20 satellites plus spares operating on two planes with 453 inclination at an altitude of 10,300 km. The satellites provide global telecommunications services for individual user terminals. The satellite control is similar to LEOs.

2.1.2.1. Little LEOs

2.1.3. Next generation LEOs There are many constellation programs planned to replace and improve the existing LEOs. The predicted growth in tra$c capacity requires more powerful systems. Satellites with longer lifetimes and more available electrical power will use electric propulsion for station keeping and also for orbit raising to save launch costs. This will lead to maneuvers with very long duration at very low thrust levels.

E Paging, message transfer (very low data rates in L-band and/or UHF): Mobile communications networks for low data rate transmission. The satellites (e.g. Orbcom, LEO One, etc.) are very small and also serve as relays using UHF with good propagation. Low pointing accuracy allows for simple attitude control (e.g. gravity gradient). 2.1.2.2. Big LEOs/MEOs (Iridium, Globalstar, ICO-P, Ellipso) E Iridium } independent mobile communications network: The Iridium network consists of 66 satellites plus spares on six nearly polar orbit planes (i"863) with a height of 780 km. It is an independent system with on board switching and routing of communications links for voice service. The user cell-phone-satellite link is in the L-band and the inter-satellite link is in the Kaband. Six phased array antennas connect the adjacent satellites in plane and neighboring planes to form a worldwide tra$c net. E Globalstar } mobile communications network controlled by gateways: The Globalstar system consists of 48 satellites plus spares on eight orbit planes with a 523 inclination at an altitude of 1,400 km. The system switching and routing of communications links for mainly voice services are performed by the gateway ground stations. The user cell-phone-satellite link is in the L-band and the gateway link is in the C-band. The user antenna link for receive and transmit is formed by eight beams. The power level is tra$c dependent. The satellites are nadir pointed with yaw steering.

2.1.4. Mega LEOs (Teledesic, Sky Bridge) E Teledesic: Very big satellite system consisting of 288 satellites in 12 orbital planes at a height of 1,350 km. The system will provide high data rate broadcast services, video, teleconferencing to "xed locations worldwide. The satellites provide electrical power in the range of 6000 W and electric propulsion is envisaged. Satellite control, station keeping and orbit maintenance are fully autonomous. 2.1.5. HEO constellations. High data rate service (Ka-band), augmenting and connecting xber optic rings E Pentriad: The Pentriad system of nine high power satellites (plus spares) will be operated from 3 linked highly elliptical orbits (&Molniya' type). The system can provide broadband satellite service, including multicasting and a &direct to home' service in the northern hemisphere. From the quasistationary active arc the areas covered in North America, Europe and North Asia can be served from high elevation angles. Primary broadband channels of 155 Mbps can be grouped to channel virtually up to 3.8 Gbps. The satellites will point to the center of the coverage areas with yaw

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steering. Station and orbit keeping maneuvers will be performed in the nonactive part of the orbit. 2.2. Requirements Satellite control systems' requirements can be divided into the categories performance and autonomy. The performance criterion for orbit control is the optimal use of propellant to achieve the desired orbit and maintain the accuracy of orbital elements, while on station. GEOs are injected into circular orbit from a transfer orbit by a quasi-impulsive apogee boost maneuver (circularization and inclination). In orbit, station keeping has to be performed to maintain latitude and longitude with a typical accuracy of $0.13 and $0.053, respectively. Co-positioning of three or more satellites in &one box' requires coordinated relative station keeping to avoid radio signal interference or even collision. During launch, LEOs and HEOs are "rst injected into intermediate parking orbits. The transfer to "nal 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 the 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. The performance requirements of the attitude control system are mission dependent. For GEO satellites, generally, the beam pointing shall be kept within $0.13 in pitch and roll and $0.23 in yaw. High power satellites for direct TV request $0.053 and $0.1253, respectively. Satellites with large antenna farms will need individual beam pointing control in closed loops or open loops. Intelligent control algorithms compensate for long term and periodical distortions. LEOs normally allow for pointing errors of 0.5}13 and 3}53 for simple satellites. HEO satellites have similar requirements to GEOs, but pointing is 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 de"ned orientation to the sun for proper power and thermal conditions. Satellite communications service providers are demanding on-board autonomy to reduce operational costs, minimize outages of communication channels, and to increase overall reliability of the network. Autonomous procedures for failure identi"cation, recon"guration and recovery are necessary to meet the above demands. The autonomy requirements are di!erent for single satellites and for LEO constellations.

413

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

3. Control systems, challenges and solutions 3.1. GEO satellites The improvement of AOCS performance for geosynchronous communications satellites is illustrated in Fig. 3 for the representative example of the Spacebus AOCS. Key features of the AOCS 3000B of e.g. EUTELSAT III are: E Attitude control by momentum wheels in all three axis in Normal Mode. Advantage: Extension of satellite lifetime by three years. E Station keeping in normal mode under normal mode attitude control. Advantages: Double failure tolerance with 14 thrusters, easy adaptation to electrical thrusters. E Momentum management during station keeping. Advantage: Extension of satellite lifetime. E On-board failure management. Advantages: Autonomous operation up to two weeks, double failure tolerance for gyros, sun sensors and thrusters, survival of Earth sensor double failure. E Modular architecture, featuring ADA software and P1750 processors. Advantage: Easy exchange and up-grading of software modules. The block diagram of the AOCS 3000B is shown in Fig. 10, the autonomous failure management is described in Section 4.1.

Fig. 3. Performance of the Spacebus AOCS.

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E Flexible structures: The critical parts are the large #exible solar generators and antenna re#ectors, which lead to elastic modes within the control system bandwidth. The moments of inertia (MOI) of the solar generators about the spacecraft center of mass is typically 3 to 4 times the MOI of the spacecraft body around roll and yaw axes for this type of satellite. The elastic modes are very weakly damped. Typical values for elastic eigenfrequencies ("xed base) are: Normal bending:

First mode Second mode In-plane bending: First mode Torsion: First mode

0.06}0.08 Hz 0.36}0.48 Hz 0.20}0.30 Hz 0.45}0.55 Hz

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

Fig. 4. Wheel and sensor arrangement on EUTELSAT III.

The following list gives an overview of the AOCS 3000's components. The wheel and sensor arrangement is presented in Fig. 4. For more details see (Surauer, Bittner, Fichter & Fischer, 1993). Sensors E Two-axis Earth sensors (2). E Two-axis Sun sensors, distributed in the S/C x/z-plane (6 heads). E Three-axis gyro units (2). Actuators E Momentum wheels and reaction wheels. E Bi-propellant thrusters, 10 N, for acquisition and station keeping (14). On-board computer unit (OBCU) E Processor modules P1750 (Performance), Ma31750 (Plessey), SPARC/ERC 32 (MHS/TEMIC). E Plug-in PROM. E Safeguard and recon"guration modules. E Equipment interface. 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.

E Low acceleration propellant sloshing: At launch, about 50 to 60% of total satellite mass is made up of propellant. After the "nal orbit position has been reached, the remaining fuel for station keeping is still considerable (around 30% of total satellite mass). Depending on the fuel-tank structure, the dynamics of the liquids, the fuel sloshing, exhibits nearly undamped oscillations. The lower modes of the highly non-linear #uid dynamic equations are approximated by spring/mass or pendulum systems. The sloshing masses 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 spacecraft acceleration, which also determines the orientation of the propellant during thruster burns. During transients, sloshing amplitudes as high as 0.5 m may be encountered during station keeping. The "rst propellant eigenvalues are around 0.1 Hz for apogee boost and about 0.02 Hz for station keeping maneuvers. Propellant sloshing and #exible 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. E Pointing accuracy: Attitude pointing accuracy has to be maintained under severe sensor noise conditions during on-orbit normal modes. Transients are critical at the beginning and end of station keeping maneuvers and yaw performance in the Sun/Earth co-linearity region. E Co-positioning: The increasing demand for transponder capacity in distinct orbital positions has led to the co-positioning of three and more satellites in one orbital box of ($0.13 in North/South and ($0.053 in East/West directions. The box size of 70 by 35 km

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415

Table 3 Technical data of the Globalstar constellation No. of orbital planes No. of satellites per plane Orbital altitude Orbital period Orbit inclination

8 6 1414 km 113 min 523

Table 4 Attitude accuracy requirements

Control error (3) Sensor error (3)

Fig. 5. Illustration of the Globalstar system.

Roll

Pitch

0.2 0.25

0.2 0.25

E Magnetometer. E GPS receiver for navigation purposes, time reference, and for backup attitude sensing. Actuators

demands that the relative positions of the satellites are controlled to avoid collision. To do this with minimum ground intervention requires on-board autonomous orbit position determination and control. E De-orbiting: Every spacecraft has to be transferred to graveyard orbits with 99.3% probability of success at the end of life (12 to 15 yr) or in case of spacecraft failure. In Section 4, methods and procedures to improve onboard autonomy in particular for GEO satellites will be discussed. 3.2. LEO constellations Fig. 5 illustrates the Globalstar constellation, a typical LEO constellation for communication purposes. Table 3 gives the corresponding technical data. The attitude accuracy requirements (3p) of the Globalstar satellites in the roll and pitch axis is shown in Table 4. The required attitude accuracy in the yaw axis is 1.23, which also includes sensor errors, when Sun sensors or magnetometer are used as yaw sensors. The following list gives an overview of the control system hardware. More detailed information is given in Alexander, BruK derle, Groeger, Schrempp and Widmann (1997). Sensors E Two-axis Earth sensor. E Two-axis Sun sensors, distributed in the spacecraft's x/z-plane.

E Four momentum wheels. E Two magnetic torquers for total momentum control. E Five mono-propellant thrusters, 1 N, for orbit correction and acquisition modes. The on-board processing electronics (OBPE) is based on a 1750A CPU, 64k RAM, 128k PROM. The OBPE is fully redundant. Using Globalstar as a representative example, the following paragraphs describe two distinct features of constellation satellites, their impact on control system design, and future challenges. Available power: The available power on-board 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 two-axis solar array drive, i.e. the solar arrays are not only rotated, but also tilted. An 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 di$cult to reject when magnetic torquers, i.e. restricted, timevarying torque capability, are used as attitude control

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E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

Fig. 6. Schematic of Globalstar yaw-steering.

actuators (in addition to one momentum wheel in the orbit normal axis). Control approaches and optimization techniques to this control problem are described in Bals, Fichter and Surauer (1996). When yaw-steering is applied, the satellite body has to periodically perform large-angle yaw slews, according to a reference yaw pro"le that depends on the Earth}satellite}Sun con"guration (see Fig. 6). This is done with internal control torques generated by four redundant wheels. The control task becomes di$cult when the three-axis attitude tracking has to be performed with only two-axis attitude measurements. Solutions to this problem are described in Fichter, Surauer and Zentgraf (1996) and BruK derle, Fichter, Lange, Furumoto and Rodden (1996). Minimum equipment conxguration: In order to reduce the overall control system costs of a satellite constellation, it is important to minimize the equipment and the recurring costs. The Globalstar satellites are equipped with Earth sensors, Sun sensors, and magnetometers, but no gyros are implemented. All attitude control modes, including acquisition modes, operate completely without gyros. Appropriate estimation and control algorithms were developed to compensate for the lack of direct rate measurements. They are based on the directional measurements of the Sun sensor and its derivatives. The principles of these algorithms are described in Surauer, Zentgraf and Fichter (1996). Future challenges: The attitude control system of current constellation satellites operate fairly autonomous. However, there is a need for further development and for further cost reduction in the following areas: E implementation of autonomous position control and constellation keeping, E low cost spaceborne GPS receiver for navigation purposes, E exploitation of intersatellite links for control purposes, E implementation of functional redundancies for reliable de-orbiting.

Fig. 7. &Pentriad' (9 satellites in 3 planes, 5 active arcs): Observer on Earth "xed system.

3.3. HEO satellites Fig. 7 shows the orbits of the &Pentriad' HEO satellite system (inclination: 64.33, apogee/perigee: 43,000/ 3000 km). To avoid the two degrees-of-freedom solar generators, yaw-steering, similar to Globalstar is performed to point solar panels in the direction of the Sun. Unlike 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 references, an accurate knowledge of the satellite's orbit and position is needed on-board. Pointing to coverage area centers requires pitch and roll maneuvers in addition to the yaw steering. The control system hardware is similar to Globalstar. CCA pointing accuracy on the active arc is similar to GEO requirements, while the relaxed pointing accuracy on the non-active arc is determined by the station keeping maneuvers. For this type of orbit, disturbance torque magnitudes vary signi"cantly, depending on orbital altitude and spacecraft orientation. This e!ect can be used to optimize momentum management and momentum/reaction wheel sizing. The attitude measurement system has to be adapted to cover the range from perigee to apogee altitude. Experience gained with the three axis stabilized transfer of GEOs can be used in the design of this system. 3.4. Trends in implementation Performance, quality and cost are design drivers for future communications satellites. To achieve this requires

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4. On-board autonomy

Fig. 8. On-board AOCS Software Modularity.

The 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, while reducing operational cost. Operational safety is increased by an improved failure detection, isolation, and recovery (FDIR) system, implemented in the on-board computer. The FDIR system of Spacebus 3000B is explained in detail in Section 4.1. The reduction of nominal operational costs 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 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 modi"cations. When orbit determination is implemented, autonomous orbit control can then be realized with existing on-board equipment and software adaptation only. 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 geostationary orbit, GPS receivers or even GPS/GLONASS receivers can also be used for orbit determination. The next section brie#y describes the corresponding navigation techniques. Alternatively, a second approach for (semi) autonomous orbit determination is described for geostationary applications, based on the existing sensor con"guration of Spacebus 3000B. This leads to a limited autonomy period. The advantage is that only equipment that is already available on board is used. 4.1. Autonomous failure management

Fig. 9. Central processing node * architecture.

a simplicity of design, the multiple (re)-use of hardware and software, and the potential for growth and change by modularity. The hierarchical, modular architecture underlying AOCS software design at DSS is illustrated in Fig. 8. Hierarchy and modularity of software design simpli"es the handling and adaptation of the large number of AOCS modes (typically more than 20). Concerning hardware and on-board processing, similar considerations led to the decomposition of the physical and functional entities of Fig. 9. By optimizing the distribution of the functions and interfaces of the central processing node, the costs, mass, and power can be further reduced, while computing performance increases to accommodate the new functions, which are required for the AOCS projects in planning.

For geosynchronous communications satellites, constraints on outage time are very severe. Typically, less than a total of 60 min outage time accumulated over an operational life of 15 yr 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 h at any time. To meet these requirements failure identi"cation and recovery have to be performed on board with a minimal interruption of communication links. The FDIR System of the Spacebus 3000B is based on the hierarchical classi"cation of failures, which facilitates identi"cation and allows for recovery with minimum intervention of AOCS functions. In addition to the classical concept of (usually dual) equipment redundancy, functional redundancy is used, e.g. for back-up modes. Fig. 10 shows an example of AOCS hardware for present communications satellites. Safeguard Memory (SGM) with a failure history register (spacecraft and equipment), a Recon"guration Module (RM) with

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E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

Fig. 10. Schematic of Spacebus 3000B AOCS.

processor watchdog and status relays, and independent high priority control lines (HPC1/2) are the hardware on which the FDIR is based, independently from the AOCS On-Board Computer Unit (OBCU). The FDIR software functions are decentralized to the maximum possible extend and hierarchically structured in four levels: (1) Level zero failures: The identi"cation and recovery is handled locally at equipment level (BITE), no interruption of Normal Mode operation, status message to ground. (2) Level one failures: The detection of equipment failures by software criteria in OBCU and recovery by switching to the redundant unit. (3) Level two failures: The watchdog monitoring of OBCU and initiation of recon"guration sequences by RM in three steps: Step 1: Warmstart (Software Reset), Step 2: Coldstart (Power Supply OFF/ON), Step 3: Recon"guration to a new prede"ned con"guration. (4) Level three failures: Hardware monitoring of critical equipment, such as the propulsion or power systems and the recon"guration of AOCS by pre-established sequences stored in the RM. To support failure analysis on the ground, the occurrence and the status of

the failure is stored in the failure history register, which is part of the SGM. Only if all of the above measures fail, the system is switched to the ultimate sun-pointing Safe Mode to secure the survival of the spacecraft and to provide su$cient time for analysis and recovery. To restart Normal Mode operation, the satellite has to go through the full acquisition sequence, which may take up several minutes in the worst-case scenario. By applying this minimum intervention strategy for handling satellite anomalies, most failures can be repaired without losing pointing and therefore without outage and interruption of service. 4.2. Orbit determination of geostationary satellites using GPS signals 4.2.1. Visibility conditions The GPS satellite constellation (Fig. 11) nominally consists of 24 satellites. The semi-major axis is about 26,500 km which corresponds to an orbital period of about half a day. Since the geostationary orbit (semimajor axis 42,200 km) is above the GPS orbit, and GPS signals are radiated towards the Earth, only GPS signals of satellites beyond Earth can be received. This situation is shown in Fig. 12.

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

419

Fig. 11. GPS constellation.

Fig. 13. Visibility, longitude 703, receiver antenna gain 10 dB.

Fig. 12. Geometry GPS}GEO.

The geometrical con"guration of the GPS-satellites, GEO satellite, and the Earth leads to very poor GPS visibility conditions, seen from the geostationary orbit. Fig. 13 shows the number of visible GPS satellites over one orbit revolution, based on the following assumptions: E receiver antenna gain of 10 dB; GPS satellite half beam width of up to 213 (GPS satellite antenna pattern), E carrier-to-noise ratio threshold of 37 dB Hz for signal acquisition, E arti"cially increased Earth radius (1000 km), in order to avoid tropospheric and ionospheric e!ects. Fig. 14 shows the accumulated time interval, when more than two GPS satellites are visible simultaneously, as a function of the geostationary longitude. It can be seen, that the visibility conditions vary signi"cantly. 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.

Fig. 14. Visibility from GEO.

From the visibility plots it is clear that there are less than four satellites available at one time instant, and a direct kinematic navigation solution is not possible. Therefore Kalman "ltering has to be applied. Most publications in the past have proposed Kalman "ltering 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. In other words, a relatively precise receiver clock is required for reasonable navigation accuracy. However, it would be more favorable to achieve good navigation results without stringent requirements on the receiver clock. In Averin, Vinogradov, Ivanov and Salischev (1996), an alternative approach is described based on the single

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di!erences of pseudo-range measurements. It eliminates the receiver clock bias and requires only the modeling of well-known orbit dynamics. This leads to relatively accurate navigation results without the stringent requirements on the receiver clock. 4.2.2. Single diwerence method The pseudo-range measurement between the geostationary satellite and the GPS satellite is given by the following: p ""r !r "#c ) b #f . (1) G G%.1 12 12 G A linearization with respect to a reference value (best estimation) yields





*r 12 #* f , *p "p( !p "[ l) , 1] ) G G G G G c ) *b 12 where



(2)



2 r !rL 12 l) "! G%.1 (3) G "r !rL " G%.1 12 is the unit vector (line-of-sight direction) from the GEO satellite to the GPS satellite. In order to apply Kalman "ltering, the pseudo-range measurement has to be related to the state of the underlying dynamics. This can be expressed as



Fig. 15. Performance analysis, reference case.

1

*p(k)"C(k)U(k, 0)*x(0)# : ) c ) *b(k)#* f (k)

(4)

1

where the measurement matrix is given by



l) 2  C(k)" $

0 0 0

l) 2 K

0 0 0

$

$



$ .

(5)

*p(k) is a (n;1) vector, when n pseudo-ranges are measured at a time. For n*2 (practically n"2 or n"3 in the GEO orbit) the clock bias b can be eliminated by forming di!erences of the pseudo-range measurements. This can formally expressed by a multiplication of Eq. (4) with ¹ from the left, where



1

0

)

0

!1

0

)

)

)

)

)

)

)

)

)

0

)

0

1

!1

¹"



, dim [(n!1);n].

(6)

For the new measurement equation we have d(k)"T(k)*p(k)"T(k)C(k)U(k, 0)*x(0)#T(k)*f(k). (7) The Kalman "lter is fed by the pre-processed (di!erenced) measurements d(k).

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

solar pressure, Earth oblateness, triaxiality of the Earth, gravitation of Sun and Moon.

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 random noise with a magnitude of 5% (1p) of their nominal value. The noise measurement 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) e!ects. Receiver noise and other measurement errors are modeled as white noise. All simulation runs are based on worst-case visibility conditions, i.e. not more than two GPS satellites are visible at a time. Fig. 15 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 5 summarizes the corresponding statistics for the reference case. A deviation of 1,000 m in the X-direction corresponds to an

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

4.3. Semi-autonomous orbit determination based on Earth and Sun sensor measurements

Table 5 Statistics of the reference case Reference case

Standard deviation (m)

Mean value (m)

Orbit normal > Along track X Radial Z

51.1 182.1 99.8

0.1 !36.8 !7.4

Max. deviation (m) 152.0 878.6 505.9

Table 6 Results of the performance analysis Position error

Reference case #5% (random) Earth oblateness #5% (random) Earth triaxiality #10% (constant) solar pressure #5% (random) Sun and Moon grav.

421

Orbit normal >

Along track X

Radial Z

(%)

(m)

(%)

(m)

(%)

(m)

* 1.3

152 154

* 0.5

879 883

* 0

506 506

!4.6

145

0

879

0.6

509

12.5

171

4

913

7.3

543

6.6

162

0.7

885

3

521

angular East/West deviation of 1.5;10\3 in the geostationary orbit. Note, that a common station keeping window has a width of $0.053 in E/W direction, which corresponds approximately to $32 km. Table 6 shows the navigation accuracy when GPS signals are used for navigation. Results are given for each axis (orbit normal, along-track, radial), and for di!erent disturbance forces. The maximum error occurs in X-direction, which is caused by selective availability and solar pressure. The position deviations in percentage are computed with respect to the reference case. In Table 7 the performance of GPS navigation is compared with navigation based on the GLONASS satellites only, as well as with the combination of both (GNSS). Only position errors in the Xaxis (along track, worst-case axis) are shown.

An alternative approach to orbit determination is to exploit two-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 determine the orbit autonomously. Besides that, satellite #ight data (telemetry) shows that the sensor con"guration is sensitive to the thermal deformation of the satellite structure. A semiautonomous orbit determination will be presented, exhibiting the following key features: E propagation of the satellite's North/South motion based on exact mathematical models, E estimation of the satellite's East/West motion based on Earth and Sun sensor measurements, E calibration of the Earth/Sun sensor con"guration regarding thermal deformation of the satellite structure and sensor misalignment. 4.3.1. Sensor conxguration The sensor con"guration consists of a two-axis Earth sensor and a two-axis Sun sensor. The Earth sensor is aligned along the satellite Earth pointing #z -axis and @ furnishes information on the roll and pitch attitude ( , h) of the satellite. The Sun sensor heads (SSH) are distributed in the satellite x /z -plane (see Fig. 16) so that the @ @ direction of the Sun can always be measured in the form of a unit vector s . @ 4.3.2. Orbital motion Typically, deviations of a geostationary satellite from its nominal position must not exceed $0.053 in the East/West and $0.13 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, represented by the reference coordinate system (x , y , z ) in Fig. 17. P P P

Table 7 Position error in X-direction

Reference case #5% Earth oblateness #5% Earth triaxiality #10% solar pressure #5% Sun and Moon grav.

GPS position error X

GLONASS position error X

GNSS position error X

(%)

(m)

(%)

(m)

(%)

(m)

* 0.5 0 4 0.7

879 883 879 913 885

* 0.4 0.1 21 6

787 790 788 953 741

* 0.5 0 2 5

720 723 720 704 754

422

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427 Table 8 Dominant perturbing forces and their e!ect on the orbital motion of geostationary satellites Source

E!ect on GEO

Earth potential Zonal e!ects Tesseral e!ects Luni-solar gravitation Solar pressure Station keeping maneuvers

E/W oscillation, drift E/W drift rate N/S oscillation E/W oscillation, drift E/W and N/S accelerations

Fig. 16. Typical distribution of Earth sensor (IRES) and Sun sensor heads (SSH) in the x /z -plane of a geostationary satellite (SIRIUS2). @ @

Fig. 18. In#uence of longitudinal deviations on the angle between Sun and Earth direction. Fig. 17. Orbit reference, attitude reference and body frame (subscript r, r, b).

The resulting equations of relative motion are well known as Hill's equations. For a representation with respect to small deviations j"x /a (longitudinal deviP ECM ation), b"y /a (latitudinal deviation), r"z /a P ECM P ECM (radial deviation), the following set of equations is valid: f j$ !2u r " V ,  ma ECM f b$ #u b" W ,  ma ECM f rK #2u jQ !3u r" X , (8)   ma ECM with a the geosynchronous radius, u the frequency of ECM  the reference orbit, m the satellite mass and f the domiG nant perturbing forces on the satellite (see Table 8). 4.3.3. Measurement principle To outline the measurement principle, a "xed body satellite (and therefore attitude dependent) system (x , y , z ) with its origin in the center of mass of the @ @ @

satellite is de"ned. The measured Sun vector s is related @ to its calculated reference s by a transformation about P the small orbit and attitude deviation angles, s is known P from ephemeris data. The resulting measurement quantity y(t) corresponds to the di!erence of the nominal (c )  and the actual angle (c) between the Sun and Earth directions: y(t)"e2s !e2s "cos c!cos c . (9) @ @ P P  The e!ect of longitudinal and latitudinal deviations of the satellite from its nominal position on the measurement quantity is presented in Figs. 18 and 19, respectively. Summing up, the measurements of Earth and Sun sensors can be processed to a scalar quantity featuring: (1) attitude independence, (2) time variance, periodic with orbit frequency u ,  (3) contributions of orbit deviations j and b. 4.3.4. Orbit determination principle Fig. 20 shows the combined in#uence of orbit deviations on the angle between Earth and Sun directions. There is some ambiguity in the e!ects of longitudinal and

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

423

Fig. 19. In#uence of latitudinal deviations on the angle between Sun and Earth direction. Fig. 21. Schematic of semi-autonomous orbit determination.

4.3.5. System A: North/South propagator The propagator is based on Hill's equations and the perturbation models of Table 8. In general, propagation accuracy is limited by the following dominant factors:

Fig. 20. Combined in#uence of orbit deviation angles on the angle between Sun and Earth direction

latitudinal deviations on the scalar measurement, i.e. a deviation angle j may have the same e!ect on y(t) as latitudinal deviations. This can also be shown analytically: The East/West (j) and the North/South (b) motion of the satellite are not fully observable with measurements y(t). Measurements of the Earth and Sun sensors can only be used to estimate one of the two subsystems: the North/South motion or the East/West 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. For practical autonomous orbit determination purposes it is favorable to estimate the East/West 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 a relatively fast divergence with pure propagation. In contrast the North/South motion can be accurately propagated over long periods of time. Fig. 21 shows a schematic of semi-autonomous orbit determination. The process is divided into subsystems A (North/South propagator) and B (East/West "lter).

E initialization errors due to limited tracking accuracy: *b +$0.001!0.0023, RP?AI E accumulation of errors in maneuver modeling: "*v ")0.02"*v ". CPP ,1 Fig. 22 compares the SIRIUS2 latitudinal ground station tracking data (triangles) to the propagation results. The propagator is initialized with tracking data and no station keeping maneuver is performed. The propagation error is signi"cantly smaller than one thousands of a degree after more than a week of propagation. Its magnitude is in the range of tracking accuracy. Errors in the propagation of the North/South motion contribute to the measurements y(t) and therefore directly e!ect the observation of the East/West motion. 4.3.6. System B: East/West xlter The estimation is based on a discrete Kalman "lter algorithm (Hill's equations supported by Earth and Sun sensor measurements). The estimation accuracy is limited (in the worst case) by E sensor noise: Dominant IRES noise 3p"$0.13, E errors in the disturbance model: 10% of solar pressure (deterministic), 10% of other disturbances (stochastic), E propagation error (2% N/S maneuver modeling, tracking inaccuracy): "*b"+0.013 in six months. One advantage of the Kalman "lter is that it allows the estimator to be run in Redundancy Mode, i.e. measurements are taken from a single Sun sensor head in a limited section (1203 "eld of view) of the orbit. The accuracy of the estimation in Redundancy Mode compared to Normal Mode (3603 Sun sensor measurements) is not signi"cantly decreased.

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E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

Fig. 22. SIRIUS2 tracking data (triangle) compared to North/South propagation.

Fig. 23 represents estimation results in Redundancy Mode. During the propagation sequence (2403) the error increases to "j")0.0043. Fig. 24 shows the "lter's response to propagation errors. North/South maneuver modeling errors (six months station keeping) are assumed to be of identical orientation (very conservatively) and are represented as errors in the initial condition of the North/South propagation (*b)0.013). Note that the resulting estimation error vanishes for s "0. All simuPW lation results are based on a &cold start' operation (zero initial "lter condition) and calibrated sensors. 4.3.7. Calibration The governing equation for Earth and Sun sensor measurements is valid for calibrated sensors only, i.e. the sensor matrices are known precisely. Extensive #ight data investigation (EUTELSAT2 FM2, SIRIUS2) has shown, that the sensors are subject to constant and time varying misalignment, the latter due to thermal deformation of the satellite structure. The misalignment has the e!ect of time varying sensor matrices on the measurements. This is expressed as an error function k(t) in the governing Eq. (9): y(t)"e2s !e2s !k(t). @ @ P P

(10)

Fig. 23. Satellite longitude estimation compared to the true deviation (center) and the estimation error (bottom) related to the Sun vector x-component s (top): 1203 Sun sensor "eld of view SIRIUS2 SSH 3/6 PV con"guration.

Fig. 25 presents a typical error function k(t) for a single Sun sensor head with 1203 "eld of view (EUTELSAT2 FM2, summer solstice 1992). The top plot exhibits daily variations with a magnitude of up to $0.053, the bottom plot gives an idea of the di!erence in two subsequent error functions, i.e. the short-term reproducibility of k(t). The magnitude of k(t) is too big to ignore these sensor errors in the course of autonomous orbit determination. Calibration of the sensor con"guration, i.e. determination of the relevant k(t) is the key to successful orbit determination. At DaimlerChrysler Aerospace a longterm investigation is currently being conducted to analyze the #ight data and to identify a suitable model for the error functions. Additionally, modi"cations to future bus design will reduce the in#uences of thermal deformation. Mounting the Sun sensor head 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). 4.3.8. 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,

E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

Fig. 24. Long term in#uence of N/S propagation errors *b: latitudinal errors (top) related to resulting longitudinal estimation error and Sun vector y-component s . PW

the introduction of autonomous orbit determination increases neither system complexity nor hardware expenses. The disadvantage is that the available measurements do not contain su$cient orbit information to estimate North/South and East/West motion. A reasonable solution to this observability problem is the propagation of a stable and well modeled North/South motion. A "lter can then be designed to estimate the East/West motion, especially the satellite drift. 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 speci"c 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 of the Earth and Sun sensor con"guration presented can therefore easily be adapted for the application of star sensors for example. Summing up, the challenge in autonomous orbit determination with Earth and Sun sensor is to reduce the relative calibration error of both sensors to 0.013. Table 9 shows an error budget on this assumption. The resulting orbit determination accuracy of $0.023 in satellite longitude is believed to be su$cient to control the satellite within the position tolerance limits of $0.053.

425

Fig. 25. EUTELSAT2 FM2 summer solstice: A typical error function k(t) (top) and the di!erence in subsequent k(t) (bottom).

Table 9 Error budget: Satellite longitude Constant

Orbit

Stochastic

Sensor noise Redundant operation (1SSH) Error in N/S propagation Calibration error

* *

* 0.0053

0.0023 *

* 0.0053

0.0103 0.0053

* *

( x (columns) G

0.0053

0.0133

0.0023

Total error

0.0203

5. Standard AOCS process 5.1. Introduction For the "rst time the growing demand and worldwide competition in communications satellites leads to a need to streamline the product and its manufacturing process for mass production. To reduce costs and delivery time, while maintaining high standards in quality and performance requires a re-engineering of the process by which the complex AOCS subsystems are designed, developed, and manufactured.

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E. Gottzein et al. / Control Engineering Practice 8 (2000) 409}427

Fig. 26. Key elements of improved AOCS process.

In this section the approach developed at DaimlerChrysler Aerospace to ful"ll all of the above needs is described as an example. The main steps of the re-engineering process are: E streamlining the organization by bringing all AOCS related activities together in a single product center, E introducing standardized tools and standardized development, veri"cation and test facilities, E developing a modular on-board software and hardware architecture, E centralizing the procurement and production of components and equipment. Fig. 26 shows the key elements of the improved AOCS process, see Lange, MoK llenho!, Oesterlin and Widman (1997). In the following sections interest is focused on the technical aspects of standardization of the development and test facilities. More detailed information is also given in Eichhorn, Farnetani, Fischer and Seidl (1997). 5.2. Standardized facilities The three main facilities shown in Fig. 27 (Lange et al., 1997; Stinsho!, Fischer, Surauer & Lange, 1997), are E the Software Development Facility (SDF), E the Software Veri"cation Facility (SVF) and E the Dynamic Bench Test Facility (DBT). Following their design, the control algorithms are transferred to the Software Development Facility, where they are directly coded into #ight software in ADA by the control engineers. This allows the simultaneous functional testing of AOCS algorithms in the closed loop and the software at a very early stage of the process. In addition, sensitivity analysis and module tests are performed. For ease of use the simulation environment is built with Matrix 2+ in a modular architecture,  with de"ned interfaces to the spacecraft model and the equipment models (written in C). Adaptations, extensions or an exchange of single models can be realized easily.

Fig. 27. Standardized facilities for the AOCS process.

In the Software Veri"cation 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 quali"cation not only with the computer but also with sensors and actuators (as much as possible) in the closed control loop. For performance tests of the "rst #ight model of both GEO satellites and constellation satellites, a full set of tests is performed at the DBT facility, by using the telemetry/telecommand interface as realized in the ground station, to verify the AOCS operation procedures. For the subsequent #ight 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. 28 shows the DBT facility and Table 10 the DBT speci"c data. In all three facilities the same quali"ed spacecraft model and the same quali"ed equipment models are used. Because of this provision, adaptation and quali"cation only have to be realized once and no di!erences in reference runs due to di!erent 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, #exibility, performance and quality have been improved. New software and hardware modules can be added easily, and modi"cations and adaptations to the AOCS as well as to the test benches can be realized in an easy and controlled way.

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427

roughly divided into two areas: autonomous failure management and autonomous orbit determination and control. Both areas were described in detail. Worldwide competition and the new situation of series production not only has an impact on the design of the product itself, but also on the development and production process. A suitable approach (&standard AOCS process') was outlined.

Acknowledgements The authors would like to thank Mr. Colin Rogers from Eutelsat, Paris, for providing #ight data of the Eutelsat2 FM2 satellite.

References

Fig. 28. Dynamic bench test facility.

Table 10 Dynamic Bench Test Facility data z Accuracy z Min. velocity z Max. velocity Outer axes Inner axes z 2 Sun simulators z 2 Earth simulators

0.00013 0.000013/s 30/603/s 100/2003/s

Modi"cations, in particular to the 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 spacecraft rapid prototyping process.

6. Summary Following a brief introduction to the development of the space communication market, current and future scenarios were described and corresponding requirements for attitude and orbit control systems were derived. State-of-the art attitude control systems were outlined and future challenges for the enhancements of these systems presented. A major part of future AOCS requirements is spacecraft autonomy, which can be

Alexander, R., BruK derle, E., Groegor, E., Schrempp, W., & Widmann, H., (1997). Attitude and orbit control for globalstar. Proceedings of ESA/IFAC international workshop on spacecraft attitude and orbit control ESTEC, Noordwijk, The Netherlands (pp. 5}19). Averin, S., Vinogradov, V., Ivanov, N., & Salischev, V. (1996). Application of di!erential method for relative positioning of geostationary satellites with use of GLONASS and GPS Navigation signals. Proceedings of the xfth international conference on diwerential navigation DSNS-96, Vol. 2 (pp. 1}6). Bals, J., Fichter, W., & Surauer, M. (1996). Optimization of magnetic attitude and angular momentum control for low earth orbit satellites. Proceedings of third international conference on spacecraft guidance, navigation, and control systems ESTEC, Noordwijk, The Netherlands (pp. 559}567). BruK derle, E., Fichter, W., Lange, B., Furumoto N., & Rodden, J. (1996). Dynamic momentum bias for yaw steering. Proceedings of 13th IFAC world congress. San Francisco (pp. 103}108). Eichhorn, E., Farnetani, G., Fischer, H.D., & Seidl, J. (1997). The development, test and validation of the ADCS for the Spacebus family. Proceedings of ESA/IFAC international workshop on spacecraft attitude and orbit control ESTEC, Noordwijk, The Netherlands (pp. 295}313). Fichter, W., Surauer, M., & Zentgraf, P. (1996). Control design for generalized normal mode operation of bias momentum satellites. Control Engineering Practice, 4(10), 1355}1360. Lange, G., MoK llenho!, M., Oesterlin, W., & Widmann, H. (1997). Approach for a standard AOCS development process. Proceedings of ESA/IFAC international workshop on spacecraft attitude and orbit control ESTEC, Noordwijk, The Netherlands (pp. 275}281). Stinsho!, K., Fischer, H.D., Surauer, M., & Lange, G. (1997). The role and signi"cance of dynamic bench testing in AOCS development and veri"cation. Proceedings of ESA/IFAC international workshop on spacecraft attitude and orbit control ESTEC, Noordwijk, The Netherlands (pp. 282}293). Surauer, M., Bittner, H., Fichter, W., & Fischer, H.D. (1993). Advanced attitude and orbit control concepts for three-axis-stabilized communication and application Satellites. IFAC Symposia Series. No. 12 (pp. 159}181). Surauer, M., Zentgraf, P., & Fichter, W. (1996). Attitude control of geostationary satellites with minimal use of gyroscopes. Proceedings of third international conference on spacecraft guidance, navigation, and control systems ESTEC, Noordwijk, The Netherlands (pp. 569}575).