JOURNAL OF SPACECRAFT AND ROCKETSVol. 34, No. 5, September–October 1997

Systems Engineering Trades for the IRIDIUM® Constellation

T. P. Garrison,¤ M. Ince,† J. Pizzicaroli,‡ and P. A. Swan§

Motorola Satellite Communications Group, Chandler, Arizona 85248

The customer imperative has driven the system design for the IRIDIUM® constellation and ground supportinfrastructure. Although the goal of the system is to provide global, mobile telephone service, the two principaldrivers for the trade studies are 16-dB link margin and commercial success. The systems engineering and con-stellation architecture trade studies that determined the design of the IRIDIUM telecommunications system werecomprehensive and are continuously evaluated. The complexity of the total system drove the engineers to criticalevaluations of major factors such as complexity of payload, size/weight of spacecraft, altitude of constellation,size of individual beams, space environment (radiation belts), size of telephone (pocket size), quality of voice/data,number of telephone calls, type of commercial launch vehicles, number of satellites, orbits and inclination of or-bits, number of gateways, crosslink types, and type of modulation. The complexity of the decisions and how theytrade off to one of the two basic principles of customer satisfaction, 16-dB link margin and commercial success, isdiscussed.

Introduction

T HE IRIDIUM® satellite system consists of an interconnectednetworkof 66 satellites(plus 6 spares) in low Earth orbit (LEO)

plus the ground-basedinfrastructurerequired to command and con-trol the satellites and connect them to the public switched telephonenetwork (PSTN). Figure 1 shows the system architecture with therf connections between satellites and the Earth. The satellites orbitthe Earth at an altitudeof 421 n miles. Their orbits are inclinedat anangleof 86.4 deg with respect to the equator,providingglobal Earthcoverage.The constellationis con� gured as six planeswith 11 oper-ational satellitesand 1 spare satellite in each plane.The satellitesareequallyspacedandarrangedso that theircoveragefootprintsoverlapto provide continuous coverage by handing communications traf-� c from satellite to satellite as they pass over the Earth. The sparesatellites are � own at a slightly lower altitude (350 n miles) untilthey are needed to replace worn-out operational satellites, at whichpoint the depleted satellite is deorbited and the spare maneuveredinto the appropriate slot in the constellation.

The satellitesare launched using a combinationof three differentlaunch vehicles. The U.S. Delta II rocket is used to launch � vesatellites at a time, the Russian Proton rocket launches seven at atime, and the Chinese Long March IIC/SD launches two satellitesat a time. Each satellite weighs approximately1460 lb fully fueled.Once in orbit, two large rigid solar array panels are deployed toprovide electricalpower, and three main mission antennapanels aredeployed to enable the L-band communications traf� c.

A signi� cant feature of the IRIDIUM satellite system is the K-band rf crosslinksused to provideconnectivitybetween the satellitesthat make up the constellation.Each satellite containsfour crosslinkantennas.Two are � xed and communicatewith the satellite’s nearestneighbors in the same plane: the satellite immediately ahead and theone directly behind. The north–south in-plane crosslinks are main-tainedcontinuously.Two additionalcrosslinkantennasare gimbaledand communicate with the nearest neighbor satellites in each of thetwo adjacentplanes. As adjacentplanes convergenear the poles, theeast–west crosslinks are disconnectedand then reestablished whenthe planesdivergesuf� ciently.The crosslinkscarry communicationstraf� c as well as satellite command and telemetry data.

Presented as Paper 95-U.204 at the 46th International AstronauticalCongress, Oslo, Norway, Oct. 26, 1995; received Nov. 1, 1995; revisionreceived April 14, 1997; accepted for publication May 22, 1997. Copyrightc° 1997 by the American Institute of Aeronautics and Astronautics, Inc. Allrights reserved.

¤Manager, Launch Services, 2501 South Price Road.†Consultant, Space Systems, 2501 South Price Road.‡Manager, Space Systems, 2501 South Price Road.§Manager, System Development, 2501 South Price Road. Fellow AIAA.

Satellite Physical DescriptionThe satellite vehicle (SV) is roughly 14 ft tall, with a triangu-

lar cross section about 3 ft along each side. This con� guration wasselected to allow manifesting multiple SVs on a single launch ve-hicle. In orbit, the SV is three-axes stabilized, with its long axismaintained in an Earth-pointingorientation.Figure 2 shows the SVin its deployed con� guration. The spacecraft mass budget is shownin Table 1. The SV is composedof three main assemblies: the spacebus module, the communicationsmodule, and the main mission an-tennas. In the following sections many aspects of the space systemdesign will be discussed, and in so doing, the complexity level andsome of the unique results from the system tradeswill be illustrated.

The communications section (CS) consists of equipment to sup-port the communicationsmission of the SV. This includes a suite ofconverters, tuners, switches, and modems, which interface with themain mission antenna (MMA) to provide the subscriber link tele-phony services. Additional equipment is included to operate fourSV–SV crosslinks, four feeder links for communicationswith gate-way and control segment Earth terminals, and a secondary link forbackup command and telemetry interfaces. Seven SV computerswith supporting memory and interface equipment provide the dataprocessing and routing control functions needed to support com-munications routing and resource management, antenna and linkcontrol, and SV � ight operations.A functionalblock diagramof theIRIDIUM � ight system shown in Fig. 3 illustrates the extent thatthe satellite is under software control.

The MMA consists of three deployable phased array antennas,which providethe rf links to the subscriberequipment.Each antennapanel utilizes an array of rf patches controlled by an array panelcontroller, which is in turn controlled by the CS. A beamformernetwork distributes the modulated carriers received from the CSand forms the 48 beams, which are transmitted to the ground.

The satellite bus module is capable of supporting the mass, vol-ume, and heat dissipation requirements of the satellite and will becapable of providing the required support and rigidity during allphases of launch, transfer, and orbit insertion.The SV is made fromgraphite composite material and supports the subsystems describednext.

Thermal Control SystemThe thermal control system (TCS) maintains a controlled ther-

mal environmentfor SV subsystemsand equipment througha com-bination of passive and active controls. Active thermal control isperformed using heaters and heat pipes. Passive thermal controlemploys surface � nishes, paints, and radiators, as well as thermalblanketsto control temperatures.A battery radiatorassembly is usedto dissipate heat generated in the batteries during charging. Batterycharge control is performedby the bus control software, which runswithin the space vehicle computer (SVC).

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GARRISON ET AL. 677

Fig. 3 IRIDIUM � ight system functional block diagram.

for attitude reference during initial Earth acquisition, storage or-bit, ascent, and deboost. Seven hydrazine reaction engine assem-blies (REAs) are used for attitude control during separation Earthcapture, MWA spin up, ascent, and deboost. An electrothermal hy-drazine thruster (EHT), which has a high speci� c impulse, is usedfor ascent and deboost. The REAs are also used for station keepingand as backup to the EHT for ascent/deboost.

Ascent, station keeping, and deboost can be performed autono-mouslyor in conjunctionwith groundcommands.The approachusesa method wherein each satellite is maintained within a prescribedbox with position within the box determined by the AOCS. Theephemeris of each satellite and its neighbors are propagated on-board and are updated by the ground approximately once a week.The positions and velocities of neighboring satellites are used forcrosslink antenna pointing.

The system is fully autonomous with ground control backup ca-pability after the satellite is placed in its � nal orbit and initiallyattitude stabilized. Position and attitude are available at 90-ms in-tervals with the following accuracy: attitude (C=¡ deg 3 sigma):0.2 roll, 0.3 pitch, 0.4 yaw and position (C=¡ km 3 sigma): 6 intrack, 5 cross track.

Propulsion SubsystemThe satellite uses a blowdown monopropellant hydrazine sub-

system to provide all propulsive functions. This subsystem appliesexternal torques and forces to the satellite to perform the functionsof orbit insertion, orbit adjustment, maintenance, reaction control,and deorbit. The principle components of the subsystem include apropellant storage pressure vessel, an EHT for major delta-velocitymaneuvers, standard hydrazine REAs for attitude control and sta-tion keeping, latching torque motor and manual valves, plumbing,and telemetry instrumentation to evaluate the in-� ight performanceof the subsystem.

The propulsionsubsystemincludesa propellantloadand thrusterscapable of providing the delta-velocityincrement required to insertthe spacecraft into the designatedorbital slot and effectivelydeorbitthe spacecraft at EOL. The on-orbit functions of the subsysteminclude altitude, mean longitude, and orbit inclination maintenancefor the duration of the operationallife. The on-orbit propellant loadhas been sized to provide a 60% consumable reserve for operationalorbit maintenance functions (8-year operationalorbit capability).

The hardware implementation of the subsystem utilizes one sur-face tensionpropellanttank manifolded to the thrusters,each havinga series redundant propellant control valve and its own thermal el-ements. Thrusters are located to provide limited redundancy. Thesurface-tension device in the fuel tank ensures gas-free propellantdelivery under all acceleration environments. Propellant and pres-surant are loaded into the tank through a pair of manual valves,each with triple-redundant seals. A single latching torque motorvalve prevents propellant from � owing into the delivery manifold

until commanded open after separation from the launch vehicle.Anin-line � lter rated at 10 ¹m absolute is located downstream fromthe propellant tank outlet to prevent any � ne contaminants fromreaching the thrusters. A pressure transducer and thermistors lo-cated on the propellant tank and lines provides health and status ofthe propulsion subsystem.

When the latchingvalve is openedafter launch, the pressurantgasin the propellant tank forces the propellant through the plumbing tothe thruster inlets. Propellant feed pressure slowly decaysas propel-lant is consumed (blowdown operation) and the volume occupiedby the pressurant gas in the propellant tank expands.

Because of the low thrust of the engines, large orbit adjust ma-neuvers such as the initial orbit raising and the deorbit maneuverare accomplished with numerous long-duration EHT � rings. TheREAs are operated in pulse mode for attitude control during theEHT � rings. Operationalorbit maintenance is performedwith shortpulses on an REA, principally for drag makeup.

Each thruster valve is controlled by a drive circuit in the inte-grated bus electronics (IBE), which provides the interface betweenthe propulsion subsystem and the main payload processor.A back-ground task runningon the main processorselects the proper valvesand timing for whatever function is commanded from the groundor by the onboard attitude control task. The status of the subsystemand maneuver bookkeeping is monitored via telemetry.

Telemetry, Tracking, and Control SubsystemThe telemetry, tracking, and control (TT&C) subsystemprovides

the functional hardware required for the reception, processing, andimplementation of command data, and the collection, storing, mul-tiplexing, and transmission of satellite telemetry data. The TT&Csubsystemoperateswith or without the simultaneousfunctioningofthe mission communications system and does not degrade or inter-fere with the mission communications operations. The spacecraftantenna arrangement and the communicationshardware con� gura-tion ensure that command and telemetry functions are accessibleduring all phases of the mission. Typical functions of the TT&Csubsystemare satellite electronicequipmentredundancyswitching,multiple-point system monitoring for health and status evaluation,control of system initializationand testing, and general housekeep-ing tasks.

When the satellite is serving as a node in the network, telemetrydownlink transmissions will utilize the 19.4–19.6 GHz band, andtelecommand uplink transmissions will utilize the 29.1–29.3 GHzband. TT&C data are multiplexed into the wideband 20=30-GHzsystemcontrol facilitygroundlink (usingcrosslinksas necessaryforsatellite-to-satelliterelay). During transfer and sparing orbit opera-tions, the TT&C capabilityis providedby transmissionover narrow-bandchannelsin the frequencybands justnoted.These narrow-bandchannels use circularly polarized omnidirectionalspacecraft anten-nas to permit direct communicationswith the system control facility

678 GARRISON ET AL.

Table 3 TT&C satellite transmission characteristics

Nonnetworked Networked

Frequency band 19:4¡19:6 GHz 19.4–19.6 GHzAntenna polarization LHCP LHCP (feeder link)Modulation BPSK/PM QPSKa

Data rate 1 kbps 3.125/12.5 Mbps(total feeder link/

crosslink data rate)aQuadrature phase-shift keying.

Table 4 Command signal characteristics

Nonnetworked Networked

Frequency band 29.1–29.3 GHz 29.1–29.3 GHzAntenna polarization RHCP RHCPModulation BPSK/PM QPSKa

Data rate 1 kbps 3.125 Mbps(total uplink data rate)

aQuadrature phase-shift keying.

regardlessof the satelliteattitude.This link supportscontrolanddatacommunication during postseparationmaneuvers, orbit raising andlowering procedures, and other times when the satellite is not in thenetwork. The narrow-band TT&C communication link signal mar-gins are adequate to permit a ground station antenna of moderatebeamwidth(approximately1 deg)to receiveand transmitTT&C linksignals. This meets the requirements of initial acquisition and re-acquisitionoperations.When the narrow-bandtransmissionmode isin operation,all of the TT&C dataand controlsignalsare transmittedin digital form at a rate of 1.0 kbps each. Large signal margins andbinary phase-shift keying/pulse modulation (BPSK/PM) are usedto minimize the effects of anticipated antenna pattern irregularitiesand grating lobes, which are characteristic of the satellite omnidi-rectionalantennas.Some of the attributesof the telemetry downlinkare summarized in Table 3.

The command subsystem is designed to maintain positive con-trol of the spacecraft during all mission phases. It provides reliablecontrol during launching maneuvers and for all satellite operatingattitudes. It also maintains the orbital velocity of the satellite andcontrols housekeeping functions and communications subsystemcon� gurations.The commandmessagesare authenticatedto providesecurity, protectingthe satellite control subsystemagainst unautho-rized access.

The command transmissions received from the ground are de-modulated into a digital bit stream. When the destination satellitereceives a command from the TT&C Earth control station, the com-mand is authenticated and, if successful, executed at the speci� edtime. The ground can then verify its execution based on telemetry.Some of the attributes of the command uplink are summarized inTable 4.

Operational Lifetime and Space Segment ReliabilityThe operational lifetime and reliability of each satellite is deter-

mined by a number of factors, including solar array degradation,station keeping fuel consumption,battery degradation,and randomparts failure. Redundancy is provided on critical hardware as deter-mined to be necessary through reliability analyses and predictionsto achieve a 0.58 probability of success for a 5-year mission.

System TradesThe systems engineering and constellation architecture trade

studies that determined the design of the IRIDIUM telecommunica-tions system were comprehensive and are continuouslyevaluated.

Complexity of PayloadThe complexity of the payload in the IRIDIUM SV is a result

of the need to minimize the processing burden placed on centralground stations and to provide the greatest � exibility (and hence thegreatest service) in the routing of communications traf� c. Onboardprocessingcan route voice and data in a dynamicenvironmentwith-out the overhead of continual ground intervention.Delays are min-imized for added voice quality. Much of the complexity is provided

by state-of-the-art,application-speci�c integrated circuits and real-time software, which reduces the hardware complexity.

Muchof this complexityis allowedbecausethedesignphilosophyprovidesredundancyat the systemlevel insteadof the hardwarecon-� guration item level. Autonomous operation and dynamic resourcemanagement and routing provide constellation failure mitigation.In effect the traditional hardware redundancy is spread over manyspacecraft.

Size/Weight of SVThe primary driver in sizing the SV is the payload. Whereas

communications electronics can be packaged in many forms, theantennasize (and technology)determinesthe formfactor,alongwiththe considerationof packingmultiplespacecraftinto existinglaunchvehicle shroud envelopes. Payload electronic demand determinesthe size of energy generation hardware (solar arrays) and energystorage hardware (batteries).

Once the overall target of 16-dB L-band link margin was vali-dated, the selectionof phasedarray technologymoved the IRIDIUMSV designfroma six-sidedform to the existingtriangularform, lead-ing to ease of deployment of the L-band panels and keeping a highvehicle packing density within the launch vehicle shrouds. Corre-sponding increases in electrical power needs are handled by takingadvantage of the large spacecraft surface area exposed, allowingmaximum solar array size without complex panel folding.

The aspect ratio of the design leads to the use of composite struc-tures to provide the required stiffness for the launch environment.Composite technology keeps the weight below that which oldermetal technology would have permitted.

Type of Commercial Launch VehiclesThe major concern with the deployment of a satellite constel-

lation with a signi� cant number of nodes is the availability andreliability of suf� cient launch system assets to guarantee the com-plete deployment within a short time frame. Among the technicalconsiderations are the number of vehicles required (dependent onthe size and number of payloads to be placed in orbit), the potentialfor adding custom attach and dispense hardware, and the overallreliability of the launch system. Economic and political considera-tions are extremely important if the number of required launches ishigh: availabilityof launchopportunities(manifest slots) and abilityto overcome export barriers if nondomestic carriers are anticipated.

For a commercial venture, economic considerationsall but elim-inate the expensive, massive vehicles currently available for gov-ernment and crewed operations.The selection of the Delta II, LongMarch IIC/SD and Proton launch systems recognizes the need todeploy signi� cant numbers of satellites using affordable, commer-cially available assets that guarantee the successful deployment ofthe IRIDIUM constellation.

Size of Individual BeamsFor a communicationssystem utilizingspace nodes, the most im-

portant design requirement is the rf link performance.Recognizingthat the service provision for subscriber units located anywhere onthe surface of the globe requires multiple beams (or cells), techni-cal parameters such as overall gain requirements, maximum allow-able gain variation,polarization,and uplinkequipmentperformancemust be determined.

Hardware tradeoffs must be made between available beam gen-eration technology(phased array or waveguide/re� ector implemen-tations) and usable SV resources (area, volume, and mass distribu-tion). Other factors involved are beam shape (hexagonal, elliptical,and complex), cooperative mechanisms with adjacent space nodes(means of transitioning communications services from one space-craft to the next as the constellationmoves with respect to the user),and, most unusually, ground coverage control (the need to limitdynamic geographical coverage to satisfy political and licensingrestrictions).

Space EnvironmentThe establishment of spacecraft constellation altitude is � rst

established using communications parameters [beam coverage,

GARRISON ET AL. 679

system gain, and gain to temperature ratio (G/T)] but the uniqueradiationenvironmentof spacemust beused at a minimumas a hard-ware design constraintand, possibly, as a major driver in system de-sign. The primary concern is the use of affordable, commercial-off-the-shelfelectroniccomponentswhile deliveringacceptablesystemavailabilityand reliability performance.The IRIDIUM system alti-tude is signi� cantlybelowthe � rstmajorpeak in chargedparticleen-vironment, allowing state-of-the-artcomponents to deliver the nec-essary communicationsperformance.At higher altitudes, either theoverall mass fraction of payload electronics would decline as moreandmore shieldingwould be necessary,or communicationsserviceswould be degraded to accommodateolder, more expensive,and lessef� cient semiconductortechnologieswithin thepayloadelectronics.

Size of Telephone (Handheld or Suitcase)The size of the subscriber unit for voice communications is dic-

tated by how much technology can be economically packaged intoan acceptable form factor. Acceptability is determined by the usermarket,with familiarityandeaseof use.The tradesfor the subscriberunit directly affect the physical and functional characteristicsof theIRIDIUM SV. Technically, the size is driven by antenna capabilityand battery capability. However, the economies of scale of produc-tion in� uencewhatmight seemanobviousoutcome.Althoughspacehardware demands premiums of reliabilityand compactness(small-est mass possible), the production of millions of hand-held unitsweighs highly in the choices of system design, balanced against 66active space vehicles.And although the space vehicle function is farbeyond the notice of the customer, the handheld subscriber unit isthe ultimate interface for the system.

Quality of Voice/DataQuality of voice and data communications is the ultimate driver

of the IRIDIUM system. Extensive analyses and testing were un-dertaken to fully characterize the expected link performance of thesystem. The bit error rates in this digital system were establishedto provide adequate customer perceptions (for voice traf� c) andacceptable throughput performance (for digital modes of commu-nication via several protocols). Applying these error rates to theexpected deep fading environments led to the system requirementof 16-dB L-band link margin over theoretical minimums. Delaysexpected due to spacecraft altitude and packet routing were shownto be acceptable.

Following the development of the derived link speci� cations,total system performance was modeled to demonstrate the inter-action of all parameters and to demonstrate the lack of systemself-interferenceand the ability to transport the required volume oftraf� c.

Number of Telephone CallsThe number of telephone calls supported by the IRIDIUM con-

stellation is determined by market selection and spectrum capabil-ity. Once this target is established, all design requirements are thentraded to provide adequate margin to this number while maintain-ing an economically viable system. Dependent design parametersincludepacketroutingpaths, spectrumreuse, and gatewaypartition-ing. Independentconsiderationsare the natureof human voice com-munications(activityfactor), externalinterferenceenvironment,andmarket expectationsof user pro� les and revenue capacity.

For the spacecraft, the number of telephone calls supported isdetermined by processing load and power availability.The payloadhardwareand softwarearedesignedfor the maximumpeak load everexpected during the life of the constellation.The spacecraft powersubsystem is designed to take advantageof the peak-to-averagena-ture of the user distribution (most users are in the northern hemi-sphere land areas, with little nighttime traf� c). By balancing thepeak power capabilities with a multiorbit recharge scheme for en-ergy balance, the mass of the spacecraft is minimal to support thetraf� c load.

Altitude of ConstellationThe communicationssystem impacts of constellationaltitude are

primarily rf physical considerations (system gain and G/T) and

quality of service (delay). For a � xed payload electronics design,higher altitudemeans larger coverageareas, requiringmore antennabeams of higher gains, lower link margins as interference environ-ments grow, and possibly higher (therefore, more costly) perfor-mance requirements on the subscriber units.

Secondary considerations for altitude are economics of insertion(too high of an altitude means very costly launch services and/orhigh mass penalties), ease and speed of maintenance (replacementof older spacecraft from wearout or decayed orbits), and most im-portantly ease of deorbit.Altitude-dependentenvironmentalfactors(radiation, atomic oxygen, drag, and debris) directly affect relia-bility and availability. Eclipse duration directly affects the powersubsystem design for energy storage.

Also important are collision probabilitiesas the quantity of otherconstellations of other spacecraft represent a growing threat to on-going service. However, the IRIDIUM program mandates a zerodebris policy and will deorbit its satellites, after life, to ensure aminimal probabilityof collisionwith any other satellitewhile in theoperational constellation.

Number of Orbits and Number of SatellitesThe number of orbits (orbital planes and vehicles per plane) in

the IRIDIUM constellation is selected in concert with the beamsize to optimize the coverage. At the equator, vehicle footprintsmust cover the greatest area, with the smallest amount of overlap.As the satellites move closer to the poles, overlap increases, andthe constellation begins to shut down selected cells on individualspacecraft to minimize interference.This has the additional bene� tof lowering overall spacecraft power demand. Too much overlapresults in inef� cient use of assets, whereas too little overlap leadsto potential gaps in worldwide coverage.

Special consideration must be given to seam coverage (seam isde� ned as the area where adjacent spacecraft orbits are movingin opposite directions). This condition arises from using all space-craft continuously (as opposed to only half the orbit, ascendingor descending). Ground coverage demands different orbital spac-ing (east/west), adjustments to spacecraft hardware for differentcrosslinktrackingrates,and accommodationsin groundpatternhan-dling within mission software.

An important parameter of the constellationde� nition is the costof � lling all orbital positions.Launch vehicle capacity, both for de-ploymentof initialvehiclesand for themaintenanceas spacecraftarereplaced, is extremely expensive.Sparing philosophy,reaction timefor spacecraft failure mitigation, and desirability for partial constel-lation functioning as population rises must be weighed against theeconomic realities of limited � exibility in today’s launch servicesmarket.

Inclination of OrbitsThe IRIDIUM constellation contains satellites in circular, near-

polar orbits. The exact selection of orbital inclination is based onthe need to maximize the stability of the orbits and to maximize theclosest approach distance as satellites near the poles.

Number of GatewaysThe selection of the number of gateways handling the commu-

nications connections in the IRIDIUM system are chosen for eco-nomic reasons (control of revenue collecting operations), politicalreasons(licensingprovisions,access to PSTNs, access control), andtechnical reasons.

The prime technical driver is system capacity. Fewer gatewaysmean that each gateway must service a greater portion of the callsmade within the system.Too few and the required feeder linksband-width increases, and traf� c on intersatellite links will be overbur-dened(fewer satelliteswill be involvedin the total feederconnectionfrom constellation to the ground). Too many and more feeder linksmust be placed on the satellite to accommodate a greater numberof simultaneous feeds, as more stations requiring service would bewithin view at any given time.

Although a uniformspacingof Earth stationswould make systemdesign easier, the gateway operationsmust be aligned with popula-tion concentrationsand political boundaries.

680 GARRISON ET AL.

Crosslink TypesThe IRIDIUM system uses intersatellitelinks to provide commu-

nicationsserviceswithoutaddedgrounddelaysandbottlenecks.Theselection of the physical link design is dependent on the communi-cations architecture, interference environment, orbital geometries,and system reliability constraints.

Because of the dynamic pointing environment of crosslinks andspacecraft attitude, the use of optical crosslinks would require ex-tensive development for accurate beam pointing, whereas standardrf links allow simpler, proven hardware implementations.

Type of ModulationIn a constellation of high inclination, circular orbits, the north/

south separation between spacecraft remains constant, whereas theeast/west separation is continually changing. The use of time-di-vision multiple access architectures in the rf crosslinks, along withspectrum limitations,providesa simple half-duplexlink design thatmeets all traf� c requirements.This architecturealso eases the prob-lem of system self-interference, inasmuch as multiple spacecraftwill be in view of crosslink antennas.

Prelaunch and Launch OperationsPrelaunch and launch operations are basically the same for all

launch vehicle suppliers. The system control segment (SCS) willcoordinate the overall launch plans and launch schedules with thelaunch segment. The launch vehicle is assembled and tested beforemating with the satellitedispenser.The satellitesare then matedwiththe dispenser and then with the launch vehicle.After all stages havebeen mated, the launch system will undergo � nal pre� ight integra-tion testing,fueling,and checkout.The SCS providesthe parametersthat specify the desired parking orbit based on the constellation � llplan and schedule.

The SV constellation will be boosted to an initial separation or-bit from three separate sites, which comprise the launch segment,in Russia, China, and the United States. Each launch scenario issimilar, but the exact timeline, separation sequence, and initial or-bital characteristics may differ for each launch vehicle. The SVsare powered off during launch with the exceptionof the IBE, whichprovides the necessary interfaces to detect the separationof the SVsfrom their dispenser. In general, the primary boostingstages deliveran SV dispenser (which includes a rocket motor and maneuveringsystem) to an initial parking orbit. The dispenser then maneuversinto a � nal separationorbit and adjusts its attitude for release of theSVs. Once released, all SV operations timelines for initializationand deployment activities are nearly identical.

The SCS will nominally acquire the satellite using predictedvec-tor data received from the launch provider in the prelaunch timeframe. The SCS can update the initial acquisition vectors at thetelemetry, tracking, and control (TTAC) stations based on updatedephemeris information.This can be in the form of an injection vec-tor or real-time feedback based on vehicle performance providedby the launch segment. In cases of nonnominal launch vehicle per-formance an antenna search pattern will be initiated at the TTAC toassist in SV acquisition.

Upon release from the launch vehicle dispenser, the SV sensesthat it has separatedfrom the dispenserand applies power to the SV.After separation,the SV will autonomouslybootup its computerandbegin initializationof onboardelectronicsand systems components.The attitude control electronics will be powered on along with thecomponentssupportinginitial attitudedeterminationand propulsionsystem operations. Heaters (tank, line, and catalyst bed) are turnedon in preparation for � rst use of the propulsion system and areallowed to warm up for a periodof 15 min. The SV will use thrustersto three-axisstabilizetheattituderatesand thenacquireandmaintaincoarse Earth nadir pointing.

Mechanism deployment begins soon after completion of Earthacquisition activities. Both solar arrays are released and rotated inazimuth to allow clearance for MMA release and deployment. Allthree MMAs are then released simultaneously, after which the so-lar arrays are positioned in elevation to optimize power. The feederlink antennas are then released and uncaged. During the early orbit

phase, the secondary link supports low-rate ground communica-tions, tracking, and command activities. The secondary link will beinitialized within the � rst orbit to support initial health and statuschecks postlaunch.

Prior to initiating an ascent or orbit raising maneuver, which isrequired to place the SVs in either the storage or mission orbit, aseries of ground communications passes will be required to collectranging data in support of orbit determinationactivities.During thecourse of rangingdata collection, the SCS will assess the health andstatus of the SV.

The target orbit, storage or mission, is preprogrammed for eachSV prior to launch but may be modi� ed by the SCS prior to schedul-ing the ascent maneuver. Once an orbit solution is achieved, anephemeris update will be uplinked to the SV. Upon receipt of theupdate, the SV will be commanded to generate an ascent maneu-ver plan, which is also sent to the SCS via telemetry. The SV willautonomously schedule the ascent maneuver based on the onboardplan. The ground may overridethe ascentmaneuver should the needarise.

The SV will autonomously control the SV for attitude and deltavelocity by initiating thruster � rings based on the schedule derivedfrom the maneuverplan, to both raise the orbit and adjust inclinationas necessary. During the course of the ascent, the SV will requireperiodic updates of ephemeris data, requiring ranging and orbit de-termination activities. During the ascent phase the SV will main-tain averagethree-axisattitudedeterminationand control within thespeci� cations.

ConclusionThese complex trades leading to the design of the IRIDIUM

space system have led to a tremendously robust telecommunica-tionsnetworkfor bringinginformationto anyone,at anyplace.Thesetrades were conducted over many years with detailed engineeringanalysis supporting the decisions. The systems engineering teamsin each segment of the system, supported by the diverse team-mates across the product teams, indeed looked at all options thatseemed reasonable, even if nontraditional. The number of satel-lites, the LEO altitude, and the complexity of the payload seemedcounterintuitiveuntil the factors of voice quality and customer sat-isfaction were included in the analysis. The systems architectureof the IRIDIUM telephoningconstellationemphasized the systemsengineering trades from the very beginning of the project. In fact,the concept and preliminary trades were accomplished within thesystems engineeringdepartment.The number and diversity of tradestudies accomplished prior to the launch of the � rst satellite are re-markable.The IRIDIUM system will indeed support total customersatisfaction.

AcknowledgmentsIRIDIUM® is a registeredtrademarkand service mark of Iridium,

L.L.C. The authors wish to acknowledge the speci� c contributionsof D. Miller, D. Hegel, and D. Pelletier of Space Systems Engi-neering in the creation of this paper and to the many individualsat Motorola who have participated in the IRIDIUM space systemdesign and development. The prime contractor for the SV is Mo-torola Satellite Communications Group. The SV is subdivided intoa bus, communications payload, and MMAs. The bus, provided byLockheed Martin, consists of the satellite structure as well as sub-systems for power generation and storage, attitude control, propul-sion, and thermal control. The communications payload, providedby Motorola, consists of the electronicsincluding the K-bandanten-nas requiredto perform the communicationsmissionof the satellite.The MMA, provided by Raytheon, generates and transmits as wellas receives the L-band beams containing the subscriber communi-cations traf� c. And in conclusion, Motorola performs the � nal SVintegration, test, and launch vehicle processing. The paper that thiswas based on, “IRIDIUM Constellation Dynamics—The SystemsEngineeringTrades,” was presentedat the IAF Congress,Oct. 1995(IAF Paper 95-U.2.04).

I. E. VasAssociate Editor