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These generate electrical power and chargeonboard batteries and communicationsantennas to receive commands and sendtelemetry data from the HST. Figure 5-1 showsthe HST configuration.

The Telescope performs much like a groundobservatory. The SSM is designed to supportfunctions required by any ground astronomicalobservatory. It provides power, points theTelescope, and communicates with the OTA, SIC&DH unit, and instruments to ready anobservation. Light from an observed targetpasses through the Telescope and into one ormore of the science instruments, where the lightis recorded. This information goes to onboardcomputers for processing, and then it is eithertemporarily stored or sent to Earth in real time,via the spacecraft communication system.

The Telescope completes one orbit every 97minutes and maintains its orbital position along

HUBBLE SPACE TELESCOPE SYSTEMS

Fig. 5-1 Hubble Space Telescope – exploded view

MAGNETIC TORQUER (4)

HIGH GAIN ANTENNA (2)

SUPPORT SYSTEMS MODULE FORWARD SHELL

OPTICAL TELESCOPE ASSEMBLYSECONDARY MIRROR ASSEMBLY

SECONDARY MIRROR BAFFLE

CENTRAL BAFFLE

OPTICAL TELESCOPE ASSEMBLYPRIMARY MIRROR AND MAIN RING

FINE GUIDANCE OPTICALCONTROL SENSOR (3)

OPTICAL TELESCOPEASSEMBLY FOCALPLANE STRUCTURE

AXIAL SCIENCEINSTRUMENTMODULE(4)

SUPPORT SYSTEMS MODULEAFT SHROUD

FIXED HEAD STAR TRACKER (3)AND RATE SENSOR UNIT (3)

RADIAL SCIENCEINSTRUMENT MODULE (1)

OPTICAL TELESCOPE ASSEMBLYEQUIPMENT SECTION

SUPPORT SYSTEMS MODULEEQUIPMENT SECTION

OPTICAL TELESCOPE ASSEMBLYMETERING TRUSS

MAIN BAFFLE SOLAR ARRAY (2)

APERTUREDOOR

LIGHT SHIELD

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T he Hubble Space Telescope (HST) hasthree interacting systems:

• The Support Systems Module (SSM), an outerstructure that houses the other systems andprovides services such as electrical power,data communications, and pointing controland maneuvering

• The Optical Telescope Assembly (OTA),which collects and concentrates the incominglight in the focal plane for use by the scienceinstruments

• Eight major science instruments, four housedin an aft section focal plane structure (FPS)and four placed along the circumference ofthe spacecraft. With the exception of the FineGuidance Sensors (FGS), the ScienceInstrument Control and Data Handling (SIC&DH) unit controls all.

Additional systems that also support HSToperations include two Solar Arrays (SA).

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three axial planes. The primary axis, V1, runsthrough the center of the Telescope. The othertwo axes parallel the SA masts (V2) and theHigh Gain Antenna (HGA) masts (V3) (seeFig. 5-2). The Telescope points and maneuversto new targets by rotating about its body axes.Pointing instruments use references to theseaxes to aim at a target in space, position the SA,or change Telescope orientation in orbit.

5.1 Support Systems Module

The design features of the SSM include:

• An outer structure of interlocking shells• Reaction wheels and magnetic torquers to

maneuver, orient, and attitude stabilize theTelescope

• Two SAs to generate electrical power• Communication antennas• A ring of Equipment Section bays that

contain electronic components, such asbatteries, and communications equipment.(Additional bays are provided on the +V3side of the spacecraft to house OTAelectronics as described in para 5.2.4.)

• Computers to operate the spacecraft systemsand handle data

• Reflective surfaces and heaters for thermalprotection

• Outer doors, latches, handrails, and footholdsdesigned for astronaut use during on-orbitmaintenance.

Figure 5-3 shows some of these features.

Major component subsystems of the SSM are:

• Structures and mechanisms• Instrumentation and communications• Data management• Pointing control• Electrical power• Thermal control• Safing (contingency) system.

5.1.1 Structures and MechanismsSubsystem

The outer structure of the SSM consists ofstacked cylinders, with the aperture door on topand the aft bulkhead at the bottom. Fittingtogether are the light shield, the forward shell,the SSM Equipment Section, and the aftshroud/bulkhead – all designed and built byLockheed Martin Missiles & Space (see Fig. 5-4).

Aperture Door. The aperture door,approximately 10-ft (3 m) in diameter, coversthe opening to the Telescope’s light shield. Thedoor is made from honeycombed aluminumsheets. The outside is covered with solar-reflecting material, and the inside is paintedblack to absorb stray light.

The door opens a maximum of 105 degrees fromthe closed position. The Telescope apertureallows for a 50-degree field of view (FOV)centered on the +V1 axis. Sun-avoidance sensorsprovide ample warning to automatically closethe door before sunlight can damage theTelescope’s optics. The door begins closingFig. 5-2 Hubble Space Telescope axes

+V3

–V2

+V3

–V1+V2

+V2

–V3

+V1

SOLAR ARRAY

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when the sun is within ±35 degrees of the +V1axis and is closed by the time the sun reaches20 degrees of +V1. This takes no more than 60seconds.

The Space Telescope Operations Control Center(STOCC) can override the protective door-closing mechanism for observations that fallwithin the 20-degree limit. An example isobserving a bright object, using the dark limb(edge) of the Moon to partially block the light.

Light Shield. The light shield (see Fig. 5-4)blocks out stray light. It connects to both theaperture door and the forward shell. On theouter skin of the Telescope on opposite sides arelatches to secure the SAs and HGAs when theyare stowed. Near the SA latches are scuff plates,

HIGH GAIN ANTENNA

FORWARDSHELL

CREWHANDRAILS

APERTUREDOOR

LIGHT SHIELD

MAGNETICTORQUERS

SOLAR ARRAY

EQUIPMENTSECTION

BATTERIESAND CHARGECONTROLLER

EQUIPMENT BAY

DIGITALINTERFACE UNIT

REACTIONWHEEL ASSEMBLY

COMMUNICATION SYSTEM

COMPUTER

LOW GAINANTENNA

UMBILICALINTERFACE

LATCH PINASSEMBLY

SUN SENSORS (3)

AFTSHROUD

ACCESS DOOR K70110-503

Fig. 5-3 Design features of Support Systems Module

APERTUREDOOR

LIGHTSHIELDMAGNETIC

TORQUER (4)

HIGH-GAIN ANTENNA (2)

FORWARD SHELL

EQUIPMENTSECTION

AFT SHROUD K70110-504

Fig. 5-4 Structural components of Support SystemsModule

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large protective metal plates on struts thatextend approximately 30 in. from the surface ofthe spacecraft. Trunnions lock the Telescope intothe Shuttle cargo bay by hooking to latches inthe bay. The light shield supports the forwardLow Gain Antenna (LGA) and its communi-cations waveguide, two magnetometers, andtwo sun sensors. Handrails encircle the lightshield, and built-in foot restraints support theastronauts working on the Telescope.

Figure 5-5 shows the aperture door and lightshield. The shield is 13-ft (4 m) long, with aninternal diameter of 10-ft (3 m). It is machinedfrom magnesium, with a stiffened, corrugated-skin barrel covered by a thermal blanket.Internally the shield has 10 light baffles, paintedflat black to suppress stray light.

Forward Shell. The forward shell, or centralsection of the structure, houses the OTA mainbaffle and the secondary mirror (see Fig. 5-6).When stowed, the SAs and HGAs are latchedflat against the forward shell and light shield.Four magnetic torquers are placed 90 degrees

apart around the circumference of the forwardshell. The outer skin has two grapple fixturesnext to the HGA drives, where the Shuttle’sRemote Manipulator System can attach to theTelescope. The forward shell also has handholds,footholds, and a trunnion, which is used to lockthe Telescope into the Shuttle cargo bay.

The forward shell is 13-ft (4 m) long and 10-ft(3 m) in diameter. It is machined from alumi-num plating, with external reinforcing rings andinternal stiffened panels. The rings are on theoutside to ensure clearance for the OTA inside.Thermal blankets cover the exterior.

Equipment Section. This section is a ring ofstorage bays encircling the SSM. It containsabout 90 percent of the electronic componentsthat run the spacecraft, including equipmentserviced during extravehicular activities (EVA)by Space Shuttle astronauts.

The Equipment Section is a doughnut-shapedbarrel that fits between the forward shell and aftshroud. This section contains 10 bays forequipment and two bays to support aft trunnionpins and scuff plates. As shown in Fig. 5-7,

APERTUREDOOR

APERTUREDOOR HINGE

LIGHT SHIELD

MAGNESIUMMONOCOQUESKIN

SOLAR ARRAYLATCHES

HIGH-GAIN ANTENNAAND SOLAR ARRAYLATCH SUPPORT RING

LIGHT SHIELDASSEMBLY/FORWARD SHELLATTACH RING

INTERNALBAFFLE RINGS

HIGH GAINANTENNA LATCHES

INTEGRALLYSTIFFENED SKIN

APERTURE DOORHINGE SUPPORT

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Fig. 5-5 Aperture door and light shield

STOWEDHIGH GAINANTENNAS

MAGNETICTORQUERS

CREW AIDS

STOWEDSOLAR ARRAYS

EQUIPMENT SECTION/FORWARD SHELLINTERFACE RING

INTEGRALLYSTIFFENEDSKIN PANELS

FORWARD SHELLREINFORCINGRINGS

HIGH GAINANTENNA MAST

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Fig. 5-6 Support Systems Module forward shell

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Fig

. 5-7

Sup

port

Sys

tem

s M

odul

e E

quip

men

t Sec

tion

bays

and

con

tent

s

1 2 3 4 5

8 9 106

1

3

2

–V2

10 –V2

10

9

87

+V

3

–V3

3

4

–V2

4

+V

2

6

57

K70

110-

507

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Fig. 5-8 Support Systems Module aft shroudand bulkhead

going clockwise from the +V3 (top) position, thebays contain:

1. Bay 8 – pointing control hardware2. Bay 9 – Reaction Wheel Assembly (RWA)3. Bay 10 – SI C&DH unit4. Unnumbered trunnion support bay5. Bay 1 – data management hardware6. Bay 2 through Bay 4 – electrical power

equipment7. Unnumbered trunnion support bay8. Bay 5 – communication hardware9. Bay 6 – RWA10. Bay 7 – mechanism control hardware.

The cross section of the bays is shaped like atrapezoid, with the outer diameter (the door) –3.6-ft (1 m) – greater than the inner diameter –2.6-ft (0.78 m). The bays are 4-ft (1.2 m) wideand 5-ft (1.5 m) deep. The Equipment Section isconstructed of machined and stiffenedaluminum frame panels attached to an inneraluminum barrel. Eight bays have flathoneycombed aluminum doors mounted withequipment. In Bays 6 and 9, thermal-stiffenedpanel doors cover the reaction wheels. Aforward frame panel and aft bulkhead enclosethe SSM Equipment Section. Six mounts on theinside of the bulkhead hold the OTA.

Aft Shroud and Bulkhead. The aft shroud (seeFig. 5-8) houses the FPS containing the axialscience instruments. It is also the location of theCorrective Optics Space Telescope AxialReplacement (COSTAR) unit.

The three FGSs and the Wide Field and PlanetaryCamera 2 (WFPC2) are housed radially near theconnecting point between the aft shroud andSSM Equipment Section. Doors on the outside ofthe shroud allow shuttle astronauts to removeand change equipment and instruments easily.Handrails and foot restraints for the crew runalong the length and circumference of the

shroud. During maintenance or removal of aninstrument, interior lights illuminate thecompartments containing the scienceinstruments. The shroud is made of aluminum,with a stiffened skin, internal panels andreinforcing rings, and 16 external and internallongeron bars for support. It is 11.5-ft (3.5 m) longand 14-ft (4.3 m) in diameter.

The aft bulkhead contains the umbilicalconnections between the Telescope and theshuttle, used during launch/deployment andon-orbit maintenance. The rear LGA attaches tothe bulkhead, which is made of 2-in.-thickhoneycombed aluminum panels and has threeradial aluminum support beams.

The shroud and bulkhead support a gas purgesystem that was used to prevent contaminationof the science instruments before launch. Allvents used to expel gases are light tight. Thus,stray light is prevented from entering the OTAfocal plane.

Mechanisms. Along the SSM structure aremechanisms that perform various functions,including:

EQUIPMENTSECTION/AFT SHROUDINTERFACERING

CREW AIDS

INTEGRALLY STIFFENEDSKIN PANELS

LOW GAIN ANTENNA

HONEYCOMB LAMINATE PANELS

FLIGHT SUPPORT SYSTEMPIN SUPPORT BEAM

FLIGHT SUPPORT SYSTEM PINVENTS

ELECTRICALUMBILICALS

REINFORCING RINGS

+V2 +V1

+V3

RADIAL SCIENCE INSTRUMENT DOORS

AXIAL SCIENCE INSTRUMENT DOORS

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• Latches to hold antennas and SAs• Hinge drives to open the aperture door and

erect arrays and antennas• Gimbals to move the HGA dishes• Motors to power the hinges and latches and

to rotate arrays and antennas.

There are nine latches: four for antennas, fourfor arrays, and one for the aperture door. Theylatch and release using four-bar linkages and aredriven by stepper motors called Rotary DriveActuators (RDA).

There are three hinge drives, one for each HGAand one for the door. The hinges also use anRDA. Both hinges and latches have hex-wrenchfittings so an astronaut can manually operatethe mechanism to deploy the door, antenna, orarray if a motor fails.

5.1.2 Instrumentation andCommunications Subsystem

This subsystem provides the communicationsloop between the Telescope and the Trackingand Data Relay Satellites (TDRS), receivingcommands and sending data through the HGAsand LGAs. All information is passed throughthe Data Management Subsystem (DMS).

The HGAs achieve a much higher RF signalgain, which is required, for example, whentransmitting high-data-rate scientific data. Theseantennas require pointing at the TDRSs becauseof their characteristically narrow beam widths.On the other hand, the LGAs provide sphericalcoverage (omnidirectional) but have a muchlower signal gain. The LGAs are used for low-rate-data transmission and all commanding ofthe Telescope.

S-Band Single Access Transmitter (SSAT).The Telescope is equipped with two SSATs.“S-Band” identifies the frequency at which the

science data is transmitted, and “Single Access”specifies the type of antenna on the TDRSsatellite to which the data is sent.

High Gain Antennas. Each HGA is a parabolicreflector (dish) mounted on a mast with a two-axis gimbal mechanism and electronics to rotateit 100 degrees in either direction (see Fig. 5-9).General Electric designed and made the antennadishes. They are manufactured fromhoneycomb aluminum and graphite-epoxyfacesheets.

Each antenna can be aimed with a one-degreepointing accuracy. This accuracy is consistentwith the overall antenna beam width of overfour degrees. The antennas transmit over twofrequencies: 2255.5 MHz or 2287.5 MHz (plus orminus 10 MHz).

Low Gain Antennas. The LGAs receive groundcommands and transmit engineering data. Theyare set 180 degrees apart on the light shield andaft bulkhead of the spacecraft. Each antenna isa spiral cone that can operate over a frequencyrange from 2100 MHz to 2300 MHz.Manufactured by Lockheed Martin, the LGAs

Fig. 5-9 High Gain Antenna

HIGH GAINANTENNA

TWIN-AXISGIMBALS

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are used for all commanding of the Telescopeand for low-data-rate telemetry, particularlyduring Telescope deployment or retrieval onorbit, or during safemode operations.

5.1.3 Data ManagementSubsystem

The DMS receives communicationscommands from the STOCC and data fromthe SSM systems, OTA, and scienceinstruments. It processes, stores, and sendsthe information as requested. Subsystemcomponents are:

• DF-224 computer (will be replaced on SM3Awith the Advanced Computer)

• Data Management Unit (DMU)• Four Data Interface Units (DIU)• Three engineering/science data recorders• Two oscillators (clocks).

The components are located in the SSMEquipment Section, except for one DIU storedin the OTA Equipment Section.

The DMS receives, processes, and transmits fivetypes of signals:

1. Ground commands sent to the HST systems2. Onboard computer-generated or computer-

stored commands3. Scientific data from the SI C&DH unit4. Telescope engineering status data for

telemetry5. System outputs, such as clock signals and

safemode signals.

Figure 5-10 is the subsystem functionaldiagram.

DF-224 Computer. The DF-224 computer is ageneral-purpose digital computer for onboard

Fig. 5-10 Data Management Subsystem functional block diagram

5-9K9322-05M

engineering computations. It executes storedcommands; formats status data (telemetry);performs all Pointing Control Subsystem (PCS)computations to maneuver, point, and attitudestabilize the Telescope; generates onboardcommands to orient the SAs toward the Sun;evaluates the health status of the Telescopesystems; and commands the HGAs. TheAdvanced Computer will replace the DF-224 onSM3A and will assume its functions.

Advanced Computer. The Advanced Computeris based on the Intel 80486 microchip. It operates20 times faster and has six times as muchmemory as the DF-224.

The Advanced Computer was designed usingcommercially developed components. A batteryof mechanical, electrical, radiation and thermaltests were performed at GSFC to assure itssurvival in the space environment. A successfulflight test of the hardware was carried outaboard the space shuttle Discovery on STS-95 inOctober 1998.

The Advanced Computer is configured as threeindependent single-board computers. Eachsingle-board computer (SBC) has twomegabytes of fast static random access memoryand one megabyte of non-volatile memory.

The Advanced Computer communicates withthe HST by using the direct memory accesscapability on each SBC through the DataManagement Unit (DMU). Only one SBC maycontrol the Telescope at a time. The other SBCscan be off, in an idle state, or performing internaltasks.

Upon power on, each SBC runs a built-in self-test and then copies the operating software fromslower non-volatile memory to faster randomaccess memory. The self-test is capable ofdiagnosing any problems with the Advanced

Computer and reporting them to the ground.Fast static random access memory is used in theAdvanced Computer to eliminate wait statesand allow it to run at its full-rated speed.

The Advanced Computer measures 18.8 x 18 x 13inches (0.48 x 0.46 x 0.33 m) and weighs 70.5 lb(32 kg). It will be located in Bay 1 of the SSMEquipment Section (see Fig. 5-11).

Data Management Unit. The DMU links withthe computer. It encodes data and sendsmessages to selected Telescope units and allDMS units, powers the oscillators, and is thecentral timing source. The DMU also receivesand decodes all incoming commands, thentransmits each processed command to beexecuted.

The DMU receives science data from the SIC&DH unit. Engineering data, consisting ofsensor and hardware status readings (such astemperature or voltages), comes from eachTelescope subsystem. The data can be stored inthe onboard data recorders if direct telemetryvia a TDRS is unavailable.

Fig. 5-11 Advanced computer

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The DMU is an assembly of printed-circuitboards, interconnected through a backplate andexternal connectors. The unit weighs 83 lb (37.7kg), measures 26 x 30 x 7 in. (60 x 70 x 17 cm),and is attached to the door of EquipmentSection Bay 1 (see Fig. 5-12).

Data Interface Unit. The four DIUs provide acommand and data link between DMS andother Telescope electronic boxes. The DIUsreceive commands and data requests from theDMU and pass data or status information backto the DMU. The OTA DIU is located in the OTAEquipment Section; the other units are in Bays3, 7, and 10 of the SSM Equipment Section. Asa safeguard, each DIU is two complete units inone; either part can handle the unit’s functions.Each DIU measures 15 x 16 x 7 in. (38 x 41 x 18cm) and weighs 35 lb (16 kg).

Engineering/Science Data Recorders. The DMSincludes three data recorders that store

engineering or science data that cannot betransmitted to the ground in real time. Therecorders, which are located in EquipmentSection Bays 5 and 8, hold up to 12 billion bitsof information. Two recorders are used innormal operations; the third is a backup. Eachrecorder measures 12 x 9 x 7 in. (30 x 23 x 18cm) and weighs 20 lb (9 kg).

Solid State Recorder. During SM2 a Solid StateRecorder (SSR) was installed, replacing a reel-to-reel tape recorder. A second state-of-the-artSSR will be installed during SM3A. This digitalrecorder will replace one of the two remainingreel-to-reel recorders on HST. The Solid StateRecorders have an expected on-orbit life of atleast eight years. They can record two datastreams simultaneously, allowing both scienceand engineering data to be captured on a singlerecorder. In addition, data can be recorded andplayed back at the same time.

The SSR has no reels or tape, and no movingparts to wear out and limit lifetime. Data isstored digitally in computer-like memory chipsuntil HST’s operators at GSFC command theSSR to play it back. Although they are the samesize as the reel-to-reel recorders, the SSRs canstore over 10 times more data — 12 gigabitsversus only 1.2 gigabits for the tape recordersthey replace.

Oscillator. The oscillator provides a highlystable central timing pulse required by theTelescope. It has a cylindrical housing 4 in. (10 cm)in diameter and 9 in. (23 cm) long and weighs3 lb (1.4 kg). The oscillator and a backup aremounted in Bay 2 of the SSM EquipmentSection.

5.1.4 Pointing Control Subsystem

A unique PCS maintains Telescope pointingstability and aligns the spacecraft to point to

K70110-512

POWER SUPPLY NO. 2(CDI DC-DCCONVERTER)

TOP COVER(WITH COMPRESSION PAD)

PC BOARDSTANDARD(5 X 7 in.)

CARDGUIDE

ENCLOSURE(DIP-BRAZEDAL.)

POWER SUPPLYNO. 1(DMU DC-DCCONVERTER)

BOTTOM COVER

COAX CONNECTOR

INTERFACECONNECTOR

MATRIXCONNECTOR

HEATSHIELDPARTITION(TYP)

Fig. 5-12 Data Management Unit configuration

5-11K9322-05M

and remain locked on any target. The PCS isdesigned for pointing to within 0.01 arcsec andholding the Telescope in that orientation with0.007-arcsec stability for up to 24 hours whilethe Telescope continues to orbit the Earth at17,500 mph. If the Telescope were in LosAngeles, it could hold a beam of light on a dimein San Francisco without the beam strayingfrom the coin’s diameter.

Nominally, the PCS maintains the Telescope’sprecision attitude by locating guide stars intotwo FGSs and controlling the Telescope to keepit in the same position relative to these stars.When specific target requests requirerepositioning the spacecraft, the pointingsystem selects different reference guide starsand moves the Telescope into a new attitude.

The PCS encompasses the Advanced Computer,various attitude sensors, and two types ofdevices, called actuators, to move the spacecraft(see Fig. 5-13). It also includes the Pointing/Safemode Electronics Assembly (PSEA) and theRetrieval Mode Gyro Assembly (RMGA); bothused by the spacecraft safemode system. Seepara 5.1.7 for details.

Sensors. The five types of sensors used by thePCS are the Coarse Sun Sensors (CSS), theMagnetic Sensing System (MSS), the Rate GyroAssemblies (RGA), the Fixed Head Star Trackers(FHST), and the FGSs.

The CSSs measure the Telescope’s orientation tothe Sun. They are used to calculate the initialdeployment orientation of the Telescope,determine when to begin closing the aperturedoor, and point the Telescope in special sun-orientation modes during contingencyoperations. Five CSSs are located on the lightshield and aft shroud. CSSs also provide signalsto the PSEA, located in Bay 8 of the SSMEquipment Section.

The MSS measures the Telescope’s orientationrelative to Earth’s magnetic field. The systemconsists of magnetometers and dedicatedelectronic units that send data to the AdvancedComputer and the Safemode ElectronicAssembly. Two systems are provided. Both arelocated on the front end of the light shield.

Three RGAs are provided on the Telescope.Each assembly consists of a Rate Sensor Unit(RSU) and an Electronics Control Unit (ECU).An RSU contains two rate-sensing gyroscopes,each measuring attitude rate motion about itssensitive axis. This output is processed by itsdedicated electronics, which are contained inthe ECU. Each unit has two sets of electronics.The RSUs are located behind the SSM EquipmentSection, next to the FHSTs in the aft shroud. TheECUs are located inside Bay 10 of the SSMEquipment Section. The RGAs provide input tothe PCS to control the orientation of the Tele-scope’s line of sight and to provide the attitudereference when maneuvering the Telescope.

Four of the original six rate gyros were replacedduring the First Servicing Mission. All six rategyros are planned for replacement duringSM3A. Three of six gyroscopes are required tocontinue the Telescope science mission.

An FHST is an electro-optical detector thatlocates and tracks a specific star within its FOV.Three FHSTs are located in the aft shroudbehind the FPS, next to the RSUs. STOCC usesstar trackers as an attitude calibration devicewhen the Telescope maneuvers into its initialorientation. The trackers also calculate attitudeinformation before and after maneuvers to helpthe FGS lock onto guide stars.

Three FGSs, discussed in more detail in para 5.3,provide angular position with respect to thestars. Their precise fine-pointing adjustments,accurate to within a fraction of an arcsecond,

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K70110-513

MAGNETICTORQUERS (4)

MAGNETOMETER (2)

COARSE SUN SENSORS (2)

REACTION WHEELS (4)

COMPUTER

EQUIPMENTSECTION

SCIENCEINSTRUMENTS

FIXED HEADSTAR TRACKER (3)

COARSE SUNSENSORS (2)

RATE SENSORUNIT (3)

FINE GUIDANCESENSOR (3)

BAY 8 – POINTING CONTROL AND INSTRUMENTS– RETRIEVAL MODE GYRO ASSEMBLY– POINTING AND SAFEMODE

ELECTRONIC ASSEMBLY

BAY 9 – REACTION WHEEL (2)NO. 3 FORWARDNO. 4 AFT

BAY 10 – SCIENCEINSTRUMENTCONTROLAND DATAHANDLING

BAY 1 – DATA MANAGEMENTCOMPUTER

BAY 7 – MECHANISM CONTROL

BAY 6 – REACTION WHEELREACTION WHEEL ASSEMBLY (2)NO. 1 FORWARDNO. 2 AFT

+ V2

5

6

7

+ V38

9

10

– V2

1

– V33

4

RWA

RWA

LOOKING FORWARD2

Fig. 5-13 Location of Pointing Control Subsystem equipment

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pinpoint the guide stars. Two of the FGSsperform guide-star pointing, while the third isavailable for astrometry, the positionalmeasurement of specific stars.

Pointing Control Subsystem Software. PCSsoftware accounts for a large percentage of theflight code executed by the Hubble’s maincomputer. This software translates groundtargeting commands into reaction wheel torqueprofiles that reorient the spacecraft. All motionof the spacecraft is smoothed to minimize jitterduring data collection. The software alsodetermines Telescope orientation, or attitude,from FHST or FGS data and commands themagnetic torquer bars so that reaction wheelspeeds are always minimized. In addition, thesoftware provides various telemetry formats.

Since the Telescope was launched, majormodifications have been made to the PCS. Adigital filtering scheme, known as Solar ArrayGain Augmentation (SAGA) was incorporatedto mitigate the effect of any SA vibration or jitteron pointing stability. Software also was used toimprove FGS performance when the Telescopeis subjected to the same disturbances. Thisalgorithm is referred to as the FGS Re-CenteringAlgorithm.

Software is used extensively to increase Telescoperobustness when hardware failures areexperienced. Two additional software safemodeshave been provided. The spin-stabilized modeprovides pointing of the Telescope -V1 axis to theSun with only two of the four RWAs operating.The other mode allows Sun pointing of theTelescope without any input from the RGA;magnetometer and CSS data is used to derive allreference information needed to maintain Sunpointing (+V3 and -V1 are options).

A further software change “refreshes” the FGSconfiguration. This is achieved by maintaining

data in the Advanced Computer memory so itcan be sent periodically to the FGS electronics,which are subject to single-event upsets (logicstate change) when transitioning through theSouth Atlantic Anomaly.

Actuators. The PCS has two types of actuators:RWAs and magnetic torquers. Actuators movethe spacecraft into commanded attitudes andprovide required control torques to stabilize theTelescope’s line of sight.

The reaction wheels work by rotating a largeflywheel up to 3000 rpm or braking it toexchange momentum with the spacecraft. Thewheel axes are oriented so that the Telescopecan provide science with only three wheelsoperating. Wheel assemblies are paired, twoeach in Bays 6 and 9 of the SSM EquipmentSection. Each wheel is 23 in. (59 cm) in diameterand weighs about 100 lb (45 kg). Figure 5-14shows the RWA configuration.

Magnetic torquers create torque on the space-craft and are primarily used to manage reactionwheel speed. The torquers react against Earth’smagnetic field. The torque reaction occurs in thedirection that reduces the reaction wheel speed,managing the angular momentum.

The magnetic torquers also provide backupcontrol to stabilize the Telescope’s orbitalattitude during the contingency modes, asdescribed in para 5.1.2. Each torquer, locatedexternally on the forward shell of the SSM, is 8.3

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RWA RWA

Fig. 5-14 Reaction Wheel Assembly

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ft (2.5 m) long and 3 in. (8 cm) in circumferenceand weighs 100 lb (45 kg).

Pointing Control Operation. To point precisely,the PCS uses the gyroscopes, reaction wheels,magnetic torquers, star trackers, and FGSs. TheFGSs provide the precision reference point fromwhich the Telescope can begin repositioning.Flight software commands the reaction wheelsto spin, accelerating or decelerating as requiredto rotate the Telescope toward a new target. Rategyroscopes sense the Telescope’s angularmotion and provide a short-term attitudereference to assist fine pointing and spacecraftmaneuvers. The magnetic torquers reducereaction wheel speed.

As the Telescope nears the target area, startrackers locate preselected reference stars thatstand out brightly in that region of the sky. Oncethe star trackers reduce the attitude error below60 arcsec, the two FGSs take over the pointingduties. Working with the gyroscopes, the FGSsmake possible pointing the Telescope to within0.01 arcsec of the target. The PCS can maintainthis position, wavering no more than 0.005arcsec, for up to 24 hours to guarantee faint-object observation.

5.1.5 Electrical Power Subsystem

Power for the Telescope and scienceinstruments comes from the Electrical PowerSubsystem (EPS). The major components aretwo SA wings and their electronics, sixbatteries, six Charge Current Controllers(CCC), one Power Control Unit (PCU), andfour Power Distribution Units (PDU). Allexcept the SAs are located in the bays aroundthe SSM Equipment Section.

During the servicing mission, the Shuttle willprovide the electrical power. After deployment,the SAs again begin converting solar radiation

into electricity. Energy will be stored in nickel-hydrogen (NiH2) batteries and distributed by thePCUs and PDUs to all Telescope components asshown in Fig. 5-15. The Telescope will not bereleased until the batteries are fully charged.

Solar Arrays. The SA panels, discussed later inthis section, are the primary source of electricalpower. Each array wing has a solar cell blanketthat converts solar energy into electrical energy.Electricity produced by the solar cells chargesthe Telescope batteries.

Each array wing has associated electronics.These consist of a Solar Array Drive Electronics(SADE) unit, which transmits positioningcommands to the wing assembly; a DeploymentControl Electronics Unit, which controls thedrive motors extending and retracting thewings; and diode networks to direct theelectrical current flow.

Batteries and Charge Current Controllers.Developed for the 1990 deployment mission, theTelescope’s batteries were NASA’s first flightNiH2 batteries. They provide the observatorywith a robust, long-life electrical energy storagesystem.

Six NiH2 batteries support the Telescope’selectrical power needs during three periods:when demand exceeds SA capability, whenthe Telescope is in Earth’s shadow, and duringsafemode entry. The batteries reside in SSMEquipment Section Bays 2 and 3. These unitshave extensive safety and handling provisionsto protect the Shuttle and its astronauts. The de-sign and operation of these batteries, along withspecial nondestructive inspection of each cell,have allowed these units to be “astronaut-rated”for replacement during a servicing mission. Tocompensate for the effects of battery aging,SM3A astronauts will install a Voltage/Tem-perature Improvement Kit (VIK) on each of

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Hubble’s six batteries. The VIK provides batterythermal stability by precluding battery over-charge when the HST enters safemode, effec-tively lowering the Charge Current Controller(CCC) recharge current.

Each battery consists of 22 cells in series alongwith heaters, heater controllers, pressuremeasurement transducers and electronics, andtemperature-measuring devices and theirassociated electronics. Three batteries arepackaged into a module measuring roughly 36by 36 by 10 in. (90 x 90 x 25 cm) and weighingabout 475 lb (214 kg). Each module is equippedwith two large yellow handles that astronautsuse to maneuver the module in and out of theTelescope in space.

The SAs recharge the batteries every orbitfollowing eclipse (the time in the Earth’sshadow). The recharge current is controlled bythe CCCs. Each battery has its own CCC thatuses voltage-temperature measurements tocontrol battery recharge.

Fully charged, each battery contains more than75 amp-hours. This is sufficient energy to sustainthe Telescope in normal science operations modefor 7.5 hours or five orbits. The batteries providean adequate energy reserve for all possiblesafemode contingencies and all enhancementsprogrammed into the Telescope since launch.

Power Control and Distribution Units. ThePCU interconnects and switches current flowingamong the SAs, batteries, and CCCs. Located inBay 4 of the Equipment Section, the PCUprovides the main power bus to the four PDUs.The PCU weighs 120 lb (55 kg) and measures43 x 12 x 8 in. (109 x 30 x 20 cm).

Four PDUs, located on the inside of the door toBay 4, contain the power buses, switches, fuses,and monitoring devices for electrical power dis-tribution to the rest of the Telescope. Two busesare dedicated to the OTA, science instruments,and SI C&DH; two supply the SSM. Each PDUmeasures 10 x 5 x 18 in. (25 x 12.5 x 45 cm) andweighs 25 lb (11 kg).

Fig. 5-15 Electrical Power Subsystem functional block diagram

K70110-515

SOLAR ARRAYELECTRONICS

CONTROLASSEMBLY

POWER

SOLAR ARRAYMECHANISMS

SO

LAR

AR

RA

YS

OLA

R A

RR

AY

POWER

COMMANDS

TELEMETRY

POWERPOWER

POWER

POWERPOWER

POWER

POWER

POWER

POWER

POWER

POWERCONTROL

UNITCOMMANDS

TELEMETRY

BATTERYNO. 1

23

45

6

PWRDISTRUNIT 1

23

4

SSMSASIs

OTA

SIC&DH

TELEMETRY

ORBITER POWER(PREDEPLOYMENT)

COMMANDS

POWER

COMMANDS

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5.1.6 Thermal Control

Multilayer insulation (MLI) covers 80 percent ofthe Telescope’s exterior, and supplementalelectric heaters maintain its temperatures withinsafe limits. The insulation blankets are 15 layersof aluminized Kapton, with an outer layer ofaluminized Teflon flexible optical solar reflector(FOSR). Aluminized or silvered flexible reflectortape covers most of the remaining exterior.These coverings protect against the cold ofspace and reflect solar heat. In addition,reflective or absorptive paints are used.

The SSM Thermal Control Subsystem (TCS)maintains temperatures within set limits for thecomponents mounted in the Equipment Sectionand structures interfacing with the OTA andscience instruments. The TCS maintains safecomponent temperatures even for worst-caseconditions such as environmental fluctuations,passage from “cold” Earth shadow to “hot”solar exposure during each orbit, and heatgenerated from equipment operation.

Specific thermal-protection features of the SSMinclude:

• MLI thermal blankets for the light shield andforward shell

• Aluminum FOSR tape on the aperture doorsurface facing the sun

• Specific patterns of FOSR and MLI blanketson the exteriors of the Equipment Section baydoors, with internal MLI blankets on thebulkheads to maintain thermal balancebetween bays

• Efficient placement of equipment and use ofequipment bay space to match temperaturerequirements, such as placing heat-dissipat-ing equipment on the side of the EquipmentSection mostly exposed to orbit shadow

• Silvered FOSR tape on the aft shroud and aftbulkhead exteriors

• Radiation shields inside the aft shroud doorsand MLI blankets on the aft bulkhead and shroudinteriors to protect the science instruments

• More than 200 temperature sensors and ther-mistors placed throughout the SSM, exter-nally and internally, to monitor individualcomponents and control heater operations.

Figure 5-16 shows the location and type ofthermal protection used on the SSM. SM2observations identified degradations of all of theMLI. Additional material will be installedduring SM3A to cover some of the degradedmaterial and restore the external layer surfaceproperties. The additional material has beenlife-tested to an equivalent of 10 years.

The layer being added to the SSM EquipmentSection is a composite-coated (silicone dioxide)stainless steel layer, known as the New OuterBlanket Layer (NOBL). The light shield/forward shell material is Teflon with a scrimbacking for durability.

5.1.7 Safing (Contingency)System

Overlapping or redundant Telescope equipmentsafeguards against any breakdown.Nonetheless, a contingency or Safing Systemexists for emergency operations. It uses manypointing control and data managementcomponents as well as dedicated PSEAhardware. This system maintains stableTelescope attitude, moves the SAs for maximumSun exposure, and conserves electrical power byminimizing power drain. The Safing System canoperate the spacecraft indefinitely with nocommunications link to ground control.

During scientific observations (normal mode),the Safing System is relegated to monitorautomatically Telescope onboard functions. Thesystem sends Advanced-Computer-generated

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SILVERIZED FOSR90-DEGMULTILAYERINSULATION(MLI)

90-DEG MLI

MLIMLI

MLI

MLIMLIFOSR

LIGHTSHIELD

FORWARDSHELL

ALUMINIZEDFOSR

ALUMINIZED FOSRSILVERIZED

FOSR

AFT SHROUDEQUIPMENTSECTION

OPTICALBLACK

+V1

BLACKCHEMGLAZE

APERTUREDOOR

ALUMINIZED FLEXIBLE OPTICAL SOLAR REFLECTOR (FOSR)

malfunction before any science operations ornormal functions can be resumed.

Since deployment of the Telescope in 1990, theSafing System has seen additionalimprovements to increase its robustness tosurvive hardware failures and still protect theTelescope. Paragraph 5.1.4 describes thesefeatures.

For the modes described above, the SafingSystem operates through computer software. Ifconditions worsen, the system turns overcontrol to the PSEA in Hardware Sun PointMode. Problems that could provoke this actioninclude any of the following:

• Computer malfunction• Batteries losing more than 50 percent of their

charge• Two of the three RGAs failing• DMS failing.

If these conditions occur, the AdvancedComputer stops sending keep-alive signals. Thisis the “handshake” mechanism between theflight software and the PSEA.

In the Hardware Sun Point Mode, the PSEAcomputer commands the Telescope and turns

“keep-alive” signals to the PSEA that indicateall Telescope systems are functioning. Entry intothe Safemode is autonomous once a failure isdetected.

The Safing System is designed to follow aprogression of contingency operating modes,depending on the situation aboard theTelescope. If a malfunction occurs and does notthreaten the Telescope’s survival, the SafingSystem moves into a Software Inertial HoldMode. This mode holds the Telescope in the lastposition commanded. If a maneuver is inprogress, the Safing System completes themaneuver, then holds the Telescope in thatposition, suspending all science operations.Only ground control can return to scienceoperations from Safemode.

If the system detects a marginal electrical powerproblem, or if an internal PCS safety check fails,the Telescope enters the Software Sun PointMode. The Safing System maneuvers theTelescope so the SAs point toward the Sun tocontinuously generate solar power. Telescopeequipment is maintained within operatingtemperatures and above survival temperatures,anticipating a return to normal operations. TheSTOCC must intercede to correct the

Fig. 5-16 Placement of thermal protection on Support Systems Module

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off selected equipment to conserve power.Components shut down include theAdvanced Computer and, within two hours,the SI C&DH. Before this, a payload(instruments) safing sequence begins and, ifit has not already done so, the Telescope turnsthe SAs toward the Sun, guided by the CSSs.The PSEA removes operating power fromequipment not required for Telescopesurvival.

Once ground control is alerted to a problem,NASA management of the STOCC convenes afailure analysis team to evaluate the problemand seek the best and safest corrective actionwhile the Safing System maintains control of theTelescope.

The failure analysis team is led by a seniormanagement representative from NASA/GSFCwith the authority not only to call upon theexpertise of engineers and scientists employedby NASA or its support contractors, but alsoto draft support from any organizationpreviously affiliated with the TelescopeProject. The failure analysis team is charteredto identify the nature of the anomaly and torecommend corrective action. This recom-mendation is reviewed at a higher managementlevel of NASA/GSFC. All changes to theTelescope’s hardware and all software con-figurations require NASA Level I concurrenceas specified in the HST Level I OperationsRequirements Document.

Pointing/Safemode Electronics and RetrievalMode Gyro Assemblies. The PSEA consists of40 electronic printed-board circuits withredundant functions to run the Telescope, evenin the case of internal circuit failure. It weighs86 lb (39 kg) and is installed in the EquipmentSection Bay 8. A backup gyroscope package, theRMGA, is dedicated for the PSEA and is alsolocated in Bay 8. The RMGA consists of three

gyroscopes. These are lower quality rate sensorsthan the RGAs because they are not intendedfor use during observations.

5.2 Optical Telescope Assembly

The OTA was designed and built by the Perkin-Elmer Corporation (Raytheon Optical Systems,Inc.). Although the OTA is modest in size byground-based observatory standards and has astraightforward optical design, its accuracy –coupled with its place above the Earth’satmosphere – renders its performance superior.

The OTA uses a “folded” design, common tolarge telescopes, which enables a long focallength of 189 ft (57.6 m) to be packaged into asmall telescope length of 21 ft (6.4 m). (Severalsmaller mirrors in the science instruments aredesigned similarly to lengthen the light pathwithin the particular science instrument.) Thisform of telescope is called a Cassegrain, and itscompactness is an essential component of anobservatory designed to fit inside the Shuttlecargo bay.

Conventional in design, the OTA isunconventional in other aspects. Largetelescopes at ground-based sites are limited intheir performance by the resolution attainablewhile operating under the Earth’s atmosphere,but the HST orbits high above the atmosphereand provides an unobstructed view of theuniverse. For this reason the OTA was designedand built with exacting tolerances to providenear-perfect image quality over the broadestpossible region of the spectrum.

The OTA is a variant of the Cassegrain, called aRitchey-Chretien, in which both the mirrors arehyperboloidal in shape (having a deepercurvature than a parabolic mirror). This form iscompletely corrected for coma (an imageobservation having a “tail”) and spherical

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K70110-517

FINE GUIDANCESENSORS (3)

AXIAL SCIENCEINSTRUMENTS (4)

PRIMARYMIRROR

FOCAL PLANE(IMAGE FORMED HERE)

RADIAL SCIENCEINSTRUMENTS

SUPPORTSYSTEMSMODULE

STRAY-LIGHTBAFFLES

INCOMING LIGHT

SECONDARY MIRRORAPERTURE DOOR

Fig. 5-17 Light path for the main Telescope

aberrations to provide an aplanatic system inwhich aberrations are correct everywhere in theFOV. The only residual aberrations are fieldcurvature and astigmatism. Both of these arezero exactly in the center of the field andincrease toward the edge of the field. Theseaberrations are easily corrected within theinstrument optics. For example, in the FaintObject Camera (FOC) there is a small telescopedesigned to remove image astigmatism.

Figure 5-17 shows the path of a light ray froma distant star as it travels through the Telescopeto the focus. Light travels down the tube, pastbaffles that attenuate reflected light fromunwanted bright sources, to the 94.5-in. (2.4-m)primary mirror. Reflecting off the front surfaceof the concave mirror, the light bounces back upthe tube to the 12-in. (0.3-m)-diameter convexsecondary mirror. The light is now reflected andconverged through a 23.5-in. (60-cm) hole in theprimary mirror to the Telescope focus, 3.3 ft(1.5 m) behind the primary mirror.

Four science instruments and three FGSs sharethe focal plane by a system of mirrors. A small“folding” mirror in the center of the FOV directs

light into the WFPC2. The remaining “science”field is divided among three axial scienceinstruments, each receiving a quadrant of thecircular FOV. Around the outside of the sciencefield, a “guidance” field is divided among thethree FGSs by their own folding mirrors. EachFGS receives 60 arcmin2 of field in a 90-degreesector. Figure 5-18 shows instrument/sensorfields of view.

The OTA hosts the science instruments andFGSs in that it maintains the structural supportand optical-image stability required for theseinstruments to fulfill their functions (see Fig.5-19). Components of the OTA are the primarymirror, the secondary mirror, the FPS, and theOTA Equipment Section. Perkin-ElmerCorporation designed and built all the opticalassemblies; Lockheed Martin built the OTAequipment section.

5.2.1 Primary Mirror Assemblyand Spherical Aberration

As the Telescope was put through its paces onorbit in 1990, scientists discovered its primarymirror had a spherical aberration. The outer

5-20K9322-05M

edge of the 8-foot (2.4-m) primary mirror wasground too flat by a width equal to 1/50 thethickness of a sheet of paper (about 2 microns).After intensive investigation, the problem wastraced to faulty test equipment used to defineand measure mirror curvature. The opticalcomponent of this test equipment was slightlyout of focus and, as a result, had shown themirror to be ground correctly. After thediscovery, Ball Aerospace scientists andengineers built the Corrective Optics SpaceTelescope Axial Replacement (COSTAR). TheCOSTAR was installed during the FirstServicing Mission in December 1993 andbrought the Telescope back to its originalspecifications.

The primary mirror assembly consists of themirror supported inside the main ring, whichis the structural backbone of the Telescope, andthe main and central baffles shown in Fig. 5-20.

This assembly provides the structural couplingto the rest of the spacecraft through a set ofkinematic brackets linking the main ring to theSSM. The assembly also supports the OTAbaffles. Its major parts are:

• Primary mirror• Main ring structure• Reaction plate and actuators• Main and central baffles.

Primary Mirror. The primary mirror blank, aproduct of Corning Glass Works, is known asultralow-expansion (ULE) glass. It was chosenfor its very low-expansion coefficient, whichensures the Telescope minimum sensitivity totemperature changes. The mirror is of a“sandwich” construction: two lightweightfacesheets separated by a core, or filling, of glasshoneycomb ribs in a rectangular grid (see Fig.5-21). This construction results in an 1800-lb

Fig. 5-18 Instrument/sensor field of view

SPACE TELESCOPEIMAGING SPECTROGRAPH

OPTICAL CONTROLSENSOR (3)

FINEGUIDANCESENSOR 3

NEAR INFRAREDCAMERA ANDMULTI-OBJECTSPECTROMETER

2 DETECTORS,SEPARATE APERTURES

FINE GUIDANCE SENSOR 2

CORRECTIVE OPTICSSPACE TELESCOPEAXIAL REPLACEMENT

FINEGUIDANCESENSOR 1

FAINT OBJECTCAMERA

+V3 AXIS

WIDE FIELD AND PLANETARY CAMERA 2

INCOMING IMAGE(VIEW LOOKING FORWARD

+V1 AXIS INTO PAGE)

14.1 arcminWFC

#1#2 #3

K70110-518

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Fig. 5-19 Optical Telescope Assembly components

Fig. 5-20 Primary mirror assembly

(818-kg) mirror instead of an 8000-lb solid-glassmirror.

Perkin-Elmer (now Raytheon Optical Systems,Inc.) ground the mirror blank, 8 ft (2.4 m) indiameter, to shape in its large optics fabricationfacility. When it was close to its finalhyperboloidal shape, the mirror was transferredto Perkin-Elmer ’s computer-controlledpolishing facility.

After being ground and polished, the mirrorwas coated with a reflective layer of aluminumand a protective layer of magnesium fluorideonly 0.1 and 0.025 micrometer thick,

K70110-520

MAIN BAFFLE

PRIMARY MIRROR

MIRROR MOUNT

ACTUATOR(TYPICAL)

REACTION PLATE

CENTRALBAFFLE

MAIN RING

SUPPORT SYSTEMSMODULE ATTACHMENTBRACKETS

PRIMARY MIRROR ASSEMBLY

SECONDARYMIRRORASSEMBLY GRAPHITE EPOXY

METERING TRUSS

CENTRAL BAFFLE

SUPPORT

FINE GUIDANCESENSORS (3)

FOCAL PLANESTRUCTURE

AXIALSCIENCEINSTRUMENTS(4)

ALUMINUMMAIN BAFFLE

ELECTRONIC BOXES

PRIMARY MIRROR

MAIN RING

RADIAL SCIENCEINSTRUMENT (1)

FIXED-HEADSTAR TRACKER

K70110-519

5-22K9322-05M

respectively. The fluoride layer protects thealuminum from oxidation and enhancesreflectance at the important hydrogen emissionline known as Lyman-Alpha. The reflectivequality of the mirror is better than 70 percent at1216 angstroms (Lyman-Alpha) in theultraviolet spectral range and better than 85percent for visible light.

The primary mirror is mounted to the main ringthrough a set of kinematic linkages. Thelinkages attach to the mirror by three rods thatpenetrate the glass for axial constraint and bythree pads bonded to the back of the glass forlateral support.

Main Ring. The main ring encircles the primarymirror; supports the mirror, the main baffle andcentral baffle, and the metering truss; andintegrates the elements of the Telescope to thespacecraft. The titanium ring is a hollow box

beam 15 in. (38 cm) thick, weighing 1200 lb(545.5 kg), with an outside diameter of 9.8 ft(2.9 m) (see Fig. 5-22). It is suspended inside theSSM by a kinematic support.

Reaction Plate. The reaction plate is a wheel ofI-beams forming a bulkhead behind the mainring, spanning its diameter. It radiates from acentral ring that supports the central baffle. Itsprimary function is to carry an array of heatersthat warm the back of the primary mirror,maintaining its temperature at 70 degrees. Madeof lightweight, stiff beryllium, the plate alsosupports 24 figure-control actuators attached tothe primary mirror and arranged around thereaction plate in two concentric circles. Thesecan be commanded from the ground, ifnecessary, to make small corrections to theshape of the mirror.

Baffles. The baffles of the OTA prevent stray lightfrom bright objects, such as the Sun, Moon, andEarth, from reflecting down the Telescope tubeto the focal plane. The primary mirror assem-bly includes two of the three assembly baffles.

Attached to the front face of the main ring, theouter, main baffle is an aluminum cylinder 9 ft(2.7 m) in diameter and 15.7 ft (4.8 m) long.Internal fins help it attenuate stray light. Thecentral baffle is 10 ft (3 m) long, conical in shape,

FRONTFACESHEET

INNEREDGEBAND

LIGHTWEIGHTCORE

OUTEREDGEBAND

REARFACESHEET

MIRRORCONSTRUCTION

THE MIRROR IS MADE OF CORNING CODE 7971ULTRA-LOW EXPANSION (ULE) SILICA GLASS

K70110-521

Fig. 5-21 Primary mirror construction

RADIALI-BEAMS (15)

FIGURE CONTROLACTUATORS (24)

MAIN RING

MIRROR PADS(24)

FLEXUREMOUNTS(15)

AXIALMOUNTS(15)

INTERCOSTALRIBS (48)

CENTRALRING

CENTRAL BAFFLE BRACKETS (5)

K70110-522

Fig. 5-22 Main ring and reaction plate

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and attached to the reaction plate through a holein the center of the primary mirror. It extendsdown the centerline of the Telescope tube. Thebaffle interiors are painted flat black tominimize light reflection.

5.2.2 Secondary Mirror Assembly

The Secondary Mirror Assembly cantilevers offthe front face of the main ring and supports thesecondary mirror at exactly the correct positionin front of the primary mirror. This positionmust be accurate within 1/10,000 in. wheneverthe Telescope is operating. The assemblyconsists of the mirror subassembly, a light baffle,and an outer graphite-epoxy metering trusssupport structure (see Fig. 5-23).

The Secondary Mirror Assembly contains themirror, mounted on three pairs of alignmentactuators that control its position andorientation. All are enclosed within the centralhub at the forward end of the truss support.

The secondary mirror has a magnification of10.4X, converting the primary-mirror

converging rays from f/2.35 to a focal ratiosystem prime focus of f/24 and sending themback toward the center of the primary mirror,where they pass through the central baffle to thefocal point. The mirror is a convex hyperboloid12 in. (0.3 m) in diameter and made of Zerodurglass coated with aluminum and magnesiumfluoride. Steeply convex, its surface accuracy iseven greater than that of the primary mirror.

Ground command adjusts the actuators to alignthe secondary mirror to provide perfect imagequality. The adjustments are calculated fromdata picked up by tiny optical control systemsensors located in the FGSs.

The principal structural element of the SecondaryMirror Assembly is the metering truss, a cage with48 latticed struts attached to three rings and acentral support structure for the secondary mirror.The truss, 16 ft (4.8 m) long and 9 ft (2.7 m) indiameter, is a graphite, fiber-reinforced epoxystructure. Graphite was chosen for its highstiffness, light weight, and ability to reduce thestructure’s expansiveness to nearly zero. This isvital because the secondary mirror must stayperfectly placed relative to the primary mirror,accurate to within 0.0001 in. (2.5 micrometers)when the Telescope operates.

The truss attaches at one end to the front faceof the main ring of the Primary MirrorAssembly. The other end has a central hub thathouses the secondary mirror and baffle alongthe optical axis. Aluminized mylar MLI in thetruss compensates for temperature variations ofup to 30 degrees Fahrenheit when the Telescopeis in Earth’s shadow so the primary andsecondary mirrors remain aligned.

The conical secondary mirror subassembly lightbaffle extends almost to the primary mirror. Itreduces the stray bright-object light fromsources outside the Telescope FOV.

BASE PLATE

ACTUATOR PAIR

SECONDARY MIRROR

SECONDARYMIRROR BAFFLE

ACTUATORPAIR

CLAMP

FLEXURE

SHROUD

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Fig. 5-23 Secondary mirror assembly

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5.2.3 Focal Plane StructureAssembly

The FPS is a large optical bench that physicallysupports the science instruments and FGSs andaligns them with the image focal plane of theTelescope. The -V3 side of the structure, awayfrom the Sun in space, supports the FHSTs andRSUs (see Fig. 5-24). It also provides facilities foron-orbit replacement of any instruments andthermal isolation between instruments.

The structure is 7 ft (2.1 m) by 10 ft (3.04 m) longand weighs more than 1200 lb (545.5 kg).Because it must have extreme thermal stabilityand be stiff, lightweight, and strong, the FPS isconstructed of graphite-epoxy, augmented withmechanical fasteners and metallic joints atstrength-critical locations. It is equipped withmetallic mounts and supports for OrbitalReplacement Units (ORU) used duringmaintenance.

The FPS cantilevers off the rear face of the mainring, attached at eight flexible points that adjustto eliminate thermal distortions. The structureprovides a fixed alignment for the FGSs. It has

guiderails and latches at each instrumentmounting location so Shuttle crews can easilyexchange science instruments and otherequipment in orbit.

5.2.4 OTA Equipment Section

The Equipment Section for the OTA is a largesemicircular set of compartments mountedoutside the spacecraft on the forward shell ofthe SSM (see Fig. 5-25). It contains the OTAElectrical Power and Thermal ControlElectronics (EP/TCE) System, Fine GuidanceElectronics (FGE), Actuator Control Electronics(ACE), Optical Control Electronics (OCE), andthe fourth DMS DIU. The OTA EquipmentSection has nine bays: seven for equipmentstorage and two for support. All bays haveoutward-opening doors for easy astronautaccess, cabling and connectors for theelectronics, and heaters and insulation forthermal control.

The EP/TCE System distributes power from theSSM EPS and the OTA system. Thermostatsregulate mirror temperatures and prevent mirrordistortion from the cold of space. The electricaland thermal electronics also collect thermalsensor data for transmission to the ground.

The three FGE units provide power, commands,and telemetry to each FGS. The electronicsperform computations for the sensor andinterface with the spacecraft pointing system foreffective Telescope line-of-sight pointing andstabilization. There is a guidance electronicsassembly for each guidance sensor.

The ACE unit provides the command andtelemetry interface to the 24 actuators attachedto the primary mirror and to the six actuatorsattached to the secondary mirror. Theseelectronics select which actuator to move andmonitor its response to the command.

K70110-524

SCIENCEINSTRUMENTCONNECTORPANEL

AXIAL SCIENCEINSTRUMENT

FOCAL PLANESTRUCTURE

PRIMARY MIRRORMAIN RING

RATE SENSINGUNIT FOR RATE GYROASSEMBLIES (3)

CREWSYSTEMHANDLESFIXED-HEADSTAR TRACKERS (3)

SCIENCE INSTRUMENTLATCHING MECHANISM

V2

V1

V3

Fig. 5-24 Focal plane structure

5-25K9322-05M

Positioning commands go from the ground tothe electronics through the DIU.

The OCE unit controls the optical controlsensors. These white-light interferometersmeasure the optical quality of the OTA and sendthe data to the ground for analysis. There is oneoptical control sensor for each FGS, but the OCEunit runs all control sensors. The DIU is anelectronic interface between the other OTAelectronics units and the Telescope commandand telemetry system.

5.3 Fine Guidance Sensor

The three FGSs are located at 90-degreeintervals around the circumference of the focalplane structure, between the structure frameand the main ring. Each sensor is 5.4 ft (1.5 m)long and 3.3 ft (1 m) wide and weighs 485 lb(220 kg).

Each FGS enclosure houses a guidance sensorand a wavefront sensor. The wavefront sensorsare elements of the optical control sensor usedto align and optimize the optical system of theTelescope.

The Telescope’s ability to remain pointing at adistant target to within 0.005 arcsec for longperiods of time is due largely to the accuracy ofthe FGSs. They lock on a star and measure any

apparent motion to an accuracy of 0.0028 arcsec.This is equivalent to seeing from New York Citythe motion of a landing light on an aircraftflying over San Francisco.

When two sensors lock on a target, the thirdmeasures the angular position of a star, aprocess called astrometry. Sensor astrometricfunctions are discussed in Section 4. DuringSM2 a re-certified FGS (S/N 2001) was installedas a replacement in the HST FGS Bay 1. DuringSM3A a re-certified FGS (S/N 2002) will beinstalled in the HST FGS Bay 2.

5.3.1 Fine Guidance SensorComposition and Function

Each FGS consists of a large structure housinga collection of mirrors, lenses, servos to locatean image, prisms to fine-track the image, beamsplitters, and four photomultiplier tubes, asshown in Fig. 5-26. The entire mechanismadjusts to move the Telescope into precisealignment with a target star. Each FGS has alarge (60 arcmin2) FOV to search for and trackstars, and a 5.0 arcsec2 FOV used by the detectorprisms to pinpoint the star.

The sensors work in pairs to aim the Telescope.The Guide Star Selection System, developed bythe Science Institute, catalogs and charts guidestars near each observation target to make it

K70110-525

OPTICALCONTROL ELECTRONICS

DATA INTERFACE UNIT

LOOKING FORWARD

FINE GUIDANCEELECTRONICS

ELECTRICAL POWER/THERMAL CONTROLELECTRONICS

MULTILAYERINSULATION ON DOORS

ACTUATORCONTROL ELECTRONICS FINE GUIDANCE

ELECTRONICS

-V2

-V3J

HG

FE D

C

B

A

ACE DIU

Fig. 5-25 Optical Telescope Assembly Equipment Section

5-26K9322-05M

easier to find the target. First, one sensorsearches for a target guide star. After the firstsensor locks onto a guide star, the second sensorlocates and locks onto another target guide star.The guide stars, once designated and located,keep the image of the observation target in theaperture of the selected science instrument.

Each FGS uses a 90-degree sector of theTelescope’s FOV outside the central “science”field. This region of the FOV has the greatestastigmatic and curvature distortions. The size ofthe FGS’s FOV was chosen to heighten theprobability of finding an appropriate guide star,even in the direction of the lowest starpopulation near the galactic poles.

An FGS “pickoff” mirror intercepts theincoming stellar image and projects it into thesensor ’s large FOV. Each FGS FOV has 60arcmin2 available. The guide star of interestcan be anywhere within this field, so the FGSwill look anywhere in that field to find it.After finding the star, the sensor locks onto itand sends error signals to the Telescope,telling it how to move to keep the star imageperfectly still.

The FGS can move its line of sight anywherewithin its large FOV using a pair of star selectorservos. Each can be thought of as an opticalgimbal: One servo moves in a north-southdirection, the other east and west. They steer thesmall FOV (5 arcsec2) of the FGS detectors toany position in the sensor field. Encoders withineach servo system send back the exact coordinatesof the detector field centers at any point.

Because the exact location of a guide star maybe uncertain, the star selector servos also cancause the detector to search the region aroundthe most probable guide star position. Itsearches in a spiral pattern, starting at the centerand spiraling out until it finds the guide star itseeks. Then the detectors are commanded to gointo fine-track mode and hold the star imageexactly centered in the FOV, while the starselector servo encoders send information aboutthe position of the star to the spacecraft PCS.

The detectors are a pair of interferometers,called Koester ’s prisms, coupled tophotomultiplier tubes (see Fig. 5-27). Eachdetector operates in one axis, so two detectorsare needed. Operating on the incomingwavefront from the distant guide star, theinterferometers compare the wave phase at oneedge of the Telescope’s entrance aperture with

Fig. 5-26 Cutaway view of Fine Guidance Sensor

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STARSELECTORMIRRORS

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Fig. 5-27 Optical path of Fine Guidance Sensor

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the phase at the opposite edge. When the phasesare equal, the star is exactly centered. Any phasedifference shows a pointing error that must becorrected.

Along the optical path from Telescope todetector are additional optical elements thatturn or fold the beam to fit everything inside theFGS enclosure, and to correct the Telescope’sastigmatism and field curvature. All opticalelements are mounted on a temperature-controlled, graphite-epoxy composite opticalbench.

5.3.2 Articulated Mirror System

Analysis of the FGS on-orbit data revealed thatminor misalignments of the optical pupilcentering on Koester’s prism interferometer inthe presence of spherical aberration preventedthe FGS from achieving its optimumperformance. During the recertification of FGS(S/N 2001), fold flat #3 in the radial bay moduleoptical train was mechanized to allow on-orbitalignment of the pupil.

Implementation of this system utilized existingsignals and commands by rerouting them witha unique interface harness enhancement kit(OCE-EK) interfacing the OCE, the DIU, and theFine Guidance System/Radial Bay Module(FGS/RBM). The OCE-EK was augmented withthe Actuator Mechanism Electronics (AME) andthe fold flat #3 Actuator Mechanism Assembly(AMA) located internal to the FGS/RBM.Ground tests indicate a substantial increase inperformance of the FGS with this innovativedesign improvement.

5.4 Solar Array and JitterProblems

From the beginning, in the late 1970s, the SAs– designed by the European Space Agency and

built by British Aerospace, Space Systems –have been scheduled for replacement because oftheir power loss from radiation exposure inspace. However, as engineers put the Telescopethrough its paces in April 1990, they discoveredtwo problems: a loss of focus and images thatjittered briefly when the Telescope flew into andout of Earth’s shadow. The jitter problem wastraced to the two large SAs. Abrupt temperaturechanges, from -150 to 200 degrees Fahrenheitduring orbit, cause the panels to distort twiceduring each orbit. As a temporary fix, engineerscreated software that commanded the PCS tocompensate for the jitter automatically. Theproblem was mitigated during SM1 by thereplacement of the old arrays with new onesthat had been modified to reduce thermalswings of the bi-stems.

5.4.1 Configuration

The SAs are two large rectangular wings ofretractable solar cell blankets fixed on a two-stem frame. The blanket unfurls from a cassettein the middle of the wing. A spreader bar ateach end of the wing stretches the blanket andmaintains tension. For the replacement SAsdelivered to the Telescope during the FirstServicing Mission, the spring-loaded rollerassembly was replaced by a series of springsconnecting the spreader bar to the blanket. Thischange eliminated the jitter induced into theTelescope as it passed from eclipse (night) intosunlight (day) each orbit.

The wings are on arms that connect to a driveassembly on the SSM forward shell at one endand to the secondary deployment mechanism(blankets and bistems) on the other end. Thetotal length of the cassette, arm, and drive is15.7 ft (4.8 m) (see Fig. 5-28).

Each wing has 10 panels that roll out from thecassette. The panels are made of 2438 solar cells

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attached to a glass-fiber/Kapton surface, withsilver mesh wiring underneath, covered byanother layer of Kapton. The blankets are lessthan 500 micrometers thick, so they roll uptightly when the wings are stowed. Each wingweighs 17 lb (7.7 kg) and, at full extension, is40 ft (12.1 m) long and 8.2 ft (2.5 m) wide.

5.4.2 Solar Array Subsystems

The SA subsystems include the primary andsecondary deployment mechanisms, theirdrives, and associated electronics.

The primary deployment mechanism raises theSA mast from the side of the SSM to a standingposition perpendicular to the Telescope. Thereare two mechanisms, one for each wing. Eachmechanism has motors to raise the mast andsupports to hold it in place when erect.

An astronaut can raise the array mast manuallyif the drive power fails. Using a wrench fittingon the deployment drive, the astronaut hand-cranks the mast after releasing the latches.

Once the SA is raised, the secondarydeployment mechanism unfurls the wingblankets. Each wing has a secondary

mechanism assembly: a cassette drum to holdsolar panels, a cushion to protect the blanket,and motors and subassemblies. The assemblyrolls out the blanket, applies tension evenly sothe blankets stretch, and transfers data andpower along the wing assembly. The blanketcan roll out completely or part way. Thesecondary deployment mechanism also has amanual override (see Fig. 5-29).

A SA drive at the base of each mast rotates thedeployed array toward the Sun, turning ineither direction. Each drive has a motor thatrotates the mast on command and a brake tokeep the array in a fixed position with respectto the Telescope. The drive can move and lockthe SA into any position.

Each drive has a clamp ring that acts as a releasemechanism if opened. This allows a crewmember to jettison the entire SA if necessary.

Two electronics assemblies (boxes) – the SolarArray Deployment Electronics and the SolarArray Drive Electronics – control and monitorall functions of each SA. They provide theelectronic interface to the other Telescopesystems and generate the commands for theprimary and secondary deploymentmechanisms and the SA drive.

FITTING FORMANUAL DEPLOYMENT

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Fig. 5-29 Fitting for Solar Array manual deployment

COMPENSATOR RAIL

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SPREADERBAR

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BISTEM CASSETTE

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Fig. 5-28 Solar Array wing detail

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5.4.3 Solar Array Configurationfor Servicing Mission 3A

The Solar Array wings will remain deployedduring servicing. This will allow the Telescope’sbatteries to remain fully charged during themission and will not impact servicing activities.

5.5 Science Instrument Controland Data Handling Unit

The SI C&DH unit keeps all science instrumentsystems synchronized. It works with the DMUto process, format, temporarily store on the datarecorders, or transmit all science and engineer-ing data created by the instruments to theground. Fairchild Camera and Instrument Cor-poration and IBM built this unit.

5.5.1 Components

The SI C&DH unit is a collection of electroniccomponents attached to an ORU tray mountedon the door of Bay 10 in the SSM EquipmentSection (see Fig. 5-30). Small Remote InterfaceUnits (RIU), also part of the system, provide theinterface to individual science instruments.Components of the SI C&DH unit are the NASAStandard Spacecraft Computer (NSCC-I), twostandard interface circuit boards for the com-puter, two control units/science data formatterunits, two CPU modules, a PCU, two RIUs, andvarious memory, data, and command commu-nications lines (buses) connected by couplers.The SI C&DH components are redundant so thesystem can recover from any single failure.

NASA Computer. The NSSC-I has a CPU andeight memory modules, each holding 8,192eighteen-bit words. One embedded softwareprogram (the “executive”) runs the computer. Itmoves data, commands, and operation pro-grams (called applications) for individual sci-ence instruments in and out of the processingunit. The application programs monitor and

control specific instruments and analyze andmanipulate the collected data.

The memory stores operational commandsfor execution when the Telescope is not incontact with the ground. Each memory unithas five areas reserved for commands andprograms unique to each science instru-ment. The computer can be reprogrammedfrom the ground for future requests or forworking around failed equipment.

Standard Interface Unit. The standard in-terface board is the communications bridgebetween the computer and the CU/SDF.

Control Unit/Science Data Formatter. Theheart of the SI C&DH unit is the CU/SDF.It formats and sends all commands anddata to designated destinations such as theDMU of the SSM, the NASA computer, andthe science instruments. The unit has a mi-croprocessor for control and formattingfunctions.

Fig. 5-30 Science Instrument Control and DataHandling unit

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EXTRAVEHICULAR ACTIVITYHANDLE

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The CU/SDF receives ground commands, datarequests, science and engineering data, andsystem signals. Two examples of system signalsare “time tags,” clock signals that synchronizethe entire spacecraft, and “processor interfacetables,” or communications codes. The CU/SDFtransmits commands and requests afterformatting them so that the specific destinationunit can read. For example, ground commandsand SSM commands are transmitted withdifferent formats. Ground commands use 27-bitwords, and SSM commands use 16-bit words.The formatter translates each command signalinto a common format. The CU/SDF alsoreformats and sends engineering and sciencedata. Onboard analysis of the data is an NSSC-Ifunction.

Power Control Unit. The PCU distributes andswitches power among components of the SIC&DH unit. It conditions the power required byeach unit. For example: The computer memoryboards typically need +5 volts, -5 volts, and +12volts; the CU/SDF, on the other hand, requires+28 volts. The PCU ensures that all voltagerequirements are met.

Remote Interface Units. RIUs transmitcommands, clock and other system signals, andengineering data between the scienceinstruments and the SI C&DH unit. The RIUsdo not send science data. There are six RIUs inthe Telescope: five attached to the scienceinstruments and one dedicated to the CU/SDFand PCUs in the SI C&DH unit. Each RIU canbe coupled with up to two expander units.

Communications Buses. The SI C&DH unitcontains data bus lines that pass signals anddata between the unit and the scienceinstruments. Each bus is multiplexed: one linesends system messages, commands, andengineering data requests to the module units,and a reply line transmits requested information

and science data back to the SI C&DH unit. Acoupler attaches the bus to each remote unit.This isolates the module if the RIU should fail.The SI C&DH coupler unit is on the ORU tray.

5.5.2 Operation

The SI C&DH unit handles science instrumentsystem monitoring (such as timing and systemchecks), command processing, and dataprocessing.

System Monitoring. Engineering data tells themonitoring computer whether instrumentsystems are functioning. At regular intervals,varying from every 500 milliseconds to every 40seconds, the SI C&DH unit scans all monitoringdevices for engineering data and passes data tothe NSCC-I or SSM computer. The computersprocess or store the information. Any failureindicated by these constant tests could initiatea “safing hold” situation (see para 5.1.7), andthus a suspension of science operations.

Command Processing. Figure 5-31 shows theflow of commands within the SI C&DH unit.Commands enter the CU/SDF (bottom right inthe drawing) through the SSM Command DIU(ground commands) or the DIU (SSMcommands). The CU/SDF checks and reformatsthe commands, which then go either to the RIUsor to the NSCC-I for storage. “Time-tagged”commands, stored in the computer’s memory(top right of drawing), also follow this process.

Each command is interpreted as “real time,” asif the SI C&DH just received it. Manycommands actually are onboard storedcommands activated by certain situations. Forexample, when the Telescope is positioned fora programmed observation using the SpaceTelescope Imaging Spectrograph, that programis activated. The SI C&DH can issue certainrequests to the SSM, such as to execute a limited

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number of pointing control functions to makesmall Telescope maneuvers.

Science Data Processing. Science data can comefrom all science instruments at once. The CU/SDF transfers incoming data through computermemory locations called packet buffers. It fillseach buffer in order, switching among them asthe buffers fill and empty. Each data packet goesfrom the buffer to the NSCC-I for furtherprocessing, or directly to the SSM for storage inthe data recorders or transmission to theground. Data returns to the CU/SDF aftercomputer processing. When transmitting, theCU/ SDF must send a continuous stream ofdata, either full packet buffers or empty bufferscalled filler packets, to maintain a synchronizedlink with the SSM. Special checking codes

(Reed-Solomon and pseudo-random noise) canbe added to the data as options. Figure 5-32shows the flow of science data in the Telescope.

5.6 Space Support Equipment

The Hubble Space Telescope was designed to bemaintained, repaired, and enhanced while inorbit, extending its life and usefulness. Forservicing, the Space Shuttle will capture andposition the Telescope vertically in the aft end ofthe cargo bay, and the crew will performmaintenance and replacement tasks. The SpaceSupport Equipment (SSE) to be used during themission provides a maintenance platform to holdthe Telescope, provides electrical support of theTelescope during servicing, and provides storagefor replacement components known as ORUs.

Fig. 5-31 Command flow for Science Instrument Control and Data Handling unit

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The major SSE items used for the ServicingMission 3A are the Flight Support System(FSS) and the ORU Carrier (ORUC).Additionally, crew aids and tools will be usedduring servicing.

5.6.1 Flight Support System

The FSS provides the platform that holds theTelescope during servicing (see Fig. 5-33). TheFSS has been used in different configurations forthe HST First and Second Servicing Missions,the Solar Maximum Repair Mission, and theUpper Atmospheric Research Satellite DeployMission. The FSS consists of two majorcomponents: a horseshoe-shaped cradle and asupporting latch beam providing a structuraland electrical interface with the Shuttle. Acircular ring called the Berthing and PositioningSystem (BAPS) interfaces with the Telescope,allowing it to pivot or rotate during the mission.

The BAPS ring is pivoted down and locked forliftoff. During the mission, the ring is pivoted

up to a horizontal position. A closed-circuittelevision camera mounted to the FSS helpsastronauts guide the Telescope onto the ring.Three remote-controlled latches on the ring graband hold three towel-rack-like pins on the rearof the Telescope. A remote-controlled electricalumbilical connector on the FSS engages theTelescope, providing it with orbiter powerthrough the FSS. This power helps relieve thedrain on the Telescope’s batteries during themission. Radio communications providetelescope telemetry data and control.

Once the Telescope is berthed to the ring (seeFig. 5-34), the FSS can pivot (tilt) or rotate theTelescope. This positions the appropriate regionof the Telescope for access during extravehicularactivity. Additionally, the ring can pivot theTelescope to an appropriate attitude for orbiterreboost.

During the first EVA, crew members will installthe BAPS Support Post (BSP). The BSP providesan additional linkage to support and isolate the

FILLERPACKETS

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Fig. 5-32 Flow of science data in the Hubble Space Telescope

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Fig. 5-33 Flight Support System configuration

5.6.2 Orbital Replacement UnitCarrier

An ORUC is used to carry replacements intoorbit and to return replaced units to Earth. Thecarrier consists of a Spacelab pallet outfittedwith shelves and protective enclosures to holdthe replacement units (see Fig. 5-35). Items onthe ORUC for SM3A include (1) the AdvancedComputer and two spare VIKs in the LOPE and(2) a Solid State Recorder, an S-Band SingleAccess Transmitter, the Rate Sensor Units, andassociated flight harnesses in the COPE.

The FGS will be transported in the FGSScientific Instrument Protective Enclosure(FSIPE). A spare Advanced Computer, a spareRate Sensor Unit, and Multi-Layer Insulationpatches will be carried in the Axial ScientificInstrument Protective Enclosure (ASIPE). TheNOBL will be carried in a NOBL ProtectiveEnclosure (NPE).

All ORUs and scientific instruments are carriedwithin protective enclosures to provide them abenign environment throughout the mission.

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Telescope during EVAs and for any orbitalreboosts. The BSP will remain in position forlanding.

Astronauts remotely control all FSS mechanisms– berthing latches, umbilical connector, pivoter,BSP lock, rotator, and ring down-lock – from theorbiter’s aft flight deck, providing the crewmaximum flexibility. Besides being fullyelectrically redundant, each mechanismcontains manual overrides and backups toensure mission success and astronaut safety.

SPACETELESCOPE

BERTHING ANDPOSITIONINGRING

FLIGHT SUPPORT SYSTEM

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Fig. 5-34 Flight Support System Berthing andPositioning System ring pivoted up withTelescope berthed

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The enclosures protect the instruments fromcontamination and maintain the temperature ofthe instruments or ORUs within tight limits.

Instruments are mounted in the enclosuresusing the same manually driven latch systemthat holds instruments in the Telescope. TheASIPE and FSIPE are mounted to the pallet ona spring system that reduces the level ofvibration the instruments receive, especiallyduring liftoff and landing.

The other ORUs are carried in an additionalenclosure called the Large ORU ProtectiveEnclosure (LOPE). The enclosure also providescontamination and thermal control, though notto such stringent requirements as the SIPE. The

LOPE contains Transport Modules that aredesigned to custom fit each ORU. The transportmodules have foam or Visco-Elastic Materialthat surrounds the ORU and isolates it fromlaunch and landing vibration environments.

During the change-out process, replacedscience instruments are stored temporarily inthe ORUC. A typical change-out begins withan astronaut removing the old instrumentfrom the Telescope and attaching it to abracket on the ORUC. The astronaut thenremoves the new instrument from itsprotective enclosure and installs it in theTelescope. Finally, the astronaut places the oldinstrument in the appropriate protectiveenclosure for return to Earth.

Fig. 5-35 Orbital Replacement Unit Carrier

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The ORUC receives power for its TCS from theFSS. The carrier also provides temperaturetelemetry data through the FSS for readout inthe Shuttle and on the ground during themission.

5.6.3 Crew Aids

Astronauts perform extravehicular activitiesusing many tools to replace instruments andequipment, to move around the Telescope andthe cargo bay, and to operate manual overridedrives. Tools and equipment, bolts, connectors,and other hardware were standardized not onlyfor the Telescope but also between the Telescopeand the Shuttle. For example, grapplingreceptacles share common features.

To move around the Telescope, the crew uses225 ft of handrails encircling the spacecraft. Forvisibility, the rails are painted yellow. Inaddition, the crew can hold onto guiderails,trunnion bars, and scuff plates fore and aft.

The astronauts can install portable handholdplates where there are no permanent holds, such Fig. 5-36 Portable Foot Restraint

PITCH CONTROL

ELEVATIONCONTROL

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as on the FGS. Another tool is the Portable FootRestraint (PFR), shown in Fig. 5-36.

While the astronauts work, they use tethers tohook tools to their suits and tie replacementunits to the Telescope. Each crew member hasa ratchet wrench to manually crank the antennaand array masts if power for the mast drivesfails. A power wrench also is available if hand-cranking is too time consuming. Other handtools include portable lights and a jettisonhandle, which attach to sockets on the aperturedoor and to SA wings so the crew can push theequipment away from the Telescope.