CONTENTS
Section

Page

1
1.1
1.1.1
1.1.2
1.1.3
1.1.4
1.1.5
1.2

INTRODUCTION
Hubble Space Telescope Configuration
Optical Telescope Assembly
The Science Instruments
Support Systems Module
Solar Arrays
Computers
The Hubble Space Telescope Program

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2
2.1
2.2
2.3
2.3.1
2.3.2
2.3.3
2.3.4
2.4
2.5
2.6
2.7
2.7.1
2.7.2
2.8

HUBBLE SPACE TELESCOPE SECOND SERVICING MISSION
Reasons for Orbital Servicing
Orbital Replacement Units
Shuttle Support Equipment
Remote Manipulator System
Space Support Equipment
Orbital Replacement Unit Carrier
Second Axial Carrier
Astronaut Roles and Training
Extravehicular Crew Aids and Tools
Astronauts of the Second Servicing Mission
Servicing Mission Activities
Rendezvous with the Hubble Space Telescope
Extravehicular Servicing Activities – Day by Day
Future Servicing Plans

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3.1
3.2
3.3
3.4

HUBBLE SPACE TELESCOPE SCIENCE AND DISCOVERIES
Planets
Stellar Evolution
Galaxies and Cosmology
Summary

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4
4.1
4.1.1
4.1.2
4.2
4.2.1
4.2.2
4.2.3
4.2.4
4.2.5
4.3
4.3.1
4.3.2
4.3.3
4.3.4

SCIENCE INSTRUMENTS AND COSTAR
Near Infrared Camera and Multi-Object Spectrometer
Instrument Description
NICMOS Specifications
Space Telescope Imaging SpectrographPhysical Description
Spectra Operational Modes
STIS Specifications
Advantages Over Current Instrumentation
Observations
Faint Object Camera
Physical Description
Observation Modes
Faint Object Camera Specifications
Observations

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Section

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4.4
4.4.1
4.4.2
4.4.3
4.4.4
4.5
4.5.1
4.5.2
4.5.3
4.5.4
4.6
4.6.1
4.6.2
4.6.3
4.7

Wide Field/Planetary Camera II
Comparison of WF/PC II and Faint Object Camera
Physical Description
WF/PC II Specifications
Observations
Astrometry (Fine Guidance Sensors)
Fine Guidance Sensor Specifications
Operational Modes for Astrometry
Fine Guidance Sensor Filter Wheel
Astrometric Observations
HST Operations
Space Telescope Science InstituteSpace Telescope Operations Control Center
Operational Characteristics
Acquisition and Observation

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5
5.1
5.1.1
5.1.2
5.1.3
5.1.4
5.1.5
5.1.6
5.1.7
5.2
5.2.1
5.2.2
5.2.3
5.2.4
5.3
5.3.1
5.3.2
5.4
5.4.1
5.4.2
5.4.3
5.5
5.5.1
5.5.2
5.6
5.6.1
5.6.2
5.6.3
5.6.4

HUBBLE SPACE TELESCOPE SYSTEMS
Support Systems Module
Structures and Mechanisms Subsystem
Instrumentation and Communications Subsystem
Data Management Subsystem
Pointing Control Subsystem
Electrical Power Subsystem
Thermal Control
Safing (Contingency) System
Optical Telescope AssemblyPrimary Mirror Assembly and Spherical Aberration
Secondary Mirror Assembly
Focal Plane Structure Assembly
OTA Equipment Section
Fine Guidance Sensor
Fine Guidance Sensor Composition and Function
Articulated Mirror System
Solar Array and Jitter Problems
Configuration
Solar Array Subsystems
Solar Array Configuration for Second Servicing Mission
Science Instrument Control and Data Handling Unit
Components
Operation
Space Support Equipment
Flight Support System
Orbital Replacement Unit Carrier
Second Axial Carrier
Crew Aids

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ILLUSTRATIONS
Figure
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Page
The Hubble Space Telescope (HST) – shown in a clean room at Lockheed
Martin Missiles & Space in Sunnyvale, California, before shipment toKennedy Space Center – is equipped with science instruments andengineering subsystems designed as orbital replacement units.
Schedule of extravehicular activities
HST overall configuration
HST exploded view
Telescope science instrument specifications
Space Telescope program phase summary
Development teams for science instruments and COSTAR
Organization summary for HST program operational phase
HST data collecting network
Hubble Space Telescope Second Servicing Mission Orbital Replacement Units
Second Servicing Mission configuration
Flight Support System configuration
Orbital Replacement Unit Carrier
Second Axial Carrier
Neutral Buoyancy Simulator at Marshall Space Flight Center
The STS-82 mission has seven crew members. They are (front row, from left)
Commander Kenneth D. Bowersox, Steven A. Hawley, and Pilot Scott J.
Horowitz; (back row) Joseph R. Tanner, Gregory J. Harbaugh, Mark C. Lee, andSteven L. Smith.
Schedule of extravehicular activities
Change-out of Goddard High Resolution Spectrograph and Space Telescope
Imaging SpectrographInstallation of Fine Guidance Sensor #1
Replacement of Engineering/Science Tape Recorder #2 in Bay 8
Installation of Optical Control Electronics Enhancement Kit in Bay C
Replacement of Data Interface Unit #2 in Bay 3
Installation of Solid State Recorder in Bay 5
Replacement of Solar Array Drive Electronics #2 in Bay 7
Attaching cover to Magnetic Sensing System #2
Redeploying the Space TelescopeThis view of the planet Mars is the clearest picture ever taken from Earth,surpassed only by close-up shots sent back by visiting space probes. Hubble’s
Wide Field/Planetary Camera II took the picture on February 25, 1995, when
Mars was approximately 65 million miles (103 million km) from Earth.
A state-sized dust storm churns near the edge of Mars’ north polar cap in 1996.
The polar storm probably is a consequence of large temperature differences
between the polar ice and the dark regions to the south.
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Page
The Telescope is allowing astronomers to study Neptune’s remarkably dynamic
atmosphere with a level of detail not possible since the Voyager 2 flyby in 1989.
The planet’s predominant blue color results from absorption of red and infrared
light by its methane atmosphere.
Hubble’s Faint Object Camera imaged nearly the entire surface of Pluto as it
rotated through its 6.4-day period in late June and early July 1994. Viewing theplanet’s surface detail is as difficult as trying to read the printing on a golf balllocated 30 miles away!
Resembling an imaginary sea serpent’s head, this eerie, dark structure is acolumn of cool molecular hydrogen gas and dust that serves as an incubator for
new stars. Some of the stars are embedded in finger-like protrusions extending
from the top of the nebula. Each “fingertip” is larger than Earth’s solar system.
When a star like the Earth’s sun nears the end of its life, it expands to more than
50 times its original diameter, becoming a red giant star and ejecting a nebula.
This planetary nebula, NGC 7027, was once a star like the sun. The initial
ejections occurred episodically to produce its concentric shells.
Helix Nebula, the closest planetary nebula to Earth – 450 light-years away inthe constellation Aquarius – has thousands of tadpole-like objects (upper rightcorner) that astronomers call “cometary knots.” Each “tadpole” head is at least
twice the size of Earth’s solar system; each tail stretches 100 billion miles.
Astronomers used a combination of image-processing techniques to create thisstunning image of Eta Carinae, a pair of gas and dust clouds more than 8,000
light-years away. It is one of the highest resolution images of an extended object
ever produced by Hubble. Structures only 10 billion miles across (about thediameter of Earth’s solar system) can be distinguished.
Until they viewed galaxy NGC 4261, some 100 million light-years away,
scientists had never seen the enormous circulating disks of matter that hadformed around a suspected black hole. Astronomers calculate that the object at
the center of the disk is 1.2 billion times the mass of the sun, yet concentrated
into a region of space no larger than Earth’s solar system.
HST images of different quasar home sites include evidence of a catastrophic
collision between two galaxies traveling at about 1 million mph (top center) andevidence of a dance between two merging galaxies (bottom right).
Humankind’s most detailed optical view of the universe, called the Hubble
Deep Field, was provided by Hubble in late 1995. This image was assembled
from 342 separate exposures taken by the Wide Field/Planetary Camera II for
10 consecutive days, December 18 – 28. It covers a speck of the sky only about
the width of a dime located 75 feet away.
A “bar” of stars, dust, and gas characterizes Galaxy NGC 1365, a barred spiralgalaxy located in a cluster of galaxies called Fornax. The HST color image (right)
shows a region in NGC 1365. The black-and-white photograph from a groundbased telescope shows the entire galaxy, visible from the Southern Hemisphere.
Cutaway view of COSTAR showing its optical bench deployed and pickoff
mirrors in position to capture light from the secondary mirror
iv

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Figure
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Near Infrared Camera and Multi-Object Spectrometer (NICMOS)NICMOS optical characteristics
NICMOS optical path
NICMOS cryogen dewar
NICMOS specifications
Space Telescope Imaging Spectrograph (STIS)STIS components and detectors
STIS spectroscopic modes
Multi-Anode Microchannel Plate Array (MAMA) detector
Simplified MAMA system
STIS filter set
STIS specifications
Faint Object Camera (FOC) major subsystems
FOC optical relay system layout
Optical relay system layout at f/151
Optical relay system layout at f/75.5
FOC specifications
Wide Field/Planetary Camera (WF/PC) overall configuration
WF/PC optics design
WF/PC II imaging
WF/PC II specifications
Fine Guidance Sensor (FGS)
FGS specifications
Space Telescope ground systems overview
HST nominal orbit“Continuous-zone” celestial viewing
Using the moon as an occulting disk
HST single-axis maneuvers
Sun-avoidance maneuver
TDRS-HST contact zones
HST instrument apertures
Hubble Space Telescope – exploded view
Hubble Space Telescope axes
Design features of Support Systems Module
Structural components of Support Systems Module
Aperture door and light shieldSupport Systems Module forward shell
v

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Figure
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Page
Support Systems Module Equipment Section bays and contents
Support Systems Module aft shroud and bulkheadHigh Gain Antenna
Data Management Subsystem functional block diagram
DF-224 computer
Data Management Unit configuration
Location of Pointing Control Subsystem equipment
Reaction Wheel AssemblyElectrical Power Subsystem functional block diagram
Placement of thermal protection on Support Systems Module
Light path for the main Telescope
Instrument/sensor field of view
Optical Telescope Assembly components
Primary mirror assemblyPrimary mirror construction
Main ring and reaction plate
Secondary mirror assembly
Focal plane structure
Optical Telescope Assembly Equipment Section
Cutaway view of Fine Guidance Sensor
Optical path of Fine Guidance Sensor
Solar Array wing detail
Fitting for Solar Array manual deployment
Science Instrument Control and Data Handling unit
Command flow for Science Instrument Control and Data Handling unit
Flow of science data in the Hubble Space Telescope
Flight Support System configuration
Flight Support System Berthing and Positioning System ring pivoted up with
Telescope berthed
Orbital Replacement Unit Carrier
Second Axial Carrier
Portable Foot Restraint

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INTRODUCTION
tists to rethink some of their earlier theories
about galactic evolution. (Section 4 of thisguide contains additional information on theTelescope’s scientific discoveries.)

azing through the first crude telescopes,
Galileo, Kepler, and their contemporariesof the 17th century discovered themoon’s craters, the satellites of Jupiter, and therings of Saturn, leading the way to today’s quest
for in-depth knowledge and understanding of
the cosmos. Since its launch in April 1990,
NASA’s Edwin P. Hubble Space Telescope (HST)
has continued this historic quest, providing scientific data and photographs of unprecedented
resolution from which many new and exciting
discoveries have been made.

GThe Telescope’s mission is to spend 15 years
probing the farthest and faintest reaches of thecosmos. Crucial to fulfilling this promise is aseries of on-orbit manned servicing missions.
The First Servicing Mission took place inDecember 1993.
During these missions, astronauts perform anumber of planned repairs and maintenanceactivities to restore and upgrade the observatory’s capabilities. To facilitate this process, theTelescope’s designers configured the various
science instruments and several vital engineering subsystems as Orbital Replacement Units(ORU) – modular packages with standardized
fittings accessible to astronauts in pressurized
suits (see Fig. 1-1).

This unique observatory operates around theclock above the Earth’s atmosphere to gather
information for teams of scientists studying
virtually all the constituents of our universe,
including planets, stars, star-forming regions of
the Milky Way galaxy, distant galaxies andquasars, and the tenuous hydrogen gas lying
between the galaxies.
The Telescope can produce images of the outerplanets in Earth’s solar system that rival theclarity of those achieved by the Voyager flybys.
Astronomers have been able to resolve previously unsuspected details of star-forming
regions of the Orion Nebula in the Milky Way.
They have detected expanding gas shells
blown off by exploding stars. Using the highresolution and light-gathering power of theTelescope, scientists have calibrated the distances to remote galaxies and detected cooldisks of matter trapped in the gravitational
field of the cores of galaxies that portend thepresence of massive black holes.

Flying aboard the shuttle Discovery in February 1997, the highly experienced, expertly
trained STS-82 crew will carry out Hubble’s
Second Servicing Mission. Space-walking
astronauts will remove the Goddard High
Resolution Spectrograph (GHRS) and install
the Space Telescope Imaging Spectrograph(STIS). They will replace the Faint Object
Spectrograph (FOS) with the Near Infrared
Camera and Multi-Object Spectrometer (NICMOS). They also will install and replace othercomponents needed to maintain or augment
Telescope operations. These include:

Spectroscopic observations at ultraviolet wavelengths inaccessible from the ground havegiven astronomers their first opportunity to
study the abundance and spatial distributionof intergalactic hydrogen in relatively nearby
regions of the universe and have forced scien-• Reaction Wheel Assembly (RWA)• Fine Guidance Sensor (FGS)• Data Interface Unit (DIU)• Solar Array Drive Electronics (SADE)• Engineering/Science Tape Recorder (E/STR)• Solid State Recorder (SSR)

K70110-01M

1-1

K70110-101

Fig. 1-1

K70110-01M

The Hubble Space Telescope (HST) – shown in a clean room at Lockheed Martin
Missiles & Space in Sunnyvale, California, before shipment to Kennedy Space
Center – is equipped with science instruments and engineering subsystems
designed as orbital replacement units.
1-2

the Telescope and its science instruments. The
major elements are:

• Magnetic Sensing System (MSS) covers• Optical Control Electronics Enhancement Kit(OCE-EK).

• Optical Telescope Assembly (OTA) – twomirrors and associated structures that collect
light from celestial objects• Science instruments – devices used to analyze the images produced by the OTA• Support Systems Module (SSM) – spacecraft
structure that encloses the OTA and science
instruments• Solar Arrays.

Figure 1-2 shows the servicing activities asscheduled at press time. Section 2 provides
more details of the Second Servicing Mission.
In the last 18 years, the HST team has overcome
enormous technical obstacles to successfully
develop and launch the orbiting observatory.
This tradition continues with the Second
Servicing Mission, a challenging endeavorthat promises to significantly expand the scientific capabilities of Hubble Space Telescope.

1.1.1 Optical Telescope Assembly
The OTA consists of two mirrors, support trusses,
and the focal plane structure. The optical system
is a Ritchey-Chretien design, in which two special
aspheric mirrors serve to form focused images
over the largest possible field of view. Incoming
light travels down a tubular baffle that absorbs

1.1 Hubble Space Telescope
Configuration
Figures 1-3 and 1-4 show the overall Telescope
configuration. Figure 1-5 lists specifications for
6-hour EVA period

Task Times(stand-alone)

Unsched EVA

EVA 4

EVA 3

EVA 2

EVA 1

Priority
STIS
2:00

NICMOS
2:00

FGS
2:45

DIU-2
2:00

SADE-2
2:05

OCEK
0:40

ESTR-2
1:05

SSR
1:10

RWA
0:50

MSS
Covers0:40

1. 1st-SI {NICMOS (FOS)}

2:30

2. FGS (FGS-1)

2:55

3. 1st-DR {SSR (ESTR-1)}

1:10

4. 2nd-SI {STIS (GHRS)}

2:15

5. RWA (RWA-1)

0:50

6. DIU-2

2:00

7. 2nd-DR {ESTR-2}

1:05

8. OCE Kit

0:40

9. SADE-2

2:05

10. MSS Covers

0:40

Required Setup

Task Times

Setup

1:00 (1st day)
0:15 (nth day)

Closeup

0:30 (nth day)
1:00 (last day)Contingencies (unscheduled)

RM January 8, 1997
K70110-102

Fig. 1-2 Schedule of extravehicular activities
K70110-01M

1-3

HIGH GAIN ANTENNA (2)FORWARD SHELLSECONDARYMIRROR(2)APERTURE
DOOR
LIGHT SHIELDPRIMARY MIRROR

AFT SHROUDCOSTAR AND
SCIENCE
INSTRUMENTS
AXIAL (4)RADIAL (1)FINE GUIDANCESENSORS (3)SOLAR ARRAY (2)K70110-103

Fig. 1-3 HST overall configuration• Faint Object Camera (FOC)• Faint Object Spectrograph (FOS)• Goddard High Resolution Spectrograph(GHRS)• Corrective Optics Space Telescope Axial
Replacement (COSTAR).

stray light. The light is collected by the concave94.5-in. (2.4 m) primary mirror and converged
toward the convex secondary mirror, which isonly 12.2 in. (0.3 m) in diameter. The secondary
mirror directs the still-converging light backtoward the primary mirror and through a 24-in.
hole in its center into the Focal Plane Structure,
where the science instruments are located.

In addition to the four axial instruments, fourother instruments are mounted radially (perpendicular to the main optical axis). Theseradial science instruments are:1.1.2 The Science Instruments
The Telescope can accommodate eight science
instruments. Four are aligned with theTelescope’s main optical axis and are mounted
immediately behind the primary mirror. Before
the Second Servicing Mission, the axial science
instruments consisted of:K70110-01M• Wide Field/Planetary Camera II (WF/PC II)• Three Fine Guidance Sensors (FGS).
Faint Object Camera. This camera records
images of faint, celestial objects. It reflects light
1-4

MAGNETIC TORQUER (4)LIGHT SHIELD

HIGH GAIN ANTENNA (2)SUPPORT SYSTEMS MODULEFORWARD SHELL

APERTURE
DOOR

OPTICAL TELESCOPE ASSEMBLYSECONDARY MIRROR ASSEMBLY
MAGNETOMETER (2)SECONDARY MIRROR BAFFLECENTRAL BAFFLE
OPTICAL TELESCOPE ASSEMBLY
PRIMARY MIRROR AND MAIN RINGFINE GUIDANCE OPTICALCONTROL SENSOR (3)OPTICAL TELESCOPEASSEMBLY FOCAL PLANE
STRUCTURE

MAIN BAFFLESOLAR ARRAY (2)OPTICAL TELESCOPE ASSEMBLY
METERING TRUSSAXIAL SCIENCE
INSTRUMENT
MODULE (3)AND COSTARSUPPORT SYSTEMS MODULE
EQUIPMENT SECTION
OPTICAL TELESCOPE ASSEMBLY
EQUIPMENT SECTION

LOW GAIN
ANTENNA (2)RADIAL SCIENCE
INSTRUMENT MODULE (1)FIXED HEAD STAR TRACKER (3)AND RATE GYRO ASSEMBLYSUPPORT SYSTEMSMODULE AFT SHROUD
K70110-104

Fig. 1-4 HST exploded view

lengths. Scientists can determine the chemicalcomposition, temperature, pressure, and turbulence of the stellar atmosphere producing
the light – all from spectral data.

down one of two optical pathways. The lightenters a detector after passing through filters ordevices that can block out light from brightobjects, thus allowing better backgroundimages. The detector intensifies an image, thenrecords it much like a television camera.
Images of faint objects can be built up overlonger exposure times. The total image is translated into digital data, transmitted to Earth,and then reconstructed at the Space Telescope
Science Institute (STScI).

The FOS detects detail in very faint objects,
such as those at great distances, and light ranging from ultraviolet to red spectral bands. The
GHRS detects fine detail in somewhat brighter
objects but only in ultraviolet light.
Both spectrographs operate essentially the same.
Incoming light passes through a small aperturethen through filters and diffraction gratings,
which work like prisms. The filter or gratingused determines the range of wavelength to be
examined and in what detail. Spectrographdetectors record the strength of each wavelength band and send it back to Earth.

Spectrographs. FOS and GHRS separate
incoming light into its component wavelengths, revealing information about the atomic
composition of the light source. These spectrographs can detect a broader range of wavelengths than is possible from Earth because
there is no atmosphere to absorb certain waveK70110-01M

1-5

Hubble Space Telescope (HST)
WeightLengthDiameter
Optical system
Focal length
Primary mirror
Secondary mirror
Field of view
Pointing accuracy
Magnitude range
Wavelength range
Angular resolution
Orbit
Orbit time
Mission24,500 lb (11,110 kg)
43.5 ft (15.9 m)
10 ft (3.1 m) Light Shield and Forward Shell
14 ft (4.2 m) Equipment Section and Aft Shroud
Ritchey-Chretien design Cassegrain telescope
189 ft (56.7 m) folded to 21 ft (6.3 m)
94.5 in. (2.4 m) in diameter
12.2 in. (0.3 m) in diameter
See instruments/sensors
0.007 arcsec for 24 hours
5 mv to 29 mv (visual magnitude)
1100 to 11,000 Å
0.1 arcsec at 6328 Å
320 nmi (593 km), inclined 28.5 degrees from equator
97 minutes per orbit
15 yearsWide Field/Planetary Camera II (WF/PC II)

Faint Object Camera (FOC)
WeightDimensions

700 lb (318 kg)
3 x 3 x 7 ft (0.9 x 0.9 x 2.2 m)

WeightDimensions

Principal investigator
Contractor

F. Duccio Macchetto, STScI/ESA
European Space Agency, Dornier Systems,
and Matra-Espace Corporation
f/151, f/75.5
14, 28 arcsec2
5 to 28 mv
1150 to 6500 Å

Principal investigator
Contractor

619 lb (281 kg)
Camera: 3.3 x 5 x 1.7 ft (1 x 1.3 x 0.5 m)
Radiator: 2.6 x 7 ft (0.8 x 2.2 m)
John Trauger, NASA/JPL
Jet Propulsion Laboratory

Optical modes
Field of view
Magnitude range
Wavelength range

f/12.9 (WF), f/28.3 (PC)
4.7 arcmin2 (WP), 0.3 arcmin2 (PC)
9 to 28 mv
1200 to 10,000 Å

Optical modes
Field of view
Magnitude range
Wavelength range

Near Infrared Camera and Multi-Object Spectrometer(NICMOS)Space Telescope Imaging Spectrograph (STIS)
WeightDimensions

825 lb (374 kg)
3 x 3 x 7 ft (0.9 x 0.9 x 2.2 m)

WeightDimensions

861 lb (390 kg)
7.1 x 2.8 x 2.8 ft (2.2 x 0.88 x 0.88 m)

Principal investigator
Contractor

Dr. Bruce E. Woodgate, GSFC
Ball Aerospace
MAMA: 24.9 x 24.9 arcsec
CCD: 51 x 51 arcsec
1024 x 1024
115 - 1000 nanometers

Principal investigator
Contractor
Field of view

Dr. Rodger I. Thompson, U. of Arizona
Ball Aerospace
51.2 x 51.2 arcsec
19.2 x 19.2 arcsec
11.0 x 11.0 arcsec
3 HgCdTe arrays
256 x 256 pixels

Field of view
Pixel format
Wavelength range

Detectors

Corrective Optics Space Telescope Axial Replacement(COSTAR)

Fine Guidance Sensors (FGS)
WeightDimensions
Principal investigator
Contractor
Astrometric models
Precision
Measurement speed
Field of view
Magnitude range
Wavelength range

485 lb (220 kg)
1.6 x 3.3 x 5.4 ft (0.5 x 1 x 1.6 m)
William H. Jefferys, U. of Texas
Perkin-Elmer Corporation
Stationary and Moving Target, Scan
0.002 arcsec2
10 stars in 10 min
Access: 60 arcmin2
Detect: 5 arcsec
4 to 18.5 mv
4670 to 7000 Å

WeightDimensions
Number of parts
Mirror diameter
Contractor
Project engineer
Project scientist
Instrument manager

640 lb (290 kg)
3 x 3 x 7 ft (0.9 x 0.9 x 2.2 m)
5300 (approx)
0.81 in. (20 mm), the size of a nickel
Ball Aerospace
James H. Crocker, STScI
Holand Ford, STScIPaul Geithner, GSFC

PurposeTo correct the effects of sphericalaberration on the FOC, FOS, and GHRS

K70110-105

Fig. 1-5 Telescope science instrument specifications

K70110-01M

1-6craft’s pointing system to keep the Telescope
pointed accurately at a target when one of thescience instruments is being used to take dataand (2) to act as a science instrument. Whenfunctioning as a science instrument, two of
the sensors lock onto guide stars, and thethird measures the brightness and relative
positions of stars in its field of view. These
measurements, referred to as astrometry, arehelping to advance knowledge of the distances and motions of stars, and may be useful in detecting planetary-sized companionsof other stars.

COSTAR. Mechanical arms on COSTAR position corrective mirrors in front of the aperturesof the FOC, the FOS, and the GHRS.
Investigations into the spherical aberration of
the primary mirror revealed the specific errorsthat caused the fault. This knowledge allowed
optical experts to design and fabricate small
pairs of corrective mirrors that successfully
refocused the light reflected from the primary
mirror before it entered the three axial science
instruments. The first mirror (M1) of each pairintercepts light from the Telescope’s secondary
mirror and reflects that light into a second mirror (M2), which in turn reflects light into thethree axial science instruments. The second
mirror of each pair is shaped to cancel out thespherical aberration from the primary mirror.
Such refocusing has increased the amount oflight energy entering the axial science instruments. This increase in optical performancehelped to restore the Telescope’s capability tonear-original expectations.

1.1.3 Support Systems Module
The SSM encloses the OTA and the science
instruments like the dome of an Earth-based
observatory. It also contains all of the structures, mechanisms, communications devices,
electronics, and electrical power subsystems
needed to operate the Telescope.
This module supports the light shield and an
aperture door that, when opened, admits
light. The shield connects to the forward shellon which the Solar Arrays (SA) and High
Gain Antennas (HGA) are mounted.
Electrical energy from the 40-ft (12-m) SAs
charges the spacecraft batteries to power all
HST systems. Four antennas, two high gain
and two low gain, send and receive information between the Telescope and the Space
Telescope Operations Control Center(STOCC). All commanding is done throughthe Low Gain Antennas (LGA).

Wide Field/Planetary Camera II. WF/PC II
is an electronic camera that records images at
two magnifications. The WF/PC II was developed by the Jet Propulsion Laboratory (JPL)team in Pasadena, California. They also built
the first WF/PC. When Hubble’s mirror was
found to be flawed, the science team immediately began working on an optical correctionthat could be built into WF/PC II. The new
design incorporated an optical correction by
the refiguring of relay mirrors in the opticaltrain of the cameras. Each relay mirror is polished to a prescription that compensates for theincorrect figure on HST’s primary mirror.
Small actuators fine-tune the positioning ofthese mirrors on orbit.

Behind the OTA is the Equipment Section, aring of bays that house the batteries and mostof the electronics, including the computerand communications equipment. At the rearof the Telescope, the aft shroud contains thescience instruments.

Fine Guidance Sensors. The three FGSs have
two functions: (1) to provide data to the spaceK70110-01M

1-7

1.1.4 Solar Arrays
The SAs provide power to the spacecraft. They
are mounted like wings on opposite sides of
the Telescope, on the forward shell of the SSM.
Each array stands on a 4-ft mast that supports
a retractable wing of solar panels 40 ft (12.2 m)long and 8.2 ft (2.5 m) wide.
The SAs are rotated so the solar cells face thesun. Each wing’s solar cells absorb the sun’senergy and convert that light energy into electrical energy to power the Telescope andcharge the spacecraft’s batteries, which are partof the Electrical Power Subsystem (EPS).
Batteries are used when the Telescope moves
into Earth’s shadow during each orbit.
Prior to the First Servicing Mission, as theTelescope orbited in and out of direct sunlight, the resulting thermal gradients caused
oscillation of the SAs that induced jitter in theTelescope’s line of site. This in turn caused
some loss of fine lock of the FGSs during
science observations. New SAs installed
during the First Servicing Mission withthermal shields over the array masts minimized
the effect.

1.1.5 Computers
The Hubble Telescope Data Management
Subsystem (DMS) contains two computers: theDF-224 flight computer and the Science
Instrument Control and Data Handling (SIC&DH) unit. The DF-224 performs onboard
computations and also handles data and command transmissions between the Telescope
systems and the ground system. The SI C&DH
unit controls commands received by the science instruments, formats science data, andsends data to the communications system fortransmission to Earth.
K70110-01M

1-8

During the First Servicing Mission, astronauts
installed a coprocessor to augment the capacity of DF-224 flight computer. The new 386coprocessor increased flight computer redundancy and significantly enhanced on-orbit
computational capability.

1.2 The Hubble Space Telescope
Program
Hubble Space Telescope represents the fulfillment of a 50-year dream and 23 years of dedicated scientific effort and political vision toadvance humankind’s knowledge of the universe. To accomplish the mission, the program
brought together an international communityof engineers and scientists, contractors, andinstitutions under the auspices of the NASA
Office of Space Science. On-orbit servicing is
the fifth phase of NASA’s program. The otherfour phases were:• Development, assembly, and test
• Launch and deployment• On-orbit verification of system and science
instrument functions• Initial operations to produce scientific data.
Marshall Space Flight Center managed the first
three phases. Figure 1-6 summarizes organizations and tasks for completed phases of theTelescope program.
The NASA Goddard Space Flight Center(GSFC) in Greenbelt, Maryland, has coordinated overall operational and servicing
mission preparations for the Telescope since
its deployment. GSFC works with Kennedy
Space Center (KSC), where the Second
Servicing Mission will be launched, and withJohnson Space Center (JSC), which operates
as Mission Control during servicing. JSC also
is responsible for Shuttle operations andPhaseOrganization

Development, assembly, and test

• Lockheed Martin Missiles & Space (formerly Lockheed Missiles & Space Company, Inc.) built or
supervised subcontract equipment development and the entire Support Systems Module (SSM);
assembled and tested complete HST.
• Hughes Danbury Optical Systems (formerly Perkin-Elmer) designed and built the Optical Telescope
Assembly, mirrors, and Fine Guidance Sensors.
• Principal investigators led teams to create original five science instruments.

Launch and deployment• Kennedy Space Center was the site of Shuttle launch.
• Johnson Space Center ran flight control and trained astronauts to deploy the Telescope.

Verification• Marshall Space Flight Center controlled functional system testing.
• Goddard Space Flight Center controlled science instrument verification testing.

Operations• Goddard Space Flight Center has overall program management during this continuous phase.
K70110-106

Fig. 1-6 Space Telescope program phase summary

oversees astronaut crew training and mission
safety.

Instrument
Faint Object
Camera (FOC)

Figure 1-7 summarizes major contributions ofscientists and companies to the developmentof the array of science instruments that will bein place following the Second Servicing
Mission. The roles of NASA centers and contractors for the fifth phase of the HST program,on-orbit servicing, are:F. D. Macchetto,
European Space
AgencyTeam
Subcontractor
Dornier Systems
British Aerospace
Matra-Espace

Near Infrared Camera Dr. Rodger I. Thompson, Ball Aerospaceand Multi-Object
University of Arizona
Spectrometer(NICMOS)(1) GSFC – Responsible for overall management of daily on-orbit operations of HST;
development, integration and test of replacement hardware, space support equipment, andcrew aids and tools.
(2) JSC – Overall mission management, flightcrew training, and crew aids and tools
responsibility.
(3) KSC – Overall management of launch andpost-orbit operations for mission hardware.
(4) Ball Aerospace – Designed, developed, andprovided the Science Instruments for the HSTFirst and Second Servicing Missions.
(5) Lockheed Martin – Provides personnel
support to GSFC to accomplish (a) the development, integration, and test of replacement
hardware and space support equipment; (b) sysK70110-01M

Principal
Investigator

Space Telescope
Imaging
Spectrograph (STIS)

Dr. Bruce E.Woodgate
Goddard Space
Flight Center

Ball Aerospace

Wide Field/
Planetary Camera II(WF/PC II)John Trauger,
Jet Propulsion
Laboratory

Jet Propulsion
Laboratory

Corrective Optics
Space Telescope
Axial Replacement(COSTAR)James H. Crocker,
Holland Ford,
STScl

Ball Aerospace

Fine Guidance
Sensors (FGS)William H. Jefferys,
University of Texas

Perkin-Elmer
Corporation

K70110-107

Fig. 1-7

Development teams for science instruments
and COSTAR

tem integration with the Space Transportation
System (STS); (c) launch and post-orbit operations; and (d) daily operations of the HST.
Major subcontractors for the Second Servicing
Mission include Hughes Danbury Optical,
Allied Signal, Jackson and Tull, Orbital
Sciences Corporation, Odetics, Honeywell,
ETA, and Hughes STX.
1-9

The HST program requires a complex networkof communications among GSFC, the Telescope,Space Telescope Ground System, and STScI.
Figure 1-8 summarizes the major organizations
that oversee the program. Figure 1-9 shows
communication links.

Function

Organization

Goddard Space Flight• Overall HST Program Management
Center• HST Project Management– Office of the Associate• Responsible overseeing all HST
Director of Flight Projects
operationsfor HST– HST Operations andGround Systems ProjectSpace Telescope
Operations Control
Center• Provides minute-to-minute spacecraft
control• Schedules, plans, and supports all
science operations when required• Monitors telemetry communications
data to the HSTSpace Telescope
Science Institute• Selects observing programs fromnumerous proposals• Analyzes astronomical data

Goddard Space Flight
Center– HST Flight Systems andServicing Project• Responsible for implementing HST
Servicing Program• Manages development of new HST
spacecraft hardware and service
instruments• Manages HST Servicing Payload
Integration and Test Program• Primary interface with the Space
Shuttle Program at JSC
K70110-108

Fig. 1-8

Organization summary for HST program
operational phase

STARLIGHT

TRACKING ANDDATA RELAYSATELLITE SYSTEM(TDRSS)DOMSATPAYLOAD OPERATIONSCONTROL CENTERSCIENCE
INSTITUTETDRSSGROUNDSTATIONSPACE TELESCOPEOPERATIONSCONTROL CENTER(STOCC)
K70110-109

Fig. 1-9 HST data collecting network
K70110-01M

1-10

HUBBLE SPACE TELESCOPE SECOND SERVICING MISSION
he Hubble Space Telescope (HST) is thefirst observatory designed for extensive
maintenance and refurbishment in orbit.
While other U.S. spacecraft have been retrievedor repaired by astronauts, none was so thoroughly designed for orbital servicing as HST.
Its science instruments and many other components were designed as Orbital Replacement
Units (ORU) – modular in construction withstandardized fittings and accessible to extravehicular astronauts. Features such as handrails
and footholds are built into the Telescope to help
astronauts perform servicing tasks in the Shuttlecargo bay as they orbit Earth at 17,000 mph.

T

ized modern astronomy. To achieve its fullpotential, it will continue to conduct many
years of extensive, integrated scientific observations, including follow-up work on its many
discoveries. The Telescope was designed withnumerous backup parts and safemode systems, but it is impossible and impractical todesign such a complex spacecraft with sufficient backups to handle every contingency
likely to occur during a 15-year mission.
Orbital servicing is the key to keeping theTelescope in optimum operating condition. The
Telescope’s orbital servicing plans address
three primary maintenance scenarios:

The Discovery cargo bay is equipped with several devices to help the astronauts: the Flight
Support System (FSS) to berth and rotate theTelescope; large, specially designed equipment
containers to house the ORUs; and a remotemanipulator arm from which astronauts canwork and be maneuvered as needed.

• Incorporating technological advances into
the science instruments. During theTelescope’s 15-year life, scientists and engineers will continue to upgrade the science
instruments. For example, when theTelescope was launched in 1990, work
already had begun on an improved Wide
Field/Planetary Camera (WF/PC). When aflaw in the primary mirror was discovered
later that year, Hubble planners redesigned
the new WF/PC’s mirrors to compensate for
the Telescope’s aberrated images and accelerated the development of that science instrument so it would be ready for the First
Servicing Mission in December 1993.
• Normal degradation of components.
Servicing plans take into account the need forroutine replacement of some items, for example, restoring HST system redundancy andlimited-life items such as tape recorders andReaction Wheels.
• Random equipment failure or malfunction.
Given the enormous scientific potential of theTelescope and the investment in designing,developing, building, and putting it into
orbit, NASA must be able to respond to andcorrect unforeseen problems that arise fromrandom equipment failures or malfunctions.

The Second Servicing Mission will benefit fromlessons learned on NASA’s previous on-orbit
repair missions, ranging from the 1984 Solar
Maximum repair mission to the 1993 HST First
Servicing Mission. NASA has incorporated
these lessons in detailed planning and training
sessions for astronauts Cdr. Kenneth Bowersox, Lt. Col. Scott Horowitz, Col. Mark Lee,
Gregory Harbaugh, Steven Smith, Joseph
Tanner, and Steven Hawley. All of NASA’splanning and the skills of the astronauts will be
put to the test during the 10-day Second
Servicing Mission in February. Four extravehicular activity (EVA) days are scheduled for theactual servicing.

2.1 Reasons for Orbital Servicing
The Hubble Telescope is an invaluable international scientific resource that has revolutionK70110-02M

2-1

NASA’s Space Shuttle program provides aproven system for transporting crews of fully
trained servicing technicians to the Telescope.
Originally, planners considered using the
Shuttle to return the Telescope to Earthapproximately every 5 years for maintenance,
but the idea was rejected for both technical andeconomic reasons. Returning it to Earth would
entail a significantly higher risk of contaminating or damaging delicate components.
Ground servicing would require an expensive
clean room and support facilities, including alarge engineering staff, and the Telescope
would be out of action for a year or more – along time to suspend scientific observations.
With on-orbit servicing by Shuttle astronauts,
most maintenance and refurbishment can beaccomplished within two weeks – a brief interruption to scientific operations – without theadditional facilities and staff needed forground servicing.

2.2 Orbital Replacement Units
Modularity. Engineers studied various technical and human factors criteria to simplify
Telescope maintenance. Because of the limited
time available for an EVA astronaut to makerepairs and the limitations in mobility and dexterity in the EVA environment, designers have
simplified the maintenance tasks by planningentire components for replacement. The modular ORU concept is key to successfully servicing the Telescope on orbit.
The typical ORU is a self-contained box installed
and removed using fasteners and connectors.
The ORUs range from small fuses to phonebooth-sized science instruments weighing
more than 700 lb (318 kg). Figure. 2-1 shows the
ORUs for the Second Servicing Mission.
Standardization. In addition to modularity, a
key feature that simplifies on-orbit repair is
ACRONYMS
DIU

Data Interface Unit

E/STR

Engineering andScience Tape
Recorder

FGS

Fine Guidance
SensorMSS

Magnetic Sensing
System

NICMOS Near-Infrared
Camera and
MultI-Object
Spectrometer
OCE-EK Optical Control
Electronics
Enhancement Kit
RWA

Reaction Wheel
Assembly

SADE

Solar Array Drive
Electronics

SSR

Solid State
RecorderSTISSpace Telescope
Imaging
Spectrograph
K70110-201

Fig. 2-1 Hubble Space Telescope Second Servicing Mission Orbital Replacement Units
K70110-02M

2-2

standardized bolts and connectors. ORU components are held in place by captive bolts with
7/16-in., double-height hex heads. To remove
or install them, astronauts require only a 7/16in. socket fitted to a power tool or manual wrench.

2.3 Shuttle Support Equipment
To assist the astronauts in servicing theTelescope, the Discovery will carry into orbit
16,000 pounds of hardware and Space Support
Equipment, including the Remote Manipulator
System (RMS), FSS, ORU Carrier (ORUC), andSecond Axial Carrier (SAC).

Although most ORUs require these fasteners,
some do not. When reliability assessments andground testing indicated that additional components may need replacing, some components were selected as replaceable units after
their design had matured. These exceptions
add a greater variety of fasteners to the servicing requirements, including noncaptive 5/16in.-head bolts and connectors without wingtabs. Despite these exceptions, the high level ofstandardization among units reduces the number of tools needed for the servicing mission
and also simplifies astronaut training.

2.3.1 Remote Manipulator System
The Discovery Remote Manipulator System(RMS), also known as the robotic arm, will be
used extensively during the Second Servicing
Mission. The astronaut operating this device
from inside the cabin is designated the intravehicular activity (IVA) crew member. The RMS
will be used to capture, berth, and release theTelescope; transport new components, instruments, and extravehicular activity (EVA)
astronauts between worksites; and provide a
temporary work platform for one of the EVA
astronauts.

Accessibility. To be serviced in space,
Telescope components must be seen andreached by an astronaut in a bulky pressure
suit, or they must be within range of the appropriate tool. Therefore, most ORUs are mounted
in equipment bays around the perimeter of thespacecraft. To access these units, astronauts
simply open a large door that covers the appropriate bay.

2.3.2 Space Support Equipment
Ground crews will outfit the cargo bay of
Discovery with three major assemblies essential
for the servicing mission: the FSS, ORUC, andSAC. Figure 2-2 shows a cargo bay view ofthese assemblies.

Crew Aids. Handrails, foot restraint sockets,and tether attachments are essential to efficient, safe on-orbit servicing. In anticipation ofservicing missions, 31 foot restraint sockets
and 225 ft of handrails were designed into theTelescope. The foot restraint sockets andhandrails greatly increase the mobility and stability of the EVA astronauts, giving them safe,
conveniently located worksites near ORUs.
Other crew aids such as portable lights, special
tools, installation guiderails, handholds, andportable foot restraints (PFR) also ease accessto the components being serviced.
K70110-02M

Flight Support System. The FSS is a maintenance platform used to berth the HST in the
cargo bay after the Discovery crew has rendezvoused with and captured the Telescope(see Fig. 2-3). The platform was adapted fromthe FSS first used for the 1984 Solar Maximum
repair mission. This system has a U-shaped
cradle that spans the rear of the cargo bay. The
platform has a circular berthing ring with threelatches that secure the Telescope to the cradle.
The berthing ring can rotate the Telescope
2-3

MULTI-MISSION
ORU PROTECTIVE
ENCLOSURE(MOPE)LARGE ORU
PROTECTIVE
ENCLOSURE(LOPE)

SMALL ORU
PROTECTIVE
ENCLOSURE(SOPE)ORBITALREPLACEMENTUNIT CARRIER(ORUC)HST
PORTABLE
HUBBLEFOOT
SPACE
RESTRAINTS TELESCOPE
(PFR)
(HST)

ASIPE

FSIPE

ORBITERDOCKINGSYSTEM
(ODS)

ASIPE

TOOLS
STOWAGEASSEMBLY(TSA)SECOND AXIALCARRIER (SAC)
INCHES050

100 150

CONTINGENCY ORU
PROTECTIVE
ENCLOSURE(COPE)
200

250 300 350

FORWARDFIXTUREMANIPULATORFLIGHT
FOOT RESTRAINT SUPPORT(MFR)
SYSTEM(FSS)
ASIPE – AXIAL SCIENTIFICOUTLINE
INSTRUMENT PROTECTIVE
ENCLOSURE
FSIPE – FINE GUIDANCE SENSORSCIENTIFIC INSTRUMENTPROTECTIVE ENCLOSURE
AFTFIXTURE

SOLAR
ARRAY
DEPLOYED 0°

K70110-202

Fig. 2-2 Second Servicing Mission configuration

almost 360 degrees (175 degrees clockwise orcounterclockwise from center) to give EVA astronauts access to every side of the Telescope.

it has provisions for safe transport of ORUs
to and from orbit (see Fig. 2-4). In thisassembly, the Fine Guidance Sensor (FGS) #1
and STIS are stored in a Science Instrument
Protective Enclosure (SIPE). The OCE-EK
primary and backup units, a primary andbackup Solid State Recorder (SSR),
Engineering/Science Tape Recorder (E/STR)#2, Solar Array Drive Electronics (SADE) #2,
a primary and backup Solid State Recorder(SSR) “T” harness, and the E/STR delay harness reside in a Small ORU Protective
Enclosure (SOPE). The Reaction Wheel
Assembly (RWA) is contained in a Large
ORU Protective Enclosure (LOPE) on the
ORUC shelf.

The FSS also pivots to lower or raise theTelescope as required for servicing or reboosting. The FSS’s umbilical cable provides powerfrom Discovery to maintain thermal control of
the Telescope and the ORUC, and permits
ground engineers to test and monitor
Telescope systems during the mission.

2.3.3 Orbital Replacement Unit Carrier
The ORUC is centered in Discovery’s cargobay. A Spacelab pallet modified with a shelf,
K70110-02M

2-4

FRONT VIEW

MAIN UMBILICALBERTHING LATCH (3)CCTV
ROTATORBACKUP UMBILICALDOWNLOCKPORTABLE FOOTRESTRAINTAUXILIARYSI HANDLE

FMDM-AC
FMDM-B

ORUC I/F
CONNECTOR
BRACKET/HANDRAILMSBVELCRO HARNESSTIE-DOWN (4 PLACES)
PDSU-1BAPS SUPPORT POST
(WITH LATCH MECHANISM)

HANDRAILSFWD

LOW GAIN ANTENNA COVER

+Z°
+X°
+Y°

K70110-203

Fig. 2-3 Flight Support System configuration

2.3.4 Second Axial Carrier

NICMOS and a complement of ORUs. The SAC
structure (see Fig. 2-5) is based on the Goddard
Pallet Assembly, a multimission Shuttle payload carrier built to support Shuttle-based
HST servicing activities. For the First
Servicing Mission, the Pallet Assembly wasconfigured as the Solar Array Carrier, with a
Solar Array Support Structure (SASS) designed
to transport HST Solar Arrays.

The SAC is being flown on the Second
Servicing Mission to transport and protect the

For the Second Servicing Mission, the Pallet
Assembly has been modified and is the base

The protective enclosures, their heaters, andthermal insulation control the temperature of
the new ORUs and provide an environment
equivalent to that inside the Telescope. Struts
and springs between the enclosures and thepallet protect the ORUs from the loads generated at liftoff.

K70110-02M

2-5

LARGE ORBITAL REPLACEMENTUNIT (ORU) PROTECTIVE ENCLOSURE(LOPE)
FINE GUIDANCE SENSOR (FGS)SCIENTIFIC INSTRUMENTPROTECTIVE ENCLOSURE(FSIPE)SMALL ORU PROTECTIVE
ENCLOSURE (SOPE)

ASIPE

AXIAL SCIENTIFICINSTRUMENTPROTECTIVE
ENCLOSURE(ASIPE)FORWARDFIXTUREAFT FIXTURE
POWER REGULATORJUNCTION UNIT
FGS PICKOFFMIRROR COVER

K70110-204

Fig. 2-4 Orbital Replacement Unit Carrier

structure of the SAC. The SASS has been
replaced with an Axial Scientific Protective
Enclosure (ASIPE) to transport into orbit the
new NICMOS. The Contingency ORU Protective
Enclosure (COPE) and a new ORU Protective
Enclosure, called the Multimission ORU
Protective Enclosure (MOPE), are used tohouse and thermally protect the ORUs, also
mounted on the Pallet Assembly.

automobile shock absorbers are used. Eachisolator consists of a spring and a magnetic
damper. The magnetic damper converts themechanical vibration energy from the launch
event into heat energy, reducing the loading on
NICMOS. The SAC incorporates eight isolators, configured in four sets of two each. Eachisolator is attached to the Pallet Assembly andalso to the ASIPE at its base or end plate.

The SAC spans the cargo bay and structurally
ties into the Shuttle at five points, four
longeron and one keel trunnion. To reducelaunch and landing vibration loads transmitted to the NICMOS, active isolators similar to

The SAC ASIPE is based on the ORUC ASIPE,
with design modifications to be flown as astand-alone enclosure. Aside from the unique
NICMOS/FOS latch configuration, all EVA
and instrument interfaces are the same for both

K70110-02M

2-6

AFT

LID
LATCH

AXIAL SIPE

AXIAL SIPE HINGEAXIAL SIPE LID
COPE
HANDRAILSPFR EXTENDER
MOPE LATCH
TOWER VENTSCREENSTRUNNION

STS PFRMOPE
DRAG STRUTEPDSU“Z” ISOLATOR“X” ISOLATORKEEL ASSEMBLY

FORWARD

LID LATCH

AXIAL SIPE
MOPE

AXIAL SIPE LID

MOPE LATCHHANDRAILSTRUNNION

TOWER VENT SCREENSSTS PFRAXIAL SIPE SAFETY BARDRAG STRUT“Z” ISOLATOR“X” ISOLATORCOPELOOKING UPWITNESSMIRROR BOX

BISTEM BRACE (8 PLACES)

K70110-205

Fig. 2-5 Second Axial Carrier
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2-7

ASIPEs. The ASIPE is thermally controlledusing a heater system designed to precondition
NICMOS prior to EVA changeout operations.
ORUs transported on the SAC for the Second
Servicing Mission are two DIUs in the MOPE
and two RSUs and an ECU in the COPE. The
MOPE will also carry electrical harnesses forthe DIUs, spare fuse plugs and push in-pullout (PIP) pins, and crew aids and tools. The
crew aids and tools are the Pistol Grip Tool, thePower Ratchet Tool and its controller, a DIU
portable handle, and the NICMOS cryo ventline. In addition to the ORUs, the COPE willhouse the MSS Covers, MLI Repair Kit, spareelectrical harnesses, and a JSC-provided tool
called the PFR Attachment Device.
Within the MOPE and COPE, transport modules protect ORUs from launch and landing
vibration loading. These transport modules aredesigned to custom fit each ORU using foam tosurround the ORU. Transport modules foreach ORU are integrated into the OPEs as amodule unit, simplifying system integration.
The SAC receives power directly from the
Orbiter through the Enhanced PowerDistribution and Switching Unit (EPDSU). The
EPDSU is modular avionics for Shuttle payloads designed to be readily configured for
unique payload power, command, and telemetry requirements. Custom applications are easily serviced using the EPDSU. An EPDSU, configured for the FSS, is also available for backup
to the FSS PDSU.

Center (MSFC) in Huntsville, Alabama; andGoddard Space Flight Center (GSFC) in
Greenbelt, Maryland. Although there has been
extensive cross-training, each crew member
also has trained for specific tasks.
Training for Mission Commander Bowersox
and Pilot Horowitz focused on rendezvous
and proximity operations, such as retrieval anddeployment of the Telescope. The two astronauts rehearsed these operations using JSC’s
Shuttle Mission Simulator, a computer-supported training system. In addition, theyreceived IVA training – helping the EVA astronauts into suits and monitoring their activities
outside the Discovery cabin.
The five Mission Specialists also received specific training, starting with classroom instruction on the various ORUs; tools and crew aids
available to assist them; Space Support
Equipment (SSE) such as the RMS (the roboticarm); and the FSS. Principal operator of therobotic arm is Mission Specialist Hawley, who
also performs IVA duties. The alternate operator is Pilot Horowitz.
Hawley and Horowitz trained together specifically for capture and redeployment of theTelescope, rotation and pivoting of theTelescope on FSS, and related contingencies.
These operations were simulated with JSC’s
Manipulator Development Facility, which
includes a mockup of the robotic arm and asuspended helium balloon with dimensions
and grapple fixtures similar to those of theactual Telescope.

2.4 Astronaut Roles and Training
To prepare for the Second Servicing Mission,
the seven-member Discovery crew trained
extensively at NASA’s Johnson Space Center(JSC) in Houston, Texas; Marshall Space Flight
K70110-02M

EVA crew members work in teams of two in
the cargo bay doing hands-on HST servicing.
To train for this important role, EVA astronauts
Lee, Harbaugh, Smith, and Tanner logged
many days of training in MSFC’s Neutral
2-8

In the weightlessness of space, the tiniest shove
can set 800-lb instruments, such as NICMOS or
STIS, into motion.

Buoyancy Simulator Facility (NBSF). In this40-ft (12-m)-deep water tank (see Fig. 2-6)
pressure-suited astronauts and their equipment were made neutrally buoyant, a condition that simulates weightlessness.

To simulate the delicate on-orbit conditions,models of the instruments are placed on pads
above a stainless steel floor and floated on athin layer of pressurized gas, giving crew
members an opportunity to practice carefully
nudging the instruments into their proper locations. In another room at JSC, the astronauts
worked in a vacuum chamber where temperatures varied from plus to minus 200 degrees
Fahrenheit, just as they do in orbit. Astronauts
also used virtual reality technologies in theirtraining. With this kind of ultrarealistic simulation, astronauts wearing virtual reality devices“see” themselves next to the Telescope as theirpartners maneuver them into position with therobotic arm.

Underwater mockups of the Telescope, FSS,
ORUs, ORUC, RMS, and the Shuttle cargo bay
enabled the astronauts to practice entire EVA
servicing procedures in simulated weightlessness. Such training activities help the astronauts to efficiently use the limited number ofdays (four) and duration (six hours) of each
EVA period during the servicing mission.
Additional underwater training took place atthe Weightless Environmental Training Facility(WETF) at JSC in Houston.
Other training aids at JSC helped recreate
orbital conditions for the Discovery astronauts.

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Fig. 2-6 Neutral Buoyancy Simulator at Marshall Space Flight Center
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2-9

2.5 Extravehicular Crew Aids and Tools
Astronauts servicing HST use two different
kinds of foot restraints to counteract theirweightless environment. When anchored in a
Manipulator Foot Restraint (MFR), an astronaut can be transported from one worksite to
the next with the robotic arm. With the PFR, anastronaut establishes a stable worksite by
mounting the restraint to any of 31 different
receptacles strategically placed on theTelescope or 22 receptacles on the ORUC, the
FSS, and the tool box.
In addition to foot restraints, EVA astronauts
have more than 150 tools and crew aids attheir disposal. Some of these are standard
items from the Shuttle’s toolbox; others are
unique to this servicing mission. All tools aredesigned to be used by astronauts wearing
pressurized gloves.
The most commonly used ORU fasteners arebolts with 7/16-in., double-height hex heads.
These bolts are used with three different kindsof fittings: J-hooks, captive fasteners, and keyhole fasteners. To replace a unit, the astronauts
use a 7/16-in. extension socket on a ratchet
wrench. Extensions up to 2 ft long are available
to extend an astronaut’s reach. Multisetting
torque limiters prevent over-tightening of fasteners or latch systems.
For units with bolts or screws that are not captive in the ORU frame, astronauts use tools fitted with socket capture fittings so that nothing
floats away in the weightless space environment. To grip fasteners in hard-to-reach areas,
astronauts use wobble sockets or flex-tip
screwdrivers.
Some ORU electrical connectors require special
tools. For these components, astronauts loosen
K70110-02M

standard D-connectors by carefully removing
them with a special gripper tool after releasing
two slotted screws. If connectors have no wingtabs, astronauts use another special gripper to get
a firm hold on the rotating sleeve of the connector.
Portable handles have been attached to manylarger ORUs. These handles greatly facilitate
removal or installation. Other tools and crew
aids used during the servicing mission are tool
caddies (carrying aids), tethers, transfer bags,
and protective covers for instrument mirrors,
umbilical connectors, and the Low Gain
Antenna (LGA). When astronauts work within
the Telescope’s aft shroud area, they must
guard against optics contamination by using
special tools that will not outgas or shed particulate matter.

2.6 Astronauts of the Second Servicing
MissionNASA carefully selected and trained the Second
Servicing Mission STS-82 crew (see Fig. 2-7).
All are experienced astronauts who have flown
missions involving either extensive EVA or thecapture and deployment of various types ofspacecraft. Their unique set of experiences andcapabilities makes them eminently qualified forthis challenging assignment.
Brief biographies of the STS-82 astronauts
follow.
Kenneth Bowersox: NASA Astronaut(Commander, USN). Kenneth Bowersox of
Bedford, Indiana, is the Commander of theSecond Servicing Mission. He received a bachelor of science degree in aerospace engineeringfrom the United States Naval Academy and amaster of science degree in mechanical engineering from Columbia University. This former
test pilot has logged more than 3,000 hours of
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K70110-02M

2-11

Fig. 2-7 The STS-82 mission has seven crew members. They are (front row, from left) Commander Kenneth D. Bowersox, Steven A.
Hawley, and Pilot Scott J. Horowitz; (back row) Joseph R. Tanner, Gregory J. Harbaugh, Mark C. Lee, and Steven L. Smith.

K70110-207

flight time. He joined the astronaut corps in1988. A three-flight veteran, Bowersox has
logged over 40 days in space. He flew as pilot
on STS-50 in 1992 and STS-61 in 1993, and commanded the spacecraft on STS-73 in 1995.
Scott J. “Doc” Horowitz, Ph.D.: NASA
Astronaut (Lieutenant Colonel, USAF).
Discovery pilot Scott Horowitz was born in
Philadelphia, Pennsylvania, but considers
Thousand Oaks, California, to be his hometown. He received a bachelor of science degree
in engineering from California State University
at Northridge in 1978; a master of science
degree in aerospace engineering from Georgia
Institute of Technology in 1979; and a doctorate
in aerospace engineering from the Georgia
Institute of Technology in 1982. Horowitz
served as pilot on STS-75 in 1996 and has
logged over 377 hours in space. The principal
payloads on STS-75 were the reflight of theTethered Satellite System (TSS) and the third
flight of the United States Microgravity
Payload (USMP-3).
Mark C. Lee: NASA Astronaut (Colonel,USAF). Payload Commander Mark Lee of
Viroqua, Wisconsin, received a bachelor of science degree in civil engineering from the U.S.
Air Force Academy in 1974 and a master of science degree in mechanical engineering from
Massachusetts Institute of Technology in 1980.
A veteran of three space flights, Lee has traveled over 9 million miles, going around theworld 368 times and spending over 550 hours
in orbit. He flew on STS-30 in 1989, STS-47 in
1992, and STS-64 in 1994. During STS-64, helogged EVA hours totaling 6 hours, 51 minutes.
He is currently the Chief of the Astronaut
Office EVA/Robotics Branch.
Gregory J. Harbaugh: NASA Astronaut.
Mission Specialist Gregory Harbaugh of
K70110-02M

Willoughby, Ohio, received a bachelor of science degree in aeronautical and astronauticalengineering from Purdue University in 1978
and a master of science degree in physical science from University of Houston-Clear Lake in
1986.
A veteran of three space flights, Harbaugh haslogged a total of 578 hours, 23 minutes inspace. He flew first as a Mission Specialist
aboard the Space Shuttle Discovery on STS-39
April 28 – May 6, 1991. Harbaugh then served
as Flight Engineer (Mission Specialist) aboard
Space Shuttle Endeavour on STS-54, January 13– 19, 1993. From June 27 to July 7, 1995,
Harbaugh flew as the Flight Engineer (Mission
Specialist) on a seven-member (up) eight-member (down) crew on Space Shuttle mission STS71. This was the first docking with the Russian
Space Station Mir, and involved an exchange ofcrews.
Steven L. Smith: NASA Astronaut. Mission
Specialist Steven Smith was born in Phoenix,
Arizona, but considers San Jose, California, tobe his hometown. He received a bachelor of science degree in electrical engineering, a masterof science degree in electrical engineering, anda master’s degree in business administrationfrom Stanford University.
Smith served as a mission specialist aboard theSpace Shuttle Endeavour on Mission STS-68 in
1994. His responsibilities during the 11-day
flight were split between Shuttle systems,
Space Radar Lab 2 (SRL-2, the flight’s primary
payload), and several experiments located in
the crew cabin.
From November 1994 until March 1996, Smith
was one of three astronauts assigned to duties
at the Kennedy Space Center as members of
the astronaut support team. The team was
2-12

responsible for Space Shuttle prelaunch vehicle
checkout, crew ingress and strap-in prior tolaunch, and crew egress post landing.
Joseph R. “Joe” Tanner: NASA Astronaut.
Mission Specialist Joseph Tanner received abachelor of science degree in mechanical engineering from the University of Illinois.
Tanner started working for NASA JSC in 1984
as an aerospace engineer and research pilot.
His primary flying responsibilities involved
teaching the astronaut pilots Space Shuttle
landing techniques in the Shuttle Training Aircraft
and instructing the pilots and mission specialists in the T-38. He has accumulated more than
7,000 hours in military and NASA aircraft.
Tanner flew aboard the Space Shuttle Atlantis
on the STS-66, November 3 – 14, 1994, performing the Atmospheric Laboratory for
Applications and Science-3 (ATLAS-3) mission.
ATLAS-3 was the third in a series of flights tostudy the Earth’s atmosphere composition andsolar effects at several points during the sun’s
11-year cycle.

In June 1990 Dr. Hawley left the Astronaut
Office to assume the post of Associate Directorof NASA’s Ames Research Center in Mountain
View, California. In August 1992 he returned to
JSC as Deputy Director of Flight Crew
Operations. In February 1996 Dr. Hawley
was returned to astronaut flight status andnamed to the crew of the HST Second
Servicing Mission.

2.7 Servicing Mission Activities
As the Discovery approaches the Hubble Space
Telescope in orbit February 1997, the sevenperson crew will begin one of the most ambitious servicing missions in the history of spacetravel. Four days of EVA activities are scheduled, with a fifth day available if needed. Each
EVA session is scheduled for six hours.

2.7.1 Rendezvous With the Hubble
Space Telescope

Steven A. Hawley, Ph.D.: NASA Astronaut.
Mission Specialist Steven Hawley of Salina,
Kansas, received bachelor of arts degrees inphysics and astronomy (graduating with highest distinction) from the University of Kansas
in 1973 and a doctor of philosophy in astronomy and astrophysics from the University ofCalifornia in 1977.

Discovery will rendezvous with the Hubble inorbit 316 nautical miles (506 km) above theEarth. As Discovery approaches the Telescope,
Commander Bowersox will control thethrusters to avoid contaminating HST with
propulsion residue. During this approach the
Shuttle crew will be in close contact with
Mission Control at JSC. In concert with theSpace Telescope Operations Control Center(STOCC) at GSFC, Mission Control will command HST to stow the High Gain Antennas(HGA) and close the aperture door.

Dr. Hawley was selected as a NASA astronautin January 1978. A veteran of three spaceflights, he served as a mission specialist on
STS-41D in 1984, STS-61C in 1986, and STS-31
in 1990. As a Mission Specialist on STS-31, Dr.
Hawley used the RMS to deploy the Hubble
Space Telescope.

As the distance between Discovery and HST
decreases to approximately 200 ft (60 m), the
STOCC ground crew will command HST to
perform a final roll maneuver to position itself
for grappling. The SAs will remain fully
deployed parallel to the optical axis of theTelescope.

K70110-02M

2-13

When Discovery and HST achieve the proper
position, Mission Specialist Hawley will operate the robotic arm to grapple the Telescope.
Using a camera mounted at the base of the FSS
platform in the cargo bay, he will maneuver
HST to the FSS, where the Telescope will beberthed and latched.
Once the Telescope is secured, the crew will
remotely engage the electrical umbilical andswitch Hubble from internal power to external
power from Discovery. The pilot also will
maneuver the Shuttle so that the SAs face thesun to recharge the Telescope’s six onboard
nickel-hydrogen (NiH2) batteries.

2.7.2 Extravehicular Servicing
Activities – Day by Day
Figure 2-8 shows the schedule for four six-hour
EVA servicing periods. These time spans areplanning estimates; the schedule will be modified as needed as the mission progresses.
During most of the EVAs, HST will be vertical
relative to Discovery’s cargo bay. Four EVA mission specialists will work in two-person teamson alternate days. One team is Mark Lee andSteve Smith; the other is Greg Harbaugh and
Joe Tanner.
One astronaut, designated EV1, accomplishes
primarily the free-floating portions of the EVA
tasks. He can operate from a PFR or can accomplish the tasks free floating. The other astronaut, EV2, works from an MFR mounted on
Discovery’s robotic arm (RMS), removing andinstalling the ORUs on the Hubble. EV1 assists
EV2 in removal of the ORUs and installation of
the replaced units in the Second Servicing
Mission carriers. Inside Discovery’s aft flightdeck, the off-shift EVA crew members and thedesignated RMS operator assist the EVA team

6-hour EVA period
Task Times(stand-alone)

Unsched EVA

EVA 4

EVA 3

EVA 2

EVA 1

Priority
STIS
2:00

NICMOS
2:00

FGS
2:45

DIU-2
2:00

SADE-2
2:05

OCEK
0:40

ESTR-2
1:05

SSR
1:10

RWA
0:50

MSS
Covers0:40

1. 1st-SI {NICMOS (FOS)}

2:30

2. FGS (FGS-1)

2:55

3. 1st-DR {SSR (ESTR-1)}

1:10

4. 2nd-SI {STIS (GHRS)}

2:15

5. RWA (RWA-1)

0:50

6. DIU-2

2:00

7. 2nd-DR {ESTR-2}

1:05

8. OCE Kit

0:40

9. SADE-2

2:05

10. MSS Covers

0:40

Required Setup

Task Times

Setup

1:00 (1st day)
0:15 (nth day)

Closeup

0:30 (nth day)
1:00 (last day)Contingencies (unscheduled)

RM January 8, 1997
K70110-102

Fig. 2-8 Schedule of extravehicular activities
K70110-02M

2-14

by reading out the procedures and operating
the RMS.

ments, and installation and deployment of the
BAPS Support Post (BSP).

At the beginning of the first EVA day (thefourth day of the mission), the first team of
EVA astronauts suit up and pass through the
Discovery airlock into the cargo bay. To prevent
themselves from accidentally floating off, they
attach safety tethers to a cable running along
the cargo bay sills.

The BSP is required to dampen the vibrationthat the servicing activities will induce into thedeployed SAs. Prior to its installation on theinitial day, HST is commanded to an 85-degree
pivot by the IVA team. The two center push inpull out (PIP) pins are installed each day andremoved each night in the event that the
Shuttle must make an emergency return toEarth. EV1 removes the BPS from its stowage
position in the cradle of the FSS and hands theforward end to EV2 who installs his end to the
BAPS ring with a PIP pin. EV1 then installs theaft end of the BPS to the FSS cradle with a PIP pin.
Finally the BPS is commanded to its 90-degree
limit and the two center PIP pins are installed.

EV1 accomplishes a variety of specific tasks to
prepare for that day’s EVA servicing activities.
These include removing the MFR from its
stowage location on the RMS grapple fixture;
installing the Low Gain Antenna Protective
Cover (LGA PC); deploying the Translation
Aids (TA); and removing the BAPS Support
Post from its stowage location and installing iton the FSS with the assistance of EV2.
Meanwhile, EV2 brings out of the airlock the
Crew Aids and Tools (CAT) that will be
attached to the handrails of the RMS. He transfers the CAT to EV1 to secure on the Second
Axial Carrier (SAC) ASIPE, opens the contamination mirror box lid, and reconfigures the
SAC aft PFR. He installs the CAT handrail to
the RMS and the color television camera(CTVC) on the MFR. Next he is moved to theBerthing and Positioning System (BAPS) Postinstallation worksite by the IVA RMS operator
to install the BAPS Post forward end on the BAPS.
EVA Day 1: Change-out of Goddard High
Resolution Spectrograph with Space Telescope
Imaging Spectrograph, and Faint Object
Spectrograph with Near Infrared Camera andMulti-Object Spectrometer. During EVA Day
1, Lee and Smith are scheduled to install two
new axial scientific instruments. They begin
with the initial setup, which includes installation of the LGA protective cover, TA deployK70110-02M

When the BPS is installed and other initial
setup tasks are completed, the crew starts the
specific tasks for the SI change-outs. EV1
deploys the ORUC Aft Fixture, which temporarily secures the old instrument while the new
instrument is installed. EV1 and EV2 translate
and open the HST V3 aft shroud doors.
Working as a team on the change-out of the
GHRS, EV1 demates and secures the GHRS
Repair Kit and HST side of the electrical anddata connectors and the ground strap. EV2
then loosens the A latch and grasps the GHRS
EVA handles. EV2 applies downward pressure while EV1 loosens the B latch. To assist
the crew in loosening and tightening the A
and B latches, indicator lights show the statusof the latches. When the latches are loosened,
EV2 slowly begins to remove the GHRS fromthe HST while EV1 provides clearances anddirection calls.
When the GHRS is removed, EV1 does a bay
inspection and EV2 translates to the aft fixture,
2-15

stows the GHRS into the aft fixture using the
GHRS EVA handles, and secures the instrument with a PIP pin. EV1 then translates to the
ORUC Axial Scientific Instrument Protective
Enclosure (ASIPE), opens the ASIPE lid, anddisconnects the ASIPE ground strap. Bothastronauts then position themselves at thelatches. EV1 again loosens the A latch and positions himself on the STIS EVA handles. EV2
loosens the B latch while EV1 applies slight aftpressure to keep STIS against the B latch. Withlatches disengaged, EV2, with clearance instructions from EV1, removes STIS from the ASIPE.
While the IVA team moves EV2 into position
for STIS insertion into the HST, EV1 closes theASIPE lid and latches one latch. When EV1 andEV2 are in position, EV2 begins insertion of the
STIS into the HST (see Fig. 2-9). Again, clearance and directional instructions by EV1 arecritical to successful insertion of the STIS. Afterthe STIS is inserted, EV1 tightens the B latch. In
parallel to the A latch tightening by EV2, EV1
mates the ground strap and the electrical anddata connectors. When EV1 has mated the connectors, IVA gives MCC a “go” for the STIS
aliveness test.
Next, EV2 translates to the aft fixture andremoves the GHRS from the fixture. While EV2
is translating with the GHRS to the ASIPE
worksite, EV1 opens and secures the ASIPE lid.

K70110-209

Fig. 2-9 Change-out of Goddard High Resolution Spectrograph and Space Telescope Imaging Spectrograph
K70110-02M

2-16

Again with instructions, EV2 inserts the GHRS
into the ASIPE. EV1 tightens the B latch andEV2 tightens the A latch. EV1 then installs theground strap and removes the GHRS Repair
Kit from the GHRS and secures it where it will
be stowed in the Contingency ORU Protective
Enclosure (COPE) on the SAC, after NICMOS
is installed and the HST aft shroud doors areclosed.
The change-out of the FOS, and replacement by
NICMOS, uses the same scenario as the GHRS
and STIS. EV2 and EV1 translate to the HST
where EV2 loosens the A latch and applies
downward pressure while EV1 loosens the B
latch. EV2 with EV1 instructions removes the
FOS. While EV2 installs FOS into the aft fixture,
EV1 translates to the SAC ASIPE, opens the lid,
and removes the ground strap from NICMOS.
With EV1 and EV2 at the SAC ASIPE, EV2loosens the A latch and EV1 loosens the B latch.
EV2 carefully removes NICMOS with EV1
instructions. While EV2 translates to HST, EV1
closes the ASIPE lid and latches one latch.
At the HST, EV2 inserts NICMOS with assistance from EV1. As before, EV1 tightens the B
latch and EV2 tightens the A latch. EV1 mates
the NICMOS ground strap and electrical anddata connectors while EV2 removes the cryo
vent cover and translates to the Multiple ORU
Protective Enclosure (MOPE), opens the lid,
removes the NICMOS cryo vent line from its
stowage location, and stows the cryo vent
cover. After EV1 mates all of the electrical connectors, IVA gives MCC a “go” for the NICMOS aliveness test. EV2 closes the lid andsecures five latches, routes the vent line under
and to the center guide rail, installs the NICMOS cryo vent line on the HST cryo vent hole,and secures the line to the guide rail standoff.
EV1 opens the NICMOS capillary line plug,routes the vent line through the HST aft bulkK70110-02M

head handrail, and attaches the vent line to NICMOS. EV1 then opens the cryo vent line valve
and secures the vent line to the handrail.
EV2 completes close-out photographs. Then
EV1 and EV2 close and secure the doors. Next,
EV1 opens the SAC ASIPE lid so that EV2 canremove the FOS from the ORUC aft fixture.
EV1 then closes the aft fixture and stores the
GHRS repair kit in the COPE. EV2 inserts the
FOS into SAC ASIPE with assistance from EV1.
When the FOS is fully inserted, EV1 tightens
the B latch and EV2 tightens the A latch. In parallel to the A latch tightening, EV1 installs theground strap. EV1 closes and secures theASIPE lid and lid latches. EV1 then configures
the SAC aft PFR to launch configuration.
Meanwhile, EV2 prepares the MFR handhold
for return to the airlock. For the daily close-out,
EV1 removes the center pins on the BPS andretracts the TAs while EV2 prepares the CAT
installed on the MFR handrail for return into
the airlock. Additionally, EV1 releases the MFRsafety tether from the grapple fixture for contingency Earth return and releases the lower
CTVC cable. After the completion of EVA Day
1, both astronauts return to the airlock with theDay 1 CAT installed on the MFR handrail.
EVA Day 2: Change-out of Fine Guidance
Sensor #1 and Engineering/Science Tape
Recorder #2, and Installation of Optics
Control Electronics Enhancement Kit. During
EVA Day 2, EVA astronauts Harbaugh andTanner are scheduled to change out a malfunctioning Fine Guidance Sensor (FGS) in position
number 1 (see Fig. 2-10) and theEngineering/Science Tape Recorder (E/STR) in
position number 2 (see Fig. 2-11). To increase
the capability of the FGS, installation of the
Optical Control Electronics Enhancement Kit(OCE-EK) also is scheduled.

2-17

K70110-210

Fig. 2-10 Installation of Fine Guidance Sensor #1

Fewer daily setup tasks are required for EVA
Day 2 than for EVA Day 1. EV2 still exits theairlock with the EVA Day 2 required CAT
installed on the MFR handrail. EV1 reconnects
the safety strap on the MFR and connects the
CTVC cable to the RMS end effector. EV2 theninstalls the MFR handrail and the CTVC on the
RMS. EV1 simultaneously deploys the TAs,
installs the BPS center PIP pins, and deploys
the aft fixture for FGS use. When EV2 has completed his daily set, he gets the PFR and installs
it in the HST PFR socket for EV1 FGS changeout use. EV1 and EV2 open and secure the FGS
bay doors, and EV2 installs four guide studs on
the FGS. While EV1 demates the FGS connectors and ground strap, EV2 translates to theK70110-02M

ORUC forward fixture to retrieve an FGS handhold. EV2 installs the FGS handhold andloosens the A latch. The FGS is now ready to be
removed.
With EV1 in the PFR and EV2 holding the
FGS handhold, the team removes the FGS as
EV1 gives clearance instructions. To ensure a
successful installation of the replacement
FGS, the team partially removes the old FGS
and then reinstalls it without latching thelatches or mating connectors for mass handling evaluation.
After this practice insertion, EV2 removes the
FGS and stows it on the aft fixture. EV1 then
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K70110-211

Fig. 2-11 Replacement of Engineering/Science Tape Recorder #2 in Bay 8

conducts a bay inspection. Next, EV1 translates
to the ORUC FGS Scientific Instrument
Protective Enclosure (FSIPE), opens the lid,
latches the lid in the open position, anddemates and secures the ground strap. In parallel, EV2 retrieves the other FGS handholdfrom the ORUC forward fixture and positions
himself for installation on the replacement FGS
in the FSIPE.
After installing the handhold on the FGS, EV2loosens the A latch. EV2, with instructions andassistance from EV1, removes the FGS from the
FSIPE and translates to the HST with the FGS.
EV1 closes the FSIPE lid and engages one lid
latch. He then ingresses the PFR mounted on
K70110-02M

the HST and prepares to remove the FGS mirror cover. EV2 presents the FGS to EV1 so that
the FGS mirror cover can be removed. EV1
tethers to the cover, releases the slide lock lever,and operates the handle to remove the mirror
cover. When the mirror cover is removed, IVA
repositions the EV2 and MFR so that the FGS
can be installed. EV2 inserts the FGS with assistance from EV1. EV2 tightens the A latch,
removes the FGS handhold, and stows it temporarily on the MFR. EV1 installs the groundstrap and electrical connectors. After the electrical connectors are mated, IVA informs MCC
to perform an aliveness test. EV2 returns to the
FGS bay, takes the close-out photographs, closes
and latches the doors, and repositions the scuff
2-19

plates to launch configuration. He thenretrieves the HST PFR and stows it for later
installation. EV1 stows the mirror cover on theORUC.
EV1 translates to the FSIPE and opens the lid.
EV2 stows the FGS handrail in the forward fixture, then translates to the ORUC aft fixture,
removes the FGS, and translates to the FSIPE
with the FGS. EV1 assists EV2 during the FGS
insertion into FSIPE. After full insertion, EV2
tightens the A latch while EV1 mates theground strap. EV1 releases the FSIPE on-orbit
latches, closes and secures the lid, and closes
the aft fixture. At this time EV2 retrieves the
PFR that was used for the FGS change-out andreinstalls it on the Shuttle bay adaptive payload carrier.
EV2 then translates to Bay 8 for the change-outof the E/STR #2. With assistance from EV1,
EV2 opens and secures the bay door. EV2 thendemates the three connectors, ensuring that thepower-on connector P1 is prevented fromtouching structure ground. He secures the P1
connector behind the Pointing/SafemodeElectronics Assembly (PSEA) cable harnesses.
After removing the four keyhole fasteners, EV2removes the E/STR and carefully transports itto the Small ORU Protective Enclosure (SOPE).
EV1 translates ahead of EV2 and opens the
SOPE lid and the E/STR #2 transport module.
EV2 hands EV1 the old E/STR and retrieves
the new E/STR from its transport module. EV1
installs the E/STR into the E/STR #2 transportmodule and secures the transport module lid
latch. EV1 then retrieves the timing cable andcloses the SOPE lid and engages three latches.
He then translates to Bay 3 to assist E/STR
installation.
Meanwhile, EV2 transports the E/STR up to Bay
8, being very careful not to exceed the E/STR’s
K70110-02M

angular acceleration limits. EV2 then installs
the E/STR and torques the four keyhole fasteners. EV1 transfers the timing cable to EV2 who
installs it in line with the HST and E/STR sidesof the P2 connector. The other connectors are
mated, close-out photographs are taken, andthe bay door is secured via its six J-hooks.
HST is rotated to the -V3 side in preparation for
the OCE-EK installation. As part of the OCEEK preparation, EV1 changes places with EV2
on the RMS (see Fig. 2-12). When the crew
member swap and HST rotation have been
completed, EV1 on the RMS translates to the
Optical Telescope Assembly (OTA) Bay C door
and opens the door via the three J-hooks. EV2
has returned to the SOPE, opened the three lid
latches, retrieved the OCE-EK from its stowage
pouch, and closed the SOPE lid latches.
While EV2 translates to Bay C, EV1 demates
the P6 and P10 connectors from the OCE. EV1
then mates the OCE-EK J/P6 A and J/P10A
connectors with the OCE J6 and J10 and HST
P6 and P10 connectors. When this is completed,
EV1 removes the connector cover from the
OCE J13 connector and mates the OCE-EK P13
connector to the OCE. The close-out photographs
are taken and the Bay C door is secured with its
three J-hooks.
For the EVA Day 2 daily close-out, EV2removes the center pins on the BPS, inspects
the FSS main umbilical mechanism, and retracts
the TAs while EV1 prepares the CAT installedon the MFR handrail for return to the airlock.
Additionally, EV1 releases the MFR safety tether
from the RMS in the event of contingency Earthreturn. Both astronauts return to the airlock
with the MFR handrail with CAT installed.
EVA Day 3: Change-outs of Data Interface
Unit #2 and Solid State Recorder #1 for
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K70110-212

Fig. 2-12 Installation of Optical Control Electronics Enhancement Kit in Bay C

Engineering/Science Tape Recorder #1 andReaction Wheel Assembly #1. The daily setup
for EVA Day 3 is almost identical to EVA Day 2
except that EV1 starts the day in the MFR. EV1
still exits the airlock with the required CAT and
CTVC installed on the MFR handrail andtransfers them to EV2. EV2 reconnects the safety
strap on the MFR/RMS and connects the
CTVC to the RMS end effector. EV1 theninstalls the MFR handrail and the CTVC on the
RMS. EV2 simultaneously deploys the TA andinstalls the BPS center PIP pins. EV1 also
retrieves the DIU gender changers and DIU
portable ORU handle after EV2 opens theMultimission ORU Protective Enclosure(MOPE).
K70110-02M

EV1 translates to SSM Bay 3 and operates the
six J-hooks to open the door. With assistancefrom EV2, EV1 demates the powered-on P1,
P2, and P3 connectors and mates them withgender changers to eliminate the possibility ofthem touching structure ground and shorting
out. The gender changers are secured out of theworksite. EV1 then demates the remaining 15
connectors. Two of the 20 connectors are notused on DIU #2 and have connector covers
installed on them. The DIU portable ORU handle is installed on the DIU to enable the EVA
crew members to handle and tether to it before
the six 7/16-in. noncaptive fasteners are loosened to remove the DIU. When the DIU isremoved, EV1 translates on the RMS to the
2-21

MOPE where EV2 has opened the lid andraised the tool board that holds the replacement DIU in its transport module. EV2 secureshimself and receives the old DIU from EV1
who, in turn, opens and removes the replacement DIU. EV2 installs the old DIU in thetransport module, removes the DIU portable
ORU handle, closes the transport module, andthen lowers the tool board to its stowed position. EV2 then closes and secures one MOPE
lid latch, and translates to the Bay 3 worksite toassist with the DIU installation.
While EV2 is at the MOPE worksite, EV1 translates to the Bay 3 DIU #2 worksite (see Fig. 2-13).

EV1 installs the DIU using the two center alignment marks as aids and tightening the four
captive fasteners. Once the DIU is installed by
the fasteners, EV1 mates the 18 connectors to
the harness extensions mounted on the replacement DIU. When the connectors are mated,
EV1 removes the cable restraint device andtransfers it to EV2. EV1 then takes the closeout photographs and closes and secures thebay door via the six J-hooks.
After the DIU #2 change-out is completed, EV1
and EV2 swap positions on the RMS, and HST
is rotated approximately 110 degrees to alignthe SSR and RWA worksite forward. The next

K70110-213

Fig. 2-13 Replacement of Data Interface Unit #2 in Bay 3
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2-22

change-out task is replacement of the E/STR#1 with a newer technology SSR #1. EV2 translates from the swap position to SSM Bay 5
where he opens the door via six J-hooks. EV2
demates the three connectors, being careful not
to allow the powered-on P1 connector to touch
structure ground. He secures the P1 connector
behind the cable harness along the side wall,
demates the four keyhole fasteners, andremoves the E/STR.
While EV2 translates to the SOPE with the
E/STR, EV1 translates to the SOPE via the HST
and Shuttle handrails. EV1 opens the SOPE lid
latches and the lid. EV2 installs the E/STR
directly into the E/STR #1 transport module,

then removes the SSR from its transport module and translates back to the Bay 5 worksite.
EV1 secures the transport module, removes the“T” harness from its stowage pouch, and closes
and latches the SOPE lid. He then translates to
Bay 5 to assist in SSR installation.
In parallel to closure of the SOPE by EV1, EV2
installs the SSR into the #1 position (see Fig. 2-14).
The SSR #1 is mounted via the same four keyhole bolts as E/STR #1. The SSR #1 to E/STR#3 “T” harness is installed from the SSR P4
connector to a “T” at the E/STR #3 P2 connector. The “T” harness provides increased capability to the E/STR and HST. EV2 then mates
the two remaining connectors and takes theK70110-214

Fig. 2-14 Installation of Solid State Recorder in Bay 5
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2-23

close-out photographs. With the assistance ofEV1, EV2 closes the door and secures it withthe six J-hooks.
EV2 then translates to the RWA worksite in Bay
6. The bay doors are opened after the six J-hookfasteners are operated. Four connectors must
be demated for removal of the RWA: two small
heater connectors and two RWA electronic box
connectors. In addition, three keyhole fasteners
that secure the RWA to its isolators must beloosened for RWA removal. Once the fasteners
are removed, EV2 removes the RWA and translates to the ORUC Large ORU Protective
Enclosure (LOPE). EV1 opens the LOPE lid byoperating the lid J-hooks. EV2, with assistancefrom EV1, stows the old RWA out of the way andremoves the replacement RWA from the LOPE.
With the new RWA secure, EV1 and EV2 install
the old RWA into the LOPE via the three keyhole fasteners. EV1 closes the LOPE lid while
EV2 translates to Bay 6. EV2 installs the RWA
using the three keyhole fasteners. After tightening the fasteners, EV2 mates the two heater
connectors and two electronic box connectors.
Close-out photographs are taken and the bay
door is closed and secured by the six J-hooks.
For the Day 3 EVA close-out, EV1 removes thecenter pins on the BPS, inspects the FSS mainumbilical mechanism, and retracts the TAs
while EV2 prepares the CAT installed on the
MFR handrail for return into the airlock. EV2
releases the MFR safety tether from the RMS in
the event of contingency Earth return. During
the daily close-out, the HST is reboosted to ahigher altitude. After completion of EVA Day 3
close-outs, both astronauts return to the airlock
with the MFR handrail with CAT installed.
EVA Day 4: Change-out of Solar Array Drive
Electronics (SADE) #2 and installation of theK70110-02M

Magnetic Sensing System (MSS) Covers.
EVA Day 4 has a daily setup similar to that ofthe prior EVA days. EV1 starts the day in the
MFR. EV1 exits the airlock with the required
CATs installed on the MFR handrail and theCTVC and hands them to EV2. EV2 reconnects
the safety strap on the MFR/RMS and connects
the CTVC cable to the RMS end effector. EV1
then installs the MFR handrail and the CTVC
to the RMS. EV2 simultaneously deploys theTAs and installs the BPS center PIP pins.
Upon completion of the daily setup, EV2 andEV1 translate to SSM Bay 7 and open the door.
Both astronauts then translate to the SOPE andopen the lid. EV2, on the MFR, removes the
SADE #2 from its transport module. While EV2translates back to Bay 7, EV1 closes the SOPE
lid, engages one latch, then translates to Bay 7
to assist EV2. EV2 secures the new SADE #2 tothe door and begins the demate of the old
SADE connectors. As each old SADE connector
is demated, it is mated to the corresponding
replacement SADE harness extension.
After the replacement SADE harnesses havebeen mated, EV2 demates the six noncaptive
5/16-in. fasteners that secure the SADE. The
fasteners are installed into the fastener retention block (FRB) mounted on the new SADE.
After the old SADE is removed and secured on
the MFR, the new SADE with mated connectors is installed into the SADE #2 position
where it is secured in place by its two alignment pins and four captive 5/16-in. fasteners(see Fig. 2-15). The FRB is mounted on the old
SADE for “T” fastener installation. The closeout photographs are taken and the bay door isclosed and secured. EV2 translates with SADE
#2 and installs it in the SOPE transport module. EV2 and EV1 then close and secure the
SOPE lid and latches. After the SOPE closure,the BPS center PIP pins are removed to enable2-24

K70110-216

Fig. 2-15 Replacement of Solar Array Drive Electronics #2 in Bay 7

the EVA crew members to gain access to the
MSS for installation of the MSS covers. Before
they can reach the MSS, HST must be pivoted
and rotated. The task is very similar to the HSTFirst Servicing Mission task when the MSS sensor covers were manufactured on orbit. In parallel to HST positioning, both crew members
translate to the Contingency ORU Protective
Enclosure (COPE), open the COPE lid, remove
the MSS covers from their stowage brackets,and close and secure the COPE lid and latches.
The IVA RMS operator translates the MFR to
the MSS worksites almost 40 ft above
Discovery’s bay. With assistance from EV2,
EV1 installs the MSS covers on one and thenthe other MSS (see Fig. 2-16). After the MSS
K70110-02M

covers are installed and close-out photographs
are taken, the crew returns to the Shuttle bay.
For final close-out, EV2 stows the TAs, removes
and stows the LGA-PC, inspects the FSS mainumbilical mechanism, and does a final sweepof the bay to ensure that everything is in placefor Earth return. EV1 installs all the CAT on the
MFR handrail for return to the cabin anddemates and stows the MFR from the RMS. On
the way back to the cabin, EV1 closes andsecures the contamination mirror box lid. In
parallel to the final close-out, the HST is
reboosted to a higher altitude. When the tasks
are complete, both crew members enter the airlock with the CAT-laden MFR handrail.
2-25

K70110-216

Fig. 2-16 Attaching cover to Magnetic Sensing System #2

EVA Contingency Day. A contingency day
exists in which crew members can accomplish
any required or contingency activities. This
will be decided by the mission support management and personnel at MCC.
Day 5: Redeploying the Hubble Space
Telescope. The day following the scheduled
servicing tasks, and any contingency EVAs,
will be devoted to the redeployment of the
HST into Earth orbit (see Fig. 2-17).
The SAs are slewed to the sun to generate electrical power for the Telescope and to charge thebatteries, and the High Gain Antennas (HGA)
are commanded to their deployed position.
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When the battery charging is complete, theRMS crew member guides the robotic arm toengage HST’s grapple fixture. The groundcrew commands Hubble to switch to internal
power. This accomplished, crew members
command Discovery’s electrical umbilical todemate from Hubble and open the berthinglatches on the FSS. If any Telescope
appendages fail to deploy properly, two mission specialists can make a final EVA effort tomanually override any faulty mechanisms.

2.8 Future Servicing Plans
As the Hubble Space Telescope enters the next
century, other enhancements are planned. A
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K70110-217

Fig. 2-17 Redeploying the Space Telescope

third-generation instrument, the Advanced
Camera for Surveys (ACS), will greatly
enhance the Telescope’s imaging capabilities.
Shuttle astronauts plan to install the camera on
Hubble during the third servicing mission
scheduled for 1999.

ence instruments. The unique characteristics
and dramatically improved efficiencies of the
ACS will anchor the Hubble science program at
the turn of the 21st century, exploiting the fullpotential of the telescope to serve the needs ofthe science community.

ACS is truly an advanced camera, with predicted performance improvements one to twoorders of magnitude over existing Hubble sci-

Periodic upgrades and servicing will ensurethat the Telescope continues to yield remarkable advances in our knowledge of the universe.

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2-27

HUBBLE SPACE TELESCOPE SCIENCE AND DISCOVERIES
new golden era of space exploration anddiscovery began April 24, 1990, with thelaunch and deployment of NASA’s
Hubble Space Telescope (HST). During
nearly seven years of operation, Hubble’s
rapid-fire rate of unprecedented discoveries
has invigorated astronomy. Not since theinvention of the telescope nearly 400 years ago
has our vision of the universe been so revolutionized over such a short stretch of time.

A

As the 12.5-ton Earth-orbiting observatory
looks into space unburdened by atmosphericdistortion, new details about planets, stars, andgalaxies come into crystal clear view. Hubble
has helped confirm some astronomical theories, challenged others, and has often come up
with complete surprises for which theories do
not yet even exist.
The Hubble was designed to provide threebasic capabilities:• High angular resolution – the ability to image
fine detail• Ultraviolet performance – the ability to produce ultraviolet images and spectra• High sensitivity – the ability to detect very
faint objects.
This unique, powerful observatory has produced a vast amount of information and asteady stream of images that have astounded
the world’s astronomical and scientific community. The Telescope has looked at more than
10,000 objects in the sky, taken over 100,000
exposures, produced more than 2.5 trillion
bytes of science data, and given rise to over
1,000 scientific reports and research papers byastronomers from more than 35 countries.
In addition to the thousands of observations
already undertaken by astronomers using theTelescope, over a thousand new observing proK70110-03M

3-1

posals each year are being received fromastronomers around the world. In fact, the seventh observing cycle was over-subscribed by afactor of three.
The Telescope is extremely popular because it
allows scientists to get their clearest view everof the cosmos and to obtain information on thetemperature, composition, and motion of celestial objects by analyzing the radiation emitted
or absorbed by the objects.
Results of HST observations are being reported
regularly in scientific papers presented atmeetings of the American Astronomical
Society and other major scientific conferences.
Although the Hubble’s dramatic findings todate are too numerous to be described fully inthis Media Reference Guide, the following
paragraphs highlight some of the significant
astronomical discoveries and observations inthree basic categories: planets, stellar evolution, and galaxies and cosmology.

3.1 Planets
HST has proved well suited for study ofobjects in the solar system. Figure 3-1 shows aview of the planet Mars that is the clearest picture ever taken from Earth, surpassed only byclose-up shots sent back by visiting spaceprobes. The picture was taken on February 25,
1995, when Mars was approximately 65 million miles (103 million km) from Earth.
Because it is spring in Mars’ northern hemisphere, much of the carbon dioxide frost
around the permanent water-ice cap has sublimated, and the cap has receded to its core ofsolid water-ice several hundred miles across.
The abundance of wispy white clouds indicates that the atmosphere is cooler than thatseen by visiting space probes in the 1970s.
Morning clouds appear along the planet’s

WF/PC II

Fig. 3-1

K70110-301

This view of the planet Mars is the clearest picture ever taken from Earth, surpassed only
by close-up shots sent back by visiting space probes. Hubble’s Wide Field/Planetary
Camera II took the picture on February 25, 1995, when Mars was approximately 65 million
miles (103 million km) from Earth.

Figure 3-2 shows a state-sized dust storm
churning near the edge of Mars’ north polar
cap in 1996. The polar storm probably is a consequence of large temperature differences
between the polar ice and the dark regions tothe south, which are heated by the springtime
sun. The increased sunlight also causes the dry

western (left) limb. These form overnight
when the planet’s temperatures plunge and waterin the atmosphere freezes to form ice crystals.
The distinct advantage of Hubble for planetary astronomers is that it provides superb
image clarity any time an observer calls for it.
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3-2

WF/PC II

K70110-302

Fig. 3-2

A state-sized dust storm churns near the edge of Mars’ north polar cap in 1996. The polarstorm probably is a consequence of large temperature differences between the polar ice andthe dark regions to the south.

dust storms. Such research is an important prerequisite for any manned expedition to Mars.
A similar systematic study of Jupiter’s powerful weather systems also is under way.

ice in the polar caps to sublime and shrink. HST
will monitor the atmosphere of Mars for several
years to study weather patterns and try to understand the events that trigger periodic, planetwide
K70110-03M

3-3

The distant planet Neptune is yielding its
secrets to the HST. The Telescope is allowing
astronomers to study Neptune’s remarkably
dynamic atmosphere with a level of detail notpossible since the Voyager 2 flyby in 1989.
Building on Voyager’s initial discoveries, HST
is revealing that Neptune’s atmospherechanges over just a few days.

3.2 Stellar Evolution
Looking beyond the solar system, the Hubble
has provided hundreds of astounding images
and answered some long-standing questions
about the birth and death of stars.
The eerie, dark structure shown in Fig. 3-5,
resembling an imaginary sea serpent’s head, isa column of cool molecular hydrogen gas (twoatoms of hydrogen in each molecule) and dustthat serves as an incubator for new stars. Someof the stars are embedded in finger-like protrusions extending from the top of the nebula.
Each “fingertip” is somewhat larger than
Earth’s solar system.

The predominant blue color of Neptune (seeFig. 3-3) results from absorption of red andinfrared light by Neptune’s methane atmosphere. Clouds elevated above most of themethane absorption appear white, while thevery highest clouds tend to be yellow-red as
seen in the bright feature at the top. Neptune’s
powerful equatorial jet – where winds blow atnearly 900 mph – is centered on the dark blue
belt just south of the equator. Farther south, thegreen belt indicates a region where the atmosphere absorbs blue light.

Ultraviolet light from nearby hot stars is slowly
eroding the pillar – a process called photoevaporation. As the pillar erodes, small globules of especially dense gas buried within thecloud are uncovered. These globules have been
dubbed “EGGs” – an acronym for evaporating
gaseous globules. The shadows of the EGGs
protect gas behind them, resulting in the fingerlike structures at the top of the cloud.

Figure 3-4 shows the never-before-seen surfaceof Pluto, which orbits at the dim outer reachesof the solar system nearly 3 billion miles (5 billion km) from the sun. Pluto is two-thirds thesize of the Earth’s moon but 12,000 times farther away. Viewing surface detail is as difficult
as trying to read the printing on a golf balllocated 30 miles away!

Forming inside some of the EGGs are embryonic stars – stars that abruptly stop growing
when the EGGs are uncovered and separated
from the larger reservoir of gas from which
they were drawing mass. Eventually the stars
emerge as the material within the EGGs collapses onto the star or forms a protoplanetary
disk around it. The EGGs are found in the
Eagle Nebula, a star-forming region 7,000
light-years distant in the constellation Serpens.

Hubble’s Faint Object Camera (FOC) imaged
nearly the entire surface of Pluto as it rotatedthrough its 6.4-day period in late June andearly July 1994. The images, made in blue light,
show that Pluto is an unusually complex
object, with more large-scale contrast than any
planet except Earth. Pluto probably has evenmore contrast and perhaps sharper boundaries
between light and dark than are shown here,
but Hubble’s resolution (just like early telescopic
views of Mars) tends to blur edges and blend
together small features sitting inside larger ones.
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As stars are born, they also die. When a starlike the Earth’s sun nears the end of its life, it
expands to more than 50 times its original
diameter, becoming a red giant star. Then its
outer layers are ejected into space, exposing
3-4

K70110-303

WF/PC II

Fig. 3-3

The Telescope is allowing astronomers to study Neptune’s remarkably dynamic atmosphere with a levelof detail not possible since the Voyager 2 flyby in 1989. The planet’s predominant blue color results fromabsorption of red and infrared light by its methane atmosphere.

K70110-304

Faint Object Camera

Fig. 3-4

K70110-03M

Hubble’s Faint Object Camera imaged nearly the entire surface of Pluto as it rotated through its 6.4-day
period in late June and early July 1994. Viewing the planet’s surface detail is as difficult as trying to readthe printing on a golf ball located 30 miles away!
3-5

WF/PC II

Fig. 3-5

K70110-305

Resembling an imaginary sea serpent’s head, this eerie, dark structure is a column of coolmolecular hydrogen gas and dust that serves as an incubator for new stars. Some of thestars are embedded in finger-like protrusions extending from the top of the nebula. Each“fingertip” is larger than Earth’s solar system.

The planetary nebula NGC 7027 (see Fig. 3-6)
was once a star like the sun. The HST photo
reveals that the initial ejections occurred
episodically to produce the concentric shells.
This culminated in a vigorous ejection of all theremaining outer layers, producing the brightthe small, extremely hot core of the star, whichcools off to become a white dwarf. Although astar like the sun can live for up to 10 billion
years before becoming a red giant and ejecting
a nebula, the actual ejection process takes only
a few thousand years.
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3-6

WF/PC II

K70110-306

Fig. 3-6

When a star like the Earth’s sun nears the end of its life, it expands to more than
50 times its original diameter, becoming a red giant star and ejecting a nebula.
This planetary nebula, NGC 7027, was once a star like the sun. The initial ejections occurred episodically to produce its concentric shells.

HST also has revealed that the gases ejected in
the death throes of a star can collide after thousands of years, with dramatic results. The

inner regions. At this later stage, the ejection
was nonspherical, and dense clouds of dust
condensed from the ejected material.
K70110-03M

3-7

Helix Nebula, the closest planetary nebula to
Earth at 450 light-years away in the constellation Aquarius, provides a good example (seeFig. 3-7). Hubble captured an image of thousands of tadpole-like objects, in the upper

right corner, that astronomers call “cometaryknots.” Each “tadpole” head is at least twice
the size of Earth’s solar system; each tailstretches 100 billion miles, 1,000 times the distance from the Earth to the sun.

K70110-307

WF/PC II

Fig. 3-7K70110-03M

Helix Nebula, the closest planetary nebula to Earth – 450 light-years away in theconstellation Aquarius – has thousands of tadpole-like objects (upper right corner)that astronomers call “cometary knots.” Each “tadpole” head is at least twice thesize of Earth’s solar system; each tail stretches 100 billion miles.
3-8

come to a violent end. The supermassive star
Eta Carinae was the site of a giant outburst
seen on Earth about 150 years ago, when it
became one of the brightest stars in the southern sky. Although the star released as muchvisible light as a supernova explosion, it survived the outburst. Somehow, the explosion
produced two polar lobes and a large, thinequatorial disk, all moving outward at about
1.5 million mph. Figure 3-8 shows a stunning
image of Eta Carinae: a huge, billowing pair ofgas and dust clouds. Using a combination ofimage processing techniques, astronomers

Astronomers theorize that the gaseous knots,
each several billion miles across, resulted froma collision between gases. A doomed star spews
hot gas from its surface, which collides with thecooler gas it had ejected 10,000 years before.
The collision fragments the smooth cloud surrounding the star into smaller, denser droplets,
like dripping paint. Astronomers expect thegaseous knots to eventually dissipate into thecold blackness of interstellar space.
While some stars end their lives with the
puffing of their outer layers into space, some

WF/PC II

Fig. 3-8K70110-03M

K70110-308

Astronomers used a combination of image-processing techniques to create this stunning
image of Eta Carinae, a pair of gas and dust clouds more than 8,000 light-years away. It is
one of the highest resolution images of an extended object ever produced by Hubble. Structures
only 10 billion miles across (about the diameter of Earth’s solar system) can be distinguished.
3-9

created one of the highest resolution images ofan extended object ever produced by Hubble.
Even though Eta Carinae is more than 8,000
light-years away, structures only 10 billion
miles across (about the diameter of Earth’s
solar system) can be distinguished. Dust lanes,tiny condensations, and strange radial streaks
all appear with unprecedented clarity.

3.3 Galaxies and Cosmology
Galaxies are the largest assemblages of stars in
the universe. In a galaxy, billions of stars arebound together by the mutual pull of gravity.
The sun resides in the Milky Way galaxy.
Studying galaxies falls into the realm of cosmology, the study of the evolution of the uni-verse on the largest scale. By looking at the distribution of galaxies in space, Edwin P. Hubble
discovered that the universe is expanding.
Hubble found that all galaxies in all directions
are receding from Earth, with those farther
away receding the fastest. Investigations since
Hubble’s time have increased the types ofgalaxies known. Strange, unusually active
galaxies and faint, blue, odd-shaped galaxies
have been discovered.
Astronomers have long thought that active
galaxies are powered by black holes in their
nuclei. However, until they viewed galaxy
NGC 4261, some 100 million light-years away,
they had never seen the enormous circulating
disks of matter that have formed around a suspected black hole. Figure 3-9 shows a strikingly

K70110-309

WF/PC II

Fig. 3-9K70110-03M

Until they viewed galaxy NGC 4261, some 100 million light-years away, scientists had neverseen the enormous circulating disks of matter that had formed around a suspected black
hole. Astronomers calculate that the object at the center of the disk is 1.2 billion times themass of the sun, yet concentrated into a region of space no larger than Earth’s solar system.
3-10

geometric disk – 800 light-years wide and containing enough mass to make 100,000 stars likethe sun – that was first identified in Hubble
observations in 1992. A Hubble image taken in
1995 reveals for the first time structure in thedisk, which may be produced by waves or
instabilities.
By measuring the speed of gas swirling aroundthe black hole, astronomers calculate that theobject at the center of the disk is 1.2 billion
times the mass of the sun, yet concentrated
into a region of space no larger than Earth’s
solar system. This discovery is giving
astronomers a ringside seat to bizarre, dynamic processes that may involve a titanic collision
and a runaway black hole. Studying this relatively nearby galaxy could shed light on how
far more distant active galaxies and quasars
produce energy.
Discovered only 33 years ago, quasars areamong the most baffling objects in the universe
because of their apparently compact size andprodigious energy output. Quasars pour out
100 to 1,000 times as much light as a galaxy
containing 100 billion stars.
Astronomers theorize that a super massive
black hole – gobbling up stars, gas, and dust – is the“engine” powering a quasar. Most astronomers
agree that an active black hole is the only credible
explanation for how quasars can be so compact,
variable, and powerful. Nevertheless, conclusive
evidence has been elusive because quasars are sobright they mask any details of their environment.
Quasars reside in a variety of galaxies, fromnormal to highly disturbed. When seen
through ground-based telescopes, these compact, enigmatic light sources are unresolved
points of light like stars, yet they are billions oflight-years away and several hundred billion
times brighter than normal stars.
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Figure 3-10 shows HST images of different
quasar home sites. All the sites must provide
the fuel to power these unique light beacons.
Astronomers believe that a quasar turns onwhen a massive black hole at the nucleus of agalaxy feeds on gas and stars. As the matter
falls into the black hole, intense radiation isemitted. The black hole stops emitting radiation once it consumes all nearby matter. Then it
needs debris from a collision of galaxies oranother process to provide fuel. The column ofimages on the left represents normal-looking
galaxies; the center, colliding galaxies; and theright, peculiar galaxies.
• Top left: Quasar PG 0052+251 is 1.4 billion
light-years away, at the core of a normal spiral galaxy. Astronomers are surprised to findhost galaxies, such as this one, that appear
undisturbed by the strong quasar radiation.
• Bottom left: Quasar PHL 909 is 1.5 billion
light-years away, at the core of an apparently
normal elliptical galaxy.
• Top center: Here is evidence of a catastrophic
collision between two galaxies traveling at
about 1 million mph. The debris from the collision may be fueling quasar IRAS04505-2958,
which is 3 billion light-years from Earth.
Astronomers believe that a galaxy plunged
vertically through the plane of a spiral galaxy,ripping out its core and leaving the spiralring (at the bottom of the picture).
• Bottom center: HST has captured quasar PG
1012+008, located 1.6 billion light-years away,
merging with a bright galaxy (the object justbelow the quasar). Although the galaxy andquasar are 31,000 light-years apart, theswirling wisps of dust and gas surrounding
them provide strong evidence for an interaction between them. The compact galaxy on

3-11

WF/PC II

K70110-310

Fig. 3-10 HST images of different quasar home sites include evidence of a catastrophic collision
between two galaxies traveling at about 1 million mph (top center) and evidence of adance between two merging galaxies (bottom right).

the left of the quasar also may be beginning
to merge with the quasar.
• Top right: Hubble has captured a tidal tailof dust and gas beneath quasar 0316-346,
located 2.2 billion light-years from Earth. The
peculiar-shaped tail suggests that the hostgalaxy has interacted with a passing galaxy
that is not in the image.
• Bottom right: HST has provided evidence ofa dance between two merging galaxies. The
galaxies may have orbited each other several
times before merging, leaving distinct loopsof glowing gas around quasar IRAS
13218+0552. The quasar is 2 billion lightyears away. The elongated core in the centerof the image may comprise the two nuclei ofthe merging galaxies.
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Humankind’s deepest, most detailed opticalview of the universe was provided by the HSTin late 1995 (see Fig. 3-11). The image, called
the Hubble Deep Field (HDF), was assembled
from 342 separate exposures taken by the Wide
Field/Planetary Camera II (WF/PC II) for 10
consecutive days, December 18 – 28.
Representing a narrow, keyhole view stretching to the visible horizon of the universe, the
HDF image covers a speck of the sky only
about the width of a dime located 75 feet away.
The tiny region in the constellation Ursa Major
was chosen because it appeared to be one ofthe darkest, emptiest places on the sky.
Although the field is a very small sample of theheavens, it is considered representative of thetypical distribution of galaxies in spacebecause the universe, statistically, looks largely

3-12

WF/PC II

K70110-311

Fig. 3-11 Humankind's most detailed optical view of the universe, called the Hubble Deep Field,
was provided by Hubble in late 1995. This image was assembled from 342 separate
exposures taken by the Wide Field/Planetary Camera II for 10 consecutive days,
December 18 – 28. It covers a speck of the sky only about the width of a dime located
75 feet away.

the same in all directions. After counting thegalaxies in the HDF, astronomers have revised
upward their estimate of the total number ofgalaxies in the universe by a factor of four.
K70110-03M

Gazing into this small field, HST uncovered abewildering assortment of some 1,500 galaxies
at various stages of evolution. Most of thegalaxies are so faint (nearly 30th magnitude or3-13

about 4 billion times fainter than can be seenby the human eye) that they have never before
been seen by even the largest telescopes. Somefraction of these galaxies probably dates backto nearly the beginning of the universe.
Scattered within the HDF are several dozen
galaxies that astronomers believe exhibit characteristics making them appear more distant
than any seen previously. Six of the galaxies
appear to be more distant than the farthest
quasars, the current distance record holders.
These galaxies are so far away that they may
have existed when the universe was less than
10 percent of its present age. If this early galaxy
population can be confirmed through further
observations, it means that such galaxies
would have formed remarkably early in thehistory of the universe, only a few hundred
million years after the Big Bang. The images
also give an estimate of how many galaxies
were forming at this time in the very early
universe.
The Hubble Space Telescope is helping test
and refine the Big Bang theory, which forms
the foundation of modern cosmology. The
theory states that the present universe,
including all matter and space, exploded
outward from a single point at the beginningof time.
One of the most important areas of inquiry that
HST allows astronomers to probe is the valueof the Hubble Constant, and a determination of
the age and size of the universe. The Hubble
Constant is the ratio of the recession velocities
of galaxies to their distance in the expanding
universe. Estimated from the Hubble Constant,the age of the universe currently is thought tobe between 9 billion and 14 billion years. Moreprecise measurements of the Hubble Constant
may narrow this range.
K70110-03M

Variable stars called Cepheids are the mostaccurate and reliable tool that astronomers use
to measure the distance to other galaxies. Oneof the major goals of the Hubble program has
been to detect Cepheids in galaxies much farther away than has been possible from theground. In achieving this goal, Hubble has significantly increased the accuracy of the calculation of the size of the universe. Combining
our knowledge of the size of the universe withthe rate of expansion results in an estimate forthe age of the universe.
Cepheids are pulsating stars that become alternately brighter and fainter in periods ranging
from 10 to 50 days. Astronomers have known for
more than 50 years that the periods of these stars
precisely predict their total luminous power,
which allows their distance to be measured.
Figure 3-12 shows Galaxy NGC 1365, a rotating
system of gas and stars similar to the MilkyWay. The HST color image on the right is aregion in NGC 1365, a barred spiral galaxy
located in a cluster of galaxies called Fornax. A
barred spiral galaxy is characterized by a “bar”
of stars, dust, and gas across its center. The
black-and-white photograph from a groundbased telescope shows the entire galaxy, which
is visible from the Southern Hemisphere.
Hubble astronomers who have been measuring
the distance to the Fornax cluster have estimatedit to be 60 million light-years from Earth. The
astronomers arrived at their preliminary estimate by using Cepheids. The line of small blue
dots in the color image shows the formation ofstars in the galaxy’s spiral arm, making them
ideal targets for the discovery of Cepheids.
Astronomers have discovered about 50
Cepheids in the galaxy. They also have used theFornax cluster to calibrate and compare many

3-14

K70110-312

WF/PC II

Fig. 3-12 A “bar” of stars, dust, and gas characterizes Galaxy NGC 1365, a barred spiral galaxy
located in a cluster of galaxies called Fornax. The HST color image (right) shows a regionin NGC 1365. The black-and-white photograph from a ground-based telescope shows theentire galaxy, visible from the Southern Hemisphere.

secondary distance methods. While Cepheids
are accurate distance markers for galaxies atintermediate distances, secondary methods areneeded to measure distances to galaxies much
farther away. An accurate value for the Hubble
Constant depends on reliable secondary distance methods. One secondary method relates
the total luminosity of a galaxy to the rate at
which the galaxy is spinning, the Tully-Fisher
relation. Another secondary method uses a
special class of exploding star known as a typeIa supernova. This phase of the Hubble Constantresearch will be completed within two years.

3.4 Summary
The Hubble Space Telescope has establisheditself as a premier astronomical observatory
K70110-03M

3-15

that continues to make dramatic observations
and discoveries at the forefront of astronomy.
Following the successful 1993 First Servicing
Mission, and the full restoration of Hubble’s
unsurpassed optical capability, the Telescope
was able to achieve all of its original objectives.
Among a long list of achievements, Hubble has:
• Improved our knowledge of the size and ageof the universe• Provided decisive evidence of the existenceof supermassive black holes at the centers ofgalaxies• Clearly revealed the galactic environments inwhich quasars reside• Detected objects with coherent structure(protogalaxies) close to the time of the originof the universe• Provided unprecedentedly clear images andspectra of the collision of Comet ShoemakerLevy 9 with Jupiter• Detected a large number of protoplanetarydisks around stars• Clearly elucidated the various processes bywhich stars form
• Provided the first map of the surface of Pluto

K70110-03M• Routinely monitored the meteorology ofplanets beyond the orbit of Earth• Made the first detection of an ultraviolet highenergy laser in Eta Carinae.
After the Second Servicing Mission, the Hubble
Space Telescope will view the universe anew withsignificantly expanded scientific capabilities.

3-16

SCIENCE INSTRUMENTS AND COSTAR

S

cience instruments currently operatingin the Hubble Space Telescope (HST) are
the Faint Object Camera (FOC), the Faint
Object Spectrograph (FOS), the Goddard High
Resolution Spectrograph (GHRS), and the Wide
Field/Planetary Camera II (WF/PC II). In
addition, the three Fine Guidance Sensors (FGS)
have a scientific role as astrometric instruments.

instrument through more advanced detectors
and more stringent contamination control. It
also incorporated built-in correction optics.
Development of the COSTAR, designed specifically to correct the optics for the remaining axial
instruments (FOC, FOS, and GHRS), started in1990 after the aberration was discovered. A setof optics was deployed into the region near the
focal plane to intercept the light that normally
would be sensed by these instruments andreplace it with light corrected for sphericalaberration (see Fig. 4-1).

The FOC, FOS, and GHRS are located parallelto the Telescope’s optical axis so that incoming
images fall into their entrance apertures. These
instruments share the same dimensions –roughly the size and shape of a telephonebooth – so they fit into the focal plane structureinterchangeably.

The FOC will be the only instrument using
COSTAR optical correction after the Second Servicing Mission. Both new instruments, the Near
Infrared Camera and Multi-Object Spectrometer(NICMOS) and the Space Telescope Imaging
Spectrograph (STIS), have corrective optics
incorporated into their designs.

The WF/PC II and the three FGSs are placed
just forward of the focal plane structure at aright angle to the optical axis. These four radial
instruments rely on pickoff mirrors positioned
in the optical path to deflect part of the incoming light into their respective entrances.

Descriptions of each science instrument andpossible astronomical targets are included inthis section. The astrometric function of the
FGSs also is discussed.

The Telescope’s optical system was designed toprovide light in the focal plane with a resolution close to the limit specified by the laws ofnature. This so-called “diffraction limit” is aphysical property of light that depends only on
the size of the primary mirror and the wavelength of the light, and defines the minimum
size of the image of a point-like object on the
focal plane. The presence of a spherical aberration on the primary mirror discovered shortly
after deployment of HST in 1990 hampered thepotential resolution and motivated development of corrective optics for the instruments.

4.1

NICMOS is a second-generation instrument to
be installed on the HST during the Second Servicing Mission. NICMOS will provide infrared
imaging and limited spectroscopic observationsof astronomical targets between 1.0 and 2.5 microns. NICMOS will extend HST’s capabilities
into the near-infrared by using a 5-year cryogenic dewar system. High-resolution images
will be made for detailed analyses of:During the First Servicing Mission, the Corrective Optics Space Telescope Axial Replacement(COSTAR) replaced the High Speed Photometer,
and the WF/PC II replaced the original camera(WF/PC I). WF/PC II significantly improved
ultraviolet performance over that of the original
K70110-04M

Near Infrared Camera andMulti-Object Spectrometer•Protostellar clouds, young star clusters, andbrown dwarfs•Obscured active galaxy nuclei•Temporal changes in planetary atmospheres.
4-1

Fig. 4-1 Cutaway view of COSTAR showing its optical bench deployed and pickoff mirrors
in position to capture light from the secondary mirror

NICMOS will provide an improved determination of the distance scale to the galaxies of the
Coma cluster and beyond and will view distantquasi-stellar objects that are not observable atvisible wavelengths. NICMOS also will provide
deep, high-resolution images of obscured galactic centers and star formation regions as well as
detailed imaging in infrared bands of planetaryatmospheres, much like the infrared weather
images of Earth.

field of view simultaneously. Figure 4-3 lists thecameras and their optical characteristics.
Light entering the instrument entrance aperture
falls on a flat folding mirror and is redirected
to a spherical mirror. It is then re-imaged on thecorrective mirror, which is mounted to an offset pointing mechanism. This mirror corrects the
HST spherical aberration and also has a cylindrical deformation to correct for astigmatism inthe optical path.

4.1.1 Instrument Description

The corrected image then is relayed to a threemirror field-dividing assembly, which divides
the light into three separate, second-stage optical paths (see Fig. 4-4). In addition to the fielddividing mirror, each second-stage optic uses a
two-mirror relay set and a folding flat mirror.

NICMOS is an all-reflective imaging systemwith near–room-temperature foreoptics thatrelay images to three focal plane cameras contained in a cryogenic dewar system (seeFig. 4-2). Each camera covers the same spectral
band of 0.8 to 2.5 microns with a different magnification and an independent filter wheel. Eachcamera views a different segment of the HST
K70110-04M

The field-dividing mirrors are tipped to divide
the light rays by almost 4.5 degrees. The tip
4-2

Fig. 4-2 Near Infrared Camera and Multi-Object Spectrometer (NICMOS)

allows physical separation for the two-mirror
relay sets for each camera and its field of view.
The curvature of each mirror allows therequired degree of freedom to set the exit pupil
at the cold mask placed in front of the filter
wheel of each camera.

STAGE 2SECONDRELAY MIRROR

EXIT PUPIL
COLD MASKFIELD-DIVIDINGASSEMBLYMIRROR(FIELD STOP)

A corrected image is produced in the center of
the Camera 1 field mirror. The remaining mirrors of Camera 1 are confocal parabolas with

FIELD
OFFSET
MECHANISM

CORRECTOR(PUPIL)NICMOS Optical Characteristics

Parameter
Total field (arcsec)

Camera 1
11.0

Camera 2
19.2

Camera 3
51.2

Pixel size (arcsec)

0.043

0.075

0.20

Magnification

3.33

1.91

0.716

f number

80

45.7

17.2

IMAGE
PLANE

FILTER
WINDOW

FIRST RELAY MIRROR
PUPIL ADJUSTMECHANISM

HST
IMAGE
PLANE

Z
0

500 mm
K70110-404

K70110-403

Fig. 4-3 NICMOS optical characteristics
K70110-04M

Fig. 4-4 NICMOS optical path4-3

offset axes to relay the image into the dewar
with the correct magnification and minimalaberration.

at 270K. The electronics maintains a controlled
thermal environment, partly through radiators
mounted to the outboard enclosure panels. A
combination of active proportional heaters,
selective surface finishes, and multilayer insulation (MLI) maintains the enclosure thermal
environment.

Cameras 2 and 3 have different amounts ofastigmatism because their fields are at different
off-axis points from Camera 1. To correct theresidual astigmatism, one of the off-axis relay
mirrors in Camera 3 is a hyperbola and one of
the relay mirrors in Camera 2 is an oblate ellipsoid. Camera 2 also allows a coronagraphic
mode by placing a dark spot on its field dividing mirror. During this mode the HST is maneuvered so that the star of observation falls within
the Camera 2 field dividing mirror and becomesocculted for coronagraphic measurements.

The cryogen dewar contains solid-subliming
nitrogen, which maintains the three cameras
housed on a cold bench within its well at 58K.
Surrounding the dewar are three thermal
shields interspersed with MLI: the vapor-cooled
shield (VCS), the thermoelectric-cooled inner(TECI) shield, and the thermoelectric-cooled
outer (TECO) shield (see Fig. 4-5).

All the detectors are 256 x 256-pixel arrays of
mercury cadmium telluride (HgCdTe) with
40-micron pixel-to-pixel spacing. An independent, cold filter wheel is placed in front of eachcamera and is rotated by room-temperature
motors placed on the external access port of thedewar.

The VCS is the innermost shield and is cooledby vented nitrogen vapor. The TECI and TECO
are each cooled by a set of two thermoelectriccoolers, which rejects heat to radiators isolated
from the enclosure and supported to prevent
enclosure heat loads from passing to the dewar
and bench.

A multilevel, flat field illumination system corrects detector nonuniformities. The light source
and associated electronics are located in theelectronics section at the rear of the instrument.
Infrared energy is routed to the optical systemusing a fiber bundle. The fiber bundle illuminates the rear of the corrector mirror, which ispartially transparent and fits the aperture fromthe fiber bundle. The backside of the element iscoarsely ground to produce a diffuse source.

The dewar system uses 124 layers of MLI madeof double-aluminized mylar with polyester net
spacer material. The blankets are distributed for
maximum effectiveness: 17 layers between the
dewar outer shell and the TECO, 17 layers between the TECO and TECI, 50 layers between
the TECI and VCS, and 40 layers between the
VCS and solid nitrogen tank.
The solid cryogen tank is suspended by six
fiberglass/epoxy straps attached to girth ringson the vacuum shell. The three thermal shields
are structurally and thermally attached at intermediate points along the straps.

The instrument structural enclosure houses all
operating components in two individual compartments: the optics compartment and the electronics compartment. A graphite-epoxy opticalbench is kinematically mounted within theenclosure and physically separates the two compartments. The foreoptics and the cryogenic
dewar mount to the bench and are maintained
K70110-04M

The dewar uses three plumbing lines. The
dewar is loaded with liquid nitrogen (LN 2)
through a single, combined vent and fill line
using a no-vent fill process. On orbit, this line
4-4

Fig. 4-5 NICMOS cryogen dewar

also serves as the vent path for nitrogen vapor.
The other two lines form a coolant loop. Cold,gaseous helium is circulated through this loop
to precool the dewar prior to loading, freeze theloaded LN2 to a solid, and periodically supercool the solid N2 during ground processing tokeep it solid until on-orbit operations begin. To
reduce the conductive path, these lines are notvapor cooled.

wheels. The drive shafts provide adequate
torsional rigidity for accurately positioning thefilter in the optical path while maintaining low
thermal conductivity.

4.1.2 NICMOS Specifications
Three detector cables and three detector clock
cables route electrical signals from the cryogen
tank to the hermetic connector at the vacuum
shell. The cables consist of small-diameter, stainless steel wire mounted to a polymeric carrier
film. Shielding is an aluminized polyester film
incorporated into drain wires. The cables areadequately shielded to minimize noise andcrosstalk between channels and also have low
thermal conductivity to minimize parasitic heatloads. In addition, two unshielded cables connect to thermal sensors used during fill and foron-orbit monitoring. These cables also containthe leads for accelerometers mounted on the

The optical paths penetrate the dewar in threeplaces. Each camera port consists of an external
vacuum shell window, an internal heat blocking window, and a cold mask to prevent thedetectors from seeing warm structure. Eachcamera has an independently controlled filter
wheel. The filter wheels, mounted on the vaporcooled shell, are turned by warm stepper
motors mounted on the vacuum shell. Graphiteepoxy, thin-walled tubes are used for the drive
shafts connecting the warm motors to the cold
K70110-04M

4-5

solid nitrogen tank that are used during
preflight testing. Heat leak into the solid nitrogen cryogen through the cable is 1.1 milliwatts.
In addition to processing signals from andcontrolling the detectors, the electronics prepare
the data for transmission to the HST computer,
respond to ground commands through the HST,
and control the operation of the instrument.
NICMOS uses an on-board 80386 microprocessor with 16 megabytes of memory for instrument operation and data handling. Two systems
are provided for redundancy. The detector control electronics subsystem also includes a microprocessor dedicated to the operation of the focal
plane array assemblies. Two microprocessors
are provided for redundancy.

Fig. 4-7 Space Telescope Imaging Spectrograph (STIS)

capability to HST spectroscopy. The two dimensions can be used either for “long slit” spectroscopy, where spectra of many different points
across an object are obtained simultaneously, orin an echelle mode to obtain more wavelength
coverage in a single exposure. STIS also can take
both UV and visible images through a limited
filter set.

Figure 4-6 shows the specifications for
NICMOS.

4.2

Space Telescope Imaging
Spectrograph

Designed to replace many of the current capabilities of the GHRS and some of the FOS, STIS
has some additional enhanced capabilities. It
will cover a broader wavelength range with
two-dimensional capability; it can image; itadds a coronagraph capability; it can provide
objective prism spectra in the intermediate UV;
and it has a high time-resolution capability in
the UV.

STIS was developed under the direction of theprincipal investigator, Dr. Bruce E. Woodgate
jointly with Ball Aerospace (see Fig. 4-7). It is
generally considered to be the most complex
scientific instrument built for space science.
The spectrograph was designed to be versatileand efficient, taking advantage of modern
technologies to provide a new two-dimensional

STIS carries it own aberration correcting optics
and will not require the use of COSTAR.

Near Infrared Camera and Multi-Object
Spectrometer (NICMOS)
WeightDimensions
Principal investigator
Contractor
Field of view

Detectors

4.2.1 Physical Description

861 lb (391 kg) in flight configuration
7.1 x 2.8 x 2.8 ft. (2.2 x 0.88 x 0.88 m)
Dr. Rodger I. Thompson, U. of Arizona
Ball Aerospace
51.2 x 51.2 arcsec
19.2 x 19.2 arcsec
11.0 x 11.0 arcsec
3 HgCdTe arrays
256 x 256 pixels

STIS has been designed to fit into the axial bay
behind the HST main mirror, replacing the
GHRS. It is therefore of comparable size andweight to GHRS. Externally, the instrument
measures 7.1 ξ 2.9 ξ 2.9 ft (2.2 ξ 0.98 ξ 0.98m)
and weighs 825 lb (374 kg). Internally, STIS
consists of a carbon fiber optical bench, which

K70110-406

Fig. 4-6 NICMOS specifications
K70110-04M

4-6

supports the dispersing optics and threedetectors.

detectors and mechanisms are controlled by anonboard computer.

STIS has been designed to work in threedifferent wavelength regions, each with its own
detector. Some redundancy is built into thedesign with overlap in the detector response
and backup spectral modes. To select a wavelength region or mode, a single mechanism,
called the mode selection mechanism (MSM), isused. The MSM has 21 optical elements: 16 firstorder gratings (six of which are order-sorting
gratings used in the echelle modes), an objective prism, and four mirrors. The optical benchsupports the input corrector optics, focusing
and tip/tilt motions, the input slit and filter
wheels, and the MSM.

Each of the three detectors has been optimized
for a specific wavelength region. Band 1, from
115 to 170 nm, uses a Multi-AnodeMicrochannel Plate Array (MAMA) with acesium iodide (CsI) photocathode. Band 2, from165 to 310 nm, also uses a MAMA but with acesium telluride (CsTe) photocathode. Bands 3
and 4, covering the wavelengths from 305 to
555 nm and 550 to 1000 nm, use the samedetector, a charge-coupled device (CCD). Figure 4-8 shows the instrument schematically.
Entrance Apertures. After the light beam passes
through the corrector, it enters the spectrographthrough one of several slits. The slits aremounted on a wheel, and the slit or entrance
aperture can be changed by wheel rotation.

Light from the HST main mirror is firstcorrected and then brought to a focus at the slit
wheel. After passing through the slit, it iscollimated by a mirror onto one of the MSM
optical elements. A computer selects themode and wavelength. The MSM rotates andnutates to select the correct optical element,
grating, mirror, or prism, and points the beamalong the appropriate optical path to the correct
detector.

The first-order spectral imaging modes canselect slits 50 arcsec long and from 0.05 to
2 arcsec wide. Three slits have occulting bars
that can block out a bright star in the field. Fourslits are tilted to an angle of 45 degrees for planetary observations.
For echelle spectroscopy, 16 slits ranging inlength from 0.10 to 1 arcsec are available. The
slit length is short to control the height of theechelle spectra in the image plane and avoid
spectral order overlap. The echelle slits havewidths of 0.05, 0.10, 0.12, 0.20, and 0.5 arcsec.

In the case of first-order spectra, a first-order
grating is selected for the wavelength anddispersion. The beam then is pointed to acamera mirror, which focuses the spectrum onto
the detector, or goes directly to the detector
itself.

There also are camera apertures of 50 ξ 50 and
25 ξ 25 arcsec. Some of the apertures haveocculting bars incorporated. The telescope canbe positioned to place bright stars behind theocculting bars to allow viewing and observationof faint objects in the field of view. In addition,
there is a special occulting mask or coronagraph, which is a finger in the aperture that canbe positioned over a bright star to allow

For an echelle spectrum, an order-sorting
grating that directs the light to one of the four
fixed echelle gratings is selected, and the dispersed echellogram is focused via a camera
mirror onto the appropriate detector. The detectors are housed at the rear of the bench, so theycan easily dissipate heat through an outer panel.
The optical bench is thermally controlled. The
K70110-04M

4-7

Fig. 4-8 STIS components and detectors

examination of any faint material nearby. It canbe thought of as a simulation of a total eclipse
on a nearby star. This mode is particularly useful to search for faint companion stars or planetary disks around stars.

inclination of the axis of the wheel or tilt of the
optic to select the wavelength range and point
the dispersed beam to the corresponding
detector. One of three mirrors can be selected totake an image of an object. The objective prismis used exclusively with the Band 2 MAMA (seeFig. 4-9). Of the 16 gratings, six are cross dispersers that direct dispersed light to one of thefour echelle gratings for medium- and highresolution modes.

Mode Selection Mechanism. The mode selection mechanism (MSM) is a rotating wheel thathas 16 first-order gratings, an objective prism,
and four mirrors mounted to it. The MSM axis
is a shaft with two inclined outer sleeves, one
sleeve fitting inside the other. The sleeves areconstructed so that rotation of one sleeve rotates
a wheel to orient the appropriate optic into thebeam. Rotation of the second sleeve changes theK70110-04M

Multi-Anode Microchannel Plate Array
Detectors. For UV modes, two types of MAMA
detectors are employed on STIS. A photocathode optimizes each detector to its wavelength
4-8

Band
Wavelength Range (nm)Detector

Band 1
115-170
Det #1(MAMA/CsI)

Band 2
165-310
Det #2(MAMA/Cs2Te)

Band 3305-555
Det #3(CCD)

Band 4550-1000
Det #3(CCD)

Low resolution
spectral imaging(first order)

Mode number
Resolving power (l/Dl)
Slit length (arcsec)
Exposures/band

1.1
770-1,130
24.9
1

2.1
415-73024.9
1

3.1
445-77051.1
1

4.1
425-68051.1
1Medium resolution
spectral imaging(first order scanning)

Mode number
Resolving power (l/Dl)
Slit length (arcsec)
Exposures/band

1.2
8,600-12,800
29.7
11

2.2
7,500-13,900
29.7
18

3.2
4,340-7,730
51.1
10

4.2
3,760-6,220
51.1
9Medium resolution
echelleMode number
Resolving power (l/Dl)
Exposures/band

1.3
37,0001

2.3
23,900-23,100
2

––

––

High resolution
echelleMode number
Resolving power (l/Dl)
Exposures/band

1.4
100,000
3

2.4
100,000
6––––

Objective
spectroscopy(prism)Mode number
Resolving power (l/Dl)
Field of view (arcsec)––

2.5 (115-310 nm)
930-26 (at 120-310)
29.7 x 29.7

––

––

Mode

K70110-409

Fig. 4-9 STIS spectroscopic modes

region. Each detector’s photocathode provides
maximum sensitivity in the wavelength region
selected, while it rejects visible light notrequired for the observations. The Band 1
MAMA uses a CsI cathode, which is sensitive
to the wavelength region from 115 to 170 nm
and very insensitive to wavelengths in the visible. Similarly, the Band 2 MAMA uses a CsTe
photocathode, which is sensitive to the wavelength region from 120 to 310 nm.

electrons, and so the process continues,
producing a cascade of a million electrons at theend of the pore.
In the Band 1 tube, shown in Fig. 4-10, UV
photons enter and hit the CsI photocathode thatis deposited on the front surface of the MCP.
The cathode produces an electron when aphoton hits it and the electron is accelerated into
the MCP pores. The MCP amplifies the numberof electrons, which fall as a shower onto theanode array as they leave the MCP.

The heart of each MAMA detector is a microchannel plate (MCP) – a thin disk of glass
approximately 1.5 mm thick and 5 cm in diameter that is honeycombed with small(12.5-micron) holes or pores. The front and backsurfaces are metal coated. With a voltage
applied across the plate, an electron entering
any pore is accelerated by the electric field, andit eventually collides with the wall of the pore,giving up its kinetic energy to liberate two or
more secondary electrons. The walls are treated
to enhance the secondary electron production
effect. The secondary electrons continue down
the pore and collide with the wall to emit more
K70110-04M

The anode array is a complex finger-likepattern. When electrons strike certain anodes, asignal is sent to the computer memory
indicating the position and time of arrival of thephoton. Figure 4-11 shows the detection schemein simplified form.
The anode array has been designed so that only
132 circuits are required to be able to read out
all 1024 x 1024 pixels. As the MAMA records thearrival of each photon, it can provide a time
sequence. For instance, if an object is varying intime, like a pulsar, the data can be displayed to
4-9

Fig. 4-10 Multi-Anode Microchannel Plate Array (MAMA) detector

show if there is any periodicity. Similarly, tocreate an image, the data must be integrated in
the computer memory before it is displayed.
The MAMA data is recorded to a time resolution of 125 microseconds.
In the case of Band 2, the CsTe photocathode isdeposited on the inside surface of the front window as a semi-transparent film. Photons passthrough the window, and some are stopped in
the cathode film where they generate electrons,
which are amplified and detected in the samemanner as the Band 1 detector.
When used in the normal mode, each detector
has 1024 ξ 1024 pixels, each 25 ξ 25 microns
K70110-04M

square. However, data received from the anodearray can be interpolated to give a higher resolution, splitting each pixel into four 12.5 ξ 12.5
micron pixels. This is known as the high-resolution mode; however, data taken in this modecan be transformed to normal resolution if
required. The high-resolution mode provides
higher spatial resolution for looking at finestructural details of an object and ensures fullsampling of the optical images and spectra.
Charge-Coupled Detector. The STIS CCD wasdeveloped with GSFC and Ball input at
Scientific Imaging Technologies (SITe). Fabricated using integrated circuit technology, thedetector consists of light-sensitive picture

4-10

Fig. 4-11 Simplified MAMA systemelements (pixels) deposited onto a thin wafer ofcrystalline silicon. Each element is 21 ξ
21 microns. The elements are arranged 1024 toa row, in 1024 columns. The 1024 ξ 1024 format
contains 1,048,576 pixels.
Each element acts as a small capacitance. As
light falls on a pixel, it liberates electrons, which
effectively charge the capacitance. The numberof electrons stored is then proportional to theintensity or brightness of the light received. The
charge in each pixel can be read out by applying
a small voltage across the chip.
The CCD is most sensitive to red light, but the
STIS chip has been enhanced through what is
known as a “backside treatment” to provide a
usable sensitivity in the near-ultraviolet. The
CCD is sensitive from approximately 200 nm to
the near-infrared at 1000 nm. The violet
extension allows the CCD to overlap with theBand 2 MAMA sensitivity and can serve as a
backup detector.
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To reduce thermionic noise generated in the
CCD, the detector is integrated into a housing
and cooled to below -80ϒC. The cooling is provided by a thermoelectric cooler, which isbonded onto the back of the CCD. The heat
extracted from the CCD is dissipated througha radiative cooling panel on the outside of STIS.
The housing has a front window made fromfused silica and is kept close to 20ϒC, the designtemperature of the optical bench.
The CCD can make exposures ranging from
0.1 seconds to 60 minutes. In space, above
Earth’s protective atmosphere, radiation fromcosmic rays is higher than at Earth’s surface.
CCDs are sensitive to cosmic rays, which can
produce large numbers of electrons in the pixels.
For this reason, two shorter exposures of up to
1 hour are made and comparison of the frames
allows cosmic ray effects to be subtracted.
The CCD is a 16-bit device, allowing a dynamic
range from 1 to 65,535 to be recorded. The

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dynamic range can be further extended bychanging the gain. The gain is commandable for
1 electron/bit to 8 electrons/bit.
Another useful feature is called binning, inwhich pixels are merged on the chip. Typically,binning is 2 ξ 2 for imaging, making the pixels
larger, which can reduce the noise per pixel andincrease sensitivity at the cost of resolution. Binning can be used to look for extended faintobjects such as galaxies. Another binningapplication is in the long-slit mode on extended
faint objects. In this mode, binning along the slitby 1 ξ 4, for instance, would maintain the spectral resolution but sum the spectra from different parts of an object seen along the slit toincrease signal or detectivity.
Finally, to increase the CCD’s performance at
low light levels, the chip has incorporated aminichannel. The main problem with readingout a signal from a CCD is that a charge generated by light must be dragged across the chip,
through all its adjacent pixels, to be read out of
one corner. In so doing, the charge meets spurious defects in each pixel that add noise.
Because the noise can be very path dependent,
the minichannel ion implant is designed torestrict the path taken at low signal levels toimprove CCD performance.

bands using long slits. However, as dispersion
increases, not all of the spectrum falls on thedetector. Obtaining an entire spectral range may
require moving the spectrum and taking
another image. Figure 4-9 indicates the numberof exposures to cover the whole wavelength
range.
The medium-resolution echelle spectroscopy
with R~24,000 uses short slits and is available
in the UV only. Band 2 requires two exposures
to cover the whole wavelength region.
High-resolution echelle spectroscopy, with
R~100,000, uses short slits and is available in the
UV only. Both Bands 1 and 2 require multiple
exposures.
Objective spectroscopy, R~26 (at 300 nm) and
930 (at 121 nm), is available using the Band 2
detector only. This mode uses a prism is usedinstead of a grating. The prism dispersion, unlike a grating, is not uniform with wavelength.
The low-resolution gratings and the prism also
can provide imaging spectroscopy of emission
line objects such as planetary nebulae, supernova remnants, or active galaxies.

Figure 4-9 shows the spectral operationalmodes. Each instrument mode is described by
two numbers, W and R, where W refers to thewavelength range and R to the resolving power.

Imaging Operational Modes. STIS can be used
to acquire an image of an object in UV or visible
light. To do this, an open aperture is selected
and a mirror placed in the beam by the MSM.
The instrument has nine filters that can beselected (see Fig. 4-12). The cameras for the CCD
and the MAMAs have different magnification
factors. The field of view is 25 x 25 arcsec for
the MAMAs and 50 x 50 arcsec for the CCD.

The low-resolution, or spectral-imaging mode,is R~500 to 1,000, and can be carried out in all
four bands using a long slit. The medium
resolution mode, R~5,000 to 10,000, is a spectral
imaging mode that can be carried out in all four

Target Acquisition. Normally an object isacquired using the CCD camera with a 50 ξ 50
arcsec field. Two short exposures are taken to
enable subtraction of cosmic rays. The HST
FGSs have a pointing accuracy of ±2 arcsec, and4.2.2 Spectra Operational Modes

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4-12

Central or
CutoffWavelength

FWHM
(nm)

Peak
Transmission(%)

Emission line
(Lyman α)

122 nm8.5

10

Cutoff (SrF2 crystal)>128 nm

N/A

90

Cutoff (crystalline
quartz)>148 nm

N/A

85

Continuum

182 nm49

40

Emission line (C III)

191 nm15

15

Continuum

270 nm22.8

72

Emission line (Mg II)

280 nm

5.6

65

Emission line (O II)

373 nm7.1

53

Emission line (O III)

501 nm0.6

73Filter Type4.2.4 Advantages Over Current
Instrumentation
STIS has several advantages over current
instrumentation:(1) Two-dimensional detectors (1024 x 1024 pixels), allowing both long-slit spectroscopy and anechelle format in the UV, give wide simultaneous wavelength coverage. Current HST
detectors are not two-dimensional.

K70110-412

Fig. 4-12 STIS filter set

the target usually is easily identifiable in thefield. Once identified, an object is positioned via
small angle maneuvers to the center of the chosen science mode slit position. Two more exposures are made, then the calibration lamp is
flashed through the slit to confirm the exact slit
position. A further peak up on the image is thenperformed. Acquisition can be expected to takeup to approximately 20 minutes.
Data Acquisition. The MAMAs take data in thehigh-resolution mode. For normal imaging andspectroscopy, the data will be integrated in the
onboard computer and stored in this format on
the solid-state recorders for later downlink. The
MAMAs also have a time-tag mode, where eachphoton is stored individually with its arrival
time and location (x,y,t). The data is stored in a
16-Mb memory and as the memory fills, thedata is dumped into the onboard recorder. The
time-tag mode has a time resolution of
125 microseconds.

(2) The UV detectors have a 100 to 200 times
lower background – per resolution element –than GHRS. Additionally, the two-dimensional
detector capability enables a good estimate of
the lower background to be made for background subtraction, which in turn provides
improved sensitivity.
(3) Use of a CCD detector to improve visible
efficiency and wavelength coverage out to1000 nm. This compares with coverage out to700 nm of the FOS.
(4) Solar blind imaging in the UV.
(5) Coronagraphic capability for spectroscopy
and imaging.
(6) Wide-field spectroscopy for spectra of many
objects in a field at once.

Space Telescope Imaging Spectrograph (STIS)
WeightDimensions
Principal Investigator
Prime Contractor
Field of View
Pixel Format
Wavelength Range

825 lb (374 kg)
3 x 3 x 7 ft (0.9 x 0.9 x 2.2 m)
Dr. Bruce E. Woodgate, GSFC
Ball Aerospace
MAMA 24.9 x 24.9 arcsec
CCD 51 x 51 arcsec
1024 x 1024
115 to 1000 nanometers

4.2.3 STIS Specifications

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Figure 4-13 shows STIS specifications.
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Fig. 4-13 STIS specifications
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4.2.5 Observations
Scientists using STIS will focus their science on
many areas, including:•Search for massive black holes by studying
star and gas dynamics around the centers ofgalaxies•Measurement of the distribution of matter inthe universe by studying quasar absorption
lines•Use of the high sensitivity and spatial resolution of STIS to study stars forming in distantgalaxies•Mapping – giving fine details of planets,
nebulae, galaxies, and other objects•Coronagraphic capability may enable it to image Jupiter-sized planets around nearby stars.
STIS also can provide physical diagnostics, suchas chemical composition, temperature, density,
velocity of rotation or internal mass motions inplanets, comets, stars, interstellar gas, nebulae,
stellar ejecta, galaxies, and quasars.
Studies of Black Holes in Centers of Galaxies.
A black hole was discovered in the center of
M87 using FOS. The black hole’s mass provides
the gravity required to hold in orbit gas andstars that are rapidly rotating about the galactic nucleus. From the spectra of stars surrounding the nucleus, a measure of how rapidly thevelocity changes from one side of the nucleus
to the other can be determined by measuring
the Doppler shift. STIS’s long-slit mode is particularly well suited for this type ofmeasurement because all the spatial positions
required can be measured along the slit in asingle exposure. With STIS, a study of black
holes can be made easily for many galaxies andcompared from one galaxy to another.
Abundances and Dynamics of the Intergalactic
Medium. STIS is well suited to observe theK70110-04M

spectra of distant galaxies and quasars.
Absorption lines from intervening material in
the spectra of these objects give a measure of the
dynamics and abundances of specific elements.
However, because these objects are distant, we
effectively are looking back in time to an earlier
stage of the universe when the chemical composition was different from that seen in thevicinity of our sun today. Measuring these differences can provide important clues to how theuniverse has evolved over time.
Abundances in the Interstellar Medium. The
STIS high-resolution mode is particularly well
suited to measurements of the spectral absorption lines created in the interstellar medium andseen against distant O- and B-type stars. From
the Doppler shift, a measure of gas speed is
obtained. Temperature and density and chemical composition also can be measured. Theinterstellar medium is thought to play an important role in when star formation occurs ingalaxies. Current theories point to hot material
being expelled from supernovas into a galaxy’s
surrounding halo. Later, after cooling, thematerial returns to the galactic plane and eventually forms new stars. How quickly the interstellar medium circulates can be a clue to whenstar formation occurs.
Search for Protoplanetary Disks. STIS’s
coronagraphic mode can be used to image
nearby stars and search for protoplanetarydisks. These observations can provide complementary data to NICMOS, which will be able tosense warm gas around young stars. Theseobservations could shed light on how planets
form, what type of stars have planetary disks,and how quickly the disks evolve into planets.

4.3

Faint Object Camera

More than the other instruments, the FOC fully
uses the optical resolution of the Telescope to

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record objects in deep space much more clearly
than before. The FOC, with COSTAR, can detect
celestial sources to 28th visual magnitude. Currently, only a ground-based observatory, the10-meter Keck telescope, can detect objects that
faint but with far less resolution. The FOC studies star formations, examines galaxies and faintobjects like quasars, and could locate planets
outside our solar system.

4.3.1 Physical Description
The FOC was designed by the European Space
Agency (ESA) and built by Dornier Systems inWest Germany, Matra in France, and British
Aerospace Space Systems Ltd. (BAe) inEngland. Its physical dimensions are 3 ξ 3 ξ 7 ft
(0.9 ξ 0.9 ξ 2.2 m), roughly the size of a telephone booth; it weighs about 700 lb (318 kg).
Figure 4-14 shows the FOC’s four major
subsystems. The load-carrying structure assembly houses the optical elements and the photondetector head elements. The optomechanicalassembly contains an optical bench that
supports the main optical and mechanicalequipment. The electronic bay assembly holds
the data-processing and data-handling equipment and system control. The photon-detector
assembly consists of the photon-detection system, data-processing electronics, and power
supply for the detectors.
After the Telescope is positioned correctly, lightfrom an object is focused on the FOC aperture.
This light is channeled down one of two opticalpathways. At the end of the pathway, eachphoton enters the camera’s detection device.
The recorded image is placed in a science datastore. The data is transmitted to Earth, where itis processed by computer into an image of thecelestial object.
Optical System. The optical system consists of
two independent systems, using two different
apertures and focal ratios, f/75.5 and f/151.
These are similar to the f-stops on conventional
cameras except for the larger ratios.

Fig. 4-14 Faint Object Camera (FOC) major subsystems
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4-15

As shown in Fig. 4-15, light is reflected from a
primary concave mirror to a secondary convex
mirror on a different plane, then onto a folding
mirror, and finally onto the detector tube forenhancing and processing. Both optical pathways have filter wheels to isolate critical wavelength beams for specific studies. The folding
mirrors increase the focal length without
increasing physical length, move to adjust thefocus, and correct for astigmatism in theTelescope optic mechanism.
Each optic system aperture has a shutter thatremains closed, completely blocking out light
until the camera is needed for astronomical
observations.
The f/151 Optics System. This system produces
the Telescope’s best resolution by increasing its
focal ratio of f/37 fourfold. There is a tradeoff,
however. The f/151 optical system has a narrow
field of view at best (22 arcsec 2), but it allows
the FOC to distinguish between two objects
only 0.05 arcsec apart. This is a crucial requirement for studying stars within a distantglobular cluster; they appear so close together
that ground instruments cannot distinguishbetween them.
The f/151 relay has a special feature to block
unwanted light from the field of view – twocoronagraphic fingers in the aperture. The

Fig. 4-15 FOC optical relay system layout
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fingers can block the light of a bright star so thata faint target close to a bright one can bedetected with less intrusive light. This is usefulwhen observing a bright star that has a faintcompanion suspected of being a planet.
Another feature in the f/151 optical path is aseries of four filter wheels, which contain48 filters, prisms, and other optical devices. For
example, astronomers can send the beamthrough a magnesium-fluoride prism to emphasize the far-ultraviolet end of the spectrum.
Figure 4-16 shows the complete f/151 opticallayout.
The f/75.5 Optics System. This system has awider field of view – 28 arcsec2 – but less resolution than the f/151 optics system. The f/75.5
can be used for imaging and spectrography. The
f/75.5 optics system has an overall structure
identical to that of the f/151 and shares thesame calibration device. The separate components of the f/75.5 system are the reflecting
optics, the two filter wheels, and a detector likethe f/151 detector.
The f/75.5 system also has a 12.5-arcsec-longspectroscopic slit in its aperture. A mirror
rotated into the optical path diverts lightcoming through the slit onto a diffraction
grating. This grating breaks the light into aspectrum of the composite wavelengths. The
dispersed light reflects into the detector.
The spectrographic slit has a limited spectral
range: from 3600 to 5400 angstroms for the firstorder to 1200 to 1800 angstroms for the thirdorder. The detectors can resolve spectral lines
five angstroms apart. This spectral resolution
quality falls between that of the FOS and the
GHRS. The FOC spectrograph’s value is in
making spectrographic observations acrossspatially extended objects, such as galaxies, withthe full angular resolution of the FOC.

4-16

Fig. 4-16 Optical relay system layout at f/151

The f/75.5 optics (see Fig. 4-17) also include two
filter wheels, which can be used with either themain optics or the spectrographic optics for specialized viewing. For example, the filters canblock out certain wavelengths or select very
specific wavelength ranges.

Various formats and filters can change thewavelength range. Occultation mode uses thecoronagraphic fingers to block light from abright object so astronomers can study the surrounding faint background.

4.3.2 Observation Modes

4.3.3 Faint Object Camera
Specifications

The FOC has four observation modes:
acquisition, imaging, occultation, and spectrographic.
The acquisition mode precedes any observations using the camera’s coronagraphic or spectrographic devices. A special exposure of atarget is processed by the Science Instrument
Command and Data Handling (SI C&DH) unit
to find the target within the camera’s field ofview. Then the Telescope is repointed to place
the light onto the coronagraphic finger or spectrographic slit.
Imaging mode uses either of the optics systems
to take a direct-image measurement of a target.
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Figure 4-18 shows specifications for the FOC.
Note:
•Focal ratio can be selected as commandedfrom the ground.
•Fields of view listed are the maximum foreach focal ratio.
•The wavelengths studied usually are a bandwithin the maximum range. For example, one
study may concentrate on the band from 1200
to 1800 angstroms.
•Specific signal-to-noise ratios depend on
observation ranges and exposure times, but
indicate the camera’s limits.

4-17

Fig. 4-17 Optical relay system layout at f/75.5

4.3.4 Observations
The FOC can be used by itself or with other
Telescope science instruments for important
observations, including:•Studying interstellar gas clouds for evidence
relating to star formation and evolution•Measuring the distance of the farthest galaxies and quasars•Examining globular clusters and galaxies, normal and irregular•Providing high-resolution observations ofvarious solar system phenomena.
Stellar Evolution. Stellar formation begins
when gas clouds, formed by microscopic dust

Faint Object Camera
WeightDimensions
Principal investigator
Contractor
Optical modes
Field of view
Magnitude range
Wavelength range

700 lb (318 kg)
3 x 3 x 7 ft (0.9 x 0.9 x 2.2 m)
F. D. Macchetto, ESA/STScI
Dornier Systems and Matra Espace Corp
f/151, f/75.5
14, 28 arcsec 2
5 to 28 m v
1150 to 6500 angstroms
K70110-418

Fig. 4-18 FOC specifications
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and gas, are buffeted by shock waves fromnearby newly formed and internally
combusting stars; from supernova explosions;or from the whirling movement of the arms ofspiral galaxies. This violent motion collapses thegas clouds onto themselves. Growing hotter andheavier, the clouds increasingly compress tohigher density and temperatures, ignite insustained nuclear reactions, and become
protostars.
Galaxies and Quasars. Galaxies fill the universe. Many questions arise when studying
galaxies; for example, are the bright centers
filled with clumps of stars massed together? The
centers of some galaxies radiate more energy
than the surrounding galaxy itself. Of particularinterest is the center of elliptical galaxy
NGC4486, also known as Messier 87. Scientists
have found what appears to be a massive blackhole at its center.
The FOC also will concentrate on known
quasars to see if surrounding galaxies are hidden by the quasars’ tremendous luminosity. The
masking ability of the camera may help discover these faint galaxies. Because many

4-18

quasars may have formed shortly after the Big
Bang, they and their galaxies may reveal
information about the early composition anddevelopment of the universe.
Examining Solar System Objects. With the
FOC, astronomers can observe planets, moons,asteroids, and other bodies in the solar system
more clearly than from ground-based observatories. Scientists obtain data similar in detail tothat obtained by the Voyager spacecraft when
it was 5 days from encounter. However, theFOC enables astronomers to examine a planetary object for years instead of the brief time
dictated by Voyager’s flyby trajectories.
Direct observation of Mars, for example, will
advance scientific understanding of the planet’s
geography and weather, such as wind patterns,
dust storm velocities, white-cloud formation,and seasonal polar-cap movement. Close-up
views of Jupiter, Saturn, Neptune, Uranus, andPluto will expand our knowledge of the atmospheric composition and the dynamics on theseouter planets. The camera’s angular resolution
allows examination of small asteroids and planetary moons in greater detail than ever before
to determine accurate dimensions, axis orientation, orbits, and composition. The FOC alsostudies comets using its spectrographic optics
and a variety of filters.

4.4

Wide Field/Planetary
Camera II

The most often used science instrument is the
WF/PC. It records two-dimensional images at
two magnifications through a selection of 48
color filters covering a spectral range from farultraviolet to visible and near-infrared wavelengths. It provides pictorial views of the celestial universe on a grander scale than any other
instrument flown to date. Like its predecessor
WF/PC I, WF/PC II was designed and built at
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NASA’s Jet Propulsion Laboratory (JPL), which
is operated by the California Institute ofTechnology. Professor James A. Westphal ofCaltech was the principal investigator for
WF/PC I. Dr. John T. Trauger of JPL is the principal investigator for WF/PC II.
WF/PC I, the first-generation instrument, waslaunched with the Telescope in 1990 and functioned flawlessly. The second-generation
instrument, WF/PC II, was already under construction when the Hubble Telescope waslaunched. Its original purpose was to provide
a backup for WF/PC I with certain enhancements, including an upgraded set of filters,
advanced detectors, and improved UV performance. With modifications introduced since
1990, WF/PC II also provided built-in compensation for the improper curvature of theTelescope’s primary mirror so as to achieve theoriginally specified imaging performance of theTelescope in the WF/PC II field of view.
WF/PC II has four CCD cameras arranged torecord simultaneous images in four separate
fields of view at two magnifications.
In three Wide Field Camera fields, each detector
picture element (pixel) occupies 1/10th arcsec,
and each of the three detector arrays covers asquare 800 pixels on a side (or 80 arcsec, slightly
more than the diameter of Jupiter when it isnearest the Earth). The Telescope is designed toconcentrate 70 percent of the light of a star
image into a circle 0.2 arcsec (or two Wide Field
Camera pixels) in diameter. This three-fieldcamera (which operates at a focal ratio of f/12.9)
provides the greatest sensitivity for thedetection of faint objects. Stars as faint as 28th
magnitude are detectable in the longest
anticipated exposures of about 28 hours (22ndmagnitude is almost 1,000 million times fainterthan can be seen with the naked eye).

4-19

The Planetary Camera provides a magnification
about 2.2 times larger, in which each pixel
occupies only 0.046 arcsec, and the single square
field of view is only 36.8 arcsec on a side. It
operates at a focal ratio of f/28.3. Originallyincorporated for studying the finest details ofbright planets, the Planetary Camera actuallyprovides the optimum sampling of the Telescope’s images at visible wavelengths and isused (brightness permitting) whenever the finest possible spatial resolution is needed, evenfor stars, stellar systems, gaseous nebulae, andgalaxies.

astronomers to choose the instrumental
characteristics best suited to a particularobservational task.

With its two magnifications and its built-in
correction for the Telescope’s spherical aberration, WF/PC II can resolve the fine details andpick out bright stellar populations of distantgalaxies, perform precise measurements of thebrightness of faint stars, and study the characteristics of stellar sources even in crowded areas
such as globular clusters – ancient swarms of as
many as several hundred thousand stars that
reside within a huge spherical halo surrounding the Milky Way and other galaxies – that
could not be studied effectively with WF/PC I
because of the aberration. WF/PC II’s highresolution photography of the planets within
our solar system allows continued studies of
their atmospheric composition as well as
discovery and study of time-varying processeson their surfaces.

The 0.1-arc second resolution of the Wide Field
Camera is too coarse for some purposes because
the finest details cannot be resolved with it or
because the positions of star images cannot be
pinpointed with sufficient accuracy. An example
arises in the study of stars gravitationally boundto each other in orbits about their common center of mass. Such studies enable astronomers todetermine the relative masses of the orbitingstars, which are of fundamental importance instudying the physics of stars. In many instances,
positional accuracies approaching 1/1,000th
arcsec are needed. In this case, the high magnification of the f/28.3 Planetary Camera or theeven greater magnifications of the f/75.5 or
f/151 cameras would be used.

4.4.1 Comparison of WF/PC II
and Faint Object Camera
For the same reasons that a photographer
benefits from having a selection of lenses of different focal lengths, or possibly a zoom lenswith an adjustable focal length, astronomers
using the HST benefit from having the choiceof several different magnifications offered by
WF/PC II and the FOC. This selection allows
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For example, in the quest to measure the rate of
expansion of the universe, it is crucial to discover as many Cepheid variable stars in clusters of galaxies as possible. Because Cepheids
are rare and randomly scattered over the sky,
this task calls for the largest possible field ofview if it is to be completed efficiently. Therefore, such studies will utilize the Wide Field
Camera, with its focal ratio of f/12.9.

The price paid for high magnification is that thefield of view, in terms of area, varies inversely
as the square of the magnification. Similarly, theexposure time required to reach a given limiting magnitude or faintness varies directly as thesquare of the magnification. It would be necessary to use a longer focal length if the object
being studied is too bright for the Wide Field
Camera. Because of these various factors, theastronomer’s choice of which magnification touse often is a compromise.

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4.4.2 Physical Description
The WF/PC occupies one of four radial bays inthe focal plane structure of the HST. The otherthree radial bays support the FGSs, which areused primarily for controlling the pointing of
the Telescope. The three other science instruments and COSTAR occupy axial bays.
The WF/PC II’s field of view of is located at thecenter of the Telescope’s field of view, where thetelescopic images are nearly on axis and least
affected by residual aberrations (field curvature
and astigmatism) that are inherent in the
Ritchey-Chretien design.
Because the focal plane is shared by the other
instruments, WF/PC II is equipped with a flat
pickoff mirror located about 18 in. ahead of the
focal plane and tipped at almost 45 degrees tothe axis of the Telescope. The pickoff mirror is
attached to the end of a stiff truss, which is rigidly fastened to WF/PC’s precisely locatedoptical bench. The pickoff mirror reflects the
portion of the Telescope’s focal plane belonging
to WF/PC II into a nearly radial direction inwhich it enters the front of the instrument,
allowing light falling on other portions of the
focal plane to proceed without interference.
WF/PC II is shaped somewhat like a piece ofpie, the pickoff mirror lying at the point of thewedge, with a large, white-painted cylindrical
panel 2.6 ft (0.8 m) high and 7 ft (2.2 m) wideat the wide end. The panel forms part of thecurved outer skin of the Support Systems Module (SSM) and radiates away the heat generated
by WF/PC’s electronics. The instrument is heldin position by a system of latches and isclamped in place by a threaded fastener at theend of a long shaft that penetrates the radiatorand is accessible to the astronauts.
K70110-04M

WF/PC II weighs 619 lb (281 kg). The cameras
comprise four complete optical subsystems,
four CCDs, four cooling systems using thermoelectric heat pumps, and a data-processing system to operate the instrument and send data to
the SI C&DH unit. Figure 4-19 shows the overall configuration of the instrument.
Optical System. The WF/PC II optical system
consists of the pickoff mirror, an electricallyoperated shutter, a selectable optical filter
assembly, and a four-faceted reflecting pyramid
mirror used to partition the focal plane to thefour cameras. Light reflected by the pyramid
faces is directed by four “fold” mirrors into eachof four two-mirror relay cameras. The relays
re-image the Telescope’s original focal plane
onto the four detector arrays while providing
accurate correction for the spherical aberrationof the Telescope’s primary mirror. Figure 4-20
shows the light path from the Telescope to thedetectors.
As in an ordinary camera, the shutter is used tocontrol the exposure time, which can range fromabout 1/10th second to 28 hours. Typical exposure times are 45 minutes, about the timerequired for the Telescope to complete half an
orbit.
WF/PC II’s pickoff mirror and three of its four
fold mirrors are equipped with actuators thatallow them to be controlled in two axes (tip andtilt) by remote control from the ground. The
actuators ensure that the spherical aberration
correction built into WF/PC II is accuratelyaligned relative to the Telescope in all fourchannels.
The Selectable Optical Filter Assembly (SOFA)
consists of 12 independently rotatable wheels,
each carrying four filters and one clear opening,
for a total of 48 filters. These can be used singly

4-21

Fig. 4-19 Wide Field/Planetary Camera (WF/PC) overall configuration

Fig. 4-20 WF/PC optics design
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4-22

or in certain pairs. Some of the WF/PC II’s
filters have a patchwork of areas with differing
properties to provide versatility in the measurement of spectral characteristics of sources.
WF/PC II also has a built-in calibration channel,
in which stable incandescent light sources serve
as references for photometric observations.
Charge-Coupled Detectors. A CCD is a device
fabricated by methods developed for the manufacture of integrated electronic circuits. Functionally, it consists of an array of light-sensitive
picture elements (pixels) built upon a thin waferof crystalline silicon. The light-sensitive elements are controlled by complex electronic circuits also built onto the wafer. The circuits
include low-noise amplifiers to strengthen signals that originate at the light sensors. As lightfalls upon the array, photons of light interact
with the sensor material to create small electrical charges (electrons) in the material. The
charge is very nearly proportional to the number of photons absorbed. The built-in circuits
read out the array, sending a succession of signals that will allow later reconstruction of thepattern of incoming light on the array. Figure 4-21 illustrates the process.

Fig. 4-21 WF/PC II imaging
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The CCDs used in WF/PC II consist of 800 rows
and 800 columns of pixels, 640,000 pixels in eacharray. The pixels can be thought of as tinysquares side by side, 15 microns (about 6/10,000in.) on a side. Their sensitivity to light is greatest at near-infrared and visible wavelengths, but
in WF/PC it is extended to the UV by coatingthem with a thin fluorescent layer that converts
UV photons to visible ones.
To achieve a very low-noise background that
does not interfere with measurements of faintastronomical light sources, the CCDs must be
operated at a low temperature, approximately-50 to -70ϒC (-8 to -130ϒF). This is accomplishedby an electrically operated solid-state coolingsystem that pumps heat from the cold CCDs to
the warmer external radiator by means of heat
pipes. The radiator faces away from the Earthand sun so that its heat can be effectively
radiated into the cold vacuum of space.
CCDs are much more sensitive to light thanphotographic film and many older forms ofelectronic light sensors. They also have finer
resolution, better linearity, and ability to convert
image data directly into digital form. As a result,
CCDs have found many astronomical and commercial applications following their early
incorporation in WF/PC I.
Processing System. A microprocessor controls
all of WF/PC II’s operations and transfers data
to the SI C&DH unit. Commands to control
various functions of the instrument (including
filter and shutter settings) are sent by radio
uplink to the Telescope in the form of detailed
encoded instruments originated at the Space
Telescope Science Institute (STScI) in Baltimore,
Maryland. Because the information rate of theTelescope’s communication system is limited,
the large amount of data associated with evenone picture from WF/PC II is digitally

4-23

tape-recorded during the CCD readout. The
data then is transmitted at a slower rate via acommunications satellite that is simultaneously
in Earth orbit.

Wide Field/Planetary Camera II
WeightDimensions

4.4.3 WF/PC II Specifications

Principal investigator
Contractor
Optical modes
Field of view

Figure 4-22 shows the WF/PC II specifications.

Magnitude range
Wavelength range

K70110-422

4.4.4 Observations

Fig. 4-22 WF/PC II specifications

The WF/PC II can perform several tasks whileobserving a single object. It can focus on an
extended galaxy and take a wide-field pictureof the galaxy, then concentrate on the galaxy
nucleus to measure light intensity and take photographic close-ups of the center. In addition,
the WF/PC II can measure while other instruments are observing.
Specific applications of this camera range fromtests of cosmic distance scales and universe
expansion theories to specific star, supernova,
comet, and planet studies. Important searches
will be made for black holes, planets in otherstar systems, Mars atmospheric storms, and theconnection between galaxy collisions and starformation.

4.5

619 lb (281 kg)
Camera: 3.3 x 5 x 1.7 ft (1 x 1.3 x 0.5 m),Radiator: 2.6 x 7 ft (0.8 x 2.2 m)
John Trauger, Jet Propulsion Laboratory
Jet Propulsion Laboratory
f/12.9 (WF), f/28.3 (PC)
4.7 arcmin2 (WP),
0.3 arcmin2 (PC)
9 to 28 m v
1200 to 10,000 angstroms

their apertures, as shown in Fig. 4-23. (See para
5.3 for more details.)

4.5.1 Fine Guidance Sensor
Specifications
Figure 4-24 shows FGS specifications.

4.5.2 Operational Modes for
Astrometry
Once the two target-acquisition FGSs lock onto
guide stars, the third sensor can perform
astrometric operations on targets within thefield of view set by the guide stars’ positions.
The sensor should be able to measure stars as
faint as 18 apparent visual magnitude.

Astrometry (Fine Guidance
Sensors)

When two FGSs lock on guide stars to provide
pointing information for the Telescope, the third
FGS serves as a science instrument to measure
the position of stars in relation to other stars.
This astrometry helps astronomers determine
stellar masses and distances.
Fabricated by Hughes Danbury Optical Systems, Inc., the sensors are in the focal planestructure, at right angles to the optical path of
the Telescope and 90 degrees apart. They have
pickoff mirrors to deflect incoming light into
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4-24

Fig. 4-23 Fine Guidance Sensor (FGS)to classify the stars being observed. The wheel
has a clear filter for guide-star acquisition andfaint-star (greater than 13 apparent visual magnitude) astrometry; neutral-density filter for
observation of nearby bright stars; and two colored filters for estimating a target’s color(chemical) index, increasing contrast between
close stars of different colors, or reducing background light from star nebulosity.

Fine Guidance Sensor
WeightDimensions

485 lb (220 kg)
1.6 x 3.3 x 5.4 ft(0.5 x 1 x 1.6 m)

Contractor

Perkin-Elmer Corporation
William H. Jefferys

Principal investigator
Astrometric modes
Precision

Stationary and moving target, scan
0.002 arcsec2

Measurement speed

10 stars in 10 minutes

Field of view

Access: 60 arcmin2
Detect: 5 arcsec
4 to 18.5 mv
4670 to 7000 angstroms

Magnitude range
Wavelength range

4.5.4 Astrometric Observations

K70110-424

Fig. 4-24 FGS specifications

There are three operational modes for
astrometric observations: position, transferfunction, and moving-target. Position mode
allows the astrometric FGSs to calculate theangular position of a star relative to the guidestars. Generally, up to 10 stars will be measured
within a 20-minute span.
In the transfer-function mode, sensors measure
the diameter of a stellar target, either throughdirect analysis of a single-point object or byscanning a diffuse target. Examples of the latterinclude solar system planets, double stars visually closer together than 0.02 arcsec, and targets
surrounded by nebulous gases.
Binary stars visually separated by more than
0.02 arcsec can produce astrometricmeasurements of stellar masses, leading to
information on stellar gravity in the evolutionof star and planetary systems.
In moving-target mode, sensors measure a rapidly moving target relative to other targets when
it is impossible to precisely lock onto the moving target; for example, measuring the angular
position of a moon relative to its parent planet.

4.5.3 Fine Guidance Sensor Filter
Wheel
Each FGS has a filter wheel for astrometric measurement of stars with different brightness andK70110-04M

Astronomers measure the distance to a star bycharting its location on two sightings from Earthat different times, normally 6 months apart. The
Earth’s orbit changes the perceived (apparent)
location of the nearby star, and the parallax
angle between the two locations can lead to an
estimate of the star’s distance. Because stars are
so distant, the parallax angle is very small,requiring a precise field of view to calculate theangle. Even with the precision of the FGSs,
astronomers cannot measure distances by theparallax method beyond nearby stars in ourgalaxy.
An important goal of the FGS astrometry projectis to obtain improved distances to fundamental
distance calibrators in the universe, for instanceto the Hyades star cluster. This is one of thefoundations of the entire astronomical distancescale. An accurate distance to the Hyades would
make it possible for astronomers to infer accurate distances to similar stars that are too distant for the direct parallax method to work.
Astronomers have long suspected that somestars might have a planetary system like oursun’s. Unfortunately, the great distance of stars
and the faintness of any possible planet makeit very difficult to detect such systems directly.
It may be possible to detect a planet by observing nearby stars and looking for the subtle
gravitational effects that a planet would have on

4-25

the star that it is orbiting. Observations of
Proxima Centauri (the nearest star to the sun)and Barnard’s star now are being conducted in
the hope that a planetary companion may befound.
Astronomers use the FGS in two modes ofoperation to investigate known and suspectedbinary star systems. Their observations lead tothe determination of the orbits and parallaxesof the binary stars and therefore to the massesof these systems. For example, 40 stars in theHyades cluster were observed with the FGS. Tenof the targets were discovered to be binary starsystems and one of them has an orbital periodof 3.5 years.
Other objects, such as nearby M dwarf stars
with suspected low mass companions, are being
investigated with the FGS with the hope of
improving the mass/luminosity relationship at
the lower end of the main sequence.

4.6

HST Operations

Space Telescope mission operations are of two
types:•Science operations that observe celestial
objects and gather data
•Engineering operations that calibrate, test, andmaintain the Telescope’s overall performance.
Science and engineering operations often coincide and interact. For example, a science instrument may observe a star and calibrate incoming wavelengths against standards developed
during scientific verification.
Mission operations are carried out by the Telescope ground system, which consists of the
STScI and Space Telescope Operations Control
Center (STOCC). The STScI oversees science
operations. It hosts astronomers, evaluates andK70110-04M

chooses observation programs, schedules theselected observations, generates an overall mission timeline and command sequences, andstores and analyzes data from the Telescope.
Meanwhile, the STOCC makes day-to-day
operational decisions through the Payload
Operations Control Center (POCC) at GSFC. It
interacts with the STScI to receive daily mission
schedules, to process quick-look data anddisplays, and to manage the science data.
Figure 4-25 details the ground functions.

4.6.1 Space Telescope Science
Institute
Located in Baltimore, Maryland, the STScI is
responsible to GSFC for the science programs on
the HST. It is operated by the Association of
Universities for Research in Astronomy(AURA), a consortium of 20 United States universities that operates several national facilities
for astronomy.
The STScI solicits and reviews observation proposals and selects observations to be carried out.
It schedules observations and assists guest
observers in their work; generates an integrated
science and engineering timeline to support all
spacecraft activities, including any special engineering tests; and provides the facilities andsoftware to reduce, analyze, archive, and distribute Telescope data.
Additionally, STScI generates an integrated science and engineering timeline to support all
spacecraft activities, including any special engineering tests. It also monitors the Telescope andscience instruments for characteristics that could
affect science data collection, for example,
instrument performance quality, pointing inaccuracies, and Telescope focus.
Scientific Goals. The STScI helps conduct thescience program to meet the overall scientific4-26

Fig. 4-25 Space Telescope ground systems overview

goals of the Telescope program, set by theInstitute and NASA in consultation with
AURA’s Space Telescope Institute Council andcommittees representing the international astronomical community.
Institute Software. Computer hardware andsoftware play an important role in STScI work.
The Science Operations Ground System (SOGS)
is a data-management and scheduling computer
system developed by TRW under contract toNASA. The Guide Star Selection System (GSSS)
created the guide star catalog used by the pointing control subsystem. Science Data Analysis
Software (SDAS) provides analytical tools for
astronomers studying processed data. STScIdeveloped the GSSS and SDAS.
K70110-04M

SOGS handles mission planning and scheduling, observation support, software support, androutine data processing. Together, these programs perform the computations needed to run
the science operations on the Telescope.
GSSS provides reference stars and other brightobjects so the FGSs can point the Telescope
accurately. This system selects guide stars thatcan be located unambiguously in the sky when
the sensors point the Telescope. The guide starcatalog has information on 20 million celestial
objects, created from 1,477 photographic survey
plates covering the entire sky.
After SOGS collects, edits, measures, andarchives science data, observers can use SDAS
to analyze and interpret the data.

4-27

Selecting Observation Proposals. Astronomers
worldwide may use the Telescope. Any scientist may submit a proposal to the STScI outlining an observing program and describing thescientific objectives and instruments required.
The STScI selects observations by evaluatingthese requests for technical feasibility, conducting peer reviews, and choosing the highest
ranked proposals. Because individual astronomers and astronomy teams are expected to submit many more proposals than can possibly be
accepted, a team approach is encouraged. The
final decision rests with the STScI director,
advised by a review committee of astronomers
and scientists from many institutions.
First priority for observations goes to principal
investigators – the scientists and astronomers
who designed the instruments. Principal investigators and their contributions are:
• Rodger Thompson, University of Arizona(NICMOS)• Bruce E. Woodgate, GSFC (STIS)• F. Macchetto, ESA/STScI (FOC)• J. Trauger, NASA JPL (WF/PC II)• William H. Jeffreys, University of Texas (FGS).
Scheduling Selected Observations. The primary scheduling consideration is availability ofa target, limited by environmental and straylight constraints – for example, a faint object thatmust be observed when the Telescope is in
Earth’s shadow. The schedule takes into consideration system limits, observations that use
more than one instrument, and required time
for special observations.
Data Analysis and Storage. STScI archives analyzed data. Computer resources include the
SDAS and other selected computer facilities.
SOGS data processing receives science datafrom the Packet Processing Facility (PACOR) at
K70110-04M

GSFC, then automatically formats it and verifies its quality. It calibrates data to remove
instrument interference such as variation in thedetector’s sensitivity across the data field. Thenthe software places the data on archive
computer tapes from which the data can beformatted into printed reports, printer plots,
and photographic prints and negatives. The
STScI processes all data within 24 hours after
receipt.
The SDAS package is written so that observers
can interact to process images received by the
SOGS software. In addition, individual
observers will be encouraged to bring their own
data analysis software and reference data.
STScI is responsible for storing the massive
amount of data collected by the Telescope. A
database management system records the location and status of data as it pours into the storage banks. Observers and visiting astronomers
can easily retrieve the stored data for examination or to use data manipulation procedures
created on STScI computers.
ESA provides approximately 15 percent of the
STScI staff and operates its own data analysis
facility in Garching, Germany.
In addition to science data, computers store
engineering data. This is important for developing more efficient use of the Telescope systems
and for adjusting Telescope operations based onengineering findings, for example, if an instrument provides unreliable data in certain
temperature ranges.

4.6.2 Space Telescope Operations
Control Center
STOCC runs day-to-day spacecraft operations
through the POCC and PACOR. In addition,
STOCC works with the NASA Communications

4-28

Network (NASCOM) and the Tracking andData Relay Satellite System (TDRSS).

data. STScI alerts POCC to that need, and theincoming data can be processed for theobserver.

POCC uses mission-control facilities at GSFC
built specifically for HST operations.

Another important part of STOCC is PACOR
processing. When data arrives from NASCOM
for science handling, PACOR reformats thedata, checks for noise or transmission problems,and passes each packet of data along with adata quality report.

STOCC has three major operational responsibilities: spacecraft operations, including
sending commands to the spacecraft; telemetry
processing; and off-line support.
Some commands may pass to the Telescopefrom POCC, based on specific observation
objectives. POCC translates those goals into
commands, then passes them along to theTelescope through the communications
network. The commands may re-orient theTelescope, open an instrument aperture, turn on
the instrument’s detectors, request a certain
filter or grating setting, or monitor data goingfrom the detector to the instrument’s
transmission device and out through the SI
C&DH unit.
Most specific spacecraft operations are commanded by software embedded in theTelescope’s onboard computer. The POCC
orders the computer to activate the storedcommands.
Telemetry comes into POCC from the communication satellites and provides information thatreflects the equipment operation status or spacecraft operation. For example, telemetry canshow that voltage going to the Rate Gyro
Assembly indicates a power surge, which could
disable a gyroscope and affect the stability of theTelescope. POCC defuses the voltage surge orswitches to a backup gyro until the problem issolved. In many cases, consultation between
POCC and STScI is necessary, particularly if thedata affects an ongoing observation.
An important STScI function is to support
observers requiring a “quick-look” analysis of
K70110-04M

Support for STOCC comes from the Spaceflight
Tracking and Data Network (STDN), composedof TDRSS and NASCOM.
TDRSS has two communications relay satellites
130 degrees apart, with a ground terminal at
White Sands, New Mexico. There is a small“zone of exclusion” where Earth blocks theTelescope signal to either of the satellites, but up
to 91 percent of the Telescope’s orbit is withincommunications coverage. Tracking andData Relay Satellites (TDRS) receive and send
both single-access (science data) and multipleaccess (commands and engineering data)channels.
NASCOM leases domestic satellites for commercial communications, such as televisiontransmission. These satellites pass data from
TDRSs directly to STOCC.

4.6.3 Operational Characteristics
Three major operational factors affect thesuccess of the Telescope: orbital characteristics
for the spacecraft, its maneuvering characteristics, and communications characteristics forsending and receiving data and commands.
Orbital Characteristics. The orbit of theTelescope is approximately 320 nmi (593 km). It
maintains an orbit between a minimum
operating altitude of approximately 200 nmi(368 km) – to keep above altitudes where4-29

atmospheric drag comes into noticeable play –and the deployment orbit. The orbit inclines at
a 28.5-degree angle from the equator because
the Shuttle launch was due east from Kennedy
Space Center. The chosen orbit puts the sun in
the Telescope orbital plane so that sunlight falls
more directly on the Solar Arrays. In addition,
the orbit is high enough that aerodynamic dragfrom the faint atmosphere at that level will notdecay the Telescope’s orbit to below theminimum operating altitude.
The Telescope completes one orbit every 97 minutes, passing into the shadow of the Earth during each orbit. The time in shadow varies from28 minutes to 36 minutes. The variation during
a nominal 30-day period is between 34.5 and36 minutes in shadow. If Earth blocks an object
from the Telescope, the Telescope re-acquires theobject as the spacecraft comes out of Earth’s
shadow. Faint-object viewing is best while theTelescope is in Earth’s shadow. Figure 4-26
shows the nominal orbit.
The Telescope orbit is tracked by the TDRSS,
which plots the spacecraft’s orbit at least eighttimes daily and sends the data to the Flight
Dynamics Facility at GSFC. This helps predict
future orbits, although some inaccuracy in predicting orbital events such as exit from Earth’s
shadow is unavoidable. The environmental

elements with greatest effect on the Telescope’s
orbit are solar storms and other solar activities.
These thicken the upper atmosphere andincrease the drag force on the Telescope, accelerating the orbit decay rate considerably.
Celestial Viewing. The Telescope is pointed
toward celestial targets as a normal orientation
to expose instrument detectors for up to
10 hours, if needed. A continuous-viewing zone
exists, parallel to the orbit plane of the Telescope
and up to 18 degrees on either side of the northand south poles of that orbital plane (seeFig. 4-27). Otherwise, celestial viewing depends
on how long a target remains unblocked by
Earth.
The amount of shadow time available for faintobject study also affects celestial observations.
Shadow time for an observation varies with thetime of year and the location of the target relative to the orbit plane. Astronomers use a geometric formula to decide when in a given period
a target will be most visible while the Telescopeis in shadow.

Fig. 4-26 HST nominal orbit
K70110-04M

Fig. 4-27 “Continuous-zone” celestial viewing
4-30

Other sources affecting celestial viewing are
zodiacal light and integrated or backgroundstarlight. These affect viewing with certain
instruments, such as the light-intensity sensitive
FOC.
Solar System Object Viewing. Solar systemobjects also are affected by the factors mentioned for celestial viewing. In addition, theTelescope works with imprecise orbit parameters for itself and objects such as the outer planets and comets. For example, Neptune’s center
may be off by 21 km when the sensors try tolock onto it because the Telescope is changing
its position in orbit, which affects the pointing
direction toward nearby objects. However, mostsolar system objects are so bright the Telescopeneeds only a quick snapshot of the object to fix
its position. Tracking inaccuracies are more
likely to cause a blurred image if they occur
with long-exposure observations of dim targets.
The Telescope’s attitude roll also may affect theview of the object and require a maneuver thatrolls the spacecraft more than the 30-degree
limit (for example, to place the image into aspectrographic slit aperture).
Tracking interior planets (Mercury and Venus)
with the Telescope places the sun within theTelescope opening’s 50-degree sun-exclusionzone. To minimize danger from this exposure,
the Telescope views these objects using Earth toblock (occult) the sun, that is, after the sunset
shadow falls on the Telescope.

which is in shadow, precedes the illuminatedpart of the moon (see Fig. 4-28).
Natural Radiation. Energetic particles from different sources bombard the Telescope continuously as it travels around the Earth. Geomagnetic shielding blocks much of the solar andgalactic particle radiation. When the Telescope
passes through the South Atlantic Anomaly(SAA), a “hole” in Earth’s magnetic field,
charged particles can enter the Telescope andstrike its detectors, emitting electrons and producing false data.
The Telescope passes through SAA for segmentsof eight or nine consecutive orbits, then has nocontact with it for six or seven orbits. Each encounter lasts up to 25 minutes. In addition, SAA
rotates with Earth, so it occasionally coincides
with the Telescope as the spacecraft enters
Earth-shadow observation periods. Careful
scheduling minimizes the effects of the anomaly,
but it has some regular impact.
Solar flares are strong pulses of solar radiation,accompanied by bursts of energetic particles.
Earth’s magnetic field shields the lower magnetic latitude regions, such as the Telescope’s
orbit inclination, from most of these charged

Lunar Occultation Viewing. FGSs look awaywhen the Telescope approaches within
10 degrees of the bright moon. By overriding thecontrols that protect the sensor star selector servos, the ground control team can use the moon
as an occulting object for an observation. The
moon likely would be between its new-moon
and quarter-moon phases so the occulting edge,
K70110-04M

4-31

Fig. 4-28 Using the moon as an occulting disk

particles. NASA regularly monitors the flares,
and the Telescope can stop an observation until
the flares subside. The greatest possible physicaldanger is to the crew during extravehicular
activity, which would be halted until the flares
subside.
Maneuver Characteristics. The Telescopechanges its orientation in space by rotating its
reaction wheels, then slowing them. The momentum change caused by the reaction moves
the spacecraft at a baseline rate of 0.22 degree
per second or 90 degrees in 14 minutes. Figure 4-29 shows a roll-and-pitch maneuver.
When the Telescope maneuvers, it takes a few
minutes to lock onto a new target and accumulate drift errors. This means that a larger regionof the sky must be scanned for guide stars.
One consideration with maneuvering is the danger of moving the Solar Array wings out of the
sun’s direct radiation for too long. Unprotected
portions of the SSM aft shroud could be affected
thermally. Therefore, maneuvers beyond a certain
range in angle and time are limited.

When the Telescope performs a pitch to a target
near the 50-degree sun-avoidance zone, theTelescope curves away from the sun. For
example, if two targets are opposed at
180 degrees just outside the 50-degree zone, theTelescope follows an imaginary circle of
50 degrees around the sun until it locates thesecond target (see Fig. 4-30).
Communication Characteristics. The Telescopecommunicates with the ground via TDRSS.
With two satellites 130 degrees apart inlongitude, the maximum amount of contact
time is 94.5 minutes of continuous
communication, with only 2.5 to 7 minutes in azone of exclusion out of reach of either TDRS
(see Fig. 4-31). However, orbital variations by
the Telescope and communications satellites
affect this ideal situation to widen the zone of
exclusion slightly.

-V1

V1 (ROLL) MANEUVER(VIEWING AWAY FROM SUN)
50°

SUN

-V2
V2 (PITCH) MANEUVERS(MANEUVER PLANE CONTAINS SUN)
K70110-429

Fig. 4-29 HST single-axis maneuvers
K70110-04M

Fig. 4-30 Sun-avoidance maneuver
4-32

TDRS EASTLONGITUDE 41 DEG W

TDRS WESTLONGITUDE 171 DEG W

EARTH
TDRS EASTSHADOW ZONETDRS WEST
SHADOW ZONENO COVERAGE ZONE
K70110-431

Fig. 4-31 TDRS-HST contact zones

The GSFC NCC schedules all TDRS communications. The Telescope has a general orbital
communication schedule, supplemented by
specific science requests. The Network Control
Center (NCC) prepares schedules 14 days before
the start of each mission week.
The backup communications link is the GroundSpacecraft Tracking and Data Network(GSTDN), which receives engineering data, orscience data if the High Gain Antennas (HGA)
cannot transmit to TDRSS. The longest single
contact time is 8 minutes. The limiting factor ofthis backup system is the large gap in time
between contacts with the Telescope. In practical terms, at least three GSTDN contacts are
required to read data from a filled science taperecorder – with gaps of up to 11 hours between
transmissions.

4.7

Acquisition and
Observation

The major steps in the observation process aretarget acquisition and observation, datacollection and transmission, and data analysis.
Each science instrument has an entrance
aperture, located in different portions of the
Hubble’s focal plane (see Fig. 4-32). The
different aperture positions make precise

Each HGA maintains continuous contact with
one TDRS to avoid unnecessary gaps in communication. Each antenna tracks the communication satellite, even during fine-pointing
maneuvers.
The Low Gain Antennas provide at least 95 percent orbital coverage via a TDRS for the minimum multiple-access command rate used.
K70110-04M

4-33

Fig. 4-32 HST instrument aperturespointing a sometimes-lengthy procedure for the
FGSs. In addition to the small apertures inwhich the FGSs must center the target, time isrequired to reposition the Telescope – anestimated 18 minutes to maneuver 90 degrees,
plus the time the sensors take to acquire theguide stars. If the Telescope overshoots its target, the Fixed Head Star Trackers may have to
make coarse-pointing updates before theTelescope can use the FGSs again.
To increase the probability of a successfulacquisition, Telescope flight software allows theuse of multiple guide-star pairs to accountfor natural contingencies that might affect a
guide-star acquisition – such as a guide star
being a binary star and preventing the
FGSs from getting a fine lock on the target.
Therefore, an observer can submit a proposalthat includes a multiple selection of guidestar pairs. If one pair proves too difficult toacquire, the sensors can switch to the alternate
pair. However, each observation has a limited
total time for acquiring and studying the target.

K70110-04M

If the acquisition process takes too long,
the acquisition logic switches to coarse-trackmode for that observation to acquire the guidestars.
Three basic modes are used to target a star.
Mode 1 points the Telescope, then transmits acamera image, or spectrographic or photometricpseudo-image, to STOCC. Ground computers
make corrections to precisely point theTelescope, and the coordinates pass up throughthe DF-224 computer. Mode 2 uses onboard
facilities, processing information from the larger
target apertures, then aiming the Telescope toplace the light in the chosen apertures. Mode 3
uses the programmed target coordinates in thestar catalog or updated acquisition information
to re-acquire a previous target. Called blindpointing, this is used mostly for generalized
pointing and for the WF/PC II, which does not
require such precise pointing. Mode 3 relies
increasingly on the updated guide-star
information from previous acquisition attempts,
stored in the computer system.

4-34

HUBBLE SPACE TELESCOPE SYSTEMST

he Hubble Space Telescope (HST) has
three interacting systems:

generate electrical power and charge onboard
batteries and communications antennas to receive commands and send telemetry data fromthe HST. Figure 5-1 shows HST configuration.

•The Support Systems Module (SSM), an outer
structure that houses the other systems andprovides services such as electrical power,
data communications, and pointing control
and maneuvering•The Optical Telescope Assembly (OTA), whichcollects and concentrates the incoming light in
the focal plane for use by the science
instruments•Eight major science instruments, four housedin an aft section focal plane structure (FPS)
and four placed along the circumference of thespacecraft. With the exception of the Fine
Guidance Sensors (FGS), all are controlled by
the Science Instrument Control and Data
Handling (SI C&DH) unit.

The Telescope performs much like a groundobservatory. The SSM is designed to support
functions required by any ground astronomical
observatory. It provides power, points the Telescope, and communicates with the OTA, SI
C&DH unit, and instruments to ready an observation. Light from an observed target passes
through the Telescope and into one or more of
the science instruments, where the light isrecorded. This information goes to onboard
computers for processing, then it is either temporarily stored or sent to Earth in real time, via
the spacecraft communication system.
The Telescope completes one orbit every 97 minutes and maintains its orbital position along
three axial planes. The primary axis, V1, runs

Additional systems that also support HST
operations include two Solar Arrays (SA). These

Fig. 5-1 Hubble Space Telescope – exploded view
5-1

through 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 maneuvers
to 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.

• Reflective surfaces and heaters for thermal

5.1• Structures and mechanisms

protection• Outer doors, latches, handrails, and footholds
designed for astronaut use during on-orbit
maintenance.
Figure 5-3 shows some of these features.
Major component subsystems of the SSM are:Support Systems Module• Instrumentation and communications

The design features of the SSM include:• Data management• Pointing control• An outer structure of interlocking shells• Electrical power• Reaction wheels and magnetic torquers to• Thermal control

maneuver, orient, and attitude stabilize theTelescope• Two SAs to generate electrical power• Communication antennas• A ring of Equipment Section bays that containelectronic components, such as batteries, andcommunications equipment. (Additional bays
are provided on the +V3 side of the spacecraft
to house OTA electronics as described inpara 5.2.4.)• Computers to operate the spacecraft systems
and handle data

•Safing (contingency) system.

5.1.1 Structures and Mechanisms
Subsystem
The outer structure of the SSM consists ofstacked cylinders, with the aperture door on top
and the aft bulkhead at the bottom. Fitting
together 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, covers the
opening to the Telescope’s light shield. The door
is made from honeycombed aluminum sheets.
The outside is covered with solar-reflecting
material, and the inside is painted black toabsorb stray light.

+V3

+V1

SOLAR ARRAY

–V2

+V3

The door opens a maximum of 105 degrees fromthe closed position. The Telescope aperture
allows for a 50-degree field of view (FOV) centered on the +V1 axis. Sun-avoidance sensors
provide ample warning to automatically closethe door before sunlight can damage theTelescope’s optics. The door begins closing
when the sun is within ±35 degrees of the +V1

+V2

–V1

+V2

–V3

K70110-502

Fig. 5-2 Hubble Space Telescope axes
5-2

Fig. 5-3 Design features of Support Systems Module

MAGNETICTORQUER (4)LIGHT
SHIELDaxis and finishes closing by the time the sunreaches 20 degrees of +V1. This takes no morethan 60 seconds.

APERTURE
DOOR

HIGH-GAIN ANTENNA (2)The Space Telescope Operations Control Center(STOCC) can override the protective doorclosing 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.

FORWARD SHELL

EQUIPMENTSECTION

AFT SHROUDLight 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 they
are stowed. Near the SA latches are scuff plates,
large protective metal plates on struts that
extend approximately 30 in. from the surface ofK70110-504

Fig. 5-4 Structural components of Support Systems
Module5-3the spacecraft. Trunnions lock the Telescope into
the Shuttle cargo bay by hooking to latches inthe bay. The light shield supports the forward
Low Gain Antenna (LGA) and its communications waveguide, two magnetometers, and two
sun sensors. Handrails encircle the light shield,
and built-in foot restraints support the astronauts working on the Telescope.

next to the HGA drives, where the Shuttle’s
Remote Manipulator System can attach to theTelescope. The forward shell also has
handholds, footholds, and a trunnion, which isused to lock the Telescope into the Shuttle cargobay.
The forward shell is 13 ft (4 m) long and 10 ft(3 m) in diameter. It is machined from aluminum 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.

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, corrugatedskin barrel covered by a thermal blanket. Internally the shield has 10 light baffles, painted flat
black to suppress stray light.

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

Forward Shell. The forward shell, or central
section of the structure, houses the OTA main
baffle 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 fixtures

The Equipment Section is a doughnut-shapedbarrel that fits between the forward shell and aftshroud. This section contains 10 bays for equipment and two bays to support aft trunnion pins
and scuff plates. As shown in Fig. 5-7, goingclockwise from the +V3 (top) position, the bays
contain:

APERTURE
DOOR
LIGHT SHIELD
APERTURE DOORHINGE SUPPORTINTEGRALLYSTIFFENED SKIN
HIGH GAINANTENNA LATCHESHIGH GAINANTENNA MAST

APERTURE
DOOR HINGEFORWARD SHELL
REINFORCINGRINGS

MAGNESIUM
MONOCOQUESKINSTOWEDHIGH GAINANTENNASMAGNETICTORQUERSCREW AIDSINTERNAL
BAFFLE RINGS

LIGHT SHIELDASSEMBLY/
FORWARD SHELL
ATTACH RINGSTOWEDSOLAR ARRAYSSOLAR ARRAYLATCHESINTEGRALLYSTIFFENED
SKIN PANELSHIGH-GAIN ANTENNAAND SOLAR ARRAYLATCH SUPPORT RING
K70110-505

Fig. 5-5 Aperture door and light shield

EQUIPMENT SECTION/FORWARD SHELL
INTERFACE RING
K70110-506

Fig. 5-6 Support Systems Module forward shell
5-4

+V2

5

6

4

3

7

–V3

+V3

2

8

1

9

–V2

10

5-5
5

4

3

2

1

10

9

8

7

6

K70110-507

.

1.
2.
3.
4.
5.
6.

Bay 8 – pointing control hardware
Bay 9 – Reaction Wheel Assembly (RWA)
Bay 10 – SI C&DH unit
Unnumbered trunnion support bay
Bay 1 – data management hardware
Bay 2 through Bay 4 – electrical power
equipment
7. Unnumbered trunnion support bay
8. Bay 5 – communication hardware
9. Bay 6 – RWA
10. Bay 7 – mechanism control hardware.

RADIAL SCIENCE INSTRUMENT DOORSAXIAL SCIENCE INSTRUMENT DOORS+V3+V2

EQUIPMENTSECTION/
AFT SHROUDINTERFACE
RING

+V1

REINFORCING RINGSCREW AIDSINTEGRALLY STIFFENEDSKIN PANELSLOW GAIN ANTENNAELECTRICAl
UMBILICALSThe cross section of the bays is shaped like a
trapezoid, 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) wide and5 ft (1.5 m) deep. The Equipment Section is constructed of machined and stiffened aluminum
frame panels attached to an inner aluminum
barrel. Eight bays have flat honeycombed aluminum doors mounted with equipment. In Bays
6 and 9, thermal-stiffened panel doors cover thereaction wheels. A forward frame panel and aftbulkhead enclose the SSM Equipment Section.
Six mounts on the inside of the bulkhead holdthe OTA.

HONEYCOMB LAMINATE PANELSVENTSFLIGHT SUPPORT SYSTEMPIN SUPPORT BEAM
FLIGHT SUPPORT SYSTEM PIN
K70110-508

Fig. 5-8 Support Systems Module aft shroud
and bulkhead

instruments. The shroud is made of aluminum,with a stiffened skin, internal panels and reinforcing rings, and 16 external and internal longeron bars for support. It is 11.5 ft (3.5 m) longand 14 ft (4.3 m) in diameter.
The aft bulkhead contains the umbilical connections between the Telescope and the shuttle,
used during launch/deployment and on-orbit
maintenance. The rear LGA attaches to the bulkhead, which is made of 2-in.-thick honeycombed aluminum panels and has three radialaluminum support beams.

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

The shroud and bulkhead support a gas purge
system used to prevent contamination of the science instruments before launch. All vents used
to expel gases are light-tight; that is, no straylight can enter the OTA focal plane.

The three FGSs and the Wide Field/Planetary
Camera II (WF/PC) are housed radially near theconnecting point between the aft shroud andSSM Equipment Section. Doors on the outsideof the shroud allow shuttle astronauts toremove and change equipment and instruments
easily. Handrails and foot restraints for the crew
run along the length and circumference of the
shroud. During maintenance or removal of an
instrument, interior lights illuminate thecompartments containing the science

Mechanisms. Along the SSM structure aremechanisms that perform various functions,
including:•Latches to hold antennas and SAs•Hinge drives to open the aperture door anderect the arrays and antennas
5-6•Gimbals to move the HGA dishes•Motors to power the hinges and latches andto rotate the arrays and antennas.
There are nine latches: four for the antennas,
four for the arrays, and one for the aperture
door. They latch and release using four-bar linkages and are driven by stepper motors called
Rotary Drive Actuators (RDA).
There are three hinge drives, one for each HGA
and one for the door. The hinges also use an
RDA. Both hinges and latches have hex-wrench
fittings so an astronaut can manually operate
the mechanism to deploy the door, antenna, or
array if a motor fails.

Fig. 5-9 High Gain Antenna

5.1.2 Instrumentation andCommunications Subsystem

Each antenna can be aimed with a 1-degree
pointing accuracy. This accuracy is consistent
with the overall antenna beam width of over4 degrees. The antennas transmit over two frequencies: 2255.5 MHz or 2287.5 MHz (plus orminus 10 MHz).

This subsystem provides the communications
loop between the Telescope and the Tracking
and Data Relay Satellites (TDRS), receiving
commands and sending data through the HGAs
and LGAs. All information is passed throughthe Data Management Subsystem (DMS).

Low Gain Antennas. The LGAs receive groundcommands and transmit engineering data. They
are on the light shield and aft bulkhead of thespacecraft, set 180 degrees apart. Each antenna
is a spiral cone that can operate over a frequency range from 2100 MHz to 2300 MHz.
Manufactured by Lockheed Martin, the LGAs
are used for all commanding of the Telescope
and for low-data-rate telemetry, particularly
during Telescope deployment or retrieval on
orbit or during safemode operations.

The HGAs achieve a much higher RF signal
gain, 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 spherical
coverage (omnidirectional) but have a much
lower signal gain. The LGAs are used for
low-rate-data transmission and all commandingof the Telescope.

5.1.3 Data Management
Subsystem

High Gain Antennas. Each HGA is a parabolicreflector (dish) mounted on a mast with a twoaxis gimbal mechanism and electronics to rotate
it 100 degrees in either direction (see Fig. 5-9).
General Electric designed and made the antenna
dishes. They are manufactured from honeycomb aluminum and graphite-epoxy facesheets.

The DMS receives communications commands
from the STOCC and data from the SSM systems, OTA, and science instruments. It processes, stores, and sends the information asrequested. Subsystem components are:5-7• DF-224 computer

DF-224 Computer. The DF-224 computer is a
general-purpose digital computer for onboard
engineering computations. It executes storedcommands; formats status data (telemetry); performs all Pointing Control Subsystem (PCS)
computations to maneuver, point, and attitude
stabilize the Telescope; generates onboard commands to orient the SAs toward the sun; evaluates the health status of the Telescope systems;
and commands the HGAs.

• Data Management Unit (DMU)• Four Data Interface Units (DIU)• Three engineering/science tape recorders(E/STR)• Two oscillators (clocks).
The components are located in the SSM Equipment Section, except for one DIU stored in the
OTA Equipment Section.
The DMS receives, processes, and transmits five
types of signals:
1.
2.
3.
4.
5.

The DF-224 is configured with three central processing units (CPU), two of which are backups;
six memory units, with up to 48K words total;
three input/output units, with two as backups;
and six power converter units assigned withoverlapping functions for reliability. The DF-224
measures 1.5 x 1.5 x 1 ft (0.4 x 0.4 x 0.3 m) andweighs 110 lb (50 kg). It is located in Bay 1 of
the SSM Equipment Section (see Fig. 5-11).

Ground commands sent to the HST systems
Onboard computer-generated or computerstored commands
Scientific data from the SI C&DH unit
Telescope engineering status data fortelemetry
System outputs, such as clock signals andsafemode signals.

During the First Servicing Mission, the DF-224
was augmented with an additional computer

Figure 5-10 is the subsystem functional diagram.

Fig. 5-10 Data Management Subsystem functional block diagram
5-8

Data Interface Unit. The four DIUs provide acommand and data link between DMS andother Telescope electronic boxes. The DIUs
receive commands and data requests from the
DMU and pass data or status information backto the DMU. The OTA DIU is located in the OTA
Equipment Section; the other units are in Bays
3, 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
18 cm) and weighs 35 lb (16 kg).
Engineering/Science Tape Recorders. The DMS
includes three tape recorders that store engineering or science data that cannot be transmitted to the ground in real time. The recorders,which are located in Equipment Section Bays 5
and 8, hold up to 1 billion bits of information.
Two recorders are used in normal operations;the third is a backup. Each recorder measures
12 x 9 x 7 in. (30 x 23 x 18 cm) and weighs 20 lb(9 kg).

Fig. 5-11 DF-224 computer

called a coprocessor. Its design was based on theIntel 80386 microchip. The coprocessor provided eight new memory units for the DF-224(with 1 MB of RAM) and a substantial increasein processing power.
Data Management Unit. The DMU links with
the DF-224. It encodes data and sends messages
to selected Telescope units and all DMS units,
powers the oscillators, and is the central timing
source. The DMU also receives and decodes all
incoming commands, then transmits eachprocessed command to be executed.
The DMU is an assembly of printed-circuitboards, interconnected through a backplate andexternal connectors. The unit weighs 83 lb(37.7 kg), measures 26 x 30 x 7 in. (60 x 70 x
17 cm), and is attached to the door of Equipment Section Bay 1 (see Fig. 5-12).
The DMU receives science data from the SI
C&DH unit. Engineering data, consisting of
sensor and hardware status readings (such astemperature or voltages), comes from each Telescope subsystem. The data can be stored in the
onboard E/STRs if direct telemetry via a TDRS
is unavailable.

Fig. 5-12 Data Management Unit configuration
5-9

Solid State Recorder. A state-of-the-art Solid
State Recorder (SSR) will be installed during theSecond Servicing Mission in February 1997. Thisnew type of digital recorder will replace one of
the current, reel-to-reel recorders on HST. The
SSR was developed at Goddard Space Flight
Center, Greenbelt, MD. This new recorder has
an expected on-orbit life of at least eight years
and will add to the efficiency of operations on
HST.

in San Francisco without the beam straying
from the coin’s diameter.

Unlike the 1970s-style reel-to-reel recorders, theSSR has no reels, no tape, and no moving parts
to wear out and limit lifetime. Data is digitally
stored in computer-like memory chips until
HST’s operators at GSFC command the SSR toplay it back.

The PCS encompasses the DF-224 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.

Although it is about the same size as a reel-toreel recorder, it can store over 10 times moredata. The SSR stores 12 gigabits of data, while
the tape recorder it replaces can store only
1.2 gigabits.
Oscillator. The oscillator provides a highly
stable central timing pulse required by theTelescope. It has a cylindrical housing 4 in.
(10 cm) in diameter and 9 in. (23 cm) long andweighs 3 lb (1.4 kg). The oscillator and a backup
are mounted in Bay 2 of the SSM Equipment
Section.

5.1.4 Pointing Control Subsystem
A unique PCS maintains Telescope pointing
stability and aligns the spacecraft to point toand remain locked on any target. The PCS isdesigned for pointing to within 0.01 arcsec andholding the Telescope in that orientation with
0.007-arcsec stability for up to 24 hours while
the Telescope continues to orbit the Earth at
17,000 mph. If the Telescope were in Los
Angeles, it could hold a beam of light on a dime

Nominally, the PCS maintains the Telescope’s
precision attitude by locating guide stars into
two FGSs and controlling the Telescope to keep
it in the same position relative to these stars.
When specific target requests require repositioning the spacecraft, the pointing system selects
different reference guide stars and moves theTelescope into a new attitude.

Sensors. The five types of sensors used by the
PCS are the Coarse Sun Sensors (CSS), theMagnetic Sensing System (MSS), the Rate Gyro
Assemblies (RGA), the Fixed Head Star Trackers(FHST), and the FGSs.
The CSSs measure the Telescope’s orientation to
the sun. They are used to calculate the initial
deployment orientation of the Telescope, determine when to begin closing the aperture door,
and point the Telescope in special sunorientation modes during contingency operations. Five CSSs are located on the light shield
and aft shroud. CSSs also provide signals to the
PSEA, located in Bay 8 of the SSM Equipment
Section.
The MSS measures the Telescope’s orientation
relative to Earth’s magnetic field. The system
consists of magnetometers and dedicated electronic units that send data to the DF-224 computer and the Safemode Electronic Assembly.
Two systems are provided. Both are located on
the front end of the light shield.

5-10Fig. 5-13 Location of Pointing Control Subsystem equipment
5-11

Three RGAs are provided on the Telescope.
Each assembly consists of a Rate Sensing Unit
(RSU) and an Electronics Control Unit (ECU).
An RSU contains two rate sensors, each measuring attitude rate motion about its sensitive axis.
This output is processed by its dedicated electronics, which are contained in the ECU. Each
unit has two sets of electronics. The RSUs arelocated behind the SSM Equipment Section,next to the FHSTs in the aft shroud. The ECUs
are located inside Bay 10 of the SSM Equipment
Section. The RGAs provide input to the PCS tocontrol the orientation of the Telescope’s line ofsight and to provide the attitude reference when
maneuvering the Telescope.
Three of the original six rate gyros were replaced during the First Servicing Mission. Originally, three gyroscopes were required to continue the Telescope mission; however, softwarethat permits science operations to continue with
only two gyroscopes operating has been developed and put into service.
An FHST is an electro-optical detector thatlocates and tracks a specific star within its FOV.
Three FHSTs are located in the aft shroud behind the FPS, next to the RSUs. STOCC uses star
trackers as an attitude calibration device when
the Telescope maneuvers into its initial orientation. The trackers also calculate attitude information before and after maneuvers to help the
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,
pinpoint the guide stars. Two of the FGSs perform guide-star pointing, while the third isavailable for astrometry, the positional measurement of specific stars.

Pointing Control Subsystem Software. PCS
software accounts for approximately 85 percentof the flight code executed by the DF-224 computer. This software translates ground targeting
commands into reaction wheel torque profiles
that reorient the spacecraft. All motion of thespacecraft is smoothed to minimize jitter during data collection. The software also determines Telescope orientation, or attitude, fromFHST or FGS data and commands the magnetic
torquer bars so that reaction wheel speeds arealways minimized. In addition, the software
provides various telemetry formats.
Since the Telescope was launched, major
modifications have been made to the PCS. A
digital filtering scheme, known as Solar Array
Gain Augmentation (SAGA), was incorporated
to mitigate the effect of any SA vibration or jitteron pointing stability. Software also was used to
improve FGS performance when the Telescopeis subjected to the same disturbances. Thisalgorithm is referred to as the FGS Re-Centering
Algorithm.
Software was used extensively to increase Telescope robustness when hardware failures areexperienced. Two additional software
safemodes have been provided. The spinstabilized mode provides pointing of the Telescope -V1 axis to the sun with only two of thefour RWAs operating. The other mode allows
sun pointing of the Telescope without any input
from the RGA; magnetometer and CSS data isused to derive all reference information needed
to maintain sun pointing (+V3 and -V1 are
options).
A further software change “refreshes” the FGS
configuration. This is achieved by maintaining
data in the DF-224 plated wire memory so it canbe sent periodically to the FGS electronics,
which are subject to single-event upsets (logic

5-12state change) when transitioning through theSouth Atlantic Anomaly.
A

RW

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 large
flywheel up to 3000 rpm or braking it toexchange momentum with the spacecraft. The
wheel axes are oriented so that the Telescope
can provide science with only three wheels
operating. Wheel assemblies are paired, twoeach in Bays 6 and 9 of the SSM Equipment Section. Each wheel is 23 in. (59 cm) in diameter
and weighs about 100 lb (45 kg). Figure 5-14
shows the RWA configuration.
Magnetic torquers create torque on the spacecraft and are primarily used to manage reaction
wheel speed. The torquers react against Earth’s
magnetic field. The torque reaction occurs in the
direction that reduces the reaction wheel speed,
managing the angular momentum.
The magnetic torquers also provide backup
control to stabilize the Telescope’s orbital attitude during the contingency modes, as
described in para 5.1.2. Each torquer, locatedexternally on the forward shell of the SSM, is
8.3 ft (2.5 m) long and 3 in. (8 cm) in circumference and weighs 100 lb (45 kg).
Pointing Control Operation. To point precisely,
the PCS uses the gyroscopes, reaction wheels,
magnetic torquers, star trackers, and FGSs. The
FGSs provide the precision reference point fromwhich the Telescope can begin repositioning.
Flight software commands the reaction wheels
to spin, accelerating or decelerating as required
to rotate the Telescope toward a new target. Rate

RW

A

K70110-514

Fig. 5-14 Reaction Wheel Assembly

gyroscopes sense the Telescope’s angularmotion and provide a short-term attitude reference to assist fine pointing and spacecraft
maneuvers. The magnetic torquers reduce reaction wheel speed.
As the Telescope nears the target area, star
trackers locate preselected reference stars thatstand out brightly in that region of the sky. Oncethe star trackers reduce the attitude error below
60 arcsec, the two FGSs take over the pointing
duties. Working with the gyroscopes, the FGSs
make possible pointing the Telescope to within
0.01 arcsec of the target. The PCS can maintainthis position, wavering no more than
0.005 arcsec, for up to 24 hours to guarantee
faint-object observation.

5.1.5 Electrical Power Subsystem
Power for the Telescope and science instruments
comes from the Electrical Power Subsystem(EPS). The major components are two SA wings
and their electronics, six batteries, six Charge
Current Controllers (CCC), one Power Control
Unit (PCU), and four Power Distribution Units(PDU). All except the SAs are located in the bays
around the SSM Equipment Section.
During the servicing mission, the Shuttle will
provide the electrical power. After deployment,
the SAs are extended and begin converting solar
radiation into electricity. Energy will be storedin nickel-hydrogen (NiH2) batteries and distributed by the PCUs and PDUs to all Telescope

5-13

components as shown in Fig. 5-15. The Telescope will not be released until the batteries are
fully charged.
Solar Arrays. The SA panels, discussed later inthis section, are the primary source of electrical
power. Each array wing has a solar cell blanket
that converts solar energy into electrical energy.
Electricity produced by the solar cells charges
the Telescope batteries.
Each array wing has associated electronics.
These consist of a Solar Array Drive Electronics (SADE) unit, which transmits positioning
commands to the wing assembly; a Deployment
Control Electronics Unit, which controls thedrive motors extending and retracting thewings; and diode networks to direct the electrical current flow.
Batteries and Charge Current Controllers.
Developed for the 1990 deployment mission, theTelescope’s batteries were NASA’s first flight
NiH2 batteries. They provide the observatory

with a robust, long-life electrical energy storage
system.
Six NiH 2 batteries support the Telescope’s
electrical power needs during three periods:
when demand exceeds SA capability, when theTelescope is in Earth’s shadow, and during
safemode entry. The batteries reside in SSM
Equipment Section Bays 2 and 3. These units
have extensive safety and handling provisions
to protect the Shuttle and its astronauts. The
design and operation of these batteries, alongwith special nondestructive inspection of eachcell, have allowed these units to be “astronaut
rated” for replacement during a servicing
mission.
Each battery consists of 22 cells in series alongwith heaters, heater controllers, pressure measurement transducers and electronics, and temperature-measuring devices and their associatedelectronics. Three batteries are packaged into a
module measuring roughly 36 by 36 by 10 in.
(90 x 90 x 25 cm) and weighing about 475 lb

Fig. 5-15 Electrical Power Subsystem functional block diagram
5-14

(214 kg). Each module is equipped with twolarge yellow handles that astronauts use tomaneuver the module in and out of theTelescope in space.
The SAs recharge the batteries every orbit
following eclipse (the time in the Earth’s
shadow). The recharge current is controlled by
the CCCs. Each battery has its own CCC thatuses voltage-temperature measurements to control battery recharge.
Fully charged, each battery contains more than75 amp-hours. This is enough energy to sustain
the Telescope in normal science operations
mode for 7.5 hours or five orbits. The batteries
provide an adequate energy reserve for all possible safemode contingencies and all enhancements programmed into the Telescope sincelaunch.
Power Control and Distribution Units. The
PCU interconnects and switches current flowing among the SAs, batteries, and CCCs.
Located in Bay 4 of the Equipment Section, thePCU provides the main power bus to the four
PDUs. The PCU weighs 120 lb (55 kg) andmeasures 43 x 12 x 8 in. (109 x 30 x 20 cm).
Four PDUs, located on the inside of the door to
Bay 4, contain the power buses, switches, fuses,
and monitoring devices for electrical power distribution to the rest of the Telescope. Two buses
are dedicated to the OTA, science instruments,
and SI C&DH; two supply the SSM. Each PDU
measures 10 x 5 x 18 in. (25 x 12.5 x 45 cm) andweighs 25 lb (11 kg).

5.1.6 Thermal Control
Multilayer insulation (MLI) covers 80 percent of
the Telescope’s exterior, and supplemental electric 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 silverized flexible
reflector tape covers most of the remaining
exterior. These coverings protect against thecold of space 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 Section
and structures interfacing with the OTA andscience instruments. The TCS maintains safe
component 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 SSM
include:• MLI thermal blankets for the light shield andforward shell• Aluminum FOSR tape on the aperture door
surface facing the sun• Specific patterns of FOSR and MLI blankets on
the exteriors of the Equipment Section bay
doors, with internal MLI blankets on the bulkheads to maintain thermal balance between
bays• Efficient placement of equipment and use of
equipment bay space to match temperature
requirements, such as placing heat-dissipating
equipment on the side of the Equipment Section mostly exposed to orbit shadow• Silverized FOSR tape on the aft shroud and aftbulkhead exteriors• Radiation shields inside the aft shroud doors
and MLI blankets on the aft bulkhead andshroud interiors to protect the science
instruments

5-15

• More than 200 temperature sensors andthermistors placed throughout the SSM, externally and internally, to monitor individualcomponents and control heater operations.
Figure 5-16 shows the location and type of thermal protection used on the SSM.

5.1.7 Safing (Contingency)
SystemOverlapping or redundant Telescope equipment
safeguards against any breakdown. Nonetheless, a contingency or Safing System exists for
emergency operations. It uses many pointing
control and data management components aswell as dedicated PSEA hardware. This system
maintains stable Telescope attitude, moves theSAs for maximum sun exposure, and conserves
electrical power by minimizing power drain.
The Safing System can operate the spacecraftindefinitely with no communications link toground control.
During scientific observations (normal mode),
the Safing System is relegated to monitor automatically Telescope onboard functions. The
system sends DF-224-generated “keep-alive”
signals that indicate all Telescope systems arefunctioning to the PSEA. Entry into the
Safemode is autonomous once a failure is
detected.
The Safing System is designed to follow a
progression of contingency operating modes,
depending on the situation aboard the Telescope. If a malfunction occurs and does notthreaten the Telescope’s survival, the Safing System moves into a Software Inertial Hold Mode.
This mode holds the Telescope in the last position commanded. If a maneuver is in progress,
the Safing System completes the maneuver, thenholds the Telescope in that position, suspending all science operations. Only ground control
can return to science operations from Safemode.
If the system detects a marginal electrical power
problem, or if an internal PCS safety check fails,
the Telescope enters the Software Sun Point
Mode. The Safing System maneuvers the Telescope so the SAs point toward the sun to continuously generate solar power. Telescope
equipment is maintained within operating temperatures and above survival temperatures,
anticipating a return to normal operations. The
STOCC must intercede to correct the malfunction before any science operations or normal
functions can be resumed.

Fig. 5-16 Placement of thermal protection on Support Systems Module
5-16Since deployment of the Telescope in April 1990,
the Safing System has seen additional improvements to increase its robustness to survive hardware failures and still protect the Telescope.
These additional system features are describedin para 5.1.4.
For the modes described above, the Safing
System operates through computer software. If
conditions worsen, the system turns over control to the PSEA in Hardware Sun Point Mode.
Problems that could provoke this action include
any of the following:• Computer malfunction• Batteries losing more than 50 percent of theircharge• Two of the three RGAs failing• DMS failing.
If these conditions occur, the DF-224 stops
sending keep-alive signals. This is the “handshake” mechanism between the flight softwareand the PSEA.
In the Hardware Sun Point Mode, the PSEA
computer commands the Telescope and turns
off selected equipment to conserve power. Components shut down include the DF-224 computer and, within two hours, the SI C&DH.
Before this, a payload (instruments) safing
sequence begins and, if it has not already doneso, the Telescope turns the SAs toward the sun,guided by the CSSs. The PSEA removes operating power from equipment not required forTelescope survival.
Once ground control is alerted to a problem,
NASA management of the STOCC convenes afailure analysis team to evaluate the problem
and seek the best and safest corrective action
while the Safing System maintains control of theTelescope.

The failure analysis team is led by a senior
management representative from NASA/GSFC
with the authority not only to call upon theexpertise of engineers and scientists employedby NASA or its support contractors, but also todraft support from any organization previously
affiliated with the Telescope Project. The failure
analysis team is chartered to identify the natureof the anomaly and to recommend corrective
action. This recommendation is reviewed at ahigher management level of NASA/GSFC. Allchanges to the Telescope’s hardware and allsoftware configurations require NASA Level I
concurrence as specified in the HST Level I
Operations Requirements Document.
Pointing/Safemode Electronics and Retrieval
Mode Gyro Assemblies. The PSEA consists of
40 electronic printed-board circuits withredundant functions to run the Telescope, evenin the case of internal circuit failure. It weighs
86 lb (39 kg) and is installed in the Equipment
Section Bay 8. A backup gyroscope package, theRMGA, is dedicated for the PSEA and is also
located in Bay 8. The RMGA consists of threegyroscopes. These are lower quality rate sensors
than the RGAs because they are not intended
for use during observations.

5.2

Optical Telescope Assembly

The OTA was designed and built by the PerkinElmer Corporation (now Hughes Danbury
Optical Systems). Although the OTA is modest
in size by ground-based observatory standards
and has a straightforward optical design, its
accuracy – coupled with its place above theEarth’s atmosphere – renders its performancesuperior.
The OTA uses a “folded” design, common tolarge telescopes, which enables a long focal
length of 189 ft (57.6 m) to be packaged into a

5-17

small telescope length of 21 ft (6.4 m). (Several
smaller mirrors in the science instruments aredesigned similarly to lengthen the light pathwithin the particular science instrument.) This
form of telescope is called a Cassegrain, and its
compactness is an essential component of anobservatory designed to fit inside the Shuttlecargo bay.
Conventional in design, the OTA is unconventional in other aspects. Large telescopes atground-based sites are limited in their performance by the resolution attainable while operating under the Earth’s atmosphere, but theTelescope orbits high above the atmosphere andprovides an unobstructed view of the universe.
For this reason the OTA was designed and builtwith exacting tolerances to provide near-perfectimage quality over the broadest possible regionof the spectrum.
The OTA is a variant of the Cassegrain, called a
Ritchey-Chretien, in which both the mirrors are
hyperboloidal in shape (having a deeper curvature than a parabolic mirror). This form iscompletely corrected for coma (an image observation having a “tail”) and spherical aberrations

to provide an aplanatic system in whichaberrations are correct everywhere in the FOV.
The only residual aberrations are field curvature
and astigmatism. Both of these are zero exactly
in the center of the field and increase toward theedge of the field. These aberrations are easily
corrected within the instrument optics. For
example, in the Faint Object Camera (FOC)
there is a small telescope designed 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 up
the tube to the 12-in. (0.3-m)-diameter convex
secondary mirror. The light is now reflected andconverged through a 23.5-in. (60-cm) hole in the
primary 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

Fig. 5-17 Light path for the main Telescope
5-18

light into the WF/PC. The remaining “science”
field is divided among three axial science instruments, each receiving a quadrant of the circular FOV. Around the outside of the science field,
a “guidance” field is divided among the three
FGSs by their own folding mirrors. Each FGS
receives 60 arcmin2 of field in a 90-degree sector. Figure 5-18 shows instrument/sensor fieldsof view.
The OTA hosts the science instruments and
FGSs in that it maintains the structural support
and optical-image stability required for these
instruments to fulfill their functions (seeFig. 5-19). Components of the OTA are the primary mirror, the secondary mirror, the FPS, andthe OTA Equipment Section. Perkin-Elmer
Corporation designed and built all the opticalassemblies; Lockheed Martin built the OTA
equipment section.

5.2.1 Primary Mirror Assemblyand Spherical Aberration
As the Telescope was put through its paces onorbit in 1990, scientists discovered its primary
mirror had a spherical aberration. The outeredge of the 8-foot (2.4-m) primary mirror was
ground too flat by a width equal to 1/50 the
thickness of a sheet of paper (about 2 microns).
After intensive investigation, the problem wastraced to faulty test equipment used to define
and measure mirror curvature. The optical component of this test equipment was slightly out
of focus and, as a result, had shown the mirror
to be ground correctly. After this discovery, Ball
Aerospace scientists and engineers built theCorrective Optics Space Telescope Axial
Replacement (COSTAR). The COSTAR was
installed during the First Servicing Mission in
December 1993 and brought the Telescope backto its original specifications.

Fig. 5-18 Instrument/sensor field of view
5-19

Fig. 5-19 Optical Telescope Assembly components

The primary mirror assembly consists of themirror supported inside the main ring, which isthe structural backbone of the Telescope, andthe main and central baffles shown in Fig. 5-20.
This assembly provides the structural coupling
to the rest of the spacecraft through a set ofkinematic brackets linking the main ring to theSSM. The assembly also supports the OTA
baffles. Its major parts are:
• Primary mirror• Main ring structure• Reaction plate and actuators• Main and central baffles.

Fig. 5-20 Primary mirror assembly
5-20Primary Mirror. The primary mirror blank is a
product of Corning Glass Works known as
ultralow-expansion (ULE) glass. It was chosen
for its very low-expansion coefficient, which
ensures the Telescope minimum sensitivity totemperature changes. The mirror is of a “sandwich” construction: two lightweight facesheets
separated by a core, or filling, of glass honeycomb ribs in a rectangular grid (see Fig. 5-21).
This construction results in an 1800-lb (818-kg)
mirror instead of an 8000-lb solid-glass mirror.
The mirror blank, 8 ft (2.4 m) in diameter, wasground to shape by Perkin-Elmer (now Hughes
Danbury Optical Systems) in its large optics
fabrication facility. When it was close to its final
hyperboloidal shape, the mirror was transferred
to Perkin-Elmer’s computer-controlled
polishing facility.
After being ground and polished, the mirror
was coated with a reflective layer of aluminum
and a protective layer of magnesium fluoride
only 0.1 and 0.025 micrometer thick, respectively. The fluoride layer protects the aluminumfrom oxidation and enhances reflectance at theimportant hydrogen emission line known as
Lyman-Alpha. The reflective quality of the mirror is better than 70 percent at 1216 angstroms(Lyman-Alpha) in the ultraviolet spectral range
and better than 85 percent for visible light.
The primary mirror is mounted to the main ring
through a set of kinematic linkages. The linkages attach to the mirror by three rods thatpenetrate the glass for axial constraint and by
three pads bonded to the back of the glass forlateral support.
Main Ring. The main ring encircles the primary
mirror; supports the mirror, the main baffle andcentral baffle, and the metering truss; and integrates the elements of the Telescope to theFig. 5-21 Primary mirror construction

spacecraft. 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 (2.9m) (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 main
ring, spanning its diameter. It radiates from acentral ring that supports the central baffle. Its
primary 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 also
supports 24 figure-control actuators attached tothe primary mirror and arranged around thereaction plate in two concentric circles. Thesecan be commanded from the ground, if necessary, to make small corrections to the shape ofthe mirror.

5-21

actuators that control its position andorientation. All are enclosed within the central
hub at the forward end of the truss support.

Fig. 5-22 Main ring and reaction plate

Baffles. The baffles of the OTA prevent straylight from bright objects, such as the sun, moon,and Earth, from reflecting down the Telescopetube to the focal plane. The primary mirror
assembly 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. The
central baffle is 10 ft (3 m) long, conical in shape,
and attached to the reaction plate through a holein the center of the primary mirror. It extends
down the centerline of the Telescope tube. The
baffle interiors are painted flat black tominimize light reflection.

The secondary mirror has a magnification of
10.4X, converting the primary-mirror converging rays from f/2.35 to a focal ratio systemprime focus of f/24 and sending them backtoward the center of the primary mirror, where
they pass through the central baffle to the focal
point. The mirror is a convex hyperboloid 12 in.
(0.3 m) in diameter and made of Zerodur glass
coated with aluminum and magnesium fluoride. Steeply convex, its surface accuracy is evengreater than that of the primary mirror.
Ground command adjusts the actuators to alignthe secondary mirror to provide perfect image
quality. The adjustments are calculated fromdata picked up by the optical control system’s
tiny sensors located in the FGSs.
The principal structural element of theSecondary Mirror Assembly is the metering
truss, a cage with 48 latticed struts attached tothree rings and a central support structure forthe secondary mirror. The truss, 16 ft (4.8 m)

5.2.2 Secondary Mirror Assembly
The Secondary Mirror Assembly cantilevers off
the front face of the main ring and supports thesecondary mirror at exactly the correct positionin front of the primary mirror. This position
must be accurate within 1/10,000 in. whenever
the Telescope is operating. The assembly consists of the mirror subassembly, a light baffle,
and an outer graphite-epoxy metering truss
support structure (see Fig. 5-23).
The Secondary Mirror Assembly contains themirror, mounted on three pairs of alignment
5-22

Fig. 5-23 Secondary mirror assemblylong and 9 ft (2.7 m) in diameter, is a graphite,
fiber-reinforced epoxy structure. Graphite waschosen for its high stiffness, light weight, andability to reduce the structure’s expansivenessto nearly zero. This is vital because the secondary mirror must stay perfectly placed relative tothe 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 Mirror Assembly. The other end has a central hub that houses
the secondary mirror and baffle along the optical axis. Aluminized mylar MLI in the truss
compensates for temperature variations of up to
30 degrees Fahrenheit when the Telescope is in
Earth’s shadow so the primary and secondary
mirrors remain aligned.
The conical secondary mirror subassembly lightbaffle extends almost to the primary mirror. It
reduces the stray bright-object light fromsources outside the Telescope FOV.

5.2.3 Focal Plane Structure
Assembly
The FPS is a large optical bench that aligns theimage focal plane of the Telescope with the science instruments and FGSs and physically supports the science instruments and FGSs. The -V3side of the structure, away from the sun inspace, supports the FHSTs and RSUs (seeFig. 5-24). It also provides facilities for on-orbit
replacement of any instruments and thermal
isolation between instruments.
The structure is 7 ft2 (2.1 m) by 10 ft (3.04 m)long and weighs more than 1200 lb (545.5 kg).
It is made of graphite-epoxy, augmented withmechanical fasteners and metallic joints atstrength-critical locations, because it must haveextreme thermal stability and be stiff,lightweight, and strong. The FPS has metallic

Fig. 5-24 Focal plane structure

mounts and supports for Orbital Replacement
Units (ORU) used during maintenance.
The FPS cantilevers off the rear face of the main
ring, attached at eight flexible points that adjust
to eliminate thermal distortions. The structure
provides a fixed alignment for the FGSs. It has
guiderails and latches at each instrument
mounting location so Shuttle crews can easily
exchange science instruments and otherequipment in orbit.

5.2.4 OTA Equipment Section
The Equipment Section for the OTA is a large
semicircular set of compartments mounted outside the spacecraft on the forward shell of the
SSM (see Fig. 5-25). It contains the OTA Electrical Power and Thermal Control Electronics(EP/TCE) System, Fine Guidance Electronics(FGE), Actuator Control Electronics (ACE),
Optical Control Electronics (OCE), and thefourth DMS DIU. The OTA Equipment Section
has nine bays, seven for equipment storage andtwo for support. All bays have outwardopening doors for easy astronaut access, cablingand connectors for the electronics, and heaters
and insulation for thermal control.

5-23

Fig. 5-25 Optical Telescope Assembly Equipment Section

The EP/TCE System distributes power from the
SSM EPS and the OTA system. Thermostats
regulate mirror temperatures and prevent
mirror distortion from the cold of space. The
electrical and thermal electronics also collectthermal sensor data for transmission to theground.
The three FGE units provide power, commands,
and telemetry to each FGS. The electronics perform computations for the sensor and interface
with the spacecraft pointing system for effective
Telescope line-of-sight pointing and stabilization. There is a guidance electronics assemblyfor each guidance sensor.
The ACE Unit provides the command andtelemetry interface to the 24 actuators attachedto the primary mirror and to the six actuators
attached to the secondary mirror. These electronics select which actuator to move and monitor its response to the command. Positioning
commands go from the ground to the electronics through the DIU.
The OCE Unit controls the optical control
sensors. These white-light interferometers
measure the optical quality of the OTA and send
the data to the ground for analysis. There is one
optical control sensor for each FGS, but all control sensors are run by the OCE Unit.

The DIU is an electronic interface between theother OTA electronics units and the Telescopecommand and telemetry system.

5.3

Fine Guidance Sensor

The three FGSs are located at 90-degreeintervals around the circumference of the focal
plane structure, between the structure frame
and 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 sensor
and a wavefront sensor. The wavefront sensors
are elements of the optical control sensor used
to align and optimize the optical system of theTelescope.
The Telescope’s ability to remain pointing at a
distant target to within 0.005 arcsec for long
periods of time is due largely to the accuracy of
the 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 City
the motion of a landing light on an aircraft
flying over San Francisco.
When two sensors lock on a target, the third
measures the angular position of a star, a
process called astrometry. Sensor astrometricfunctions are discussed in Section 4.

5-24

5.3.1 Fine Guidance Sensor
Composition and Function
Each FGS consists of a large structure housing
a collection of mirrors, lenses, servos to locate
an image, prisms to fine-track the image, beam
splitters, and four photomultiplier tubes, as
shown in Fig. 5-26. The entire mechanism
adjusts to move the Telescope into precise alignment with a target star. Each FGS has a large(60 arcmin) FOV to search for and track stars
and a 5.0 arcsec 2 FOV used by the detector
prisms to pinpoint the star.
The sensors work in pairs to aim the Telescope.
The Guide Star Selection System, developed by
the Science Institute, catalogs and charts guidestars near each observation target to make iteasier to find the target. First, one sensor
searches for a target guide star. After the first
sensor locks onto a guide star, the second sensor locates and locks onto another target guidestar. The guide stars, once designated andlocated, keep the image of the observation target
in the aperture 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 greatest
astigmatic and curvature distortions. The size of
the FGS’s FOV was chosen to heighten the probability of finding an appropriate guide star, evenin the direction of the lowest star populationnear the galactic poles.
An FGS “pickoff” mirror intercepts the incoming stellar image and projects it into the sensor’s
large FOV. Each FGS FOV has 60 arcmin2 available. The guide star of interest can be anywhere
within this field, so the FGS will look anywherein that field to find it. After finding the star, thesensor locks onto it and sends error signals to
the Telescope, telling it how to move to keep thestar image perfectly still.
The FGS can move its line of sight anywhere
within its large FOV using a pair of star selector servos. Each can be thought of as an opticalgimbal: One servo moves in a north-south
direction, the other east and west. They steer thesmall FOV (5 arcsec) of the FGS detectors to any
position in the sensor field. Encoders withineach servo system send back the exact coordinates of the detector field centers at any point.
Because the exact location of a guide star may
be uncertain, the star selector servos also cancause the detector to search the region aroundthe most probable guide star position. It
searches in a spiral pattern, starting at the center and spiraling out until it finds the guide starit seeks. Then the detectors are commanded to
go into fine-track mode and hold the star image
exactly centered in the FOV, while the star
selector servo encoders send information about
the position of the star to the spacecraft PCS.

Fig. 5-26 Cutaway view of Fine Guidance Sensor

The detectors are a pair of interferometers,
called Koester’s prisms, coupled to photomultiplier tubes (see Fig. 5-27). Each detector
5-25

a 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 with
the 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 innovative
design improvement.

5.4
Fig. 5-27 Optical path of Fine Guidance Sensor

operates in one axis, so two detectors areneeded. Operating on the incoming wavefront
from the distant guide star, the interferometerscompare the wave phase at one edge of the Telescope entrance aperture with the phase at the
opposite edge. When the phases are equal, thestar is exactly centered. Any phase differenceshows a pointing error that must be corrected.
Along the optical path from Telescope to detector are additional optical elements that turn or
fold the beam to fit everything inside the FGS
enclosure, and to correct the Telescope’s astigmatism and field curvature. All optical elements
are mounted on a temperature-controlled,graphite-epoxy composite optical bench.

5.3.2 Articulated Mirror System
Analysis of the FGS on-orbit data revealed thatminor misalignments of the optical pupil centering on Koester’s prism interferometer in thepresence of spherical aberration prevented the
FGS from achieving its optimum performance.
During the recertification of FGS (S/N 2001),
fold flat #3 in the radial bay module opticaltrain was mechanized to allow on-orbit alignment of the pupil.
Implementation of this system utilized existing
signals and commands by rerouting them with

Solar Array and Jitter
Problems

From the beginning, in the late 1970s, the SAs –designed by the European Space Agency andbuilt by British Aerospace, Space Systems –have been scheduled for replacement because of
their power loss from radiation exposure inspace. However, as engineers put the Telescopethrough its paces in April 1990, they discovered
two 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 Fahrenheit
during orbit, cause the panels to distort twice
during each orbit. As a temporary fix, engineers
created software that commanded the PCS tocompensate for the jitter automatically. The
problem was remedied during the First
Servicing Mission by the replacement of the oldarrays with new ones that had been modified
to alleviate the jitter problem.

5.4.1 Configuration
The SAs are two large rectangular wings ofretractable solar cell blankets fixed on a twostem 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 SAs
delivered to the Telescope during the First

5-26

Servicing Mission, the spring-loaded roller
assembly was replaced by a series of springs
connecting the spreader bar to the blanket. Thischange eliminated the jitter induced into theTelescope as it passed from eclipse (night) into
sunlight (day) each orbit.

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

The wings are on arms that connect to a drive
assembly on the SSM forward shell at one end
and to the secondary deployment mechanism(blankets and bistems) on the other end. The
total length of the cassette, arm, and drive is
15.7ft (4.8 m) (see Fig. 5-28).

An astronaut can raise the array mast manually
if the drive power fails. Using a wrench fitting
on the deployment drive, the astronaut handcranks the mast after releasing the latches.

Each wing has 10 panels that roll out from thecassette. The panels are made of 2438 solar cells
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 up
tightly when the wings are stowed. Each wing
weighs 17 lb (7.7 kg) and, at full extension, is
40 ft (12.1m) 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.

Once the SA is raised, the secondary deployment mechanism unfurls the wing blankets.
Each wing has a secondary mechanism assembly: a cassette drum to hold solar panels, a cushion to protect the blanket, and motors and subassemblies. The assembly rolls out the blanket,
applies tension evenly so the blankets stretch,and transfers data and power along the wingassembly. The blanket can roll out completely
or part way. The secondary deployment mechanism also has a manual override (see Fig. 5-29).
A SA drive at the base of each mast rotates the
deployed 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 respect
to 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 crew member to jettison the entire SA if necessary.

Fig. 5-28 Solar Array wing detailTwo electronics assemblies (boxes) – the Solar
Array Deployment Electronics and the Solar
Array Drive Electronics – control and monitor
all functions of each SA. They provide the electronic interface to the other Telescope systems
and generate the commands for the primaryand secondary deployment mechanisms andthe SA drive.
5-27

units, two CPU modules, a PCU, two RIUs, andvarious memory, data, and command communications 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 8,192
eighteen-bit words. One embedded software
program (the “executive”) runs the computer. It
moves data, commands, and operation programs (called applications) for individual science instruments in and out of the processing
unit. The application programs monitor andcontrol specific instruments and analyze andmanipulate the collected data.

Fig. 5-29 Fitting for Solar Array manual deployment

5.4.3 Solar Array Configurationfor Second Servicing
MissionThe Solar Array wings will remain deployedduring servicing. This will allow the Telescope’s
batteries to remain fully charged during themission and will not impact servicing activities.

5.5

Science Instrument Control
and Data Handling Unit

The SI C&DH unit keeps all science instrument
systems synchronized. It works with the DMU
to process, format, temporarily store on the taperecorders, or transmit all science and engineering data created by the instruments to theground. Fairchild Camera and Instrument
Corporation and IBM built this unit.

The memory stores operational commands for
execution when the Telescope is not in contact
with the ground. Each memory unit has fiveareas reserved for commands and programs
unique to each science instrument. The computer can be reprogrammed from the groundfor future requests or for working around failed
equipment.

5.5.1 Components
The SI C&DH unit is a collection of electronic
components attached to an ORU tray mounted
on the door of Bay 10 in the SSM Equipment
Section (see Fig. 5-30). Small Remote Interface
Units (RIU), also part of the system, provide theinterface to individual science instruments.
Components of the SI C&DH unit are the NASA
Standard Spacecraft Computer (NSCC-I), twostandard interface circuit boards for the computer, two control units/science data formatter
5-28

Fig. 5-30 Science Instrument Control and Data
Handling unit

Standard Interface Unit. The standard interface
board is the communications bridge between
the computer and the CU/SDF.
Control Unit/Science Data Formatter. The heartof the SI C&DH unit is the CU/SDF. It formats
and sends all commands and data to designateddestinations such as the DMU of the SSM, theNASA computer, and the science instruments.
The unit has a microprocessor for control andformatting functions.
The CU/SDF receives ground commands, data
requests, science and engineering data, and system signals. Two examples of system signals are“time tags,” clock signals that synchronize theentire spacecraft, and “processor interface
tables,” or communications codes. The CU/SDF
transmits commands and requests after
formatting them so that the specific destination
unit can read. For example, ground commands
and SSM commands are transmitted with different formats. Ground commands use 27-bitwords, and SSM commands use 16-bit words.
The formatter translates each command signal
into a common format. The CU/SDF also reformats and sends engineering and science data.
Onboard analysis of the data is an NSSC-I
function.
Power Control Unit. The PCU distributes andswitches power among components of the SI
C&DH unit. It conditions the power required byeach unit. For example: The computer memoryboards, typically need +5 volts, -5 volts, and+12 volts; the CU/SDF, on the other hand,requires +28 volts. The PCU ensures that all
voltage requirements are met.
Remote Interface Units. RIUs transmit
commands, clock and other system signals, andengineering data between the science instruments and the SI C&DH unit. The RIUs do notsend science data. There are six RIUs in theTelescope: five attached to the science instruments and one dedicated to the CU/SDF and
PCUs in the SI C&DH unit. Each RIU can becoupled with up to two expander units.
Communications Buses. The SI C&DH unit
contains data bus lines that pass signals anddata between the unit and the science instruments. Each bus is multiplexed: one line sends
system messages, commands, and engineering
data requests to the module units, and a reply
line transmits requested information andscience data back to the SI C&DH unit. A
coupler 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 instrument
system monitoring (such as timing and system
checks), command processing, and dataprocessing.
System Monitoring. Engineering data tells themonitoring computer whether instrument
systems are functioning. At regular intervals,
varying from every 500 milliseconds to every
40 seconds, the SI C&DH unit scans all monitoring devices for engineering data and passes
data to the NSCC-I or SSM computer. The computers process or store the information. Any
failure indicated by these constant tests could
initiate a “safing hold” situation (see para 2.1.7)– 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 in
the drawing) through the SSM Command DIU(ground commands) or the DIU (SSM commands). The CU/SDF checks and reformats thecommands, which then go either to the RIUs or
to the NSCC-I for storage. “Time-tagged”

5-29

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

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. Many commands actually are onboard stored commands
activated by certain situations. For example,
when the Telescope is positioned for a programmed observation using the Space Telescope
Imaging Spectrograph, that program is activated. The SI C&DH can issue certain requests
to the SSM, such as to execute a limited number of pointing control functions to make small
Telescope maneuvers.
Science Data Processing. Science data can comefrom all science instruments at once. The CU/
SDF transfers incoming data through computer
memory locations called packet buffers. It fills
each buffer in order, switching among them as
the buffers fill and empty. Each data packet goesfrom the buffer to the NSCC-I for further
processing, or directly to the SSM for storage inthe tape recorders or transmission to theground. Data returns to the CU/SDF after computer processing. When transmitting, the CU/
SDF must send a continuous stream of data,
either full packet buffers or empty buffers called
filler packets, to maintain a synchronized linkwith the SSM. Special checking codes (ReedSolomon and pseudorandom noise) can be
added to the data as options. Figure 5-32 shows
the flow of science data in the Telescope.

5.6

Space Support Equipment

The Hubble Space Telescope was designed to be
maintained, repaired, and enhanced while inorbit, extending its life and usefulness. For servicing, the Space Shuttle will capture and position the Telescope vertically in the aft end of the
cargo bay, and the crew will perform

5-30

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

maintenance and replacement tasks. The Space
Support Equipment (SSE) to be used during themission provides a maintenance platform tohold the Telescope, provides electrical supportof the Telescope during servicing, and provides
storage for replacement components known as
ORUs.
The major SSE items used for the Second
Servicing Mission are the Flight Support System(FSS), the ORU Carrier (ORUC), and the Second
Axial Carrier (SAC). Additionally, crew aids andtools will be used during servicing.

5.6.1 Flight Support System
The FSS provides the platform that holds theTelescope during servicing (see Fig. 5-33). The
FSS has been used in different configurations for
the HST First Servicing Mission, the Solar Maximum Repair Mission, and the Upper Atmospheric Research Satellite Deploy Mission. The
FSS consists of two major components: ahorseshoe-shaped cradle and a supporting latch
beam providing a structural and electricalinterface with the Shuttle. A circular ring called
the Berthing and Positioning System (BAPS)
interfaces with the Telescope, allowing it topivot or rotate during the mission.
The BAPS ring is pivoted down locked forliftoff. During the mission, the ring is pivoted
up to a horizontal position. A closed-circuit
television camera mounted to the FSS helps
astronauts guide the Telescope onto the ring.
Three remote-controlled latches on the ring grab
and hold three towel-rack-like pins on the rearof the Telescope. A remote-controlled electricalumbilical connector on the FSS engages the Telescope, providing it with orbiter power throughthe FSS. This power helps relieve the drain on
the Telescope’s batteries during the mission.
Radio communications provide telescope
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 extravehicular
activity. Additionally, the ring can pivot the

5-31

Fig. 5-33 Flight Support System configuration

Telescope to an appropriate attitude for orbiter
reboost.
During the first EVA, crew members will install
the BAPS Support Post (BSP). The BSP provides
an additional linkage to support and isolate theTelescope during EVAs and for any orbital
reboosts. The BSP will remain in position for
landing.
Astronauts remotely control all FSS mechanisms – berthing latches, umbilical connector,
pivoter, BSP lock, rotator, and ring down-lock –5-32

Fig. 5-34 Flight Support System Berthing andPositioning System ring pivoted up with
Telescope berthedfrom the orbiter’s aft flight deck, providing thecrew maximum flexibility. Besides being fully
electrically redundant, each mechanism
contains manual overrides and backups toensure mission success and astronaut safety.

5.6.2 Orbital Replacement Unit
Carrier
An ORUC is used to carry replacements into
orbit and to return replaced units to Earth. The
carrier consists of a Spacelab pallet outfittedwith shelves and protective enclosures to holdthe replacement units (see Fig. 5-35). Items on
the ORUC for the Second Servicing Mission
include a Reaction Wheel Assembly in the

LOPE, two Solid State Recorders, one Engineering/Science Tape Recorder, one Solar Array
Drive Electronics, and associated flight harnesses in the SOPE. The FGS will be transported
in the FGS Scientific Instrument Protective
Enclosure (FSIPE) and the Space Telescope Imaging Spectrograph (STIS) in the Axial Scientific
Instrument Protective Enclosure (ASIPE). The
carrier also will have various aids for the crew
during liftoff and landing. All ORUs and scientific instruments are carried within protective
enclosures to provide them a benign environment throughout the mission. The enclosuresprotect the instruments from contamination andmaintain the temperature of the instruments or
ORUs within tight limits.

Fig. 5-35 Orbital Replacement Unit Carrier
5-33

The instruments are mounted in the enclosuresusing the same manually driven latch systemthat holds instruments in the Telescope. The
ASIPE and FSIPE are mounted to the pallet on
a spring system that reduces the level of vibration the instruments receive, especially duringliftoff and landing.
The other ORUs are carried in two additional
enclosures called the Small and Large ORU Protective Enclosures (SOPE and LOPE). These
enclosures also provide contamination andthermal control, though not to such stringent
requirements as the SIPE. The SOPE and LOPE
contain Transport Modules that are designed tocustom fit each ORU. The transport modules
have foam or Visco-Elastic Material thatsurrounds the ORU and isolates it from launchand landing vibration environments.
During the change-out process, replaced science
instruments are stored temporarily in the
ORUC. A typical changeout begins with anastronaut removing the old instrument from theTelescope and attaching it to a bracket on the
ORUC. The astronaut then removes the new

instrument from its protective enclosure andinstalls it in the Telescope. Finally, the astronaut
places the old instrument in the appropriate
protective enclosure for return to Earth.
The ORUC receives power for its TCS from the
FSS. The carrier also provides temperaturetelemetry data through the FSS for readout in
the Shuttle and on the ground during themission.

5.6.3 Second Axial Carrier
The Second Axial Carrier (SAC) is being flown
on the HST Second Servicing Mission to transport and protect the NICMOS science instrument and a complement of ORUs (see Fig. 5-36).
The SAC structure is based on the Goddard
Pallet Assembly.
The Goddard Pallet Assembly is a multimission
shuttle payload carrier built to support Shuttlebased HST servicing activities. For the HST First
Servicing Mission, the Pallet Assembly wasconfigured as the Solar Array Carrier, with a
Solar Array Support Structure designed totransport HST Solar Arrays.

Fig. 5-36 Second Axial Carrier
5-34

For the Second Servicing Mission, the Pallet
Assembly has been modified and is the basestructure of the SAC. The SASS has been
replaced with an Axial Scientific Protective
Enclosure (ASIPE) to transport onto orbit the
new HST NICMOS science instrument. The
Contingency ORU Protective Enclosure (COPE)
and a new ORU Protective Enclosure, called theMultimission ORU Protective Enclosure(MOPE), are used to house and thermally protect the ORUs, also mounted on the Pallet
Assembly.
The SAC spans the cargo bay and structurally
ties into the Shuttle at five points, four longeron
and one keel trunnion. Active isolators, similar
to automobile shock absorbers, reduce launchand landing vibration loads transmitted to the
NICMOS. Each isolator consists of a spring anda magnetic damper. The magnetic damper converts the mechanical vibration energy from thelaunch event into heat energy, reducing theloading on NICMOS. The SAC incorporates
eight isolators, configured in four sets of two
each. Each isolator is attached to the Pallet
Assembly and also to the ASIPE at its base or
end plate.

aids and tools, including the pistol grip tool,
power ratchet tool and its controller, a DIU
portable handle, and the NICMOS cryo ventline. In addition to the ORUs, the COPE willhouse the MSS covers, an MLI repair kit, spareelectrical harnesses, and a JSC-provided tool
called the PFR Attachment Device.
Within the MOPE and COPE, transportmodules protect the ORUs from launch andlanding vibration loading. These transport modules are designed to custom fit each ORU using
foam to surround the ORU. Transport modules
for each ORU are integrated into the OPEs as amodule unit, simplifying system integration.
The SAC receives power directly from the
Orbiter through the Enhanced Power Distribution and Switching Unit (EPDSU). The EPDSU
is modular avionics for Shuttle payloads
designed to be readily configured for uniquepayload power, command and telemetry
requirements. Custom applications are easily
serviced using the EPDSU. An EPDSU, configured for the FSS, is also available for backup to
the FSS Power Distribution and Switching Unit.

5.6.4 Crew Aids

The SAC ASIPE is based on the ORUC ASIPE,
with design modifications to be flown as astand-alone enclosure. Aside from the unique
NICMOS/FOS latch configuration, all EVA and
instrument interfaces are the same for both
ASIPEs. The ASIPE is thermally controlled using a heater system designed to precondition
the NICMOS prior to EVA change-out
operations.

Astronauts perform extravehicular activities
using many tools to replace instruments andequipment, to move around the Telescope andthe cargo bay, and to operate manual override
drives. Tools and equipment bolts, connectors,
and other hardware were standardized not only
for the Telescope but between the Telescope andthe Shuttle. For example, grappling receptacles
share common features.

ORUs transported on the SAC for the Second
Servicing Mission are two DIUs in the MOPE,
and two RSUs and an ECU in the COPE. The
MOPE also will carry electrical harnesses for the
DIUs, spare fuse plugs and PIP pins, and crew

To move around the Telescope, the crew uses
225 ft of handrails encircling the spacecraft. For
visibility, the rails are painted yellow. In
addition, the crew can hold onto guiderails,
trunnion bars, and scuff plates fore and aft.

5-35

The astronauts can install portable handhold
plates where there are no permanent holds, suchas on the FGS. Another tool is the Portable Foot
Restraint (PFR), shown in Fig. 5-37.

LOCK
YAW CONTROLPITCH CONTROL
A

While the astronauts work, they use tethers tohook tools to their suits and tie replacement
units to the Telescope. Each crew member has
a ratchet wrench to manually crank the antenna
and array masts if power for the mast drives
fails. A power wrench also is available if handcranking is too time consuming. Other handtools include portable lights and a jettisonA

ELEVATIONCONTROL
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Fig. 5-37 Portable Foot Restraint

handle, which attach to sockets on the aperture
door and to SA wings so the crew can push theequipment away from the Telescope.

5-36

GLOSSARY-AÅ

Angstrom

aberration

Property of an optical system that causes an image to have certain easily
recognizable flaws. Aberrations are caused by geometrical factors suchas the shapes of surfaces, their spacing, and alignments. Image problems
caused by factors such as scratches or contamination are not calledaberrations.

ACE

Actuator Control Electronics

acquisition, target

Orienting the HST line of sight to place incoming target light in an
instrument’s aperture

actuator

Small, high-precision, motor-driven device that can adjust the locationand orientation of an optical element in very fine steps, making fine
improvements to the focus of the image

aftRear of the spacecraftalignment

Process of mounting optical elements and adjusting their positions andorientations so that light follows exactly the desired path through the
instrument and each optical element performs its function as planned

altitude

Height in space

AMA

Actuator Mechanism Assembly

AME

Actuator Mechanism Electronics

aperture

Opening that allows light to fall onto an instrument’s optics

aplanatic

Image corrected everywhere in the field of view

apodizer

Masking device that blocks stray lightarcsecA wedge of angle, 1/3600th of one degree, in the 360-degree “pie” thatmakes up the sky. An arcminute is 60 seconds; a degree is 60 minutes.

ASIPE

Axial Scientific Instrument Protective Enclosure

astigmatism

Failure of an optical system, such as a lens or a mirror, to image a point
as a single point

astrometry

Geometrical relations of the celestial bodies and their real and apparent
motions

attitude

Orientation of the spacecraft’s axes relative to Earth
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AURA

Association of Universities for Research in Astronomy

axial science instruments Four instruments – the FOS, FOC, GHRS, and COSTAR – located behindthe primary mirror. Their long dimensions run parallel to the opticalaxis of the HST. The GHRS will be removed and replaced by STIS, and
FOS will be removed to make way for NICMOS.

-Bbaffle

Material that extracts stray light from an incoming image

BAPS

Berthing and Positioning System

BPSBAPS Support Post

-CC

Celsius

Cassegrain

Popular design for large, two-mirror reflecting telescopes in which theprimary mirror has a concave parabolic shape and the secondary mirror
has a convex hyperbolic shape. A hole in the primary allows the image
plane to be located behind the large mirror.

CAT

Crew Aids and Tools

CCC

Charge Current Controller

CCD

Charge-coupled device

CDI

Command data interfacechange-out

Exchanging a unit on the satellitecm

Centimeter

collimate

To straighten or make parallel two light paths

coma

Lens aberration that gives an image a “tail”

concave

Mirror surface that bends outward to expand an image

convexMirror surface that bends inward to concentrate on an image

coronographDevice that allows viewing a light object’s corona

COSTAR

Corrective Optics Space Telescope Axial Replacement

CPM

Central Processor Module
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CPU

Central Processing Unit

CTVC

Color television camera

CU/SDF

Control Unit/Science Data Formatter

CSS

Coarse Sun Sensor-Ddiffraction grating

Device that splits light into a spectrum of the component wavelengths

DIU

Data Interface Unit

DMS

Data Management Subsystem

DMU

Data Management Unit

DOB

Deployable Optical Benchdrag, atmosphericEffect of atmosphere that slows a spacecraft and forces its orbit to decay-EECA

Electronic Control Assembly

ECU

Electronics Control Unit

electron

Small particle of electricity

ellipsoid

Surface whose intersection with every plane is an ellipse (or circle)

EPDSU

Enhanced Power Distribution and Switching Unit

EPSElectrical Power Subsystem

EP/TCE

Electrical Power/Thermal Control Electronics

ESA

European Space Agency

E/STR

engineering/science tape recorders

EVA

extravehicular activity

extravehicularOutside the spacecraft; activity in space conducted by suited astronauts-FF

Fahrenheit

FGE

Fine Guidance Electronics
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FGS

Fine Guidance Sensor

FHST

Fixed Head Star Tracker

FOC

Faint Object Camera

focal plane

Axis or geometric plane where incoming light is focused by thetelescope

FOS

Faint Object Spectrograph

FOSR

Flexible optical solar reflector

FOV

Field of view

FPS

Focal plane structure

FPSA

Focal plane structure assembly

FRB

Fastener retention blockFS

Forward Shell

FSIPE

FGS Scientific Instrument Protective Enclosure

FSS

Flight Support System-GG/E

Graphite-epoxy

GE

General Electric

GGM

Gravity Gradient Mode

GHRS

Goddard High Resolution Spectrograph

GSE

Ground support equipment

GSFC

Goddard Space Flight Center

GSSS

Guide Star Selection System

GSTDN

Ground Spaceflight Tracking and Data Network-HHGA

High Gain Antenna

HRS

High Resolution Spectrograph
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HSP

High Speed Photometer

HST

Hubble Space Telescope

hyperboloidal

Slightly deeper curve, mathematically, than a parabola; shape of the
primary mirror

Hz

Hertz (cycles per second)-IIBM

International Business Machines Corporationin.

Inchesinterstellar

Between celestial objects; often refers to matter in space that is not astar, such as clouds of dust and gas

intravehicular

Inside the spacecraft

IOU

Input/output unit

IR

InfraredIV

Intravehicular

IVA

Intravehicular activity-JJPL

Jet Propulsion Laboratory

JSC

Johnson Space Center-Kk

Kilo (1000)

kB

Kilobytes

kg

Kilogramkm

Kilometer

KSC

Kennedy Space Center-LLatchMechanical device that attaches one component, such as a science
instrument, to the structure of the telescope and holds it in precisely
the right place
G-5

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lb

Pound

LGA

Low Gain Antenna

LGA PC

Low Gain Antenna Protective Cover

Light year

The distance traveled by light in 1 year, approximately 6 trillion miles

LMMS

Lockheed Martin Missiles & Space

LOPE

Large ORU Protective Enclosure

LOS

Line of sight

LS

Light Shieldluminosity

Intensity of a star’s brightness

-Mm

Meter

µm

Micrometer; one millionth of a metermm

Millimeter

MA

Multiple accessmagnitude, absolute

How bright a star appears without any correction made for its distancemagnitude, apparent

How bright a star would appear if it were viewed at a standard distance

MAMA

Multi-Anode Microchannel Plate Array

MAT

Multiple Access Transponder

MCC

Mission Control Center

MCP

Microchannel plate

metrology

Process of making extremely precise measurements of the relative
positions and orientations of the different optical and mechanicalcomponents

MFR

Manipulator Foot RestraintMHz

Megahertz

MLI

Multilayer insulation

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MOPE

Multimission ORU Protective Enclosure

MSFC

Marshall Space Flight Center

MSM

Mode Selection Mechanism

MSS

Magnetic Sensing System

MT

Magnetic torquer

MTA

Metering Truss Assembly

MTS

Metering Truss StructureM

Absolute visual magnitudem

Apparent visual magnitude-NNASA

National Aeronautics and Space Administration

NBSF

Neutral Buoyancy Simulator Facility

NASCOM

NASA Communications Network

NCC

Network Control Center

nebulaMass of luminous interstellar dust and gas, often produced after a stellar
nova

NICMOS

Near Infrared Camera and Multi-Object Spectrometernm

Nanometers

nmi

Nautical miles

nova

Star that suddenly becomes explosively bright

NSSC-I

NASA Standard Spacecraft Computer, Model-I-Ooccultation

Eclipsing one body with another

OCE

Optical Control Electronics

OCE-EKOCE Enhancement Kit

OCS

Optical Control Subsystem
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Orientation

Position in space relative to Earth

ORU

Orbital Replacement Unit

ORUC

Orbital Replacement Unit Carrier

OTA

Optical Telescope Assembly-PPACOR

Packet Processing Facility

parallax

Change in the apparent relative orientations of objects when viewed
from different positions

PCEA

Pointing Control Electronics Assembly

PCS

Pointing Control Subsystem

PCU

Power Control Unit

PDA

Photon Detector Assembly

PDM

Primary Deployment Mechanism

PDU

Power Distribution Unit

PFR

Portable Foot Restraint

photonUnit of electromagnetic energy

PIPpush in-pull out (pin)pixelSingle element of a detection device

POCC

Payload Operations Control Center

polarity

Light magnetized to move along certain planes. Polarimetricobservation studies the light moving along a given plane.

primary mirror

Large mirror in a reflecting telescope the size of which determines thelight-gathering power of the instrument

prism

Device that breaks light into its composite wavelength spectrum

PSEA

Pointing/Safemode Electronics Assembly-Qquasar

Quasi-stellar object of unknown origin or composition
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-RRAM

Random-access memory

radialPerpendicular to a plane (i.e., instruments placed at a 90-degree angle
from the optical axis of the HST)

RBM

Radial Bay Module

RDA

Rotary Drive Actuatorreboost

To boost a satellite back into its original orbit after the orbit has decayed
because of atmospheric dragreflecting telescope

Telescope that uses mirrors to collect and focus incoming lightrefracting telescope

Telescope that uses lenses to collect and focus light

resolution

Ability to discriminate fine detail in data. In an image resolution, it
refers to the ability to distinguish two objects very close together inspace. In a spectrum, it is the ability to measure closely separated
wavelengths.

resolution, spectral

Determines how well closely spaced features in the wavelength
spectrum can be detected

resolution, angular

Determines how clearly an instrument forms an image

RF

Radio frequency

RGA

Rate Gyro Assembly

Ritchey-Chretien

A modern optical design for two mirror reflecting telescopes. It is a
derivative of the Cassegrain concept in which the primary mirror hasa hyperbolic cross-section.

RIU

Remote Interface Unit

RMGA

Retrieval Mode Gyro Assembly

RMS

Remote Manipulator System

ROM

Read-only memory

RS

Reed-Solomon

RSU

Rate Sensing Unit

RWA

Reaction Wheel Assembly
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-SSA

Solar ArraySAASouth Atlantic Anomaly

SAC

Second Axial Carrier

SAD

Solar Array Drive

SADE

Solar Array Drive Electronics

SADM

Solar Array Drive Mechanism

SAGA

Solar Array Gain Augmentation

SBA

Secondary Baffle Assembly

SCP

Stored Command Processor

SDAS

Science Data Analysis Software

SDM

Secondary Deployment Mechanism

secondary mirror

In a two-mirror reflecting telescope, the secondary mirror sits in frontof the larger primary mirror and reflects light to the point at which it
will be detected and recorded by an instrument. In simple telescopes,
the secondary mirror is flat and bounces the light out the side of the
tube to an eyepiece. In more complex and larger telescopes, it is convex
and reflects light through a hole in the primary mirror.

Servicing Mission

NASA’s plan to have the Space Shuttle retrieve the HST and haveastronauts perform repairs and upgrades to equipment in space

SI

Science instrument

SI C&DHSI Control and Data Handling (subsystem)

SIPE

Science Instrument Protective Enclosure

SM

Secondary mirror

SMA

Secondary Mirror Assembly

SOFA

Selectable Optical Filter Assembly

SOGS

Science Operations Ground System

SOPE

Small ORU Protective Enclosure
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spectral devices

These include spectrographs, instruments that photograph the spectrumof light within a wavelength range; spectrometers, which measure theposition of spectral lines; and spectrophotometers, which determine
energy distribution in a spectrum.

spectrograph

Instrument that breaks light up into its constituent wavelengths andallows quantitative measurements of intensity to be made

spectrum

Wavelength range of light in an image

spherical aberration

Image defect caused by a mismatch in the shapes of the reflecting
surfaces of the primary and secondary mirrors. Light from different
annular regions on the primary mirror comes to a focus at different
distances from the secondary mirror, and there is no one position where
all of the light is in focus.

SSC

Science Support Center

SSE

Space Support Equipment

SSM

Support Systems Module

SSM-ES

SSM Equipment Section

SSR

Solid State Recorder

STDN

Space (flight) Tracking and Data Network

STINT

Standard interface

STIS

Space Telescope Imaging Spectrograph

STOCC

Space Telescope Operations Control Center

STS

Space Transportation System

STScI

Space Telescope Science Institute-TTA

Translation Aids

TAG

Two-axis gimbal

TCE

Thermal Control Electronics

TCS

Thermal Control Subsystem

TDRS

Tracking and Data Relay Satellite
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TDRSSTDRS System

TECI

Thermoelectric-cooled inner (shield)

TECO

Thermoelectric-cooled outer (shield)

telemetry

Data and commands sent from the spacecraft to ground stations

TLM

Telemetry-UUDM

Umbilical disconnect mechanism

ULE

Ultralow expansion

UV

Ultraviolet

USA

United States Army

USAF

United States Air Force

USN

United States Navy

-VV

Volt

V1, V2, V3

HST axes

VCS

Vapor-cooled shield

-WW

Watt

Wavelength

Spectral range of light in an image

WFC

Wide Field Camera

WF/PC

Wide Field/Planetary Camera. The camera currently in use is thesecond-generation instrument WF/PC II, installed during the First
Servicing Mission in December 1993. It replaced WF/PC I and was builtwith optics to correct for the spherical aberration of the primary mirror.

G-12
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