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Space Telescope
Copyright
©
SPACECRAFT ATTITUDE CONTROL
IFAC 9th Triennial World Congress
Budapest. Hung;u)'. 19R4
SPACE TELESCOPE F. S. Woj talik Deputy Project Managfr, Space Telescope Projec/ITA01, Marshall Space Flight Cen/fr, AL 35812, USA
Abstract. This paper describes the Space Telescope, a program of the National Aero nautics and Space Administration (NASA) . The Space Telescope is the largest and most powerful optical and ultraviolet astronomical observatory to be operated in space . Through the remote eyes of this telescope, astronomers will look further into space and time to produce data incl uding imagery of unequaled quality of galaxies, star systems, quasars and other objects of scientific interest. The Space Telescope will do this by observing the sky from a nominal 550 kilometer low Earth orbit via an optical system with a 2 . 4 meter diameter primary mirror . Additionally, this observatory has a precision pointing control system that is capable of maintaining a locked state on an object for extended periods accurately to within 0 . 01 arc seconds. The observatory instrument complement, optical system and pointing control system is presented. Keywords . Strapdown systems; space vehicles; satellites; stability; attitude control; observability. INTRODUCTION Viewing the celestial sky in space, outside the Earth's obscuring atmosphere, through a telescope with observatory class features is a project that astronomers have proposed and worked to achieve for many years . First, relatively small telescopes like NASA ' s Orbiting Astronomical Observatory (OAO), International Ultraviolet Explorer (IUE), and later the much larger series of three High Energy Astronomy Observatories (HEAO ' s) were orbited with support from the scientific and industrial communities with a high degree of scientific success. NASA is now past mid point in developing and launching into orbit a much larger and more long lived observatory called Space Telescope. This unmanned observatory, scheduled to become operational in the next few years, will have a sophisticated optical telescope, be capable of accommodating a variety of scientific instruments, and will operate in low Earth orbit for many years. Astronomy is concerned and challenged by fundamental and basic issues; the position, motion, constitution, history and
future of celestial bodies . With the Space Tele scope the astronomers capa~ilities for observation as a means of gathering data to resolve many astronomical concerns will be greatly improved. The Space Telescope is designed to be launched, deployed, serviced, and returned to Earth by means of the Space Shuttle Transportation System. After launch from the Kennedy Space Center, the Space Shuttle Orbiter will maneuver to attain the proper orbit and attitude, then the payload bay doors will be opened to allow the 15 meter mechanical arm of the Remote Manipulator System (RMS) to remove the Space Telescope from the payload bay. While the Space Telescope is connected to the RMS or immediately after being released into space, solar arrays and high gain antenna deployment will occur and systems functional checks will be initiated to verify that the observatory is functioning properly. At this point, the Orbiter will return to Earth and the Space Telescope mission will begin. For on orbit maintenance visits the Orbiter will maneuver into rendezvous position, the RMS will engage the free flying observatory and transfer it (with appendages retracted) into the payload bay where maintenance will proceed.
Because the observatory initial orbit altitude will gradually decay due to atmospheric drag, the Orbiter will reboost the Observatory to its initial orbit altitude during a maintenance visit. The capability to return the observatory to Earth for extensive overhaul is possible but every effort will be made to achieve repair/replacement on orbit for the fifteen or more years of expected life. SPACE TELESCOPE SYSTEM The final spacecraft and optical system design of the Space Telescope is the result of scientific objectives and requirements merged with various space oriented system operational constraints (Fig. 1) . As an example, the Space Shuttle Orbiter payload bay imposes a final limitation on the payload overall length and diameter. The envelope of the telescope once basically established, the number and types of scientific instruments planned, and the pointing accuracy and stability required impose other real limitations on the final design. The Space Telescope is a long-lifetime general purpose telescope that is by design capable of operating with a wide variety of scientific instruments at its focal plane. This observatory is a cylindrical structure with a length of 13.1 meters, a diameter of 4.27 meters that weighs about 11,000 kilograms. Two large solar arrays that can be deployed and retracted on command are capable of rotation about the observatory pitch axis (V2) to maintain a normal attitude with the sunline . The two solar arrays (wings) each measure 2.3 by 11 . 8 meters and are made of two solar cell blankets/wing. Each solar cell blanket consists of five solar panel assemblies for a total of 20 panels for the observatory. Each solar panel assembly contains 2438 cells giving a grand total for the observatory of 48,760 solar cells. The two solar arrays, six nickel-cadmium batteries, and power conditioning equipment will supply a minimum of 2 . 4 KW of orbital average electrical power after two years of operation. S The major configuration elements of the Space Telescope are the Optical Telescope Assembly (OTA), a Support Systems Module (SSM), the Scientific
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Fig. 1.
Space Telescope Configuration
Instruments (SI's) end the Solar Arrays (SA's) previously described. The OTA contains a 2.4 meter reflecting Cassegrain-type telescope housed within a meteoroid shield and sunshade. The telescope itself is a Ritchey-Chretien folded optical system with the secondary mirror located inside the prime focus. The primary mirror is fabricated of ultra-low expansion titanium silicate glass where the front and back plates are fused to a honeycomb core. Primary mirror weight is 829 kg with a clear aperture of 2.4 meters and f24 focal ratio. The mirror has an aluminum magnesium fluoride coating and is heated during operation to about 70°F to minimize variations from its original fabrication accuracy. After the aperture cover is commanded open, any entering light will illuminate the primary mirror at the rear of the telescope. The primary mirror will project the light to the smaller secondary mirror in the front end of the OTA. The light beam in turn will then be reflected back through a hole in the primary mirror to the SI's in the rear at the focal plane. The optical surface of the mirror has been measured and shown to have a figure error of less than 1/60 of a wavelength (rms) at 632.8 nm. Seventy percent of the energy from a star impinges in a radius at the focal plane of less than 25 microns at 632.8 nm, corresponding to about 0.1 arc sec. The reflection efficiency of the aluminized MgF2 coated optical surfaces, as measured, is better than 70% at 121.6 nm and 85% at 632.8 nm. With its complement of instruments the telescope will cover the range of wavelengths from the far ultraviolet to the far infrared. The field of view at the focal plane minus some structural obscuration is a circular area with a diameter of 28 arc min (Fig. 2). The central area of 22 arc min has the best resolution as realized from a Ritchey-Chretien optical design and is, therefore, used by five scientific focal plane instruments and three Fine Guidance Sensors (pseudo scientific instruments) that view three 90 degree segments of the outer portion of the field of view (FOV). The central 2.7 x 2.7 arc min area is diverted by a pick-off mirror that sends the beam
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Space Telescope f/24 Focal Plane
of light to the Wide Field/Planetary Camera instrument. Four quadrants about 6 arc min wide at the central area of the FOV are assigned to four axially oriented instruments. The Support System Module that encloses the OTA and SI's provides interfaces to the Orbiter and ground operations. Ground operations are accomplished via the Tracking and Data Relay Satellites (TDRS). The SSM also provides the essentials for proper and safe operation of the observatory. These include the electrical power and distribution system, communications, attitude and stabilization control, data management and thermal control systems. Observatory attitude control during maneuvers and pointing are the result of speed variations in the four available reaction wheels. The signals for wheel speed and rotation control are derived from a variety of sensors; a coarse sun sensor, fixed head-star trackers, gyroscopes and interferometric star trackers, The functional interactions of these devices is discussed in more detail later.
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Space Telescope SCIENCE INSTRUMENTS Five scientific instruments and one pseudo scientific instrument are planned for the Space Telescope initial mission. This complement of instruments are located at the focal plane behind the primary mirror where they are held in exact position by the focal plane structure through a set of latches. These instruments are the Wide Field/Planetary Camera (WF/PC), Faint Object Spectrograph (FOS) , High Resolution Spectrograph (RRS), High Speed Pho t ometer (HSP) , Faint Object Camera (FOC), and the Fine Guidance Sensors (FGS). As mentioned, the FGS is considered to be a pseudo scientific instrument because of its accuracy in star location. 4 It is used primarily to provide Space Telescope high pointing stability. For this function two sensors (there are three) are sufficient for locating and locking onto a target, therefore, the third, a standby unit, can be used to precisely measure the positions of other stars in the vicinity in relation to the stars on which the other two sensors are locked. These instruments will provide precision pointing accuracy and stability to the observatory, calibrate the positions of nearby and distant stars and galazies, reveal new information on the unseen companions of binary star systems, and provide better positional reference systems for compact stars. Each scientific instrument is a separate module capable of being removed and replaced in space. The Wide Field/Planetary Camera and the three Fine Guidance Sensors are located in radial positions with respect to the optical axis and the other instruments are in bays parallel to the telescope line of sight. The Wide Field/Planetary Camera operates in two modes. It has a wide field (2.63 x 2.63 arc min 2 ) capability at a focal ratio of f/12.9 that allows large areas of the sky to be observed (wide field) and a second field of view (68.7 x 68.7 arc sec 2 ) (Planetary). The detector is designed with four charged coupled detector arrays of 800 x 800 pixels each. The pixel in the wide field mode corresponds to 0.1 arc sec and in the planetary mode to 0.043 arc sec. The first mode provides the largest field of view (FOV) available on the Space Telescope at a reduced resolution while the second mode provides optics limited resolution. Sensitivity in the range of 115.0 to 1100.0 nm with photometric accuracy of 1 percent over a wide dynamic range has been verified by test. Fortyeight special purpose filters, polarizers, and gratings are included in a wheel arrangement to produce color cosmology information. The sensitivity of this instrument has been measured showing that it can achieve high quality point sources information down to 28 Mv after a one hour exposure . The High Resolution Spectrograph is intended for operation in the 110.0 to 320.0 nm region with a resolution capability of 2 x 10 4 . This resolution corresponds to 0.05 Aand will be one of the best available in astronomy. This instrument uses Digicon detectors and concave light dispersing gratings. Spectroscopic resolution, photometric accuracy and sensitivity is expected to be greatly improved over what was provided by previous space instruments including the highly successful International Ultraviolet Explorer (IUE). The Faint Object Spectrograph is a medium resolution, 10 2 to 10 3 , instrument. Digicon arrays are used in the design to detect wavelengths of 115 to 800 nm. This instrument will be capable of 30 A to 3 A detection depending on the commanded mode of operation. Polarization and time variability studies in the spectra of bright sources of the incoming light with resolution of 10 milliseconds will be routinely performed. The Faint Object Camera, developed by the European Space Agency (ESA), will complement the Wide Field/Planatory Camera and uses the full performance capability of the Space Telescope through the attainment of high angular resolution
and sensitivity in a narrow wavelength band of 120 to 500 nm. It operates in two modes which produce images of 11 x 11 arc sec 2 or 22 x 22 arc sec 2 onto vidicon tubes with magnesium fluoride intensifiers. Pixels of 25 microns equate to 0.02 and 0.04 arc sec, respectively, in the two available operational modes. The instrument records individual photons and can achieve very high angular resolution of 0.007 arc sec in very narrow fields thereby permitting post facto image reconstruction methods. As with many of the other Space Telescope scientific instruments, it also has filter wheels, each containing a variety of filters with 12 positions that can be commanded for insertion into the optical path. 1 The Faint Object Camera and other instruments will be used to observe extragalactic supergiant stars, study variable brightness stars, gather data on globular clusters, examine binary star systems, search for extrasolar planets, establish stellar masses, do detailed studies of shock fronts and condensing gas clouds, search for direct evidence that quasars might be at the center of faint galaxies, and other studies and investigations of astronomical interest. 4 POINTING AND STABILIZATION CONTROL The quality of the Space Telescope (ST) optical system, the sensitivity of its complement of scientific instruments and mission operating constraints were the principle drivers that established requirements and design architecture of the pointing and stabilization control system. The selected design has two basic operating attitude control modes; Sun Point Control (SPC) and Operational Vehicle Control (OVC). SPC is used during initial observatory deployment from the Space Shuttle Orbiter when pre-mission equipment checks are being performed and in contingency operations. Entry into the SPC mode can occur autonomously whenever one of several on-board electrical power system failure conditions is sensed or by ground command. In this mode, the observatory +V3 axis and solar array active surfaces are positioned toward the sun and maintained in this orientation under firm attitude control but within loose position limits. Maintaining the solar arrays active surfaces to be within ±S degrees normal to the sun is very adequate. Control is provided by reaction wheels, sun sensors, rate gyros and the digital computer. The computer uses output signals from the sun sensors to establish attitude orientation during orbit sunlight and rate gyro outputs to maintain proper attitude position in eclipse. Gyro-sensed body rotational rates are used to maintain the plane, formed by the V3 and V1 axes within ±S degrees of the sunline. A hardwired set of logic is also provided to accomplish the computer functions, if required in this contingency mode. Entry into the long term and normal Operational Vehicle Control (OVC) mode is accomplished by ground command only. In this mode, solar array orientation toward the sun is maintained by the computer as previously described while rotation of the observatory about the V3 and V2 axes is introduced. The function of the OVC mode is to orient and maintain the Space Telescope optical axis (+V1 axis) to targets (celestial coordinates) of scientific interest for observations lasting several minutes to several days with precision accuracy and stability. To accomplish this, the control mode is separated into two distinct phases. A course positioning phase that uses body-fixed rate integrating gyros arranged in pairs so their inputsensitive axes are in opposite directions and skewed at 31.72 degrees up from their mounting plane, fixed head star trackers that are separate and apart from the telescope optics and fine guidance sensors which derive their signals from stars imaged near the edge of the Space Telescope FOV
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F. S. Wojtalik
(Fig. 2). This course positioning system as well as the precision pointing system requires the services of the computer and reaction wheels. ThE precision pointing system utilizes the fine guidance sensors in a fine lock mode to update the stability reference of the set of rate gyros being used every second to a noise equivalent positional error of 0.0028 arc sec (rms) yielding an overall Space Telescope positioning accuracy of 0.01 arc sec with 0.007 arc sec (rms) of stability. The sequence of events in a typical acquisition of a target of interest starts with a maneuver (slew) of the telescope limited to a maximum rate of 15 degrees/min. Using the coordinates of the present known location and the new desired location plus the available reaction wheel momentum, an eigen axes rotation is computed. The observatory is commanded to follow this rotation at a specified rate. At approximately 0.4 degrees away from the maneuver termination, a forward-loop integrator within the computer logic is used to minimize rate over-shoot and other disturbance torques. Once the observatory arrives at the new desired location, pointing reference errors are expected to be less than !30 arc sec (30) and body rates diminished to less than 0.05 arc sec/sec. Since the uncertainties of position due to gyro drift and the inaccuracies (especially roll angle) of the initial direction increase with displacement, slews will be planned to be as short as possible. The uncertainty can be several arc minutes after only a 90 degree slew. During this phase of operation, the fixed head star trackers are used to determine the rough position (!10 arc-sec calibrated accuracy) as necessary. At this time, the coordinates and brightness of a candidate guide star, for the target of interest, is transferred into each of two Fine Guidance Sensors (FGS's). The three on-board FGS's operate with the observatory support system module (SSM) pointing control system (PCS) to provide on-board fine (precision) pointing control and stabilization. Two FGS's are used during each pointing and hold operation while the third unit in the complement is in standby and can perform astrometry as previously described. Guide star positional information is provided by two FGS modes; (1) a coarse measurement mode accurate to !0.1 arc-sec (rms), at an update rate of 0.025 sec per axis, and (2) a fine (precision) measurement of 0.0028 arc-sec (rms) which is the error allocated to the FGS for image stability.5 The total field of view (TFOV) of the FGS is onequarter of the outer 4 arc min annulus of the optical telescope assembly (OTA) FOV (Fig. 2). The unvignetted FOV encompasses 72.2 arc-min 2 • The TFOV is scanned when in search mode with a reticle (3.14 x 3.14 arc-sec 2 ) called an instantaneous field of view (IFOV). Two internal electromechanical star selectors are used to position the IFOV, provide encoder measurements used in calculating star position and direct the incoming light beam to illuminate the base of a Koesters prism (interferometric device). Photometric detection and measurement of the emergent beams that are formed from the incoming light by the prism provides null position information. Figure 3 illustrates the action of the two star selectors in the object space. Each star selector deflects the incoming "guide star" light beam by 7.10 arc-min in object space. Therefore, the final deflection in FGS pupil space is 6.777 degrees since the FGS magnification factor is 57.275. Object space is defined by the projection on the celestial sphere. Pupil space is collimated space at which an image of the entrance pupil is located. Objects in the IFOV lie on the circumference of a circle with a radius of 7.1 arc-min (Circle B of Fig. 3). The first star selector within an FGS commanded to locate a specific guide star images
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•
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Positioning of IFOV of TFOV by Two Star Selector Assemblies
the IFOV into lB. IB lies on a circle of radius 7.1 arc-min (Circle A) from the telescope optical axis. IB is imaged by the second star selector within the FGS onto the optical axis. The position of the objects in the IFOV on Circle B and their telescope image, IB, on Circle A are defined by the star selector rotations OB and OA. OA and OB are positive rotations about the T1 axis from the T2 axis. The position of the IFOV as determined from the vectors established by the rotations of the star selectors can be converted into spherical polar coordinates (R, e ) information that is centered about the telescope optical axis. A refinement of this positional information from the two star selectors is accomplished by the signal available from the Koesters prism. This signal provides an accurate measure of the incident wave front tilt from the null position in one direction parallel to the base of the prism. Since the degree of offset in orthogonal directions is needed for precise positioning from null, two prisms are used in each FGS. A beam splitter is employed to create two separate beams from the single beam directed from the star selectors. The two Koesters prisms are positioned mechanically so their bases are orthogonal to each other. Operation of the Koesters prism interferometer is illustrated by Fig. 4. The Koesters prism contains a 50/50 beam splitter along its central interface. A wavefront of incident light on the base of the prism is reflected internally by each side of the prism. A quarter wave retardation of the transmitted beams is experienced at the beam splitter surface. The primed letters represent the transmitted wavefront and the double primed letters the reflected wavefront. At null (no wavefront tilt) the transmitted light beams in both exit faces of the prism are of equal intensity. As an example, when a A/4 tilt is present, a' is advanced by A/8, and c" is retarded by A/8. Therefore, additive interference will occur at the upper edge of the left hand face. The left hand face will have a gradual increase in intensity from the bottom to the top. At the right hand face, c' will be retarded by A/8, and a" will be advanced by A/8. Therefore, subtractive interference will occur at the upper edge of the right hand face. The right hand face will have a gradual decrease in intensity from the bottom to the top. With the tilt used in
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Space Telescope
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,
The treatment of items presented in this paper was brief but intended to convey an overview of this challenging project and the design features of several vital systems. As with most sophisticated projects comprised of new designs that represent the frontiers of technology, numerous problems have been encountered and resolved. Designs have been developed to satisfy a very demanding set of scientific requirements. At mid-pOint in the development and fabrication of this observatory it is gratifying to report that many essential elements of hardware/software are available and verified "within specification". Noteworthy in this category are all the scientific instruments, the optical system, and much of the pointing control system. Although all systems will collectively contribute to the ultimate success of the Space Telescope, the optical, and pointing control systems contain the real technical challenge of this project. The final determination and degree to which the Space Telescope satisfies the scientific requirements will be a direct function and measure of the performance of these two unique systems.
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REFERENCES A Z
Fig. 4.
1.
Giacconi, R. (1982). Science Operations with Space Tele scope • ~Th~e~S::.tp~a~c:.:e:.....;T~e:.:l~e:.:s::;c:.:o~p~e=--O::.b::.s::.e=rv-=-- atory NASA CP-2244.
2.
O'Dell, C. R. (1982). The Space Telescope Observatory. The Space Telescope Observatory NASA CP-2244.
3.
O'Dell, C. R. (1981). The Space Telescope. Annual Reviews Inc. #0-8243-2902-3, Palo Alto, California, U.S.A.
4.
McRoberts, J. J. (1982). NASA Publication EP-166.
5.
Design Characteristics Definition Documents (1983). Marshall Space Flight Center, AL 35812.
Koester's Prism Interferometer
this example, the total integrated illumination intensity of the left hand face will be greater than on the right hand face. The total integrated illumination intensity from one side of the interferometer is measured by a photomultiplier tube (PMT). The difference between the two PMTs assigned to each Koesters prism is the phase difference average between the two portions of the angle of light incidence and results in a signal used to precisely control the observatory.5 Because the exact position of the object to be observed is known in relation to the guide stars selected for the FGS's, the observatory will maneuver to place the object of interest in the focal plane position of the scientific instrument with an accuracy of 0.01 arc-sec. CONCLUSION The Space Telescope will be a powerful tool for astronomy throughout its fifteen or more years of operational life. As a national facility with international participation, it will be available to any qualified science investigator with an appropriate program. The day-to-day observations accomplished by the Space Telescope will be planned and administered by the Space Telescope Science Institute under a contractual arrangement with the National Aeronautics and Space Administration. Therefore, scientific operations of the Space Telescope will be accomplished in much the same way as other ground-based astronomy centers are currently operated except for those differences made necessary by real-time management of a remote telescope in space.
Space Telescope.
©
SPACECRAFT ATTITUDE CONTROL
IFAC 9th Triennial World Congress
Budapest. Hung;u)'. 19R4
SPACE TELESCOPE F. S. Woj talik Deputy Project Managfr, Space Telescope Projec/ITA01, Marshall Space Flight Cen/fr, AL 35812, USA
Abstract. This paper describes the Space Telescope, a program of the National Aero nautics and Space Administration (NASA) . The Space Telescope is the largest and most powerful optical and ultraviolet astronomical observatory to be operated in space . Through the remote eyes of this telescope, astronomers will look further into space and time to produce data incl uding imagery of unequaled quality of galaxies, star systems, quasars and other objects of scientific interest. The Space Telescope will do this by observing the sky from a nominal 550 kilometer low Earth orbit via an optical system with a 2 . 4 meter diameter primary mirror . Additionally, this observatory has a precision pointing control system that is capable of maintaining a locked state on an object for extended periods accurately to within 0 . 01 arc seconds. The observatory instrument complement, optical system and pointing control system is presented. Keywords . Strapdown systems; space vehicles; satellites; stability; attitude control; observability. INTRODUCTION Viewing the celestial sky in space, outside the Earth's obscuring atmosphere, through a telescope with observatory class features is a project that astronomers have proposed and worked to achieve for many years . First, relatively small telescopes like NASA ' s Orbiting Astronomical Observatory (OAO), International Ultraviolet Explorer (IUE), and later the much larger series of three High Energy Astronomy Observatories (HEAO ' s) were orbited with support from the scientific and industrial communities with a high degree of scientific success. NASA is now past mid point in developing and launching into orbit a much larger and more long lived observatory called Space Telescope. This unmanned observatory, scheduled to become operational in the next few years, will have a sophisticated optical telescope, be capable of accommodating a variety of scientific instruments, and will operate in low Earth orbit for many years. Astronomy is concerned and challenged by fundamental and basic issues; the position, motion, constitution, history and
future of celestial bodies . With the Space Tele scope the astronomers capa~ilities for observation as a means of gathering data to resolve many astronomical concerns will be greatly improved. The Space Telescope is designed to be launched, deployed, serviced, and returned to Earth by means of the Space Shuttle Transportation System. After launch from the Kennedy Space Center, the Space Shuttle Orbiter will maneuver to attain the proper orbit and attitude, then the payload bay doors will be opened to allow the 15 meter mechanical arm of the Remote Manipulator System (RMS) to remove the Space Telescope from the payload bay. While the Space Telescope is connected to the RMS or immediately after being released into space, solar arrays and high gain antenna deployment will occur and systems functional checks will be initiated to verify that the observatory is functioning properly. At this point, the Orbiter will return to Earth and the Space Telescope mission will begin. For on orbit maintenance visits the Orbiter will maneuver into rendezvous position, the RMS will engage the free flying observatory and transfer it (with appendages retracted) into the payload bay where maintenance will proceed.
Because the observatory initial orbit altitude will gradually decay due to atmospheric drag, the Orbiter will reboost the Observatory to its initial orbit altitude during a maintenance visit. The capability to return the observatory to Earth for extensive overhaul is possible but every effort will be made to achieve repair/replacement on orbit for the fifteen or more years of expected life. SPACE TELESCOPE SYSTEM The final spacecraft and optical system design of the Space Telescope is the result of scientific objectives and requirements merged with various space oriented system operational constraints (Fig. 1) . As an example, the Space Shuttle Orbiter payload bay imposes a final limitation on the payload overall length and diameter. The envelope of the telescope once basically established, the number and types of scientific instruments planned, and the pointing accuracy and stability required impose other real limitations on the final design. The Space Telescope is a long-lifetime general purpose telescope that is by design capable of operating with a wide variety of scientific instruments at its focal plane. This observatory is a cylindrical structure with a length of 13.1 meters, a diameter of 4.27 meters that weighs about 11,000 kilograms. Two large solar arrays that can be deployed and retracted on command are capable of rotation about the observatory pitch axis (V2) to maintain a normal attitude with the sunline . The two solar arrays (wings) each measure 2.3 by 11 . 8 meters and are made of two solar cell blankets/wing. Each solar cell blanket consists of five solar panel assemblies for a total of 20 panels for the observatory. Each solar panel assembly contains 2438 cells giving a grand total for the observatory of 48,760 solar cells. The two solar arrays, six nickel-cadmium batteries, and power conditioning equipment will supply a minimum of 2 . 4 KW of orbital average electrical power after two years of operation. S The major configuration elements of the Space Telescope are the Optical Telescope Assembly (OTA), a Support Systems Module (SSM), the Scientific
2909
F. S. Wojtalik
2910
OmcAl
nLlsca.l ASSlMllT
V3
Fig. 1.
Space Telescope Configuration
Instruments (SI's) end the Solar Arrays (SA's) previously described. The OTA contains a 2.4 meter reflecting Cassegrain-type telescope housed within a meteoroid shield and sunshade. The telescope itself is a Ritchey-Chretien folded optical system with the secondary mirror located inside the prime focus. The primary mirror is fabricated of ultra-low expansion titanium silicate glass where the front and back plates are fused to a honeycomb core. Primary mirror weight is 829 kg with a clear aperture of 2.4 meters and f24 focal ratio. The mirror has an aluminum magnesium fluoride coating and is heated during operation to about 70°F to minimize variations from its original fabrication accuracy. After the aperture cover is commanded open, any entering light will illuminate the primary mirror at the rear of the telescope. The primary mirror will project the light to the smaller secondary mirror in the front end of the OTA. The light beam in turn will then be reflected back through a hole in the primary mirror to the SI's in the rear at the focal plane. The optical surface of the mirror has been measured and shown to have a figure error of less than 1/60 of a wavelength (rms) at 632.8 nm. Seventy percent of the energy from a star impinges in a radius at the focal plane of less than 25 microns at 632.8 nm, corresponding to about 0.1 arc sec. The reflection efficiency of the aluminized MgF2 coated optical surfaces, as measured, is better than 70% at 121.6 nm and 85% at 632.8 nm. With its complement of instruments the telescope will cover the range of wavelengths from the far ultraviolet to the far infrared. The field of view at the focal plane minus some structural obscuration is a circular area with a diameter of 28 arc min (Fig. 2). The central area of 22 arc min has the best resolution as realized from a Ritchey-Chretien optical design and is, therefore, used by five scientific focal plane instruments and three Fine Guidance Sensors (pseudo scientific instruments) that view three 90 degree segments of the outer portion of the field of view (FOV). The central 2.7 x 2.7 arc min area is diverted by a pick-off mirror that sends the beam
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Fig. 2.
Space Telescope f/24 Focal Plane
of light to the Wide Field/Planetary Camera instrument. Four quadrants about 6 arc min wide at the central area of the FOV are assigned to four axially oriented instruments. The Support System Module that encloses the OTA and SI's provides interfaces to the Orbiter and ground operations. Ground operations are accomplished via the Tracking and Data Relay Satellites (TDRS). The SSM also provides the essentials for proper and safe operation of the observatory. These include the electrical power and distribution system, communications, attitude and stabilization control, data management and thermal control systems. Observatory attitude control during maneuvers and pointing are the result of speed variations in the four available reaction wheels. The signals for wheel speed and rotation control are derived from a variety of sensors; a coarse sun sensor, fixed head-star trackers, gyroscopes and interferometric star trackers, The functional interactions of these devices is discussed in more detail later.
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Space Telescope SCIENCE INSTRUMENTS Five scientific instruments and one pseudo scientific instrument are planned for the Space Telescope initial mission. This complement of instruments are located at the focal plane behind the primary mirror where they are held in exact position by the focal plane structure through a set of latches. These instruments are the Wide Field/Planetary Camera (WF/PC), Faint Object Spectrograph (FOS) , High Resolution Spectrograph (RRS), High Speed Pho t ometer (HSP) , Faint Object Camera (FOC), and the Fine Guidance Sensors (FGS). As mentioned, the FGS is considered to be a pseudo scientific instrument because of its accuracy in star location. 4 It is used primarily to provide Space Telescope high pointing stability. For this function two sensors (there are three) are sufficient for locating and locking onto a target, therefore, the third, a standby unit, can be used to precisely measure the positions of other stars in the vicinity in relation to the stars on which the other two sensors are locked. These instruments will provide precision pointing accuracy and stability to the observatory, calibrate the positions of nearby and distant stars and galazies, reveal new information on the unseen companions of binary star systems, and provide better positional reference systems for compact stars. Each scientific instrument is a separate module capable of being removed and replaced in space. The Wide Field/Planetary Camera and the three Fine Guidance Sensors are located in radial positions with respect to the optical axis and the other instruments are in bays parallel to the telescope line of sight. The Wide Field/Planetary Camera operates in two modes. It has a wide field (2.63 x 2.63 arc min 2 ) capability at a focal ratio of f/12.9 that allows large areas of the sky to be observed (wide field) and a second field of view (68.7 x 68.7 arc sec 2 ) (Planetary). The detector is designed with four charged coupled detector arrays of 800 x 800 pixels each. The pixel in the wide field mode corresponds to 0.1 arc sec and in the planetary mode to 0.043 arc sec. The first mode provides the largest field of view (FOV) available on the Space Telescope at a reduced resolution while the second mode provides optics limited resolution. Sensitivity in the range of 115.0 to 1100.0 nm with photometric accuracy of 1 percent over a wide dynamic range has been verified by test. Fortyeight special purpose filters, polarizers, and gratings are included in a wheel arrangement to produce color cosmology information. The sensitivity of this instrument has been measured showing that it can achieve high quality point sources information down to 28 Mv after a one hour exposure . The High Resolution Spectrograph is intended for operation in the 110.0 to 320.0 nm region with a resolution capability of 2 x 10 4 . This resolution corresponds to 0.05 Aand will be one of the best available in astronomy. This instrument uses Digicon detectors and concave light dispersing gratings. Spectroscopic resolution, photometric accuracy and sensitivity is expected to be greatly improved over what was provided by previous space instruments including the highly successful International Ultraviolet Explorer (IUE). The Faint Object Spectrograph is a medium resolution, 10 2 to 10 3 , instrument. Digicon arrays are used in the design to detect wavelengths of 115 to 800 nm. This instrument will be capable of 30 A to 3 A detection depending on the commanded mode of operation. Polarization and time variability studies in the spectra of bright sources of the incoming light with resolution of 10 milliseconds will be routinely performed. The Faint Object Camera, developed by the European Space Agency (ESA), will complement the Wide Field/Planatory Camera and uses the full performance capability of the Space Telescope through the attainment of high angular resolution
and sensitivity in a narrow wavelength band of 120 to 500 nm. It operates in two modes which produce images of 11 x 11 arc sec 2 or 22 x 22 arc sec 2 onto vidicon tubes with magnesium fluoride intensifiers. Pixels of 25 microns equate to 0.02 and 0.04 arc sec, respectively, in the two available operational modes. The instrument records individual photons and can achieve very high angular resolution of 0.007 arc sec in very narrow fields thereby permitting post facto image reconstruction methods. As with many of the other Space Telescope scientific instruments, it also has filter wheels, each containing a variety of filters with 12 positions that can be commanded for insertion into the optical path. 1 The Faint Object Camera and other instruments will be used to observe extragalactic supergiant stars, study variable brightness stars, gather data on globular clusters, examine binary star systems, search for extrasolar planets, establish stellar masses, do detailed studies of shock fronts and condensing gas clouds, search for direct evidence that quasars might be at the center of faint galaxies, and other studies and investigations of astronomical interest. 4 POINTING AND STABILIZATION CONTROL The quality of the Space Telescope (ST) optical system, the sensitivity of its complement of scientific instruments and mission operating constraints were the principle drivers that established requirements and design architecture of the pointing and stabilization control system. The selected design has two basic operating attitude control modes; Sun Point Control (SPC) and Operational Vehicle Control (OVC). SPC is used during initial observatory deployment from the Space Shuttle Orbiter when pre-mission equipment checks are being performed and in contingency operations. Entry into the SPC mode can occur autonomously whenever one of several on-board electrical power system failure conditions is sensed or by ground command. In this mode, the observatory +V3 axis and solar array active surfaces are positioned toward the sun and maintained in this orientation under firm attitude control but within loose position limits. Maintaining the solar arrays active surfaces to be within ±S degrees normal to the sun is very adequate. Control is provided by reaction wheels, sun sensors, rate gyros and the digital computer. The computer uses output signals from the sun sensors to establish attitude orientation during orbit sunlight and rate gyro outputs to maintain proper attitude position in eclipse. Gyro-sensed body rotational rates are used to maintain the plane, formed by the V3 and V1 axes within ±S degrees of the sunline. A hardwired set of logic is also provided to accomplish the computer functions, if required in this contingency mode. Entry into the long term and normal Operational Vehicle Control (OVC) mode is accomplished by ground command only. In this mode, solar array orientation toward the sun is maintained by the computer as previously described while rotation of the observatory about the V3 and V2 axes is introduced. The function of the OVC mode is to orient and maintain the Space Telescope optical axis (+V1 axis) to targets (celestial coordinates) of scientific interest for observations lasting several minutes to several days with precision accuracy and stability. To accomplish this, the control mode is separated into two distinct phases. A course positioning phase that uses body-fixed rate integrating gyros arranged in pairs so their inputsensitive axes are in opposite directions and skewed at 31.72 degrees up from their mounting plane, fixed head star trackers that are separate and apart from the telescope optics and fine guidance sensors which derive their signals from stars imaged near the edge of the Space Telescope FOV
2912
F. S. Wojtalik
(Fig. 2). This course positioning system as well as the precision pointing system requires the services of the computer and reaction wheels. ThE precision pointing system utilizes the fine guidance sensors in a fine lock mode to update the stability reference of the set of rate gyros being used every second to a noise equivalent positional error of 0.0028 arc sec (rms) yielding an overall Space Telescope positioning accuracy of 0.01 arc sec with 0.007 arc sec (rms) of stability. The sequence of events in a typical acquisition of a target of interest starts with a maneuver (slew) of the telescope limited to a maximum rate of 15 degrees/min. Using the coordinates of the present known location and the new desired location plus the available reaction wheel momentum, an eigen axes rotation is computed. The observatory is commanded to follow this rotation at a specified rate. At approximately 0.4 degrees away from the maneuver termination, a forward-loop integrator within the computer logic is used to minimize rate over-shoot and other disturbance torques. Once the observatory arrives at the new desired location, pointing reference errors are expected to be less than !30 arc sec (30) and body rates diminished to less than 0.05 arc sec/sec. Since the uncertainties of position due to gyro drift and the inaccuracies (especially roll angle) of the initial direction increase with displacement, slews will be planned to be as short as possible. The uncertainty can be several arc minutes after only a 90 degree slew. During this phase of operation, the fixed head star trackers are used to determine the rough position (!10 arc-sec calibrated accuracy) as necessary. At this time, the coordinates and brightness of a candidate guide star, for the target of interest, is transferred into each of two Fine Guidance Sensors (FGS's). The three on-board FGS's operate with the observatory support system module (SSM) pointing control system (PCS) to provide on-board fine (precision) pointing control and stabilization. Two FGS's are used during each pointing and hold operation while the third unit in the complement is in standby and can perform astrometry as previously described. Guide star positional information is provided by two FGS modes; (1) a coarse measurement mode accurate to !0.1 arc-sec (rms), at an update rate of 0.025 sec per axis, and (2) a fine (precision) measurement of 0.0028 arc-sec (rms) which is the error allocated to the FGS for image stability.5 The total field of view (TFOV) of the FGS is onequarter of the outer 4 arc min annulus of the optical telescope assembly (OTA) FOV (Fig. 2). The unvignetted FOV encompasses 72.2 arc-min 2 • The TFOV is scanned when in search mode with a reticle (3.14 x 3.14 arc-sec 2 ) called an instantaneous field of view (IFOV). Two internal electromechanical star selectors are used to position the IFOV, provide encoder measurements used in calculating star position and direct the incoming light beam to illuminate the base of a Koesters prism (interferometric device). Photometric detection and measurement of the emergent beams that are formed from the incoming light by the prism provides null position information. Figure 3 illustrates the action of the two star selectors in the object space. Each star selector deflects the incoming "guide star" light beam by 7.10 arc-min in object space. Therefore, the final deflection in FGS pupil space is 6.777 degrees since the FGS magnification factor is 57.275. Object space is defined by the projection on the celestial sphere. Pupil space is collimated space at which an image of the entrance pupil is located. Objects in the IFOV lie on the circumference of a circle with a radius of 7.1 arc-min (Circle B of Fig. 3). The first star selector within an FGS commanded to locate a specific guide star images
IIISTAIITAIIEOUS FIELD Of VIEW OFTHEfSS
14."
TOTAL FIELD OF VIEW OF FGS
T2
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Fig. 3.
•
7.1'
7.1'
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1.771"
Positioning of IFOV of TFOV by Two Star Selector Assemblies
the IFOV into lB. IB lies on a circle of radius 7.1 arc-min (Circle A) from the telescope optical axis. IB is imaged by the second star selector within the FGS onto the optical axis. The position of the objects in the IFOV on Circle B and their telescope image, IB, on Circle A are defined by the star selector rotations OB and OA. OA and OB are positive rotations about the T1 axis from the T2 axis. The position of the IFOV as determined from the vectors established by the rotations of the star selectors can be converted into spherical polar coordinates (R, e ) information that is centered about the telescope optical axis. A refinement of this positional information from the two star selectors is accomplished by the signal available from the Koesters prism. This signal provides an accurate measure of the incident wave front tilt from the null position in one direction parallel to the base of the prism. Since the degree of offset in orthogonal directions is needed for precise positioning from null, two prisms are used in each FGS. A beam splitter is employed to create two separate beams from the single beam directed from the star selectors. The two Koesters prisms are positioned mechanically so their bases are orthogonal to each other. Operation of the Koesters prism interferometer is illustrated by Fig. 4. The Koesters prism contains a 50/50 beam splitter along its central interface. A wavefront of incident light on the base of the prism is reflected internally by each side of the prism. A quarter wave retardation of the transmitted beams is experienced at the beam splitter surface. The primed letters represent the transmitted wavefront and the double primed letters the reflected wavefront. At null (no wavefront tilt) the transmitted light beams in both exit faces of the prism are of equal intensity. As an example, when a A/4 tilt is present, a' is advanced by A/8, and c" is retarded by A/8. Therefore, additive interference will occur at the upper edge of the left hand face. The left hand face will have a gradual increase in intensity from the bottom to the top. At the right hand face, c' will be retarded by A/8, and a" will be advanced by A/8. Therefore, subtractive interference will occur at the upper edge of the right hand face. The right hand face will have a gradual decrease in intensity from the bottom to the top. With the tilt used in
2913
Space Telescope
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,
The treatment of items presented in this paper was brief but intended to convey an overview of this challenging project and the design features of several vital systems. As with most sophisticated projects comprised of new designs that represent the frontiers of technology, numerous problems have been encountered and resolved. Designs have been developed to satisfy a very demanding set of scientific requirements. At mid-pOint in the development and fabrication of this observatory it is gratifying to report that many essential elements of hardware/software are available and verified "within specification". Noteworthy in this category are all the scientific instruments, the optical system, and much of the pointing control system. Although all systems will collectively contribute to the ultimate success of the Space Telescope, the optical, and pointing control systems contain the real technical challenge of this project. The final determination and degree to which the Space Telescope satisfies the scientific requirements will be a direct function and measure of the performance of these two unique systems.
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REFERENCES A Z
Fig. 4.
1.
Giacconi, R. (1982). Science Operations with Space Tele scope • ~Th~e~S::.tp~a~c:.:e:.....;T~e:.:l~e:.:s::;c:.:o~p~e=--O::.b::.s::.e=rv-=-- atory NASA CP-2244.
2.
O'Dell, C. R. (1982). The Space Telescope Observatory. The Space Telescope Observatory NASA CP-2244.
3.
O'Dell, C. R. (1981). The Space Telescope. Annual Reviews Inc. #0-8243-2902-3, Palo Alto, California, U.S.A.
4.
McRoberts, J. J. (1982). NASA Publication EP-166.
5.
Design Characteristics Definition Documents (1983). Marshall Space Flight Center, AL 35812.
Koester's Prism Interferometer
this example, the total integrated illumination intensity of the left hand face will be greater than on the right hand face. The total integrated illumination intensity from one side of the interferometer is measured by a photomultiplier tube (PMT). The difference between the two PMTs assigned to each Koesters prism is the phase difference average between the two portions of the angle of light incidence and results in a signal used to precisely control the observatory.5 Because the exact position of the object to be observed is known in relation to the guide stars selected for the FGS's, the observatory will maneuver to place the object of interest in the focal plane position of the scientific instrument with an accuracy of 0.01 arc-sec. CONCLUSION The Space Telescope will be a powerful tool for astronomy throughout its fifteen or more years of operational life. As a national facility with international participation, it will be available to any qualified science investigator with an appropriate program. The day-to-day observations accomplished by the Space Telescope will be planned and administered by the Space Telescope Science Institute under a contractual arrangement with the National Aeronautics and Space Administration. Therefore, scientific operations of the Space Telescope will be accomplished in much the same way as other ground-based astronomy centers are currently operated except for those differences made necessary by real-time management of a remote telescope in space.
Space Telescope.