- Email: [email protected]
Low-Cost Space Structure Pointing Experiment
Copyright ~ IFAC Automatic Control in Aerospace, Palo Alto, California, USA, 1994
LOW-COST SPACE STRUCTURE POINTING EXPERIMENT S. M . SELTZER* AND PAUL S. SHIRLEY**
*Chiej Engineer, SVS Inc. . 2786 Hurricane Road. New Market. AL. 35761 * *President, SVS Inc. , 7408 Coachman Street NE. Albuquerque. NM 87109 Abstract. Because of the high cost of space programs and the specter of system failure, it is important to be as certain as possible that spacecraft perform adequately in space. Often the most important ingredient is adequate pointing and control performance. This paper describes an approach to investigate whether pointing performance can be demonstrated for spaceborne optical telescopes . The proposed experiment consists of a few representative "sparse" elements incorporated into a small satellite, dynamically emulating the actual large systems, and ejected from a Space Shuttle as a Hitchhiker payload . Key Words. Modeling, satellite control, space vehicles, structural controllability, target tracking
1.
INTRODUCTION
proven to be misleading. Accurate spaceborne performance testing for an y particular program has been prohibitively expensive. The primary objective of the LeSS experiment is to demonstrate that the necessary precision pointing and control performance for a dynamically emulated large scale optical telescope can be achieved on a low-cost, lightweight space experiment on a spacecraft that can be placed into Earth orbit by, and ejected from, a Space Shuttle arbiter. The spacecraft's payload will be comprised of a sparse representative optical payload system. The satellite will be dynamically representative of a large spaceborne optics system . Use of sparse elements minimizes the cost, weight , and size of the experiment, while the Hitchhiker approach substantially reduces launch costs. After orbital insertion. stowed sparse components will be deployed to the correct distances with representative structural dynamics. They represent a spacecraftborne large optical telescope that can be pointed and controlled with the desired accuracy. The experiment is not designed to investigate the full breadth of structural dynamics problems. It will, however, demonstrate the accurate precision pointing and control of a given class of large optical systems in space . It is only in space that one can obtain a realistic test of precision pointing and control and structural dynamics .
Because of the high cost of space programs and the specter of system failure, it is important to be as certain as possible that the spacecraft will perform adequately (or better) in space. Often the most important ingredient of a successful space program is adequate pointing and control performance. A number of approaches have been used to attempt to predict and assess the capabilities and problems associated with pointing performance. These have included ground test facilities, balloon and aircraft flight experiments, limited space experiments, pure analytical (including computer simulation) investigations, and extrapolations from previous space programs. In general, these approaches have been expensive, inaccurate, or both. Although costly, ground test performance results provide abundant information. However, numerous problems associated with the Earth's gravitational field and its atmosphere often mask the dynamic nature of the test article when it is actually in a space (or near-space) environment. A comprehensive description of the practical problems encountered in ground testing may be found in Waites and Worley, 1987. Simulated space environment testing onboard aircraft is limited by time duration and size constraints, as are balloon tests and free-fall experiments. Limited space experiments have been performed onboard the Shuttle arbiter. The results have been meaningful and sometimes inexpensive. However, extrapolating from these and other fullblown program performance results has often been
A realistic future orbiting large optical system has been selected and its characteristics defined . The key optical elements for the system are shown in in the upper left corner of Fig. 1. They represent the
157
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.
.
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Selected Sparse DEW Elements .-- ...... _-_ ....... --... .. '.
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---------------.
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'\.
.- .. --,
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Fig. I. LeSS concept development
---\\
'----i
,
optical properties of a full-scale realistic optical system. From it, a sparse collection of its optical elements (lower left corner of Fig. 1) is extracted to provide a geometrically and dynamically scalable subset of elements which bec omes the proposed space experiment (right half of Fig. 1) (Seltzer , 1993). Figure 2 illustrates the stowed experiment, as well as the deployed configuration of LCSS.
elements to the spacecraft controller during the target engagements in order to keep the steering mirrors from exceeding their dynamic range ("throw") . The optical LOS control system is hierarchical in nature with a moderate inner loop (30-Hz alignment control) and a slower outer track loop (nominally 1 to 3 Hz) providing precision LOS stabilization and target tracking. Low-frequency, open-loop offloads from the steering mirrors to spacecraft pointing are provided to prevent the mirrors from "hitting their stops" during high-rate maneuvers . Bandwidths for the various loops are constrained by the frame rate of the tracking cameras but in general are selected to minimize LOS pointing errors caused by spacecraft disturbances . Implementation of the inner alignment loop combines a relatively high-bandwidth sensor [quad cell or photopot type for light-emitting diode (LED) or target source] driving two galvo steering mirrors. The control errors are generated by the payload processor as centroid errors of the LED spots referenced to particular (boresighted) locations on the alignment sensor.
2. SYSTEM DEFINITION Based on a typical future orbiting large optical system , it is desired that the LCSS Pointing Experiment geometry have a distance of 1.93 m between the edge of the outboard primary mirror segment and the centerline of the spacecraft. A distance of 4.67 m between the sparsely popUlated primary mirror and the secondary mirror is desired. In consideration of the selected typical system's structural tlexibilities, it is desired that the first tl e xible mode of the extendible boom from the spacecraft main body to the outboard primary mirror segment be as close to 70 Hz as possible . Similarly, the first Ilexible mode of the extendible boom from the spacecraft main body to the secondary mirror should be close to 22 Hz. It is planned to simulate the effects of the major onboard disturbances with an onboard disturbance generator. Initially , the primary targets for the pointing e xperiment will be stars.
3. SATELLITE The LCSS satellite must be capable of providing nominal slew rates and accelerations for the optical payload . It is desired to be able to slew the spacecraft at a rate of 20 mrad/s, with a maximum acceleration of 0 .2 mrad/s 2 .
The technical approach for pointing and controlling the future orbiting system is to conceive of as simple a control system as possible. The two major reasons are : (1) to keep overall program costs low and (2) to improve operational reliability by decreasing complexity. A simple control algorithm , such as PID, is selected to provide and direct the "high" (relatively) authority control. If possible, control authority for the spacecraft and for the optical pointing and control will be provided by the same set of actuators . In case higher control authority is needed , design provision has been made for providing a cold gas mass expUlsion system . Because of the small size of the LCSS Pointing Experiment. a set of small reaction wheels has been selected; their electrical power requirements are small. If necessary, the dynamic effects of the Ilexible appendages will be dealt with by applying notch filters , phase stabilization, or a form of high authority/Iow authority control (HACILAC).
The system weight and geometric envelope are prescribed by the Space Transportation System (STS) Shuttle Orbiter Hitchhiker payload constraints as 204 kg and 84-cm diameter x 99-cm height. These dimensions are also within the capability of small launch vehicles such as Pegasus. The satellite structure must provide for mounting all the spacecraft subsystems and for attachment and deployment from the flight carrier. The satellite structure must be lightweight (maximum use of aluminum) to avoid a high overhead penalty; thus , the design approach is to simplify the satellite structure such that the mounting, carrier attachment, and deployment structure is one plate . In addition, since one stiff main plate is used, it also serves as the optical bench for the payload.
Spacecraft control is concerned primarily with stabilizing the platform and performing lowbandwidth pointing functions to assist in target acquisition , sensor calibration, data down linking, etc . It is coupled with the optical line-of-sight (LOS) control system. For example, the spacecraft is commanded by the target acquisition sensor as well as by the "normal" complement of spacecraft sensors (i .e. sun sensors, magnetometers, inertial gyros, etc.) for initial pointing of the payload . In addition , the optical system "oftloads" the active
The satellite equipment complement is simple and conventional. The attitude control system consists of two pitch-axis horizon scan wheels, roll and yaw reaction wheels, three magnetic torque rods, a 3-axis magnetometer, actuator drive electronics, a sun sensor set, a cold gas thruster system, an inertial attitude reference unit, a GPS receiver, and flight computer software. Thermal control is passive. The electrical power subsystem is a conventional mix of solar arrays and batteries.
159
E
o v
00
99cm (39")r------"-~I
160
analysis has shown that reasonable signal-to-noise ratios are still possible on the focal planes with fifth magnitude or brighter stars. The light that is collected is shared between the two cameras (NFOY and WFOY) by means of a spectral filter in the peak bandpass of the visible CCO array. When the time history motion of the star image on the NFOY camera (which represents the LOS of the main optical subsystem) is differenced with the WFOY spot motion, the structural response of the telescope system can be determined. The two conditions required are (1) the mechanical alignment errors between the two CC Os must be small and constant (10 mrad), and (2) the alignment of the sparse array optics relative to each other must be very good (0.4cm translation control). The issue for sensor alignment is one of maintaining a repeatable relative measure of spot motion; telescope optics alignment is required to maintain a reasonable image quality for the NFOY camera. X- and Y -axes scanning mirrors are baselined as the active elements for LOS optical control. These mirrors are to be 2.5 cm commercial galvo scanning mirrors that will be designed to work closed-loop with the NFOY and WFOY data acquisition cameras.
4. PAYLOAD The LCSS optical payload is a "sparse" primary optical system that is intended to emulate the onorbit dynamics of a large distributed optical platform. The key components of the experiment payload are described below. The Acquisition Subsystem is a small compact lens and CCO camera (array and electronics) attached to the main optical bench of the optics payload. The assembly is offset from the secondary boom optic to assure an unobstructed field-of-view. Identical cameras are selected for the various payload sensors to minimize the cost of software development. Extensive use of commercial products for the system design should minimize development costs. The sparse optical system is simple in design, containing two off-axis portions of a 4-m primary and a 20-cm secondary. The combination comprises a complementary parabolic system designed to be afocal and represents a 20x reduction. All three of the major optical elements are to be made of aluminum honeycomb, polished and optically coated to a reasonable quality finish (1110th wave). The two primary sections are each parabolic with 20-cm and 40-cm minor and major axes, respectively. Two LEOs are mounted on each element and are used by the alignment subsystem to maintain optical alignment. The secondary is con figured so that a small 5-cm optical relay lens is mounted coaxial with the primary optical subsystem to form the image used by the wide FOY (WFOY) camera for "truth" sensing. The three major payload subsystems supported by the sparse primary optics and reduction telescope are the Alignment, the WFOY, and the Narrow FOY (NFOY) Subsystems. The Alignment Subsystem consists of two diodes each embedded in the two primary elements, a CCO array and electronics, a OSP hosted by the CPU, and a suite of algorithms designed to establish and monitor the positions of both primary elements relative to predetermined positions on the NFOY focal plane array (FPA). Both primary elements are to be actuated in two dimensions (tip and tilt) with the boom arm on one of the elements providing an additional degree of freedom. The WFOV Subsystem represents the true LOS of the optical system. It is configured to look through the secondary mirror assembly by means of the coaxial 5-cm relay optic that is aligned with the main telescope LOS. This enables the camera to observe a point source image of a target star. This image, and the motion in the spot that is generated as a function of the spacecraft motion, allows for the analysis of the dynamic motion of the secondary boom under both slewing and normal tracking system operations. The NFOV Camera views an image of the target star as seen by the main optical subsystem (the sparse primary and secondary). This image will appear as elliptical because of the sparse nature of the primary.
After initial detection of the target star in the acquisition camera, centroid track errors are computed by the payload computer (using the acquisition video) and sent as pointing commands to the spacecraft controller. The spacecraft is then driven at low frequency (nominally 0.1 Hz) to center the target star in the acquisition FOY. Once the target is within a 20x20 pixel acquisition track gate (roughly 2 mrad), hand over to the WFOY camera is enabled and this sensor begins "looking" for the target (intensity above a threshold). When the target is detected in the WFOY, centroid track errors are computed (from the WFOY video) and used to close a nominal I-Hz coarse track loop using the galvo steering mirrors. The target will be centered in the WFOY focal plane, and offload commands from the steering mirrors to the spacecraft controller are initiated. As the WFOY track error becomes small (target within a 6x6 pixel track gate size), the star should become detectable on the NFOY focal plane. Collocation of the two cameras on the payload optical bench should ensure adequate alignment for this condition. When the target is detected in the NFOY, centroid errors are computed in the track processor using NFOY video, and a fine-track loop (nominally 3-Hz bandwidth) is closed using the galvo steering mirrors.
5. PROGRAM STATUS The SYS-Ied development team has been awarded a Phase II Small Business Innovative Research (SBIR) contract to proceed with the development of the LCSS Pointing Experiment at the U.S. Air Force Phillips Laboratory (PL) in Albuquerque, New Mexico. The program is being managed by the PL Space and Missiles Technologies Structures
Because it is very sparse, the telescope will be an inefficient collector of light, but radiometric 161
requirements that are placed on spaceborne optical telescope programs. The nature of the requirements placed on the LCSS optical system to measure the performance of the space borne structural system has offered an expanded opportunity to support a broader range of space-based missions. Included in these missions are astronomical measurements of closely spaced objects in space and surveillance of spacebased and ground-based objects that require highresolution capability. The demonstrations of the optical payload and structural concepts developed by the LCSS Pointing Experiment will lead to highpayoff risk reduction for future military and civil missions that require precision pointing.
Division. Assisting SVS in the development of the satellite system is Teledyne Brown Engineering of Huntsville. Work on the Phase 11 system has already begun in Albuquerque. The planned Phase Jl activities will continue through the Spring of 1995. The laboratory-based Phase 11 program will locus on the development and demonstration of the hreadboard payload at PL. Specific activities will include the demonstration of the necessary precision alignment and control, structural response studies to support the boom design and development, characterization of the optical system performance, spacecraft hierarchical controls demonstrations, and the development of a full-scale mock-up to assist in the Phase III design and flight system development.
7.
REFERENCES
6. CONCLUSION Seltzer, S.M . (1993). "Low Cost Space Structure Pointing Experiment," Final Report PL-TR-933023 prepared for USAF Phillips Laboratory Structures and Controls Directorate, Edwards AFB, CA.
The LCSS Pointing Experiment offers a costeffective approach to addressing several of the key issues that face the advanced space structures and optical systems developers. By using off-the-shelf components, along with a few representative critical "sparse" elements of the generally complex and cost! y payload system, LCSS can demonstrate that SIgnificant elements of acquisition, tracking, and pOInting can be shown to meet the stressing
Waites, H.B. and H.E. Worley (1987) "Large Space Structures Testing", AAS Paper 87-036, Proceeedings of AAS Annual Guidance and Control Conference, Keystone, CO.
162
LOW-COST SPACE STRUCTURE POINTING EXPERIMENT S. M . SELTZER* AND PAUL S. SHIRLEY**
*Chiej Engineer, SVS Inc. . 2786 Hurricane Road. New Market. AL. 35761 * *President, SVS Inc. , 7408 Coachman Street NE. Albuquerque. NM 87109 Abstract. Because of the high cost of space programs and the specter of system failure, it is important to be as certain as possible that spacecraft perform adequately in space. Often the most important ingredient is adequate pointing and control performance. This paper describes an approach to investigate whether pointing performance can be demonstrated for spaceborne optical telescopes . The proposed experiment consists of a few representative "sparse" elements incorporated into a small satellite, dynamically emulating the actual large systems, and ejected from a Space Shuttle as a Hitchhiker payload . Key Words. Modeling, satellite control, space vehicles, structural controllability, target tracking
1.
INTRODUCTION
proven to be misleading. Accurate spaceborne performance testing for an y particular program has been prohibitively expensive. The primary objective of the LeSS experiment is to demonstrate that the necessary precision pointing and control performance for a dynamically emulated large scale optical telescope can be achieved on a low-cost, lightweight space experiment on a spacecraft that can be placed into Earth orbit by, and ejected from, a Space Shuttle arbiter. The spacecraft's payload will be comprised of a sparse representative optical payload system. The satellite will be dynamically representative of a large spaceborne optics system . Use of sparse elements minimizes the cost, weight , and size of the experiment, while the Hitchhiker approach substantially reduces launch costs. After orbital insertion. stowed sparse components will be deployed to the correct distances with representative structural dynamics. They represent a spacecraftborne large optical telescope that can be pointed and controlled with the desired accuracy. The experiment is not designed to investigate the full breadth of structural dynamics problems. It will, however, demonstrate the accurate precision pointing and control of a given class of large optical systems in space . It is only in space that one can obtain a realistic test of precision pointing and control and structural dynamics .
Because of the high cost of space programs and the specter of system failure, it is important to be as certain as possible that the spacecraft will perform adequately (or better) in space. Often the most important ingredient of a successful space program is adequate pointing and control performance. A number of approaches have been used to attempt to predict and assess the capabilities and problems associated with pointing performance. These have included ground test facilities, balloon and aircraft flight experiments, limited space experiments, pure analytical (including computer simulation) investigations, and extrapolations from previous space programs. In general, these approaches have been expensive, inaccurate, or both. Although costly, ground test performance results provide abundant information. However, numerous problems associated with the Earth's gravitational field and its atmosphere often mask the dynamic nature of the test article when it is actually in a space (or near-space) environment. A comprehensive description of the practical problems encountered in ground testing may be found in Waites and Worley, 1987. Simulated space environment testing onboard aircraft is limited by time duration and size constraints, as are balloon tests and free-fall experiments. Limited space experiments have been performed onboard the Shuttle arbiter. The results have been meaningful and sometimes inexpensive. However, extrapolating from these and other fullblown program performance results has often been
A realistic future orbiting large optical system has been selected and its characteristics defined . The key optical elements for the system are shown in in the upper left corner of Fig. 1. They represent the
157
s
t t
Pt_TI
.
.
;-;:A.~.~-----"""-'
VI 00
Selected Sparse DEW Elements .-- ...... _-_ ....... --... .. '.
01
s-.o.., .......
EJ· . . .:.: ;.;r;.;,,;w· .... · ...... · .. ·
~ PI
e
S
@
P2
"-Torr<
P2 ~~.:'-~.- -_ _----c
-_.- ..
--",.
_
..... _-----
\
) ' _... -
...--. \.
---------------.
'"'"
'\.
.- .. --,
',------.
~
""
\
'\.
Fig. I. LeSS concept development
---\\
'----i
,
optical properties of a full-scale realistic optical system. From it, a sparse collection of its optical elements (lower left corner of Fig. 1) is extracted to provide a geometrically and dynamically scalable subset of elements which bec omes the proposed space experiment (right half of Fig. 1) (Seltzer , 1993). Figure 2 illustrates the stowed experiment, as well as the deployed configuration of LCSS.
elements to the spacecraft controller during the target engagements in order to keep the steering mirrors from exceeding their dynamic range ("throw") . The optical LOS control system is hierarchical in nature with a moderate inner loop (30-Hz alignment control) and a slower outer track loop (nominally 1 to 3 Hz) providing precision LOS stabilization and target tracking. Low-frequency, open-loop offloads from the steering mirrors to spacecraft pointing are provided to prevent the mirrors from "hitting their stops" during high-rate maneuvers . Bandwidths for the various loops are constrained by the frame rate of the tracking cameras but in general are selected to minimize LOS pointing errors caused by spacecraft disturbances . Implementation of the inner alignment loop combines a relatively high-bandwidth sensor [quad cell or photopot type for light-emitting diode (LED) or target source] driving two galvo steering mirrors. The control errors are generated by the payload processor as centroid errors of the LED spots referenced to particular (boresighted) locations on the alignment sensor.
2. SYSTEM DEFINITION Based on a typical future orbiting large optical system , it is desired that the LCSS Pointing Experiment geometry have a distance of 1.93 m between the edge of the outboard primary mirror segment and the centerline of the spacecraft. A distance of 4.67 m between the sparsely popUlated primary mirror and the secondary mirror is desired. In consideration of the selected typical system's structural tlexibilities, it is desired that the first tl e xible mode of the extendible boom from the spacecraft main body to the outboard primary mirror segment be as close to 70 Hz as possible . Similarly, the first Ilexible mode of the extendible boom from the spacecraft main body to the secondary mirror should be close to 22 Hz. It is planned to simulate the effects of the major onboard disturbances with an onboard disturbance generator. Initially , the primary targets for the pointing e xperiment will be stars.
3. SATELLITE The LCSS satellite must be capable of providing nominal slew rates and accelerations for the optical payload . It is desired to be able to slew the spacecraft at a rate of 20 mrad/s, with a maximum acceleration of 0 .2 mrad/s 2 .
The technical approach for pointing and controlling the future orbiting system is to conceive of as simple a control system as possible. The two major reasons are : (1) to keep overall program costs low and (2) to improve operational reliability by decreasing complexity. A simple control algorithm , such as PID, is selected to provide and direct the "high" (relatively) authority control. If possible, control authority for the spacecraft and for the optical pointing and control will be provided by the same set of actuators . In case higher control authority is needed , design provision has been made for providing a cold gas mass expUlsion system . Because of the small size of the LCSS Pointing Experiment. a set of small reaction wheels has been selected; their electrical power requirements are small. If necessary, the dynamic effects of the Ilexible appendages will be dealt with by applying notch filters , phase stabilization, or a form of high authority/Iow authority control (HACILAC).
The system weight and geometric envelope are prescribed by the Space Transportation System (STS) Shuttle Orbiter Hitchhiker payload constraints as 204 kg and 84-cm diameter x 99-cm height. These dimensions are also within the capability of small launch vehicles such as Pegasus. The satellite structure must provide for mounting all the spacecraft subsystems and for attachment and deployment from the flight carrier. The satellite structure must be lightweight (maximum use of aluminum) to avoid a high overhead penalty; thus , the design approach is to simplify the satellite structure such that the mounting, carrier attachment, and deployment structure is one plate . In addition, since one stiff main plate is used, it also serves as the optical bench for the payload.
Spacecraft control is concerned primarily with stabilizing the platform and performing lowbandwidth pointing functions to assist in target acquisition , sensor calibration, data down linking, etc . It is coupled with the optical line-of-sight (LOS) control system. For example, the spacecraft is commanded by the target acquisition sensor as well as by the "normal" complement of spacecraft sensors (i .e. sun sensors, magnetometers, inertial gyros, etc.) for initial pointing of the payload . In addition , the optical system "oftloads" the active
The satellite equipment complement is simple and conventional. The attitude control system consists of two pitch-axis horizon scan wheels, roll and yaw reaction wheels, three magnetic torque rods, a 3-axis magnetometer, actuator drive electronics, a sun sensor set, a cold gas thruster system, an inertial attitude reference unit, a GPS receiver, and flight computer software. Thermal control is passive. The electrical power subsystem is a conventional mix of solar arrays and batteries.
159
E
o v
00
99cm (39")r------"-~I
160
analysis has shown that reasonable signal-to-noise ratios are still possible on the focal planes with fifth magnitude or brighter stars. The light that is collected is shared between the two cameras (NFOY and WFOY) by means of a spectral filter in the peak bandpass of the visible CCO array. When the time history motion of the star image on the NFOY camera (which represents the LOS of the main optical subsystem) is differenced with the WFOY spot motion, the structural response of the telescope system can be determined. The two conditions required are (1) the mechanical alignment errors between the two CC Os must be small and constant (10 mrad), and (2) the alignment of the sparse array optics relative to each other must be very good (0.4cm translation control). The issue for sensor alignment is one of maintaining a repeatable relative measure of spot motion; telescope optics alignment is required to maintain a reasonable image quality for the NFOY camera. X- and Y -axes scanning mirrors are baselined as the active elements for LOS optical control. These mirrors are to be 2.5 cm commercial galvo scanning mirrors that will be designed to work closed-loop with the NFOY and WFOY data acquisition cameras.
4. PAYLOAD The LCSS optical payload is a "sparse" primary optical system that is intended to emulate the onorbit dynamics of a large distributed optical platform. The key components of the experiment payload are described below. The Acquisition Subsystem is a small compact lens and CCO camera (array and electronics) attached to the main optical bench of the optics payload. The assembly is offset from the secondary boom optic to assure an unobstructed field-of-view. Identical cameras are selected for the various payload sensors to minimize the cost of software development. Extensive use of commercial products for the system design should minimize development costs. The sparse optical system is simple in design, containing two off-axis portions of a 4-m primary and a 20-cm secondary. The combination comprises a complementary parabolic system designed to be afocal and represents a 20x reduction. All three of the major optical elements are to be made of aluminum honeycomb, polished and optically coated to a reasonable quality finish (1110th wave). The two primary sections are each parabolic with 20-cm and 40-cm minor and major axes, respectively. Two LEOs are mounted on each element and are used by the alignment subsystem to maintain optical alignment. The secondary is con figured so that a small 5-cm optical relay lens is mounted coaxial with the primary optical subsystem to form the image used by the wide FOY (WFOY) camera for "truth" sensing. The three major payload subsystems supported by the sparse primary optics and reduction telescope are the Alignment, the WFOY, and the Narrow FOY (NFOY) Subsystems. The Alignment Subsystem consists of two diodes each embedded in the two primary elements, a CCO array and electronics, a OSP hosted by the CPU, and a suite of algorithms designed to establish and monitor the positions of both primary elements relative to predetermined positions on the NFOY focal plane array (FPA). Both primary elements are to be actuated in two dimensions (tip and tilt) with the boom arm on one of the elements providing an additional degree of freedom. The WFOV Subsystem represents the true LOS of the optical system. It is configured to look through the secondary mirror assembly by means of the coaxial 5-cm relay optic that is aligned with the main telescope LOS. This enables the camera to observe a point source image of a target star. This image, and the motion in the spot that is generated as a function of the spacecraft motion, allows for the analysis of the dynamic motion of the secondary boom under both slewing and normal tracking system operations. The NFOV Camera views an image of the target star as seen by the main optical subsystem (the sparse primary and secondary). This image will appear as elliptical because of the sparse nature of the primary.
After initial detection of the target star in the acquisition camera, centroid track errors are computed by the payload computer (using the acquisition video) and sent as pointing commands to the spacecraft controller. The spacecraft is then driven at low frequency (nominally 0.1 Hz) to center the target star in the acquisition FOY. Once the target is within a 20x20 pixel acquisition track gate (roughly 2 mrad), hand over to the WFOY camera is enabled and this sensor begins "looking" for the target (intensity above a threshold). When the target is detected in the WFOY, centroid track errors are computed (from the WFOY video) and used to close a nominal I-Hz coarse track loop using the galvo steering mirrors. The target will be centered in the WFOY focal plane, and offload commands from the steering mirrors to the spacecraft controller are initiated. As the WFOY track error becomes small (target within a 6x6 pixel track gate size), the star should become detectable on the NFOY focal plane. Collocation of the two cameras on the payload optical bench should ensure adequate alignment for this condition. When the target is detected in the NFOY, centroid errors are computed in the track processor using NFOY video, and a fine-track loop (nominally 3-Hz bandwidth) is closed using the galvo steering mirrors.
5. PROGRAM STATUS The SYS-Ied development team has been awarded a Phase II Small Business Innovative Research (SBIR) contract to proceed with the development of the LCSS Pointing Experiment at the U.S. Air Force Phillips Laboratory (PL) in Albuquerque, New Mexico. The program is being managed by the PL Space and Missiles Technologies Structures
Because it is very sparse, the telescope will be an inefficient collector of light, but radiometric 161
requirements that are placed on spaceborne optical telescope programs. The nature of the requirements placed on the LCSS optical system to measure the performance of the space borne structural system has offered an expanded opportunity to support a broader range of space-based missions. Included in these missions are astronomical measurements of closely spaced objects in space and surveillance of spacebased and ground-based objects that require highresolution capability. The demonstrations of the optical payload and structural concepts developed by the LCSS Pointing Experiment will lead to highpayoff risk reduction for future military and civil missions that require precision pointing.
Division. Assisting SVS in the development of the satellite system is Teledyne Brown Engineering of Huntsville. Work on the Phase 11 system has already begun in Albuquerque. The planned Phase Jl activities will continue through the Spring of 1995. The laboratory-based Phase 11 program will locus on the development and demonstration of the hreadboard payload at PL. Specific activities will include the demonstration of the necessary precision alignment and control, structural response studies to support the boom design and development, characterization of the optical system performance, spacecraft hierarchical controls demonstrations, and the development of a full-scale mock-up to assist in the Phase III design and flight system development.
7.
REFERENCES
6. CONCLUSION Seltzer, S.M . (1993). "Low Cost Space Structure Pointing Experiment," Final Report PL-TR-933023 prepared for USAF Phillips Laboratory Structures and Controls Directorate, Edwards AFB, CA.
The LCSS Pointing Experiment offers a costeffective approach to addressing several of the key issues that face the advanced space structures and optical systems developers. By using off-the-shelf components, along with a few representative critical "sparse" elements of the generally complex and cost! y payload system, LCSS can demonstrate that SIgnificant elements of acquisition, tracking, and pOInting can be shown to meet the stressing
Waites, H.B. and H.E. Worley (1987) "Large Space Structures Testing", AAS Paper 87-036, Proceeedings of AAS Annual Guidance and Control Conference, Keystone, CO.
162