The NASA Spaceshuttle (STS) will in 1983 have the ESA Spacelab onboard, Which will carry on one of its pallets a three, axes gimballed platform called Instrument Pointing System (IPS). The IPS will serve as a prcgrammable pointing system for the scientific user with a need for precise pointing of an instrument at a stellar Object or at Earth. The pointing error will be of an order of magnitude of 2 to 15 arc sec in function of disturbances, payload properties, and DS position in the Orbiter. The user of IPS will have access at the installation interface to electrical power and to the data and command processing facilities of Spacelab (SL).

The paper presents a brief introduction to the pointing problems associated with the European part of the First Spacelab Payload. Of ten experiments which require capabilities beyond mere shuttle positioning at least two, the Grille spectrometer and the microwave remote sensing experiment require extensive on-board pointing calculations. These and several other European experiments make use of a twin head horizon sensor. The paper considers the scientific requirements for accurate pointing and discusses how these objectives can be achieved. Typical error budgets are considered, and the requirements for on-board pointing computations including crew interaction are developed. The impact of uncertainties in Shuttle and Spacelab performances is explained.

Thirteen experiments have been selected for development for the multidisciplinary Spacelab 2 mission. They include three solar physics instruments and an atmospheric physics investigation mounted on the Instrument Pointing System (IPS), a two-axis gimballed X-ray experiment, an infrared telescope scanned about a single axis, and several fixed instruments. Integration of the IPS with its instruments and integration of the entire experiment complement into a flight mission has been a most demanding, complex task which has evolved as the Shuttle carrier vehicle, Spacelab, the IPS, and the experiments have matured. Mechanical interactions between all of these elements, thermal design restrictions, Shuttle and IPS operational constraints, mission duration, pointed instrument collision avoidance, and many other factors have impacted the scientific and engineering objectives of all of the experi-ments. Not all of the impacts have been adverse, however, and four types of joint science between previously unrelated experiments have so far been generated. This paper summarizes the current state of mission planning, particularly as it relates to the achievement of the scientific and engineering objectives of the pointed experiments. It discusses some of the compromises that have been made to satisfy the diverse requirements and capabilities of the hardware, the planning to date for joint solar observations from the IPS, and some system problems which remain.

The extensive participation of the Shuttle crewmembers (Commander, Pilot, Mission Specialist, and Payload Specialist) in Shuttle/experiment pointing is required to accomplish and enhance payload pointing operations and to optimize associated data collection. For payloads requiring extensive pointing operations, the degree of mission success which can be achieved by the involvement of the crew can not be matched by a combination of remote ground operations and automation. The importance of the use of the crew's manipulative skills, scientific judgment, and flexibility in the alteration of preplanned time lines was clearly demonstrated during the Skylab Apollo Telescope Mount solar experiment operations. The payload pointing operations associated with the Spacelab Instrument Pointing System and the Annular Suspension and Pointing System Gimbal System will also extensively use these crew capabilities. Crew operations will also play a potentially important role in the use of advanced pointing systems to accurately point experiments which have an extensive interaction with the surrounding environment, such as laser experiments.

Although the United States Air Force has been conducting space missions for the last 20 years, it has lagged behind on formulation of military space doctrines. Only in recent years have the military missions of force enhancement, space defense, and space support mission been hammered out thereby giving 20 years of activity legitimacy. To a great extent, the implementation of these military missions in space depends on the possession of space systems which are the product of a superior technology base. This base, however, must be nurtured with continuing support in order to retain the technology advantage. The decade of the 80's promises much greater progress, both in terms of military space missions as well as expanded research and development activities.

The Laboratory for Astronomy and Solar Physics at the Goddard Space Flight Center currently has two astronomy experiments approved for flight on Shuttle. These experiments will each produce on the order of 1.0 Ell bits of data per flight. Blind batch processing of the total data sets would far exceed the data processing budgets. Moreover, not all of the data frames will be useable. Therefore it will be necessary to carefully select those images or spectra that contain the most useful scientific results. This can be done with an interactive display and analysis system that lets the scientist view and react with his data in a timely fashion. This paper will describe the interactive Astronomical Data Analysis Facility, IADAF, of the Laboratory for Astronomy and Solar Physics and how each of our experimenters will use this facility in conjunction with his own data analysis program.

Infrared measurements of targets and backgrounds are made from five different platforms-ground, aircraft, balloons, rocket probes, and satellite. The advantages and disadvantages of each platform are discussed, and the measurement goals that would motivate the selection of one over another are assessed.

The demonstrated pointing, tracking and offsetting characteristics of the NASA 3-meter infrared telescope are reviewed. Other special features of the telescope such as low thermal background and versatile chopping secondary are also discussed. Experience is presented as to why these design features are important when making infrared observations of astronomical objects.

In early 1971, the Harvard Smithsonian Center for Astrophysics and the University of Arizona engaged in a cooperative program to develop a balloon-borne gryo-stabilized, 102-cm telescope capable of carrying out far-infrared observations of astronomical interest above the earth's atmosphere. Since 1972 the telescope has been flown and successfully recovered a total of sixteen times. Ten of the flights have provided high quality astronomical data resulting in more than eighty hours of photometric and spectroscopic observations of numerous astronomical objects. The telescope and its modes of operation are described, including the attitude control systems and the methods of aspect determination. A brief summary will be given of the in-frared instrumentation used. The experience gained from the operation of balloon-borne infrared telescopes and from development of new infrared instrumentation has been extremely valuable when applied to space experiments, particularly the infrared telescopes planned for operation on the Space Shuttle.

Computational techniques developed for the reduction and analysis of interferometer and radiometer data from the Balloon Altitude Mosaic Measurements (BAMM) program are presented. Emphasis is given to the reduction and analysis techniques developed for mosaic detector interferometer measurements. The presentation includes discussion of techniques to quality check, edit and simulate radiance values. Statistical analyses performed on simulated radiance data is described and examples of raw and edited data and statistical parameters are given. A brief discussion of radiometer reduction and analysis techniques is presented which includes data editing; filtering and decimation; statistical parameter determination; autocovariance estimation; and Fourier Power Spectral estimation of radiance.

The photometric quality of film recorded imagery from electrographic detectors such as those proposed for use on the STARLAB and SPACE SCHMIDT telescopes and on the NRL-803 experiment 2 requires that special attention be given to the information-extraction (digitization) and data reduction and analysis techniques. These will be discussed in detail and illustrated by example. An initial attempt at ex post facto image motion compensation, which may be required for cases of inadequate pointing stabilization, and results of recent experiments in post-development autoradiographic enhancement of weakly exposed electrographic images will be described.

The Multi-Anode Microchannel Arrays (MAMAs) are a family of photoelectric, photon-counting array detectors that have been developed and qualified specifically for use in space. MAMA detectors with formats as large as 256 x 1024 pixels are now in use or under construction for a variety of imaging and tracking applications. These photo-emissive detectors can be operated in a windowless configuration at extreme ultraviolet and soft x-ray wavelengths or in a sealed configuration at ultraviolet and visible wavelengths. The construction and modes-of-operation of the MAMA detectors are briefly described and the scientific objectives of a number of sounding rocket and Space Shuttle instruments utilizing these detectors are outlined. Performance characteristics of the MAMA detectors that are of fundamental importances for operation in the Space Shuttle environment are described and compared with those of the photo-conductive array detectors such as the CCDs and CIDs.

Skylab's Apollo Telescope Mount (ATM) was this country's first full scale manned astronomical observatory in space. It was a predecessor to the currently proposed Spacelab experiment observational systems. The ATM experiment canister, which housed the eight principal solar telescopes, was three m long, over two m in diameter, and weighed approximately 11,200 kilograms. The system was designed to operate within a pointing accuracy of ± 2.5 arc-sec and within a stability accuracy of ± 2.5 arc-sec over a 15-minute observation period. Review of photographic data indicated that the Experiment Pointing Control System (EPCS), which controlled the ATM experiment canister, performed with a stability of better than 1.0 arc seconds, exceeding its design requirement by a factor of 2.5. The EPCS was a part of the Skylab Attitude and Pointing Control System (APCS) which utilized momentum exchange devices and cold gas thrusters for overall vehicle control. The EPCS operated in a nested configuration within the vehicle control system. It was designed to isolate the ATM experiments from vehicle disturbances in two axes, up/down and left/right. Roll positioning control was also available. A fine sun sensor (FSS) provided position information while rate information was obtained from canister mounted rate gyros. Torque motors on each of the dynamic axes responded to controller signals to position the experiment canister.

The Instrument Pointing Subsystem (IPS), currently under development at Dornier System, is a most versatile Spacelab subsystem providing precision pointing capabilities to any single or clustered group of scientific instruments observing inertially fixed or moving targets. The IPS comprises a three-axis gimbal system mounted to the payload aft end and a payload clamping assembly for support of the IPS mounted experiments during Orbiter launch and landing phases. The IPS control system is based on the inertial reference of a three-axis gyro package being updated by the payload mounted IPS star/sun trackers and operated in a gimbal mounted minicomputer. It enables inertial stabilization as well as slewing and target tracking operations. The functional and operational control is performed via the Spacelab S/S computer and its display and keyboard, thereby establishing a flexible and responsive operator interface in the Spacelab Module and the Orbiter AFD enabling optimized display and keyboard entries for any operational mode. The AFD accommodates in addition a hardwired panel serving for IPS emergency operations. Incorporated Experiment supporting services consist of several independent power sources, three Experiment Computer Remote Acquisition Units and direct data links to IPS and Spacelab which provide maximum flexibility for operation of independent experiment clusters and for an operational interlock between IPS and its payload.

The Spacelab Instrument Pointing Subsystem (IPS), currently under development for the European Space Agency by Dornier System, is designed as a facility-type pointing system. The IPS will be used for experiments of a mass of up to 2000kg with pointing requirements, which exceed the pointing capabilities of the Shuttle Orbiter. The nointing performance of the IPS will be in the arc-second region. In order to satisfy the various payload needs for nointing quite a number of pointing options, e.g. scanning, have been designed into the IPS software and hardware. This paper focusses on the operational capabilities of the IPS in a manner the experimenter needs to understand them, in order to utilise the IPS for his experimental work most beneficially. The operational modes will be explained, like the stellar, solar, and earth pointing mode together with their options for acquisition and pointing, experiment related pointing, manoeuvres, and contingency operations. Further, operator and user related services like error messages, telemetry and uplink commands will be detailed.

Sperry Flight Systems is developing a high stability space shuttle pointing system for NASA, Marshall Space Flight Center with potential application for the Air Force Sortie Support Experiment Orientation Subsystem. This shuttle attached gimballed system will provide high stability pointing of electro-optical experiments mounted upon it, even in the presence of shuttle disturbances such as rocket firings. Presented in this paper is a description of the system configuration and requirements, photographs of the prototype hardware, system performance predictions and a description of system simulation and test support facilities.

A need has been identified for a payload auxilliary pointing system onboard the Space Shuttle which provides sub-arcsecond stability in the Shuttle disturbance environment. This paper describes such a pointing system, presently being developed by Sperry Flight Systems for the NASA Langley Research Center. At the core of the design is a non-contacting magnetic suspension which provides a high degree of isolation between payload and carrier. Design concepts and control laws will be discussed. Also, test results from full-scale protoflight hardware and planned system refinements will be presented.

A fine pointing system has been under development which has linear isolation as well as angular control. This system uses simple flexure hindges whose flexure torques are cancelled. This allows very low un-controled natural frequencies. Relatively large motion is also achievable (± 5cm, ± 15 degrees). With inertially referenced control on both the linear and angular systems, nanoradian pointing stability is possible.

The need for a small pointing facility for Spacelab experiments weighing up to 200kg and calling for low or medium pointing stability became evident from the evaluation of surveys of potential European Spacelab payloads. The Position and Hold Mount (PHM) which covers the low stability range, would satisfy approximately 30% of all European experiments. The PHM concept combines the advantages of low development and flight-hardware fabrication cost with those of lowest user cost for this group of experiments. Due to its inherent growth capability and the experience gained during manufacture, testing and operation, it will be possible to improve the capabilities of the mount, to produce a stabilised pointing mount, at comparatively low cost. The development programme initiated by ESA for the PHM minimises development cost and schedule risk by relying on existing basic designs and on building and testing of hardware at an early stage in the development programme. The PHM performance and the basic operation capability of a stabilised pointing mount will be demonstrated early in 1981. The pointing and payload accommodation requirements of small European experiments, the design requirements for a PHM and its integration into Spacelab are analysed and the PHM hardware is described.

This paper presents an overview of the considerations which are important in Shuttle-attached pointing systems for electro-optical experiments. These considerations are described in terms of the mission characteristics which include the mission orbit, host vehicle, sensor, and target. The pointing system characteristics of interest include the pointing system configuration, tradeoffs, and interfaces. Key problem areas are discussed, and expected evolutionary pointing system improvements are outlined. The fundamental limitations which constrain the evolution of the pointing systems are also identified.

A development program is currently underway to produce a precision star sensor using imaging charge coupled device (ICCD) technology. The effort is the critical component development phase for the Air Force Multi-Mission Attitude Determination and Autonomous Navigation System (MADAN). A number of unique considerations have evolved in designing an arcsecond accuracy sensor around an ICCD detector. Three tiers of performance criteria are involved: at the spacecraft attitude determination system level, at the star sensor level, and at the detector level. Optimum attitude determination system performance involves a tradeoff between Kalman filter iteration time and sensor ICCD integration time. The ICCD star sensor lends itself to the use of a new approach in the functional interface between the attitude determination system and the sensor. At the sensor level image data processing tradeoffs are important for optimum sensor performance. These tradeoffs involve the sensor optic configuration, the optical point spread function (PSF) size and shape, the PSF position locator, and the microprocessor locator algorithm. Performance modelling of the sensor mandates the use of computer simulation programs. Five key performance parameters at the ICCD detector level are defined. ICCD error characteristics have also been isolated to five key parameters.

Fine guidance technology development for the Shuttle Infrared Telescope Facility (SIRTF) centers upon the use of a single multiple-star-tracking sensor to provide the position information necessary to produce three-axis attitude control signals for precision payload pointing. Development of this guidance technology is a joint effort of the Jet Propulsion Laboratory (JPL) and the Ames Research Center. The JPL effort, described in this paper, is concerned with the development of a Fine Guidance Sensor that employs a high-density charge-coupled imaging device for producing position information signals by using star fields. Multiple star position information produces 3-axis position error signals that are used to update inertial reference gyros. The sensor employs advanced position interpolation algorithms to enhance field-of-view resolution and to correct for optical aberrations inherent in spatially chopped star images resulting from the telescope's movable secondary mirror. Operation of the sensor is under the control of a high-performance microcomputer that provides both autonomy and flexibility in a guidance application. Details of the technology involved in the development of the Fine Guidance Sensor as well as performance data are described in the paper.

Programs in development for the near future will require pointing sensors with accuracies in the region of 1 arc-sec and sub-arc-sec stability. The Advanced X-Ray Astronomical Facility, Annular Suspension Pointing System, Shuttle Infra-Red Telescope Facility and Stereosat are representatives of these programs. This paper describes a Solid State Aspect Sensor which could support this pointing capability and the testing technology necessary to verify its performance. The sensor uses a Charge Coupled Device (CCD) array integrated with a high-performance optical assembly. Perkin-Elmer has developed in-house a solid state sensor which could be used for precision pointing and attitude estimation applications. A test facility has been established which can be used to evaluate CCD arrays and optical assemblies for their suitability and use in a star sensor. The detail design of a stellar sensor to meet 1 arc-sec pointing requirements is presented. Optical performance characteristics are discussed, as is the verification testing of a prototype optical assembly. Test data are provided. The optical assembly is a 253-mm focal length, F/1.5 catadioptric lens system. Discussion of the solid state focal plane includes focal plane requirements and centroiding algorithms, focal plane and CCD test facilities, representative test data from linear CCD arrays, and a candidate CCD array for this application. An integrated sensor test station is described. Important considerations in the development of this precision pointing sensor are also presented, along with the predicted sensor performance based upon computer simulations.

The accurate pointing of certain shuttle experiments can be achieved with the aid of an electro-optical attitude transfer system. By means of projected beams of collimated light, the angular attitude of such an experiment in one part of the Orbiter payload bay may be determined relative to a reference location elsewhere in the bay where the orientation is accurately known from a star tracker and/or MU. principles of attitude transfer by means of autocollimators and passive reflectors are reviewed, including reference to the evolution of electronic autocollimation techniques in the last 30 years. Possible equipment arrangements for shuttle configurations are suggested. Finally, an example of an hypothetical problem requiring three-axis attitude determination within the Orbiter bay, is examined in detail, with an indication of equipment requirements and of the accuracies attainable.

The atmospheric emissions photometric imaging experiment (AEPI) to be flown on Spacelab 1 is designed to study faint natural and artificial atmospheric emission phenomena. Optical emissions are imaged in the region 2150 Å to 7320 Å using a television system consisting of two optical channels, one wide-angle and one telephoto. The detection system is an image-enhanced SEC vidicon. A third optical channel images onto the photocathode of a microchannel plate photomultiplier tube that has 100 discrete anodes. Photons are counted for each discrete anode, providing a direct measure of the luminosity of an object viewed by the TV telephoto lens, albeit with low spatial resolution. The AEPI detector is mounted on a two-axis gimbal comprised of a Modified Apollo Telescope Mount Star Tracker (MAST), which provides experiment pointing over a ±40-deg x ±80-deg range, exclusive of restrictions due to the proximity of other experiments. The pointing stability is one arc minute with respect to the spacecraft coordinate system for an exposure of one second. The tracking capability is 3.5 deg/s with a stability of one arc minute. The detector and pointing system are located on the Spacelab pallet. The experiment is controlled by stored programs resident in the Dedicated Experiment Processor located in the Spacelab module.

The Spacelab-2 Small Helium-Cooled Infrared Telescope will be used to map extended astronomical sources of low surface brightness emission, to measure the Shuttle induced environment and to develop techniques for managing large volumes of superfluid helium in space. The instrument is an f/4 15.2-cm Herschelian telescope with ten photoconductor detectors in the focal plane. This paper describes the hardware and software aspects of the instrument with emphasis on mission operations. In particular, a description is given of the observing plan formulated to meet the scientific and engineering objectives, the scan drive system, the precautions in design and operation necessary to prevent the sun, moon, and earth from adversely affecting the observations, the implications of thruster firings, and the on-board experiment computer application software to control the scanning of the telescope and support on-board displays.

The High Resolution Telescope and Spectrograph (HRTS) instrument on Spacelab 2 has a stringent pointing stability duration requirement. Therefore an image drift compensation system has been developed which transmits data from solar sensors located within the HRTS to the Instrument Pointing System (IPS) control loop. Internal image drift compensation and a step-rastering capability have also been designed into the HRTS. Solar feature target selection is achieved in real time by using a manual pointing controller and an H-alpha TV slit display in closed loop with the payload specialist. The pointing stability requirement of the HRTS, based on the results of three rocket flights, is compared with the expected IPS performance.

The Grazing Incidence Solar Telescope (GRIST) described in this report is intended for flight in Spacelab and offers for the first time the combination of high spatial and spectral resolution in the XUV and EUV wavelength range. The telescope is a sector shaped Wolter type-2 paraboloid-hyperboloid mirror pair of 412 cm effective focal length, 280 cm2 aperture and 6 arcmin x 6 arcmin field of view. Spatial resolution is 1 arcsec (20 um in focal plane) defined as 50% of the image energy within this element. The telescooe may be operated in the wavelength range 9n-1700 Å. In combination with suitable focal Diane instruments a snectral resolution of better than 104 can be achieved.

The Naval Research Laboratory is developing a payload, planned for Shuttle sortie missions sponsored by the Space Test Program, for imaging and photometry of natural and artificial sources of far-ultraviolet radiation. The experiment objectives emphasize, and the instrumentation is optimized for, detection and measurement of faint, diffuse sources. The NRL-803 experiment package consists of two far-UV electrographic Schmidt cameras, a dedicated pointing platform, and a low light level TV camera for target acquisition. The dedicated platform allows the number of useable flight opportunities to be maximized, by minimizing interfaces with the Shuttle and requirements on Shuttle support systems, but allows maximum use of the available time in orbit. A fine stabilization system, based on the use of magnetic deflection for image motion compensation, is currently under development. The dedicated platform and fine stabilization systems can also be used by other small experiments in Shuttle sortie missions.

The radiance of the earth's atmosphere in the ultraviolet can be used for horizon sensing from space. In order to obtain the radiance data to assess this possibility, an experiment utilizing Space Test Program pointing equipment on shuttle is being implemented. The experiment Horizon Ultraviolet Program, (HUP), AFGL-801, will measure the limb profile brightness and its geophysical variability using the STP pointing system and the HUP experiment limb scanner to maximize the viewing opportunities. The best combination of these two scanning methods is to be sought. Scans across the horizon from a tangent altitude of about 250 km will be made with a vertical angular resolution goal of one milliradian. One requirement of the shuttle pointing system is to provide knowledge of the pointing direction equivalent to this resolution. A 20° range is to be scanned in about 10 seconds so that latitudinal variability can be evaluated. The six spectrometers covering the 1100-4000 Å range have many options for viewing including the use of selectable fixed wavelengths and the capability for wavelength scans. The feasibility of combining this relatively small experiment with other experiments on the same flight is discussed.

The objective of the Space Infrared (SIRE) Sensor program is to measure LWIR radiation of natural and man made sources in space. Measurements will be used to support development and operation of space based space surveillance systems. This paper describes the planned concept for operating SIRE as a non-deployed payload within the payload bay of the Orbiter. The operations concept is prefaced with an overview of the SIRE system addressing the sensor, space segment, ground segment and supporting elements of the Space Transportation System. This is followed by a description of operational concepts and data processing that will be used within the ground segment during flights to plan, command and evaluate SIRE operations. This responsive system provides for inflight evaluation of data and replanning of measurements as necessary to accomodate operational perturbations from the Orbiter and react to unexpected measurement results.

During the 19701s, Air Force Geophysics Laboratory field measurement programs identified several aurorally enhanced infrared emitters in the upper atmosphere including chemiluminescentinescent NO emission at 2.8 μm and vibraluminescent CO2 emission at 4.3 μm. Models generally predict a much. smoother background at 4.3 μm due to CO2 radiation trapping. To test these models, AFGL is developing an earth limb sensing experiment (ELIAS) to measure auroral structure simultaneously in three colors: 3914A, 2.8 μm and 4.3 μm. The cryogenic sensor employs three scanning five detector arrays, telescoped to intercept an 8 by 8 km scene at the earth limb with spatial and temporal resolutions of 1.5 km and 0.1 sec, respectively. A near term auroral rocket probe experiment is planned for October 1981 followed by two space shuttle flights in the mid 1980's. The current knowledge of auroral earth limb backgrounds, details of the ELIAS sensor, and the proposed measurement technique are presented.

In recent years, the Air Force Geophysics Laboratory has flown several rocketborne experiments with cryogenic interferometers to measure natural and induced infrared atmospheric emissions. AFGL is currently developing two separate shuttle payloads based on advanced versions of these rocket sensors mated to cryogenic high off-axis rejection telescopes. CIRRIS will have a spectral resolution capability of better than 1 cm-1 over the 4 - 25 μm region, and will be dedicated to measurements of infrared emissions from the earthlimb at altitudes from 30 - 300 km. CIRRIS data is expected to provide an assessment of the effects of the atmosphere on current and planned AF space systems and a comprehensive data base for atmospheric modelling. Additional CIRRIS objectives are to measure and assess the effects of shuttle contamination on other planned shuttle experiments and to obtain data on a large number of atmospheric trace species. Specifics of the CIRRIS instrument, measurement plan and capabilities are presented.

A low-cost single-axis pointing control system is currently being developed for the CIRRIS-82 Space Shuttle Program. This system is essentially self-contained and requires a minimum of interfacing with the host vehicle to perform the rudimentary step-hold maneuver patterns. A more advanced two-axis system is designed for future CIRRIS missions to perform more complex stare and scan maneuvers. Both systems use a microprocessor controller.

The Shuttle Infrared Telescope Facility (SIRTF) is a cryogenically-cooled, 1-m-class telescope that will be operated from the Space Shuttle as an observatory for infrared astronomy. Over the 2- to 200-μm band, SIRTF will be 100 to 1000 times more sensitive than existing infrared facilities. This paper discusses the scientific constraints on and the requirements for pointing and controlling SIRTF as well as several aspects of SIRTF orbital operations. The basic pointing requirement is for an rms stability of 0.25 arcsec, which is necessary to realize the full angular resolution of the 5-μm diffraction-limited SIRTF. Achieving this stability requires the use of hardware and software integral to SIRTF working interactively with the gyrostabilized Shuttle pointing-mount. The higher sensitivity of SIRTF, together with orbital and time constraints, puts a premium on rapid target acquisition and on efficient operational and observational procedures. Several possible acquisition modes are discussed, and the importance of source acquisition by maximizing the output of an infrared detector is emphasized. The ability of the pointing-mount to slew the telescope over a range of angles is discussed, including its capability for executing raster scans over limited areas. Both minimum-time scans and constant-rate scans, which may take significantly longer in some cases, are discussed.

The Shuttle Infrared Telescope Facility (SIRTF) is being designed as a 0.85 m cryogenically cooled telescope capable of a three order of magnitude improvement over currently available infrared instruments. The SIRTF requires that the image at the focal plane be stabilized to better than 0.25 arcseconds with an absolute accuracy of 1.0 arcsecond. Current pointing-mount performance simulations indicate that neither of these requirements can be met without additional stabilization. The SIRTF pointing and control system will utilize gyro outputs, star field position measurements from a focal plane fine guidance sensor, and a steerable secondary mirror to provide the necessary stabilization and pointing control. The charge coupled device fine guidance sensor tracks multiple stars simultaneously and, through the use of multistar processing algorithms in a high performance microcomputer, generates three-axis attitude errors and gyro-drift estimates to correct the pointing-mount gyros. A high-bandwidth feedforward loop, driven directly from the pointing-mount gyro package, controls the steering mirror in order to correct disturbances not compensated for by the pointing-mount control system. A prototype design for the SIRTF pointing and control system is described in detail. Performance analyses made using a digital simulation of the pointing and control system as well as experimental data obtained in laboratory and field test measurements are presented.

GIRL consists of a liquid helium cooled 50 cm telescope and four focal plane instruments dedicated to astronomical and aeronomical observations. These instruments, a detector array, a photometer/polarimeter, an Ebert-Fastie-spectrometer and a Michelson-Interferometer are under development. They make up an "infrared observatory" having high sensitivity and high spectral and spatial resolution. The full-size "thermal-model" of the GIRL cryostat has now been completed and is undergoing extensive tests. GIRL will be pointed by IPS. It is scheduled to fly on a Spacelab mission in 1986.

The Hopkins Ultraviolet Telescope (HUT) is a 90-cm f/2 prime focus telescope and spectrometer with a photon-counting microchannel plate detector designed to obtain moderate resolution (Δλ - 4 Å) spectrophotometry of faint (my ⪅ 17) objects in the far-ultraviolet spectral region (~900-1700 Å). The HUT will be mounted on the Spacelab Instrument Pointing System and will be operated by a Payload Specialist using a field acquisition TV system to assist in locating targets. The initial flight of the HUT, expected in 1984 or 1985, will be devoteg primarily to the study of quasars and active galactic nuclei, emphasizing the 900-1200 Å spectral region which is inaccessible with other existing or currently planned orbiting instruments.

This paper describes instrumentation for three astronomical investigations, planned or proposed for flight on Spacelab, which are concerned with moderate- to wide-field imagery of nonsolar astronomical objects. The first of these, the Goddard Space Flight Center's Ultraviolet Imaging Telescope (UIT), is based on a current sounding rocket instrument and has been accepted by NASA for development and tentative flight on Spacelab in the 1984-1985 time period. It has a 40 arc minute field of view and better than 2 arc second angular resolution. The second instrument, which has been subjected to a feasibility study by NASA, is an all-reflecting Schmidt telescope which has a 5° field of view and 1 to 2 arc sec resolution. The third instrument, Starlab, is based on a 1-meter-aperture telescope and also has been through a definition study. It is proposed to be a multi-instrument astronomical facility which may be initially tested in Spacelab missions, but whose long-term operation will be based on use of a longer-duration space platform. One of the major scientific instruments planned for Starlab is a direct-imaging camera having 0.5° field of view and 0.2 arc sec resolution. Because of the differing fields of view and angular resolutions of these three instruments, their pointing requirements are somewhat different. Problem areas include alignment of the instruments with external startrackers and with other instruments on the same pointing platform. For Starlab, an internal image motion compensation system is necessary to provide the required ± 0.03 arc sec image stability.

WINDSAT is a proposed future space-based global wind measuring system. A Shuttleborne experiment is proposed as a proof-of-principle demonstration before development of a full operational system. WINDSAT goals are to measure wind speed and direction to ± 1 m/s and ± 10 deg accuracy, respectively, over the entire earth from 0- to 20-km altitude with 1-km altitude resolution. The wind measuring instrument is a coherent lidar incorporating a pulsed CO2 TEA laser transmitter and a continuously scanning 1.25-m-diameter optical system. The laser fires at an 8-Hz prf and the optics performs a conical scan at 60° to nadir every 7 seconds. Each laser pulse is backscattered by aerosols in the earth's atmosphere. The original transmitted laser frequency is Doppler shifted since the aerosols are moving with the wind. The wind speed is measured by heterodyne detecting the backscattered return laser radiation and measuring the frequency shift. Each wind speed measurement must be repeated at a different look angle in order to determine wind direction. A special feature of combining a continuously rotating optical system with heterodyne detection is the requirement for active alignment or image motion compensation of the return radiation. Short-term pointing stability of 2 μrad and long-term pointing accuracy of 100 μrad is required for efficient detection and accurate wind mapping. A separate WINDSAT attitude determination and control system is required to meet these accuracies.

Wind measurement by coherent lidar imposes unique requirements on pointing and image motion compensation systems. Pointing must be accomplished with respect to an inertial frame of reference and image motion compensation must be accomplished at a pupil to maintain heterodyne efficiency. Pupil control considerations become a major optical system driver. Short-term image motion compensation accuracy of 2 microradians and absolute pointing accuracy of 100 microradians are achieved simultaneously with proper pupil control by a simple optical system. Separate servo control systems compensate for altitude and alignment perturbations and provide for fine correction of absolute pointing. S/C velocity leads to a 25 microradian offset in the absolute sample position, but the offsets for transmission and reception cancel, so the image motion compensation system need not account for velocity. Systematic attitude changes couple to the sample direction to determine the extent of image motion. A complete systems analysis will be presented.

This paper presents the results of a study to identify and define an evolutionary multi-user Lidar Instrument System to be flown aboard the Shuttle/SPacelab (Space Transportation System) starting in the mid-1980's. This Lidar System is to be capable of accommodating a wide range of scientific investigations in atmospheric science, weather and climate, and environmental quality. The range of Potential investigations is defined by, but not limited to, 26 experiment classes.

The advent of the "Space Shuttle Era" has spearheaded a new wave of thought in our approach to the exploitation of space. Through use of the Shuttle, science and applications payloads need no longer be abandoned at the scheduled end of a particular mission, or when struck with premature failure, or even when they simply become outdated through advancements in technology. Rather the option will now exist for on-orbit maintenance and/or recovery of the payload for potential reuse. The Shuttle itself can even serve as an operational base for the gathering of data. This will be accomplished primarily through the use of Spacelab and a multitude of "Spacelab Instruments," many of which are already being developed. Additionally, the Shuttle along with other members of the Space Transportation Systems. family, will allow the buildup of space structures which can be routinely maintained on-orbit, thereby allowing long-term technical and economic exploitation. One such structure being given increased consideration for use in Low Earth Orbit (LEO) is the "Space Platform." Such platforms are envisioned to have lifetimes of many years and to provide basic stability, various utilities, and on-orbit accessibility to a number of temporarily emplaced payloads. Some payloads, depending on the mission for which they are being flown, would operate from a few weeks or months to many years. This paper reports current planning efforts by NASA for these space platforms directed towards determining the technically most suitable concepts and the approaches which might be followed to evolve these platforms as a cost-effective extension of the Spacelab era.