The operation teams for the Infrared Astronomical Satellite (IRAS) included scientists from the IRAS International Science Team. The scientific decisions on an hour-to-hour basis, as well as the long-term strategic decisions, were made by science team members. The IRAS scientists were involved in the analysis of the instrument performance, the analysis of the quality of the data, the decision to reacquire data that was contaminated by radiation effects, the strategy for acquiring the survey data, and the process for using the telescope for additional observations, as well as the processing decisions required to ensure the publication of the final scientific products by end of flight operations plus one year. Early in the project, two science team members were selected to be responsible for the scientific operational decisions. One, located at the operations control center in England, was responsible for the scientific aspects of the satellite operations; the other, located at the scientific processing center in Pasadena, was responsible for the scientific aspects of the processing. These science team members were then responsible for approving the design and test of the tools to support their responsibilities and then, after launch, for using these tools in making their decisions. The ability of the project to generate the final science data products one year after the end of flight operations is due in a large measure to the active participation of the science team members in the operations. This paper presents a summary of the operational experiences gained from this scientific involvement.

The Hubble Space Telescope (herein identified as the Space Telescope) is the largest and most powerful optical and ultraviolet astronomical observatory to be operated in space. From its nominal 550-600 kilometer low Earth orbit, the Space Telescope will produce data including imagery of unequaled quality of galaxies, star systems, quasars and other objects of scientific interest. Responding to a set of demanding scientific requirements, the optical, attitude control, stabilization and other equally vital subsystems of the Space Telescope are sophisticated and represent a high level of design quality. This paper describes the systems engineering program established for the Space Telescope, a project managed by the Marshall Space Flight Center (MSFC) of the National Aeronautics and Space Administration (NASA). The planning and operational methods implemented by the systems engineering organization to ensure that the scientific requirements are satisfied, that the verification process is acceptable, and that the configuration is controlled are described.

The objective of the Hubble Space Telescope Program is to build and operate a high-performance, fully-instrumented and long-lived optical observatory in earth orbit. Key Program elements are (1) a logistic strategy of Shuttle launch/maintenance and TDRSS communications, (2) overall responsibility at the MSFC Project for developing both the flight and ground systems, and (3) scientific participation in ST development and responsibility after launch for conducting the observing program. The ST Project is a management structure with a budget and schedule for delivering an ST system capable of achieving its established scientific objectives. The Project distributes through contracts the responsibility for supplying particular components, but retains the responsibility for system-level integrity. Scientists have set the performance objectives and now support the Project during the development stage by providing (1) PI management of the five Science Instruments, (2) continuing translations of scientific requirements into engineering specifications, (3) oversight of designs and performance tests, and (4) specific expertise on technical concerns.

This paper provides a description of the Space Telescope (ST) Pointing System (PS) and an overview of the pointing operations and procedures that support celestial observations. It is aimed at indicating to the user-astronomer the ST pointing capabilities and the mechanics of insuring the presence of the correct star in an instrument aperture. The paper begins by summarizing the general requirements regarding image stability and pointing accuracy to which the ST is being designed. This is followed by a description of the PS, in which primary elements are rate gyros, star sensors, reaction wheels, and an on-board digital computer. The Fine Guidance System (FGS) is the star sensor that provides the precise, long term celestial reference and is given special emphasis, since its autonomous modes are so intimately tied to the acquisition process. A description of a nominal acquisition sequence is given, and some strategies for dealing with acquisition and pointing anomalies are discussed.

The advanced X-ray Astrophysics Facility (AXAF) is planned as a major space-born stellar x-ray observatory. It will be in the Space Telescope class, i.e., fifteen year lifetime maintainable in orbit, with orbitally replaceable instruments, all sky viewing capability, and with state-of-the-art high resolution x-ray imaging optics. The AXAF, along with the Hubble Space Telescope and the Very Large Array comprise the corner stone of U . S . astronomy capabilities through the end of this century . AXAF will weigh on the order of 20,000 to 25,000 pounds, will be launched by Space Shuttle on a dedicated mission, and will be serviced and reboosted on subsequent Shuttle missions. AXAF has undergone a Phase A conceptual definition at the Marshall Space Flight Center with assistance by the Smithsonian Astrophysical Observatory. It is now undergoing a two year long competitive Phase B effort by industry to provide a preliminary design and an overall project definition. TRW and Lockheed are conducting the definition activities. A parallel instrument definition activity will take place by the x-ray astronomy community as a result of an Announcement of Opportunity released in August of 1983. The observatory will accommodate several focal plane instruments and a smaller number of supporting non-focal plane instruments. This paper will describe the general configuration of the observatory based on the Phase A study results, and will describe the conceptual approach to instrument accommodation by means of a focal plane assembly. The instruments are viewed conceptually as being modular in design so that they may be orbitally replaceable with standard interfaces. Many issues need to be resolved with respect to standard instrument interfaces and modularity. As such, this area will constitute a major challenge during the Phase B definition, and will require significant and iterative information exchange and design activities among the industrial, government and astronomy community participants.

The Advanced X-ray Astrophysics Facility will be an X-ray observatory built around a large-area, high-resolution grazing incidence X-ray telescope. Designed to operate in space for at least 15 years, AXAF will be operated as a major national facility with the majority of the observing time set aside for guest investigators. The instruments that will detect the X-rays will be supplied to the observatory by the scientific community. A discussion of the categories of instrumentation and the interface requirements that they will most likely impose is presented.

The performance of X-ray optics at the present time is limited primarily by fabrication errors rather than by design limitations, and consequently realistic specifications and predictions must be based upon the available materials and technology as much as upon the theoretical surface equations. The AXAF baseline mirror design is used as an example of a solution to these problems. Fabrication and metrology techniques which are likely to be used for the AXAF mirrors are discussed, and estimates of final performance are made, although much more reliable estimates will be available after the completion of the Technology Mirror Assembly programs.

This paper highlights the Space Infrared Telescope Facility (SIRTF) program efforts in resolving the unique situation when the SIRTF makes the transition from a Shuttle sortie mission with liberal margins of user services to a free-flyer with a dedicated spacecraft capable of providing a modest but adequate level of services and a mission life of several years. SIRTF is a 1-m-class, cryogenically cooled, IR astronomical observatory. Since the SIRTF studies began in 1972, there have been several conceptual designs from which the familiar sortie design associated with SIRTF evolved. With the success of the Infrared Astronomy Satellite (IRAS), the SIRTF design was changed to the current, free-flying long-life astronomical observatory in the summer of 1983. The transition from a sortie mission to a free-flying, long-life observatory required a basic change: a dedicated spacecraft now provides services previously provided by the Shuttle and Spacelab. These services are (1) pointing and control, (2) communications, command, and data handling, and (3) electrical power. In addition, propulsion from Shuttle orbits to the observatory working orbit and back is required. This paper covers several alternative SIRTF designs and the development of a conceptual baseline spacecraft. These evaluations were conducted with user inputs to arrive at the level of service required of the spacecraft in power, data handling, and control. For example, details such as the need to moderate the data-handling requirements with the transition from sortie to free-flyer are covered. The data-handling capability of the Shuttle/ Spacelab is a very liberal 101° bits of storage capability and an extremely fast 300-Mb/sec communication rate. The spacecraft cannot provide these capabilities without major new com-ponent development -- a costly endeavor and one to be avoided unless justified by the user community. Also, preliminary plans considering on-orbit repair and refurbishment of the spacecraft are covered.

Proteus is a system for flying Explorer scientific experiments on two or three spacecraft that never return to Earth. Experiments are integrated into an Instrument Payload that is delivered to orbit in a Space Shuttle and mated to a Proteus spacecraft with the Shuttle's Remote Manipulator System (RMS). The Proteus system includes ground support, development tools and communications as well as the orbiting satellites. The goal of Proteus is to triple the number of Explorer missions within the current budget.

The Solar Maximum Repair Mission was launched on April 6, 1984. The mission was made possible through a spacecraft designed for on-orbit repair, the availability of spacecraft repair hardware, and the availability of the Shuttle System. Repairs resulted in the return of significant new science data and a demonstration of the capability for on-orbit repair of spacecraft. Considerations of the capability to maintain and repair spacecraft on-orbit for the future will also be discussed.

The Solar Maximum Mission (SMM) satellite, which lost fine pointing attitude control early in its orbital life, was successfully repaired in orbit a little more than four years after its original launch. This successful repair removed the subject of satellite servic-ing from the realm of discussion to that of accomplished fact and, in the opinion of the authors, presages an era during which servicing, repair and up-grading of satellites utilizing the Orbiter and, later, the Space Station will be routine.

Ladies And Gentlemen, Dr. Speer, And All Attendees At The Society Of Photo-Optical Instrumentation Engineers Conference, It Is A Pleasure To Speak To You This Afternoon On The Subject Of The Space Shuttle And The Recently Completed Solar Maximum Repair Mission.

NASA's Space Telescope (ST), the first free flying spacecraft designed for on-orbit maintenance utilizing the full Shuttle capabilities, is a new concept in space-borne astronomical observatories. Long operational life time is achieved by bringing maintenance crews to the orbital worksite to make repairs and/or replace scientific insi-_rumr,nts. When major over-hauls are required, the Shuttle can return the telescope to earth and later relaunch the refurbished telescope. These capabilities will enable Space Telescope to operate at the forefront of astronomical research for a decade or more.

The large deployable reflector (LDR) is a proposed orbital astrophysical facility of substantial aperture (about 20 m in diameter) designed to obtain astronomical observations at infrared and submillimeter wavelengths. LDR has been strongly recommended by the Astronomy Survey Committee of the National Academy of Science (the Field Committee) for implementation in the mid to late 1990's. The technology to implement such a program within acceptable cost and risk constraints does not currently exist. NASA is now in the planning stages for a 5 year technology development program that will allow the requisite technology to be developed to an adequate level to enable LDR to be implemented. A LDR workshop was held in June 1982, at the Asilomar Conference Center in Pacific Grove, California. A group of nearly 100 scientists, engineers, and technologists from government, industry, and academia convened at Asilomar for one week to look at the fundamental questions concerning LDR. The scientists looked at the scientific rationale for LDR in order to reach a consensus agreement on what scientific observations LDR should undertake, and then having done this, to develop the scientific requirements for LDR such as angular resolution, field of view, etc. The technologists and engineers then used these science requirements to produce system and subsystem requirements. These telescope requirements were then compared to the current and projected state of technology development in each area. The two groups iterated until a set of consensus science and telescope requirements were arrived at that could be implemented in the 1990's. This paper will discuss the activities that have taken place since the Asilomar Workshop and specifically discuss current plans that NASA has to develop this enabling technology to support LDR. For programs as large and ambitious as LDR, it is necessary to realistically assess the state of the applicable generic technology growth, both civilian and military, and to augment it as needed. Assessing the current technology needs of LDR and comparing these needs to the projected generic technology development is a key goal in the near term LDR activity. The current and projected state of the key LDR technologies will be described as well as how shortfalls in technology development will be satisfied.

Astigmatism and field curvature, the predominent residual aberrations in a Wolter type grazing incidence telescope, can be significantly reduced by shortening the length of the mirrors. Performance improvements as well as advantages regarding alignment of a short versus a long grazing incidence telescope are discussed.

Very Long Baseline Interferometry (VLBI) affords astronomers their highest-resolution look at the universe. By combining the outputs of two radio astronomy antennas thousands of kilometers apart to form a Michelson interferometer, angular resolution of 1/1000 of an arcsecond or better is obtained. Current VLBI observations are limited in resolution by the largest antenna-antenna separation available (one Earth diameter), and in their information content by small number of antennas in use at one time. Extension of VLBI technique to include one or more antennas in space will relieve both of these constraints, and result in marked improvement in our ability to map distant radio sources at the highest resolution. The longer baselines which result between the space antenna and the ground antennas will substantially improve the angular resolution. The rapid movement of the space antenna around the Earth yields a wide-ranging and rapidly-changing space antenna-ground antenna separation, resulting in a much more complete map of each source. Four possible ways to extend VLBI into space include: (1) a Shuttle-attached mission, perhaps associated with the NASA advanced-technology program in large space antennas; (ii) a near-Earth orbit mission of six months to one year duration, based on a space platform associated with a Space Station; (iii) a larger-orbit free-flyer (semi-major axis about 15,000 km), of two or more years duration, such as the QUASAT satellite now being considered by NASA and ESA; and (iv) lunar and/or deep-space orbits, to reach the limits of resolution set by interstellar scattering.

The feasibility of performing an on-orbit refueling (OOR) operation to extend the orbital lifetime of the NASA Goddard Space Flight Center (GSFC) Gamma Ray Observatory (GRO) was investigated by TRW, the GRO mission contractor. This study shows that an OOR capability could be integrated into the GRO operational design early in the Phase D development with only minor cost impact and no schedule impact. In this approach, the GRO OOR design would be developed to achieve operational compatibility with the JSC/STS-developed Orbital Refueling System.

In recent years, X-ray astronomy has become an integral component of solar and celestial observing programs. The development of soft X-ray imaging by grazing-incidence optics, such as those flown on Skylab and the Einstein Observatory, has established soft X-ray telescopes as a major observational tool. Above 10 keV, however there has not been the same high resolution imaging capability, since grazing-incidence optics do not function effectively at these energies.

This paper deals with the types of observations that might best be done on the Moon; it contains a summary of previous design efforts; some design considerations are set out which are based on a first iteration functional analysis; some possible development scenarios are outlined; and costs and essential support areas are discussed. Although the creation of such an observatory is technically feasible (i.e., we know how to do it), the observatory won't be built as a stand-alone research facility, but it could be developed as an ancillary facility or outpost to an industrial facility. This conclusion is based on costs. Such an undertaking could also be a vehicle for fostering international cooperation.

The special requirements imposed by planetary missions require that the spacecraft which are designed to conduct them have certain unique properties as platforms for optical instruments. This paper will provide a general overview of these properties and discuss in some detail a current state-of-the-art planetary platform, the Galileo planetary spacecraft.

A precision pointing scan platform is under development at JPL in connection with the Mariner Mark II (MMII) interplanetary spacecraft program. Unlike previous designs, e.g., the Voyager and Galileo scan platforms, the new platform concept provides for closed loop celestial or target body referenced tracking and momentum compensated articulation. The new platform concept, known as the Integrated Platform Pointing and Attitude Control System (IPPACS), is expected to maximize mission to mission inheritance of attitude and articulation control hardware and software by decoupling the platform design from the spacecraft design through momentum compensation. High accuracy science instrument pointing is provided by an optical tracker co-located with an inertial rate unit on the instrument platform.

Considerations fundamental to the design of optical systems for interplanetary platforms are discussed with regard to the degree that they require departures from ground-based optical technology. Spectral availability, radiation and thermal cycles, microgravity, and space transportation constraints are considered, and new optical technologies are reviewed for their potential impact on the planetary missions planned for the rest of this century.

In recent years there has been an increasing interest in the imaging spectrometer concept, in which imaging is accomplished in multiple, contiguous spectral bands at typical intervals of 5 to 20 nm. There are two implementations of this concept under consideration for upcoming planetary missions. One is the scanning, or "whisk-broom" approach, in which each picture element (pixel)of the scene is spectrally dispersed onto a linear array of detectors; the spatial information is provided by a scan mirror in combination with the vehicle motion. The second approach is the "push-broom" imager, in which a line of pixels from the scene is spectrally dispersed onto a two-dimensional (area-array) detector. In this approach, the scan mirror is eliminated, but the optics and focal plane are more complex. This paper will discuss the application of these emerging instrument concepts to the planetary program. Key issues are the trade-off between the two types of imaging spectrometer, the available data rate from a typical planetary mission, and the focal-plane cooling requirements. Specific straw-man conceptual designs for the Mars Geoscience/Climatoloqy Orbiter (MGCO) and the Mariner Mark II Comet Rendezvous/Asteroid Flyby (CRAF) missions will be discussed.

The slow-scan television camera being built for NASA's Galileo Jupiter Orbiter consists of a 1500-mm-focal-length telescope coupled to a camera head housing a newly developed 800-X-800-element charge-coupled device (CCD) detector based on "virtual-phase" charge transfer technology. This detector provides broadband sensitivity over 100 times that of a comparable vidicon-tube camera while also yielding improved resolution, linearity, geometric fidelity, and spectral range. The system noise floor is 30 electrons, which results in a dynamic range of about 3500. The noise floor is limited by the production of small amounts of unwanted charge within the detector due to clocking the gate voltages during image read-out. Saturation of the detector with 9000-A light, followed by a high-speed erasure cycle prior to exposing each image, stabilizes the detector quantum efficiency at its maximum level for wavelengths beyond 7000A. In the near-Jovian radiation belts, interactions of high-energy particles with the silicon CCD result in the production of unwanted charge. Special techniques have been implemented (e.g., tantalum and quartz shielding, rapid image readout, and 2 x 2 picture-element on-chip averaging) to ensure adequate signal-to-noise performance for images acquired as close to Jupiter as five planetary radii. Instabilities in the inertially pointed scan platform on which the camera is mounted will at times limit image resolution by introducing smear into the pictures.

The emerging Space Station program will provide a significant expansion in the use of specialists in orbit with optics, for the furtherance of science, civil applications and technology. The U.S. Skylab and the U.S.S.R. Salyut activities, past and present, included a broad variety of optical payloads. Space Station concepts now underway will provide interior facilities for controls and data analysis, and, exterior capabilities for instrument installations, free-flyer retrieval and servicing.

The Co-Orbiting Platform is a major element in NASA's new initiative to develop and deploy a permanent space station in the early 1990's. This separate, unmanned free flying space platform supports a variety of space payloads with standard orbital services including: mechanical attachment, pointing control, electrical power, thermal control and data communications. It enhances the overall space station architecture by offering the more controlled space environments and additional mission flexibility supplied by unmanned free flying vehicles. The co-orbiting space platform additionally offers the servicing and transportation economy made possible by the nearby space station. This paper reviews the general features of space platforms drawing on previous space platform and space station studies conducted by NASA. Co-orbital characteristics and operations are reviewed. Plat-form services, particularly those pertinent to optics payloads are discussed; these include viewing access; pointing stability and control; contamination environmental control; electrical power services; thermal control; and data handling, storage and two-way communications.

This paper describes the requirements that the polar platform will have to meet if it is to satisfy the needs of the scientific earth observation communities in the early 1990's. The Eos is a prime mission for the polar platform. The polar platform will be a large satellite in a sun synchronous orbit that could be built up from the same components that are to be used to build the Space Station and its co-orbiting freeflying satellite. It will be serviced from the Shuttle so that it will be kept in operation and up-to-date for at least a decade. An initial set of scientific requirements has been documented by a Science and Missions Requirements Working Group (S&MRWG). These scientific requirements have been translated into derived technical requirements for the polar platform. These derived requirements are presented.

Earth remote sensing requirements that may be expected for Space Station are presented. They include research requirements for more detailed spectral resolution, atmospheric effects on information extraction, and much more thorough probing of irradiance angle -view angle information content in scene radiance than has been done to date. Instrument development requirements for unmanned commercial and governmental satellite systems are described. Requirements posed by NASA global habitability research efforts and Earth data systems developments such as the Pilot Oceans Data System, the Pilot Climate Data System, and the Pilot Land Data System are considered with a view to the role of Earth satellites of many kinds as aids to better understanding of the planet and its natural systems.

The orbital servicing concepts developed for several existing and planned spacecraft are reviewed to illustrate the evolution and capabilities of Shuttle-based support services, and the potential range of future services are categorized by several criteria including transportation systems requirements, supporting facilities and equipment options, and servicing locations. Implications of Space Station-basing the services are described, along with possible implementing systems. In particular, the technique of "formation flying" by means of which spacecraft may be physically isolated from each other while maintaining reasonable propellant requirements for maneuvers between them is explored by representative examples. OTV mission support is described as an evolutionary augmentation of the basic Space Station services.

By the year 1992, the first permanently-manned U.S. Space Station will be entering its operational phase. The development of this Space Station is the next step in the evolutionary Space Transportation System concept. Initially, it will be composed of an 8-man Space Station in low inclination and low earth orbit. By the mid-1990's, the Space Station will have grown to a 12- to 16-man on-orbit facility with Station-tended platforms at low and high inclinations, and Station-based upper stage capability. This system will offer a unique capability for research and development through extensive, long-duration observations and experimentation. In addition, individual users can take advantage of the myriad of technological breakthroughs achieved in other user areas (e.g., upper stage placement techniques, data handling routines, satellite servicing methodologies, etc.). Payloads delivered to the Space Station have their own specific requirements, capabilities and limitations. Electro-optical payloads are no exception. This paper discusses the Space Station and its potential benefits to electro-optical payloads. Prospective outlooks for research and development projects on the Space Station are made for two electro-optical payload categories; i.e., astrophysical and geophysical. Examples of how the Space Station facilitates research and development for electro-optical payloads located both at the Space Station and as free-flyers are given. In conclusion, cost summaries are established for utilizing the Space Station. This includes a) analysis of the Space Station as a transportation node, b) user servicing and reconfiguration costs, and c) trade-offs between Space Station attached versus free-flying payloads.

Landsats 4 and 5, the latest in the series of unmanned earth observation satellites, are used as the space platform for two remote sensing, mechanical scanning instruments: the multispectral scanner (MSS) and the recently developed thematic mapper (TM). The primary objective of the experimental portion of the Landsat 4 and 5 missions is to assess the capability of the TM to provide improved information relative to the MSS. The higher spatial resolution of the TM over the MSS requires a higher degree of flight segment attitude stability than the earlier Landsats; therefore, a more stable, low-orbit space platform must be provided. This paper describes the orbital, electrical, mechanical, and thermal chihracteristics of Landsat 4 and 5 flight segment with special emphasis on the TM and MSS interfaces. Also described are flight segment disturbances caused by the TM and MSS scanning mirrors, motion from the Tracking and Data Relay Satellite (TDRS) antenna, solar array, and the attitude control system (ACS).

A new type of blackbody simulator design is described that is based on the principle that the projected solid angle of the aperture is constant when viewed from all points on the cavity wall surface. This design provides compact blackbody simulators with excellent spatial uniformity, and they are especially well suited for the calibration of space platform infrared detector arrays.

This paper presents the technical requirements and system design approach for an optical mounting platform used on the RCA built and flown TIROS Meteorological Satellite Series. This satellite series requires precise knowledge and control of the sensing axis for the meteorological sensors to ensure accurate weather predictions. The primary structural ele-ment used to mount the sensors and satellite attitude determination and control components is a precision optical platform. This optical platform is required to support the sensors during the satellite launch and orbit environments, to minimize the thermal gradient induced misalignments, and to provide clear fields-of-view for sensor scanning and thermal control. In addition, the optical platform was designed for minimum weight, manufacturability, and flexibility for redesign (sensor complements are not always the same from one satellite to another). The TIROS platform is a unique design which meets all these requirements, while weighing 52 pounds and supporting 230 pounds of sensors, harness, and thermal control components. Testing has demonstrated that this platform has maintained co-alignment of sensors during sine and acoustic vibration. The thermal control has been demonstrated in thermal vacuum testing and during on orbit performance. This platform has flown on four space-craft and successfully met all mission requirements.

Providing timely global weather information, including advanced warning of developing storms, is the primary function of the United States Geostationary Operational Environmental Satellite (GOES) meteorological program. To provide more complete data, two new GOES satellites are under construction. Known as GOES GH, the synchronous satellites will be operated by the National Oceanic and Atmospheric Administration (NOAA) in support of the National Weather Service and other United States and international users. Hughes Aircraft Company is designing and building GOES GH under the technical direction of the Goddard Space Flight Center of NASA. The GOES GH design uses improved technology and has greater capability than previous GOES satellites. The new satellite directly transmits day and night observations of global scale weather, hurricanes, and other more localized severe storms. It also relays processed, high resolution observation data along with weather facsimile (WEFAX) data. Weather observations are generated in the Visible and Infrared Spin Scan Radiometer (VISSR) Atmospheric Sounder (VAS). The VAS performs visible and infrared (IR) imaging as well as multispectral imaging and temperature sounding of the atmosphere. The optical scanner on the 100 rpm spinning portion of the spacecraft records the visual imagery with a resolution of 0.9 km (0.6 mile) and IR imagery with a resolution of 6.9 km (4.3 miles).

All-solid-state pushbroom sensors for Multispectral Linear Array (MLA) instruments are under development. It is planned that these pushbroom sensors will replace mechanical scanners used on current Landsat earth resources satellites, providing improved performance and operational flexibility. A buttable, four-spectral-band, linear-format charge coupled device (CCD) and a buttable, two-spectral-band, linear-format, shortwave infrared charge coupled. device (IRCCD) are being developed under NASA funding. These silicon integrated circuits may be butted end-to-end to provide multispectral focal planes with thousands of contiguous, in-line photosites. The visible CCD integrated circuit is organized as four linear arrays of 1024 pixels each. Each linear array views the scene in a different spectral window, resulting in a four-band sensor. The spectral windows are defined by integral bandpass filters. First-generation filters are interference stacks tuned to Thematic Mapper bands 1-4. The pixel center-to-center spacing of 15 pm combined with a band-to-band, along-track spacing of only 60 μm provides a compact, attractive focal plane organization. The high quantum efficiency of the backside-illuminated CCD technology provides excellent signal-to-noise performance from 0.4 μm to 0.9 μm. The backside-illuminated technology also results in unusually high MTF in the red. The shortwave infrared (SWIR) sensor is organized as two linear arrays of 512 detectors each. Each linear array is optimized for performance at a different wavelength in the SWIR band (1-3.0 μm). The actual spectral window of each band is defined by bandpass filters placed in close proximity to the chip. This dual-band infrared sensor consists of Schottky barrier detectors which are read out by CCD multiplexers. The detectors and the CCD registers are formed as one monolithic structure using standard silicon process technology. These IRCCD focal planes provide radiometric performance at 125°K. This operating temperature, and the low power dissipation of 18 μW per detector, make this sensor compatible with satellite passive cooling. The detector center-to-center spacing is 30 μm with a band-to-band spacing of 300 μm.

Next-generation sensors for earth observation will utilize Multispectral Linear Array (MLA) technology wherein linear detector arrays will scan the earth in pushbroom fashion. Compared to earlier sensors utilizing cross-track scan mirrors, MLA instruments will permit a dramatic increase in detector dwell time, which can be exploited to improve spatial, spectral and radiometric resolution. Focal-plane technology development is key to these new sensors, because detector size and performance are critical factors that influence the entire sensor design.

Recent advances in the application of lasers to remote sensing have prompted a series of feasibility studies to examine the potential for using a Doppler lidar to measure tropospheric winds around the globe. As presently envisioned, such a system would circle the earth tracing out a cycloidal pattern with its conically scanning laser beam. Each pulse, after undergoing scattering and absorption, would return to the sensor with information on the aerosol concentrations, atmospheric turbulence and the wind component along the laser beam's line-of-sight. The individual radial velocity measurements would then be combined to obtain an estimate of the horizontal u and v wind components for a specified volume of the earth's atmosphere. This paper addresses the feasibility of simulating such measurements from an airborne platform. It will briefly describe the changes in configuration required of the Marshall Space Flight Center's airborne Doppler lidar and will discuss the scalability of the meteorological phenomena to be measured.

Spacecraft glow may be defined as optical emissions originating immediately above those surfaces of an orbiting spacecraft that face into the ram direction. The glow is observed to extend from the spacecraft surface out about 20 cm and, for the Space Shuttle at its lower orbital altitudes, is bright enough to be seen by the unaided eye. The glow is brightest in the yellow and red end of the spectrum. Measurements show that the brightness rises rapidly toward the red end of the spectrum and is last seen still rising toward the infra-red (IR). Although there are no direct measurements, both extrapolation of available data as well as theoretical arguments indicate the glow is brightest in the infra-red.

The First Spacelab Mission, launched on November 28, 1983, included several investigations to study the space plasma environment from Shuttle/Spacelab. The prime emphasis of these investigations was centered on the performance of particle injections (electrons, plasmas, ions, and neutrals) from the Shuttle/Spacelab and studies of the ensuing effects on the orbiter, the near orbiter environment, and the earth's atmosphere. The results of these investigations were intended to provide information required to plan for future beam-particle, and beam-atmosphere investigation programs.

National Space Development Agency (NASDA) of Japan plans to conduct First Materials Processing Test (FMPT) with a Japanese Payload Specialist (PS) on Spacelab-J which will be launched on Jan.27, 1988. FMPT includes not only materials processing experiments but also life sciences experiments which are proposed by national institutes, universities, and others in Japan. NASDA is now developing on-board experimental system and equipments forward mission accomplishment,and continuing to select Japanese PS's suitable for FMPT program. The flight hardwares of FMPT are planned to be transported to NASA KSC by May, 1987 for level integrations, and on the other hand, Japanese PS's will be selected in Sep.,1985 to be trained in Japan and U.S. under the health care management untill the space flight for the first time as the Japanese.

Concern by top NASA management in late 1982 that the cost to accomplish Spacelab Payload Integration and Operations appeared excessive and not well understood, led to the initiation of the Spacelab Mission Implementation Cost Assessment (SMICA) study. SMICA was chartered to a "bottoms up study" to define an accurate cost model for a reference mission, and to develop an implementation plan for reducing these costs. All phases of this Spacelab mission were considered including payload mission management, experiment and mission peculiar equipment development, mission integration and ground and flight operations, and science/engineering data evaluation. Excluded were the functions and costs for the Shuttle, Spacelab Data Processing Facility, and the Tracking and Data Relay Satellite System (TDRSS). The study did establish a baseline mission cost for reference. The base-line mission payload included five new instruments and four reflight instruments. SMICA showed that a total savings of approximately 20% could be attained if the following were accomplished: 1. Compress the mission management and ground processing schedules. 2. Revise the approach to equipping, staffing, and operating the Payload Operations Control Center. 3. Change the methods of working with the experiment developers (science community). The operating philosophies and procedures recommended can serve as generic guidelines to other Spacelab mission/payload managers in reducing overall cost/manpower requirements. Attainment of maximum benefit from the assessment entails the addition of some risk, and this will be discussed briefly in the report.

The dedicated discipline laboratory (DDL) concept is a new approach for implementing Spacelab missions that involves the grouping of science instruments into mission complements of single or compatible disciplines. These complements are evolved in such a way that the DDL payloads can be left intact between flights. This requires the dedication of flight hardware to specific payloads on a long-term basis and raises the concern that the purchase of additional flight hardware will be required to implement the DDL program. However, the payoff is expected to result in significant savings in mission engineering and assembly effort. A study has been conducted recently to quantify both the requirements for new hardware and the projected mission cost savings. It was found that some incremental additions to the current inventory will be needed to fly the mission model assumed. Cost savings of $2M to 6.5M per mission were projected in areas analyzed in depth, and additional savings may occur in areas for which detailed cost data were not available.

The First Spacelab Mission began in the mid seventies as a joint NASA/ESA multidisciplinary scientific mission with the prime objective of verifying the Spacelab Design and performance and to demonstrate the Spacelab ability as a scientific laboratory. The planning for the mission established many new concepts for management and space experimentation and dealt with a whole era in space transportation. In reviewing the first mission, the management and technical approach for the overall mission design will be examined against the original concept. The highlights and accomplishments of the mission will be presented. The lessons learned and how they may be applicable to future Spacelab Missions will be discussed.

Spacelab 1 activities, using the Space-plasma Computer Analysis Network (SCAN), occurred because the network provided a solution to an immediate communication problem. For many investigators, SCAN is an ideal environment for the continued analysis of Spacelab data. SCAN currently provides access to other Spacelab investigators and co-investigators, to Spacelab and non-Spacelab correlative data bases, and to large Class VI computational facilities for modeling. SCAN accomplishes this by linking computers together at remote institutions that are used routinely by space science researchers. The networking software utilizes commercially available computer-to-computer communications. SCAN is a pilot project and a testbed for space science communication which, from the beginning, included the reseach user community in the planning, building and operation of the computer network. Expansion of the network within the next two years will include more than 12 institutions (many of whom are Spacelab investigators). This expansion includes the MSFC Mission Integration Planning System or MIPS which is responsible for generating Spacelab mission timeline and scheduling information for payload functional operations. In addition, Shuttle environment information needed for instrument development is being accumulated at GSFC and will be readily accessable over SCAN. Although the space plasma community is the prime user of SCAN, it is anticipated that along with other pilot systems (planned and operational) in the other space science fields, a unified NASA-wide communication system will ultimately develop from these pilot efforts.

The Space Transportation System (STS) now in the operational phase offers routine access to a unique laboratory environment that can be used to study microgravity science and applications phenomena in ways not possible on Earth. An overview of activities and results for some of the current space experiments and future research opportunities involving microgravity is presented. Topics include some of the observed reactions of the low gravity provided by the Space Shuttle on several materials undergoing processing. Primary emphasis is placed on the use of the microgravity environment to investigate a variety of processes with-out the complicating effects of containers, buoyancy-driven convection, sedimentation, and hydrostatic pressure. Experimentation in microgravity has given insight into the growth of crystals, uncovered some subtle interfacial mechanisms that produce phase separation in multiphase systems, and demonstrated dramatic improvements in free-flow electrophoresis and in growth of latex microspheres by seeded polymerization. Anticipated future research activities include measurements of critical phase transition phenomena with increased precision, and tests of various theories of nucleation, growth, and Ostwald ripening. Ground-based test facilities and planned space research facilities are briefly explained.

Spacelab and Space Shuttle are being utilized as an observing platform for science and applications instruments. The results obtained from these early flights, as well as those missions planned between now and 1990, will provide an excellent data base for the planning of Space Station observing instruments. Many instruments planned for use on Spacelab will be readily transitioned for flight on the Space Station. Some will carry their own pointing capabilities while others will be mounted to various gimbaling and high precision pointers. Our objectives over the next few years will be to evaluate the Shuttle and Spacelab as a science and applications observation platform and determine what similar systems or capabilities will be required to make the Space Station an effective observing platform.

The Space Shuttle Orbiter, in addition to serving as a transportation system, has proven to be a valuable test platform for science and applications space sensor systems. In low earth orbit the Shuttle can manuever into a variety of orientations for sensor pointing purposes. An astronaut, trained in the use, evaluation, and testing of the specific system, can make real-time adjustments to the system involved. In some cases, the astronaut could perform hands-on sensor repairs in the payload bay. Of extreme importance is the ability of the Shuttle to bring the sensor back to earth where laboratory assessments can continue sensor evaluation or establish cause of failure.

The need for space astrophysical observatories to support advances in basic knowledge of the universe is discussed. Observatories will require permanent presence in space for construction and assembly of large-aperture instruments and for servicing. Results are presented from a study of technology development for construction of a large-aperture instrument in space.

The introduction of passive electro-optical sensors on surface ships requires the development of a precision platform stabilization and control system which must operate in an above deck environment. This paper describes the platform control system requirements, presents performance models of the stabilization system and discusses the principal design tradeoffs. Measured performance data from a shipboard EO system is presented.

As part of Emerson's ongoing evaluation of multiple sensor platforms (both elevated and hull mounted) for land based vehicles, a packaged set of parametric programs, both deterministic and probabilistic in nature, has been developed to augment various designer modes. This set of programs has been developed under the Emerson IRAD Program, designated SSRI, Sensor Stabilization Requirements Investigation. This paper presents the overall objective of the package and describes some of the more mature modules within the package.

Automatic Fixed and Mobile Laser Range and Tracking Systems utilizing eye-safe gallium arsenide injection diode lasers provide improved range accuracies and angle position data of high and low speed ground and air targets. The system can utilize small cooperative targets to ranges of 7-10 Km and non-coopera-tive targets to ranges using low power, high pulse rate gallium arsenide transmitters in excess of 1.5 Km. The systems are eye-safe. The high data rate associated with pulse rates up to 4 Khz and improved accuracy measurements, 15 cm single shot, 2 cm averaged, permits improved tracking and velocity data of high speed targets. For applications requiring extreme single shot accuracy for high speed tracking, special 1/2 nanosecond counters can be utilized providing 8 cm accuracy on a single shot basis. The paper describes the trade-offs associated with the electro-optical design, such as beamwidth, field-of-view, pulse-rate, computer control, tracking and ranging requirements and range accuracy analysis.

This paper describes a family of airborne laser systems in use for terrain profiling, surveying, mapping, altimetry, collision avoidance and shipboard landing systems using fixed and rotary wing aircraft as the platforms. The laser altimeter has also been used in systems compatible with the Army T-16 and. T-22 carrier missiles (platform). Both pulsed gallium arsenide and Nd:YAG (neodymium-doped, yttrium-aluminum-garnet) laser rangefinders have been used for these applications. All of these systems use ACCI's advanced measurement techniques that permit range accuracies of 8 cm, single shot, 1 cm averaged, to be achieved. Pulse rates up to 4 Khz are employed for airborne profiling. This high data density rate provides 1 data point every 2" along the aircraft flight line at aircraft speed of 500 knots. Scanning modes for some applications are employed. Systems have been integrated with all current inertial navigation systems (Litton, Ferranti and Honeywell), as well as a number of microwave positioning systems. Removal of aircraft motion from the laser range measurements by use of an accelerometer is described. Flight data from a number of program performed by U.S. and Canadian Federal Agencies, in addition to those of commercial surveying and mapping companies are described.

The growing use of lasers in satellite tracking has generated a need to characterize laser damage to aerial photographic film. It would be prohibitively costly to determine damage regimes experimentally for all cases of interest. In this report, a computer model of laser damage to film is described.

This paper presents some of the design, construction, and operational issues associated with using airships (dirigibles or Zeppelins) as very large stationary platforms positioned at high altitudes. Innovative construction techniques and the application of robotics to those techniques offer the potential of cost reductions sufficient to make the concept economically feasible. Some of the operational concepts could also be adapted to small Remotely Piloted Vehicles (RPVs) or semi-autonomous vehicles.

A number of promising approaches have been investigated for controlling multiple adaptive optics elements in space-based wide field-of-view telescopes. However, the individual control algorithms, and sensing and control components have never been evaluated against one another in simulation or hardware. The Air Force is undertaking an inhouse program in the RADC Optical Systems Engineering Laboratory called OPTECAL for "OPtical TEstbed for Control ALgorithms". The OPTECAL program will combine extensive analyses with hardware tests using a state-of-the-art platform to validate this control, with emphasis on the algorithms. Individual control and sensing components and algorithms, as well as integrated control system approaches, will be evaluated on the platform. An effort was recently completed to identify the key parameters the platform should address and to provide a preliminary testbed design incorporating these parameters. The design completion and construction of the OPTECAL platform will begin late in 1984 and delivery is expected in early 1986. Preliminary control simulations and experiments are now underway at RADC in preparation for the future OPTECAL studies. Initial milestones include the establishment of a basic adaptive optics test setup, a proof-of-concept deconvolution experiment, development of an optical modeling capability, and studies evaluating "deconvolution" techniques, phase retrieval approaches, and the effects of actuator transfer function uncertainties in wide field-of-view optics control. Later milestones include fully traceable experiments on the OPTECAL platform to optimize the integrated control system performance.