ALBERT

Description: This paper addresses the design considerations and strategies for astrophysical observations as key elements of an international solar system exploration program. Emphasis is placed on the technical and programmatic challenges and opportunities associated with an evolving program of lunar-based astronomy. Both robotic and human tended facilities are discussed ranging from relatively small meter-class transit telescopes to large interferometer and filled-aperture systems.

Description: A method is presented by which measured modes and frequencies from a modal test can be used to determine the location and magnitude of damage in a space struss structure. The damage is located by computing the Euclidean distances between the measured mode shapes and the best achievable eigenvectors. The best achievable eigenvectors are the projection of the measured mode shapes onto the subspace defined by the refined analytical model of the structure and the measured frequencies. Loss of both stiffness and mass properties can be located and quantified. To examine the performance of the method when experimentally measured modes are employed, various damage detection studies using a laboratory eight-bay truss structure were conducted. The method performs well even though the measurement errors inevitably make the damage location more difficult.

Description: Following the project's first major design review, some unresolved technical issues, mainly centered on details of how to integrate Russian hardware into the U.S./international space station, remain. No 'show stoppers' were found in the review. Specific open technical issues are discussed in this article.

Description: A coupled, trajectory-based flowfield and material thermal-response analysis is presented for the European Space Agency proposed Rosetta comet nucleus sample return vehicle. The probe returns to earth along a hyperbolic trajectory with an entry velocity of 16.5 km/s and requires an ablative heat shield on the forebody. Combined radiative and convective ablating flowfield analyses were performed for the significant heating portion of the shallow ballistic entry trajectory. Both quasisteady ablation and fully transient analyses were performed for a heat shield composed of carbon-phenolic ablative material. Quasisteady analysis was performed using the two-dimensional axisymmetric codes RASLE and BLIMPK. Transient computational results were obtained from the one-dimensional ablation/conduction code CMA. Results are presented for heating, temperature, and ablation rate distributions over the probe forebody for various trajectory points. Comparison of transient and quasisteady results indicates that, for the heating pulse encountered by this probe, the quasisteady approach is conservative from the standpoint of predicted surface recession.

Description: Advanced Satellite for Cosmology and Astrophysics (ASCA) is a high-throughput X-ray astronomy observatory which is capable of simultaneous imaging and spectroscopic observations over a wide energy range 0.5-10 keV. The scientific capabilities of ASCA and some aspects related to its operation and observations are briefly described.

Description: We examine the electromagnetic (EM) bias by using retracked TOPEX altimeter data. In contrast to previous studies, we use a parameterization of the EM bias which does not make stringent assumptions about the form of the correction or its global behavior. We find that the most effective single parameter correction uses the altimeter-estimated wind speed but that other parameterizations, using a wave age related parameter of significant wave height, may also significantly reduce the repeat pass variance. The different corrections are compared, and their improvement of the TOPEX height variance is quantified.

Description: Monthly Ku band sigma(sub 0) and significant wave height (SWH) histograms from the NASA altimeter on the TOPEX/POSEIDON satellite are preseneted for January through June 1993 for three latitude bands between +/- 60 degrees. The data are compared to distributions from the Geosat mission for the same months in 1987-1989. Generally, the distributions agree quite well, although there are some seasonal/hemispherical differences. The sigma(sub 0) comparison reveals an overall bias between the two altimeters with the TOPEX sigma(sub 0) higher by about 0.7 dB, which is consistent with algorithm improvements for TOPEX. The SWH distributions show strong hemispherical/seasonal changes. The seasonal/hemispherical differences between TOPEX and Geosat are consistent for SWH and sigma(sub 0). The joint distribution of sigma(sub 0) and SWH is extremely stable friom month to month. The typical SWH is independent of sigma(sub 0) for sigma(sub 0) greater than 11.3 dB. The minimum SWH grows exponentially with wind speed. This joint distribution may be useful for understanding electromagnetic bias in altimeter measurements. Finally, altimeter data are compared to buoy values from 21 overflights of the NASA verification site near Pt. Conception, California. Wave heights agree well with an root mean square (RMS) difference of only 0.2 m. Altimeter sigma(sub 0) values are compared to buoy wind speeds. The results are consistent with the -0.7 dB sigma(sub 0) offset from the histogram comparisons.

Description: To assess the accuracy of the TOPEX altimeter data, we have reprocessed the raw altimeter waveform data using more sophisticated algorithms than those implemented in the altimeter hardware. We discuss systematic contamination of the waveform which we have observed and its effect on very long wavelength errors. We conclude that these systematic errors are responsible for a very long wavelength error whose peak-to-peak magnitude for the Ku band altimeter is of the order of 1 cm. We also examine the ability of retracked data to reduce the repeat pass variance and correct for significant wave height (SWH) and acceleration dependent errors. We find that the ground postprocessing contains SWH dependent biases which depend on the altimeter fine height correction.

Description: The routine ground processing of data from the NASA radar altimeter of TOPEX/POSEIDON includes instrument corrections for the effects of significant wave height and attitude angle changes on the altimeter's estimates of range, backscattered power, and significant wave height. This paper describes how these instrument corrections were generated and how they are applied. Detailed waveform fitting to telemetered waveform samples is use to assess the effectiveness of the corrections. There are several altimeter hardware-caused small waveform departures from the model waveforms and these departures, designated waveform 'features', are described in detailed. A consequence of the waveform features, and their positioning relationship to range rate, is that range data for ground tracks moving toward the equator may differ systematically by about a centimeter compared to range data for ground tracks moving away from the equator. The results and discussion are limited to side A of the redundant altimeter, as only side A has been operated on orbit.

Description: Results of the in-flight calibration and performance evaluation campaign for the TOPEX/POSEIDON microwave radiometer (TMR) are presented. Intercomparisons are made between TMR and various sources of ground truth, including ground-based microwave water vapor radiometers, radiosondes, global climatological models, special sensor microwave imager data over the Amazon rain forest, and models of clear, calm, subpolar ocean regions. After correction for preflight errors in the processing of thermal/vacuum data, relative channel offsets in the open ocean TMR brightness temperatures were noted at the approximately = 1 K level for the three TMR frequencies. Larger absolute offsets of 6-9 K over the rain forest indicated a approximately = 5% gain error in the three channel calibrations. This was corrected by adjusting the antenna pattern correction (APC) algorithm. AS 10% scale error in the TMR path delay estimates, relative to coincident radiosondes, was corrected in part by the APC adjustment and in part by a 5% modification to the value assumed for the 22.235 FGHz water vapor line strength in the path delay retrieval algorithm. After all in-flight corrections to the calibration, TMR global retrieval accuracy for the wet tropospheric range correction is estimated at 1.1 cm root mean square (RMS) with consistent peformance under clear, cloudy, and windy conditions.

Description: The NASA altimeter on board TOPEX/POSEIDON exploits the difference in the delays of the Ku and C band radar pulses to estimate an ionosphere correction to the range measurement. The dependence of the ionosphere correction on ocean and satellite parameters is less than 1 cm. The standard deviation of the 1-s averaged ionosphere correction depends on the height of the ocean waves and ranges from 5 to 14 mm. The accuracy of the ionosphere correction is better than 1 cm at the 1 sigma confidence level. The ionosphere correction should be averaged over 140 km (20 s) along track in order to minimize its noise without sacrificing its accuracy. Ionosphere models must achieve an independent sample spacing of 900 km or less in order to allow a single-frequency altimeter to have an ionosphere correction comparable in accuracy to that of the NASA dual-frequency altimeter.

Description: Estimates of the effectiveness of an altimetric correction, and interpretation of sea level variability as a response to atmospheric forcing, both depend upon assuming that residual errors in altimetric corrections are uncorrelated among themselves and with residual sea level, or knowing the correlations. Not surprisingly, many corrections are highly correlated since they involve atmospheric properties and the ocean surface's response to them. The full corrections (including their geographically varying time mean values), show correlations between electromagnetic bias (mostly the height of wind waves) and either atmospheric pressure or water vapor of -40%, and between atmospheric pressure and water vapor of 28%. In the more commonly used collinear differences (after removal of the geographically varying time mean), atmospheric pressure and wave height show a -30% correlation, atmospheric pressure and water vapor a -10% correlation, both pressure and water vapor a 7% correlation with residual sea level, and a bit surprisingly, ionospheric electron content and wave height a 15% correlation. Only the ocean tide is totally uncorrelated with other corrections or residual sea level. The effectiveness of three ionospheric corrections (TOPEX dual-frequency, a smoothed version of the TOPEX dual-frequency, and Doppler orbitography and radiopositioning integrated by satellite (DORIS) is also evaluated in terms of their reduction in variance of residual sea level. Smooth (90-200 km along-track) versions of the dual-frequency altimeter ionosphere perform best both globally and within 20 deg in latitude from the equator. The noise variance in the 1/s TOPEX inospheric samples is approximately (11 mm) squared, about the same as noise in the DORIS-based correction; however, the latter has its error over scales of order 10(exp 3) km. Within 20 deg of the equator, the DORIS-based correction adds (14 mm) squared to the residual sea level variance.

Description: The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a registered trademark of AT&T Bell Laboratories.

Description: The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a registered trademark of AT&T Bell Laboratories.

Description: The Flexible Spacecraft Dynamics and Control program (FSD) was developed to aid in the simulation of a large class of flexible and rigid spacecraft. FSD is extremely versatile and can be used in attitude dynamics and control analysis as well as in-orbit support of deployment and control of spacecraft. FSD has been used to analyze the in-orbit attitude performance and antenna deployment of the RAE and IMP class satellites, and the HAWKEYE, SCATHA, EXOS-B, and Dynamics Explorer flight programs. FSD is applicable to inertially-oriented spinning, earth oriented, or gravity gradient stabilized spacecraft. The spacecraft flexibility is treated in a continuous manner (instead of finite element) by employing a series of shape functions for the flexible elements. Torsion, bending, and three flexible modes can be simulated for every flexible element. FSD can handle up to ten tubular elements in an arbitrary orientation. FSD is appropriate for studies involving the active control of pointed instruments, with options for digital PID (proportional, integral, derivative) error feedback controllers and control actuators such as thrusters and momentum wheels. The input to FSD is in four parts: 1) Orbit Construction FSD calculates a Keplerian orbit with environmental effects such as drag, magnetic torque, solar pressure, thermal effects, and thruster adjustments; or the user can supply a GTDS format orbit tape for a particular satellite/time-span; 2) Control words - for options such as gravity gradient effects, control torques, and integration ranges; 3) Mathematical descriptions of spacecraft, appendages, and control systems- including element geometry, properties, attitudes, libration damping, tip mass inertia, thermal expansion, magnetic tracking, and gimbal simulation options; and 4) Desired state variables to output, i.e., geometries, bending moments, fast Fourier transform plots, gimbal rotation, filter vectors, etc. All FSD input is of free format, namelist construction. FSD is written in FORTRAN 77, PASCAL, and MACRO assembler for batch execution and has been implemented on a DEC VAX series computer operating under VMS. The PASCAL and MACRO routines (in addition to the FORTRAN program) are supplied as both source and object code, so the PASCAL compiler is not required for implementation. This program was last updated in 1985.

Description: In an effort to place payloads into orbit at the lowest possible costs, the use of air-breathing space-planes, which reduces the need to carry the propulsion system oxidizer, has been examined. As this approach would require the space-plane to fly at hypersonic speeds for periods of time much greater than that required by rockets, many factors must be considered when analyzing its benefits. The Basic Hypersonic Data and Equations spreadsheet provides data gained from three analyses of a space-plane's performance. The equations used to perform the analyses are derived from Newton's second law of physics (i.e. force equals mass times acceleration); the derivation is included. The first analysis is a parametric study of some basic factors affecting the ability of a space-plane to reach orbit. This step calculates the fraction of fuel mass to the total mass of the space-plane at takeoff. The user is able to vary the altitude, the heating value of the fuel, the orbital gravity, and orbital velocity. The second analysis calculates the thickness of a spherical fuel tank, while assuming all of the mass of the vehicle went into the tank's shell. This provides a first order analysis of how much material results from a design where the fuel represents a large portion of the total vehicle mass. In this step, the user is allowed to vary the values for gross weight, material density, and fuel density. The third analysis produces a ratio of gallons of fuel per total mass for various aircraft. It shows that the volume of fuel required by the space-plane relative to the total mass is much larger for a liquid hydrogen space-plane than any other vehicle made. This program is a spreadsheet for use on Macintosh series computers running Microsoft Excel 3.0. The standard distribution medium for this package is a 3.5 inch 800K Macintosh format diskette. Documentation is included in the price of the program. Macintosh is a registered trademark of Apple Computer, Inc. Microsoft is a registered trademark of Microsoft Corporation.

Description: The DET/MPS programs model and simulate the Direct Energy Transfer and Multimission Spacecraft Modular Power System in order to aid both in design and in analysis of orbital energy balance. Typically, the DET power system has the solar array directly to the spacecraft bus, and the central building block of MPS is the Standard Power Regulator Unit. DET/MPS allows a minute-by-minute simulation of the power system's performance as it responds to various orbital parameters, focusing its output on solar array output and battery characteristics. While this package is limited in terms of orbital mechanics, it is sufficient to calculate eclipse and solar array data for circular or non-circular orbits. DET/MPS can be adjusted to run one or sequential orbits up to about one week, simulated time. These programs have been used on a variety of Goddard Space Flight Center spacecraft projects. DET/MPS is written in FORTRAN 77 with some VAX-type extensions. Any FORTRAN 77 compiler that includes VAX extensions should be able to compile and run the program with little or no modifications. The compiler must at least support free-form (or tab-delineated) source format and 'do do-while end-do' control structures. DET/MPS is available for three platforms: GSC-13374, for DEC VAX series computers running VMS, is available in DEC VAX Backup format on a 9-track 1600 BPI tape (standard distribution) or TK50 tape cartridge; GSC-13443, for UNIX-based computers, is available on a .25 inch streaming magnetic tape cartridge in UNIX tar format; and GSC-13444, for Macintosh computers running AU/X with either the NKR FORTRAN or AbSoft MacFORTRAN II compilers, is available on a 3.5 inch 800K Macintosh format diskette. Source code and test data are supplied. The UNIX version of DET requires 90K of main memory for execution. DET/MPS was developed in 1990. A/UX and Macintosh are registered trademarks of Apple Computer, Inc. VMS, DEC VAX and TK50 are trademarks of Digital Equipment Corporation. UNIX is a registered trademark of AT&T Bell Laboratories.

Description: NASA-JPL's four-frequency telecommunication system design entails the creation and integration of a frequency-selective surface (FSS) subreflector into the high-gain antenna subsystem. The FSS design, which incorporates a periodic array of conducting elements on a kevlar/polymer composite structure, will be able to multiplex S, X, Ku, and Ka frequency-band wavelengths. Accounts are presented of the FSS's development, mechanical testing, and electrical testing.

Description: Technology is being developed to process signals from distributed sensors using distributed computations. These distributed sensors provide a new feedback capability for vibration control that has not been exploited. Additionally, the sensors proposed are of an optical and distributed nature and could be employed with known techniques of distributed optical computation (Fourier optics, etc.) to accomplish the control system functions of filtering and regulation in a distributed computer. This paper reviews a procedure for the analytic design of control systems for this application. For illustration, the procedure is applied to the problem of suppressing the vibrations of a simply supported beam. A simulator has been developed to study the effects of sensor and processing errors. An extensive study of the effects of these errors on estimation and regulation performance is presented.

Description: A reanalysis of the evolution configuration module pattern for Space Station Freedom is presented. The module pattern is to consist of one habitation module, one airlock, one cupola, two Space Shuttle Orbiter pressurized docking adapters, one assured crew return vehicle, and one pressurized logistics module. Three laboratory modules are also included. A phased plan has been developed on the basis of utilization analysis being performed and derived through actual and projected user mission requirements for the growth of pressurized elements. The evolution requirement impacts to the module pattern are outlined. Application of the growth philosophy and the physical limitations resulted in a number of core element ground rules for module pattern growth. Nodes and modules are to be grown symmetrically in order to optimize poor flight control characteristics and maximize the capability for crew dual egress between elements. Core module pattern elements are not to be grown along Y-Y axis from nodes 1 or 2 due to physical clearance problems with the thermal control system radiators operational envelope.

Description: Results are presented of a thermal design optimization study of the segmented GFRP primary reflector of the earth-orbiting Submillimeter Imager and Line Survey telescope. The paper examines the thermal requirements of the primary reflector and the thermal environment of the telescope and describes the thermal design of the primary reflector. Particular attention is given to the geometric math model and the thermal math model of the telescope. A summary for the steady-state thermal performance of the optimized design is presented, showing that the optimized design has reduced, by an order of magnitude, structural spatial temperature gradients, which were earlier shown to be the most significant obstacle in maintaining the required telescope figure accuracy.

Description: The research experience of SCOLE developed by the NASA Langley Research Center is reviewed. Particular attention is given to the advances made in control and modeling techniques, the experimental facilities that are now available, and the continuing needs to advance and to validate the technologies in controlling flexible structures.

Description: An analytic redundancy management approach for on-line component failure detection is described and illustrated using a simulation of the NASA Langley Spacecaft Control Laboratory Experiment (SCOLE) research facility. The SCOLE experimental apparatus simulated is a functional model of the Space Shuttle with a large, flexible, offset-feed antenna cantilevered from the payload bay. This approach uses a single, active, Kalman filter selected from a bank of filters, each element of which was previously designed to accommodate a specific failure condition. The residuals of this active filter are processed through a sequential probability ratio test (SPRT) filter to identify the failure state of the system and, hence, closing the failure accommodation loop, to select the active filter. Results are presented that illustrate the ability of the system to detect and recover from failures introduced in the angular rate and linear acceleration sensors on the SCOLE facility.

Description: Structural sizing and performance data are presented for two different aerobrake hexagonal heatshield panel structural concepts. One concept features a sandwich construction with an aluminum honeycomb core and thin quasi-isotropic graphite-epoxy face sheets. The other concept features a skin-rib isogrid construction with thin quasi-isotropic graphite-epoxy skins, and graphite-epoxy ribs oriented at 0, +60, and -60 deg along the panel. Linear static, linear bifurcation buckling, and nonlinear static analyses were performed to compare the structural performance of the two panel concepts and assess their feasibility for a Lunar Transfer Vehicle aerobrake application.

Description: An effort is currently being carried out by the Jet Propulsion Laboratory (JPL) to study mission feasibility and to define functional requirements for various subsystems of the Space Infrared Telescope Facility (SIRTF). As a major part of this effort, structural design requirements have been derived based on the stated mission objectives. Design concerns addressed by these requirements include the limits on mass and location of the center of gravity, launch stiffness and dynamic characteristics, design loads and analysis criteria, survivability of the TITAN IV/Centaur launch environment, thermal control for maintaining a near absolute-zero operating temperature, and helium cryogen volume and storage for a five-year mission. To illustrate how the structural design requirements can be met, a point design of the SIRTF flight hardware system was developed, modeled, and analyzed. A description of the key features of this point design, along with pertinent modeling and analysis results, are discussed in this Paper.

Description: The progress in spacecraft charging is reviewed with particular attention given to the interactions of plasma and penetrating radiation with dielectrics. Topics discussed include the charging environments, elementary charging theory, the anomalies attributed to charging or discharging phenomena, and spacecraft engineering.

Description: The objectives of the Solar Probe mission and the current status of the Solar Probe thermal shield subsystem development are described. In particular, the discussion includes a brief description of the mission concepts, spacecraft configuration and shield concept, material selection criteria, and the required material testing to provide a database to support the development of the shield system.

Description: Arcjet thrusters are soon to be used for north/south stationkeeping on commercial communications satellites. A series of tests was performed to evaluate the possible effects of these thrusters on spacecraft charging and the degradation of thermal control material. During the tests the interaction between arcjet plumes and both charged and uncharged surfaces did not cause any significant material degradation. In addition, firing an arcjet thruster benignly reduced the potential of charged surfaces to near zero.

Description: The present multidisciplinary telescope-analysis approach, which encompasses thermal, structural, control and optical considerations, is illustrated for the case of an IR telescope in LEO; attention is given to end-to-end evaluations of the effects of mechanical disturbances and thermal gradients in measures of optical performance. Both geometric ray-tracing and surface-to-surface diffraction approximations are used in the telescope's optical model. Also noted is the role played by NASA-JPL's Integrated Modeling of Advanced Optical Systems computation tool, in view of numerical samples.

Description: In this paper a distributed parameter model for the estimation of modal characteristics of NASA Mini-Mast truss is proposed. A closed-form solution of the Timoshenko beam equation, for a uniform cantilevered beam with two concentrated masses, is derived so that the procedure and the computational effort for the estimation of modal characteristics are improved. A maximum likelihood estimator for the Timoshenko beam model is also developed. The resulting estimates from test data by using Timoshenko beam model are found to be comparable to those derived from other approaches.

Description: A time-optimal open-loop control problem of flexible spacecraft in the presence of modeling uncertainty has been investigated. The results indicate that the proposed approach significantly reduces the residual structural vibrations caused by modeling uncertainty. The results also indicate the importance of proper jet placement for practical tradeoffs among the maneuvering time, fuel consumption, and performance robustness.

Description: This paper uses spatial operator techniques to develop a new algorithm for the dynamics of multibody systems with hinges undergoing prescribed motion. This algorithm is spatially recursive, and its computational complexity grows only linearly with the number of degrees of freedom in the system. Its structure is a hybrid of known recursive forward and inverse dynamics algorithms for regular multibody systems. Changes to the prescribed/nonprescribed nature of hinges can be implemented during run time since they are handled with very low overhead in the algorithm.

Description: The evolution and optimization of a real-time digital control system is presented. The control system is part of a testbed used to perform focused technology research on the interactions of spacecraft platform and instrument controllers with the flexible-body dynamics of the platform and platform appendages. The control system consists of Computer Automated Measurement and Control (CAMAC) standard data acquisition equipment interfaced to a workstation computer. The goal of this work is to optimize the control system's performance to support controls research using controllers with up to 50 states and frame rates above 200 Hz. The original system could support a 16-state controller operating at a rate of 150 Hz. By using simple yet effective software improvements, Input/Output (I/O) latencies and contention problems are reduced or eliminated in the control system. The final configuration can support a 16-state controller operating at 475 Hz. Effectively the control system's performance was increased by a factor of 3.

Description: One of the main difficulties encountered in the design and implementation of control systems for spacecraft is the interaction between the control system and the flexibility of the vehicle. This difficulty has resulted in problems for a multitude of spacecraft including the earliest U.S. spacecraft, Explorer I; the Orbiting Geophysical Observatory III; Mariner 10; Galileo; and the Hubble Telescope. Recognizing the importance of the issue, NASA has an ongoing Controls-Structures Interaction (CSI) technology program to develop the methodology to design optimally and simultaneously both the control system and the structure. The CSI program is a multicenter program involving research teams from NASA's Langley Research Center (LaRC), Jet Propulsion Laboratory (JPL), and Marshall Space Flight Center (MSFC). JPL's area of interest is in optics-class vehicles and MSFC's in astrophysics vehicles. The aim of this paper is to explain the ongoing activities at LaRC, which are of a theoretical, ground test, and flight test nature, with focus on applications to spacecraft with multiple experiment-pointing mounts, large space radiometers, Space Station Freedom (SSF), and the Space Shuttle Remote Manipulator System (RMS).

Description: A critical technology area for lightweight SiC-based telescope systems is the structural integrity and thermal stability over spaceborne environmental launch and thermal operating conditions. Note, it is highly desirable to have an inherently athermal design of both SiC mirrors and structure. SSG has developed an 8 inch diameter SiC telescope system for brassboard level optical and thermal testing. The brassboard telescope has demonstrated less than 0.2 waves P-V in the visible wavefront change over +50 C to -200 C temperature range. SSG has also fabricated a SiC truss structural assembly and successfully qualified this hardware at environmental levels greater than 3 times higher than normal Delta, Titan, and ARIES launch loads. SSG is currently developing two SiC telescopes; an 20 cm diameter off-axis 3 mirror re-imaging and a 60 cm aperture on-axis 3 mirror re-imager. Both hardware developments will be tested to flight level environmental, optical, and thermal specifications.

Description: The Vibroacoustic Payload Environment Prediction System (VAPEPS) was used to predict the stationary on-orbit environments in one of the Space Station Freedom modules. The model of the module included the outer structure, equipment and payload racks, avionics, and cabin air and duct systems. Acoustic and vibratory outputs of various source classes were derived and input to the model. Initial results of analyses, performed in one-third octave frequency bands from 10 to 10,000 Hz, show that both the microgravity and acoustic environments will be exceeded in some one-third octave bands with the current SSF design. Further analyses indicate that interior acoustic level requirements will be exceeded even if the microgravity requirements are met.

Description: In this article, the vibration control of adaptive truss structures, where the control actuation is provided by length adjustable active members, is formulated as a direct output feedback control problem. A control method named Model Truncated Output Feedback (MTOF) is presented. The method allows the control feedback gain to be determined in a decoupled and truncated modal space in which only the critical vibration modes are retained. The on-board computation required by MTOF is minimal; thus, the method is favorable for the applications of vibration control of large scale structures. The truncation of the modal space inevitably introduces spillover effect during the control process. In this article, the effect is quantified in terms of active member locations, and it is shown that the optimal placement of active members, which minimizes the spillover effect (and thus, maximizes the control performance) can be sought. The problem of optimally selecting the locations of active members is also treated.

Description: Future NASA missions such as the Great Observatories of the 21st Century, will require high dimensional stability (i.e., the system's ability to retain geometrical properties related to the system's performance) which will have to be maintained with micron to nanometer accuracy over the 5 to 10 years of mission lifetime. This paper examines the thermodynamic and other mechanisms which limit the dimensional stability of a space system. It is shown that the space system's performance will be limited below 0.1 per million dimensional stability.

Description: This paper describes a preliminary design of the thrust vector control algorithm for the interplanetary spacecraft, Cassini. Topics of discussion include flight software architecture, modeling of sensors, actuators, and vehicle dynamics, and controller design and analysis via classical methods. Special attention is paid to potential interactions with structural flexibilities and propellant dynamics. Controller performance is evaluated in a simulation environment built around a multi-body dynamics model, which contains nonlinear models of the relevant hardware and preliminary versions of supporting attitude determination and control functions.

Description: Imaging the Earth's magnetosphere from space will enable scientists to better understand the global shape of the inner magnetosphere, its components and processes. The proposed Inner Magnetosphere Imager (IMI) mission will obtain the first simultaneous images of the component regions of the inner magnetosphere and will enable scientists to relate these global images to internal and external influences as well as local observations. NASA's Marshall Space Flight Center (MSFC) is performing a concept definition study of the proposed mission. As currently envisioned, the baseline mission calls for an instrument complement of approximately seven imagers to be flown in an elliptical Earth orbit with an apogee of seven Earth Radii (RE). Several spacecraft concepts have been examined for the mission. The baseline concept utilizes a spinning spacecraft with a despun platform, the second uses a three-axis stabilized spacecraft with a spinning platform, while the third option splits the instruments onto two small satellites; a spinning spacecraft and a complementary three-axis stabilized spacecraft. This paper will address the mission objectives, the rationale for using proven spacecraft designs, and the preliminary concept definition study team results for all three options.

Description: An effort was made to increase the predicted lifetime of the SIRTF dewar by lowering the exterior shell temperature, increasing the radiated energy from the vapor cooled shields and reconfiguring the vapor cooled shields. The lifetime increases can be used to increase the scientific return from the mission and as a trade-off against mass and cost. This paper describes the configurations studied, the steady state thermal model used, the analytical methods and the results of the analysis. Much of the heat input to the outside dewar shell is radiative heat transfer from the solar panel. To lower the shell temperature, radiative cooled shields were placed between the solar panel and the dewar shell and between the bus and the dewar shell. Analysis showed that placing a radiator on the outer vapor cooled shield had a significant effect on lifetime. Lengthening the distance between the outer shell and the point where the vapor cooled shields are attached to the support straps also improved lifetime.

Description: A control law for a single-tether orbiting satellite system based on a reduced order linear adaptive control technique is presented. The main advantages of this technique are its design simplicity and the facts that specific system parameters and model linearization are not required when designing the controller. Two controllers are developed: one which uses only tension in the tether as control actuation and one which uses both tension and in-plane thrusters as control actuation. Both a sixth-order nonlinear and an 11th-order bead model of a tethered satellite system are used for simulation purposes, demonstrating the ability of the controller to manage an uncertain system. Retrieval and stationkeeping results using these nonlinear models and the linear adaptive controller demonstrate the feasibility of the method. The robustness of the controller with respect to parameter uncertainties is also demonstrated by changing the nonlinear model and parameters within the model without redesigning the controller.

Description: The glowing cloud near the ram surfaces of the Space Shuttle was observed with a hand-held, intensified spectrograph operated by the astronauts from the aft-flight-deck of the Space Shuttle. The spectral measurements were made between 400 and 800 nm with a resolution of 3 nm. Analysis of the spectral response of the instrument and the transmission of the Shuttle window was performed on orbit using earth-airglow OH Meinel bands. This analysis resulted in a correction of the Shuttle glow intensity in the spectral region between 700 and 800 nm. The data presented in this report is in better agreement with laboratory measurements of the NO2 continuum.

Description: Recent laboratory results using a refined phase resonance method and the eigensystem realization algorithm on the same test structure are reported. These methods are dissimilar modal identification techniques suitable for future large spacecraft. The theory, application approach, and results obtained for each technique are summarized and compared. Although both methods worked well in this investigation, significant differences occurred in some identified mode shapes. Comparison of independently derived modal parameters provides the means for disclosing such discrepancies in flight projects.

Description: This paper discusses practical aspects of performing on-orbit modal identification using time domain analysis of free-decay data. The effects of environmental constraints, structural characteristics, excitation, and sensing are reviewed. In a recent laboratory application, an on-orbit experiment is simulated using a limited number of excitation and measurement points. The identified modal parameters correlate well, though not uniquely, with those obtained in a complete modal survey. Practical difficulties in performing the correlation are illustrated.

Description: The potential benefits of employing controlled structures technology in spacecraft design are illustrated using an example of a Mission-to-Planet Earth multisensor geostationary platform. The platform has a precipitation mapping sensor which requires large antennas to meet the science requirement of resolution cell size and dynamic measurement range. A series approach and an integrated approach to designing a controlled structure are discussed.

Description: Major issues in system identification are summarized and recent advances are reviewed. Modal testing and system identification used in control theory are examined, and the mathematical relationships and conversions of the models appropriate to modal testing and those appropriate to modern control design methods are discussed. The importance of obtaining input and output matrices in modal testing is emphasized, and the changes that may be needed in modal testing procedures to meet the needs of the control system designer are addressed. Directions for future research are considered.

Description: This paper describes the testing aspects and the problems encountered during the free-free and fixed base modal surveys completed on the original Space Station Common Module Prototype (CMP). The CMP is a 40-ft long by 14.5-ft diameter 'waffle-grid' cylinder built by the Boeing Company and housed at the Marshall Space Flight Center (MSFC) near Huntsville, AL. The CMP modal survey tests were conducted at MSFC by the Dynamics Test Branch. The free-free modal survey tests (June '90 to Sept. '90) included interface verification tests (IFVT), often referred to as impedance measurements, mass-additive testing and linearity studies. The fixed base modal survey tests (Feb. '91 to April '91), including linearity studies, were conducted in a fixture designed to constrain the CMP in 7 total degrees-of-freedom at five trunnion interfaces (two primary, two secondary, and the keel). The fixture also incorporated an airbag off-load system designed to alleviate the non-linear effects of friction in the primary and secondary trunnion interfaces. Numerous test configurations were performed with the objective of providing a modal data base for evaluating the various testing methodologies to verify dynamic finite element models used for input to coupled load analysis.

Description: A Mars Aerial Platform (MAP) mission capable of generating thousands of very-high-resolution (20 cm/pixel) pictures of the Martian surface is considered. The MAP entry vehicle will map the global circulation of the planet's atmosphere and examine the surface and subsurface. Data acquisition will use instruments carried aboard balloons flying at nominal altitude of about 7 km over the Martian surface. The MAP balloons will take high- and medium-resolution photographs of Mars, sound its surface with radar, and provide tracking data to chart its winds. Mars vehicle design is based on the fourth-generation NTP, NEP, SEP vehicle set that provides a solid database for determining transportation system costs. Interference analysis and 3D image generation are performed using manual system sizing and sketching in conjunction with precise CAD modeling.

Description: An attitude control law is derived to realize linear unforced error dynamics with the attitude error defined in terms of rotation group algebra (rather than vector algebra). Euler parameters are used in the rotational dynamics model because they are globally nonsingular, but only the minimal three Euler parameters are used in the error dynamics model because they have no nonlinear mathematical constraints to prevent the realization of linear error dynamics. The control law is singular only when the attitude error angle is exactly pi rad about any eigenaxis, and a simple intuitive modification at the singularity allows the control law to be used globally. The forced error dynamics are nonlinear but stable. Numerical simulation tests show that the control law performs robustly for both initial attitude acquisition and attitude control.

Description: The design, analysis, and simulation of a control law that guarantees a fine pointing performance under external disturbances and model uncertainties are investigated. The authors examine the achievable performance and modeling errors for the control-structure interaction (CSI) evolutionary model structure and highlight the importance of accurate knowledge of structural frequencies and damping values for very lightly damped structures. Numerical results indicate that a limited but significant line-of-sight pointing performance robustness can be attained in spite of modeled structured and unstructured uncertainties in the system. The simulation results reported demonstrate typical performance levels, types of uncertainties and their magnitudes, and design tradeoffs for a realistic laboratory model of a large flexible space structure.

Description: A state space model for the Hubble Space Telescope under the influence of unknown disturbances in orbit is presented. This model was obtained from flight data by applying the Q-Markov covariance equivalent realization identification algorithm. This state space model guarantees the match of the first Q-Markov parameters and covariance parameters of the Hubble system. The flight data were partitioned into high- and low-frequency components for more efficient Q-Markov cover modeling, to reduce some computational difficulties of the Q-Markov cover algorithm. This identification revealed more than 20 lightly damped modes within the bandwidth of the attitude control system. Comparisons with the analytical (TREETOPS) model are also included.

Description: The topics are presented in view graph form and include the following: an adaptive model following control; adaptive control of a distributed parameter system (DPS) with a finite-dimensional controller; a direct adaptive controller; a closed-loop adaptively controlled DPS; Lyapunov stability; the asymptotic stability of the closed loop; and model control of a simply supported beam.

Description: The first general research objective was to address control design challenges of the Spacecraft Control Laboratory Experiment (SCOLE) via the two stage approach: (1) slew the whole as if it were a rigid body about one Space Shuttle body axis each time using the onboard Reaction Control System (RCS) thrusters; and (2) damp out excited vibrations. The second objective was to examine the feasibility of applying the approach to shuttle-attached flexible space structures. The following was accomplished: (1) a standard bang-bang control technique was adapted; (2) a slew rate limit was imposed in the design; and (3) slew acceleration deviation was defined as the index of slew performance degradation.

Description: Information on a modal model for the Spacecraft Control Laboratory (SCOLE) is given in viewgraph form. A partial differential equation model covering roll bending, pitch bending, torsion, shear forces, and bending moments is given.

Description: The distributed element dynamic analysis package DISTEL is used to analyze the NASA/Institute of Electrical and Electronics Engineers' Spacecraft Control Laboratory Experiment (SCOLE). In this configuration, the Space Shuttle motion is coupled to the motion of a large dish antenna through a Shuttle-deployed flexible mast of 40 m long. Due to the high asymmetry of the system, the motions about the different axes (roll, pitch, yaw) are severely coupled. A general purpose software like DISTEL is especially suited for this kind of analysis. Modal frequencies of the complete spacecraft and impulse response (modal gains) to excitations at different locations are obtained. Mode-shape plots of the deformations of the entire system are given. Finally, results obtained at NASA and at Purdue University are compared to those found by the European space technology center, ESTEC.

Description: A mathematical formulation for the control of the Spacecraft Control Laboratory Experiment (SCOLE) configuration is given. Two equivalent approaches, one using a functional equation and the other an abstract wave equation, are illustrated. Such a formulation can help in digital computer simulation to evaluate control laws, providing insight, and generating control laws.

Description: This document presents in viewgraph form techniques for the optimal design of aerospace vehicles which take into account vehicle shape, aerodynamic performance, and weight minimization. Methods include multilevel optimization with linear decomposition, multi-constraint optimization, global sensitivity matrix, and calculation of sensitivity derivatives using both first and second order derivatives. The effects of including flexibility in the optimization is discussed.

Description: The objectives of this study are listed as follows: (1) to develop Lagrange's equations of motion for the shuttle antenna configuration in orbit; (2) to modify equations using the Lagrange multiplier method to develop equations of motion for the laboratory experiment; and (3) to discuss methods for simulation and control. The equations are presented in graph form.

Description: Discussed here is a NASA program which was initiated to make direct comparisons of control laws for a mathematical problem. An experimental test item is being assembled under the cognizance of the Spacecraft Control Branch at Langley Research Center. The physical apparatus will consist of a softly supported dynamic model of an antenna attached to the Space Shuttle by a flexible beam. The control objective will include the task of directing the line of sight of the Shuttle/antenna configuration toward a fixed target, under conditions of noisy data, limited control authority, and random disturbances.

Description: The scope of this study covered steady-state, continuous-time vibration control under disturbances applied to the Space Shuttle and continuous-time models of actuators, sensors, and disturbances. Focus was on a clear illustration of the methodology, therefore sensor/actuator dynamics were initially ignored, and a finite element model of the NASA Spacecraft Control Laboratory Experiment (SCOLE) was conducted, including products of inertia and offset of reflector CM from the mast tip.

Description: The motivation was to develop a control design and analysis methodology directly applicable to design concepts of flexible spacecraft of interest the the U.S. Navy and to provide a testbed for the evaluation of large space structure control techniques developed at the Naval Research Laboratory. The topics covered include the following: (1) a list of key concepts; (2) evaluation of the Spacecraft Control Laboratory Experiment (SCOLE) model with DISCOS; (3) baseline results, line of sight error vs. time; (4) general formulation of optimization; (5) geometric interpretation, projected eigenaxis; (6) closed loop control law; and (7) future directions.

Description: Nonlinear and linear equations of motions were derived. The preliminary investigation consisted of model beam as truss structure, effects of truss structure on control design, and effects of reflector offset on control design. It was concluded that the offset of the reflector c.g. from the beam reflector attach point is dynamically significant. Also, truss effects may also significantly effect the performance of the controller if ignored. If the truss is included in the modeling of the NASA/SCOLE configuration, a practically implementable scheme is available to reduce the model order.

Description: Future utilization of space will require large space structures in low-Earth and geostationary orbits. Example missions include: Earth observation systems, personal communication systems, space science missions, space processing facilities, etc., requiring large antennas, platforms, and solar arrays. The dimensions of such structures will range from a few meters to possibly hundreds of meters. For reducing the cost of construction, launching, and operating (e.g., energy required for reboosting and control), it will be necessary to make the structure as light as possible. However, reducing structural mass tends to increase the flexibility which would make it more difficult to control with the specified precision in attitude and shape. Therefore, there is a need to develop a methodology for designing space structures which are optimal with respect to both structural design and control design. In the current spacecraft design practice, it is customary to first perform the structural design and then the controller design. However, the structural design and the control design problems are substantially coupled and must be considered concurrently in order to obtain a truly optimal spacecraft design. For example, let C denote the set of the 'control' design variables (e.g., controller gains), and L the set of the 'structural' design variables (e.g., member sizes). If a structural member thickness is changed, the dynamics would change which would then change the control law and the actuator mass. That would, in turn, change the structural model. Thus, the sets C and L depend on each other. Future space structures can be roughly divided into four mission classes. Class 1 missions include flexible spacecraft with no articulated appendages which require fine attitude pointing and vibration suppression (e.g., large space antennas). Class 2 missions consist of flexible spacecraft with articulated multiple payloads, where the requirement is to fine-point the spacecraft and each individual payload while suppressing the elastic motion. Class 3 missions include rapid slewing of spacecraft without appendages, while Class 4 missions include general nonlinear motion of a flexible spacecraft with articulated appendages and robot arms. Class 1 and 2 missions represent linear mathematical modeling and control system design problems (except for actuator and sensor nonlinearities), while Class 3 and 4 missions represent nonlinear problems. The development of an integrated controls/structures design approach for Class 1 missions is addressed. The performance for these missions is usually specified in terms of (1) root mean square (RMS) pointing errors at different locations on the structure, and (2) the rate of decay of the transient response. Both of these performance measures include the contributions of rigid as well as elastic motion.

Description: A fundamental problem facing controls-structures analysts is a means of determining the trade-offs between structural design parameters and control design parameters in meeting some particular performance criteria. Developing a general optimization-based design methodology integrating the disciplines of structural dynamics and controls is a logical approach. The objective of this study is to develop such a method. Classical design methodology involves three phases. The first is structural optimization, wherein structural member sizes are varied to minimize structural mass, subject to open-loop frequency constraints. The next phase integrates control and structure design with control gains as additional design variables. The final phase is analysis of the 'optimal' integrated design phase considering 'real' actuators and 'standard' member sizes. The control gains could be further optimized for fixed structure, and actuator saturation constraints could be imposed. However, such an approach does not take full advantage of opportunities to tailor the structure and control system design as one system.

Description: The Spacecraft Control Laboratory Experiment (SCOLE) will allow direct experimental comparison of competing control schemes for large flexible spacecraft structures. The experiment was designed to emulate the essential characteristics of a mathematical model design challenge which represents a Space Shuttle with a flexible mast and antenna attached. This experiment represents the third in a series of three flexible structure control experiments used by the Flight Dynamics and Control Division at LaRC. The key problem addressed by the facility is that flexible motion of the mast and antenna must be considered in the slewing and pointing control problems.

Description: The following topics are covered in view graph form: (1) pulse control strategy; (2) stability analysis and digital simulations; (3) digital/analog and analog/digital conversions, and analog simulation; and (4) experimental studies.

Description: Researchers simplified the analytical expression of the line of sight (LOS) error, discovered and proved the independence of Euler angle Psi, calculated attitude angles corresponding to 0 degrees and 20 degrees LOS errors, determined choices of initial alignment, tailored the slew maneuvers for LOS pointing, simulated numerically the LOS pointing slew of the Spacecraft Control Laboratory Experiment (SCOLE), and evaluated the pointing performance.

Description: The JPL Control/Structure Interaction Program is developing new analytical methods for designing micro-precision spacecraft with controlled structures. One of these, the Conceptual Design Tool, will illustrate innovative new approaches to the integration of multi-disciplinary analysis and design methods. The tool will be used to demonstrate homogeneity of presentation, uniform data representation across analytical methods, and integrated systems modeling. The tool differs from current 'integrated systems' that support design teams most notably in its support for the new CSI multi-disciplinary engineer. The design tool will utilize a three dimensional solid model of the spacecraft under design as the central data organization metaphor. Various analytical methods, such as finite element structural analysis, control system analysis, and mechanical configuration layout, will store and retrieve data from a hierarchical, object oriented data structure that supports assemblies of components with associated data and algorithms. In addition to managing numerical model data, the tool will assist the designer in organizing, stating, and tracking system requirements.

Description: Ever since the United States space program started some forty years ago, there have been many ideas on how the U.S. should proceed to explore space. Throughout the years, many innovative designs have surfaced for transfer vehicles, space stations, and surface bases. Usually the difference in designs are due to differences in mission objectives and requirements. The problem for Mars is how to choose an architecture for human travel to Mars and what kind of base construction to design for Mars that will be reliable and cost effective. Eventually, if the Space Exploration Initiative is to become a reality, NASA will have to select and fund a single mission architecture involving manned and unmanned Mars fly-by precursors, a Mars landing vehicle, and, ultimately, the plan for constructing a Mars base. The decision to commit to a single architecture is a vital one and, therefore, the design issues, the decision making process, and the analysis tools must be available to explore all of the options that are available. A large part of any space mission architecture is the Earth-to-Mars transfer vehicle. The decision on the type of transfer vehicle to design is a crucial one. The many options must take into account the constraints encountered when assembling the vehicle in earth orbit such as effective joining methods, test and evaluation methods, preventative maintenance measures, etc. Therefore, the process of trading off various designs must include every facet of that design. The on-orbit assembly/construction constraints will drive designs and architectures. This viewgraph presentation highlights the above critical issues so that designs may be evaluated from these viewpoints. Evaluating designs from the issues contained in this paper will help decision makers detect inadequate designs. Stressing these issues in the evaluation procedure will have a great impact on the decisions of future space mission transfer vehicles and consequent architectures.

Description: The deliberations of the Systems/Operations Technology Panel are summarized. The first real question that arose was to develop an understanding of what systems/operations technology is. A relativelynew discipline in the NASA technology organization, necessitates the definition of the objectives. Two objectives were established: (1) to make new things possible, and (2) to make existing capabilities cost less or work better. Making new things possible is not really applicable in the case of a space station. Both Salyut 7 and Skylab indicate that space stations are possible with existing (not necessarily new) technology. There was a concern on the part of some of the penelists that work better might mean higher performance, and that is not necessarily the case at all. Work better may mean simply to provide better service to the users of the space station at lower cost. The panel felt this to be a more realistic viewpoint. As evidenced from interaction with users (and all of the contractors found this basically to be true), the users want low cost, no schedule constraints, and no hassles.

Description: The objective of this experiment is to obtain examples of meteoroid impact damage to typical spacecraft components, and by so doing to help establish design approaches to minimize meteoroid damage effects to future spacecraft. The results of the complete inspection of the LDEF will complement and extend the data obtained from specific meteoroid experiments flying in LDEF trays.

Description: A comprehensive set of measurements about the orbiter environment are provided by the plasma diagnostics package (PDP). Ion and electron particle densities, energies, and spatial distribution functions; ion mass for identification of particular molecular ion species; and magnetic fields, electric fields and electromagnetic waves over a broad frequency range are studied. Shuttle environmental measurements will be made both on the pallet and, by use of the remote manipulator system (RMS), the PDP will be maneuvered in and external to the bay area to continue environmental measurements and to carry on a joint plasma experiment with the Utah State University fast-pulsed electron generator. Results of orbiter environment EMI measurements and S-band field strengths as well as preliminary results from wake search operations indicating wake boundary identifiers are reported.

Description: The Earth radiation budget experiment (ERBE) software development approach is described. An iterative development approach was adopted which provides for three releases or versions of the processing system, each of increasing levels of complexity and solidity. The final release of the system will be used to process the flight data. The major phases for each iterative release consist of specifications developed in concert with the science team, preliminary design, subsystem reviews, coding, subsystem code walkthroughs, system testing, system documentation, and project status review.

Description: The correlation of outgassing to the stability of the damping properties of polymer materials to be used in spacecraft structures is discussed. A test series was devised to obtain basic information from off-the-shelf damping materials. The test results could be considered as a guideline toward the application of these materials. Eight materials were selected to form a representative cross section of those polymers having both ready availability as commercial damping materials and desirable properties. A table indicates the temperatures at which peak damping occurs at 1 Hz and the type of beam specimen used in the vacuum exposure tests. These materials as a group cover the temperature range of -85 C to 38 C.

Description: During the conceptual development of advanced aerospace vehicles, many compromises must be considered to balance economy and performance of the total system. Subsystem tradeoffs may need to be made in order to satisfy system-sensitive attributes. Due to the increasingly complex nature of aerospace systems, these trade studies have become more difficult and time-consuming to complete and involve interactions of ever-larger numbers of subsystems, components, and performance parameters. The current advances of computer-aided synthesis, modeling and analysis techniques have greatly helped in the evaluation of competing design concepts. Langley Research Center's Space Systems Division is currently engaged in trade studies for a variety of systems which include advanced ground-launched space transportation systems, space-based orbital transfer vehicles, large space antenna concepts and space stations. The need for engineering analysis tools to aid in the rapid synthesis and evaluation of spacecraft has led to the development of the Interactive Design and Evaluation of Advanced Spacecraft (IDEAS) computer-aided design system. The ADEAS system has been used to perform trade studies of competing technologies and requirements in order to pinpoint possible beneficial areas for research and development. IDEAS is presented as a multidisciplinary tool for the analysis of advanced space systems. Capabilities range from model generation and structural and thermal analysis to subsystem synthesis and performance analysis.

Description: A molecular beam facility to simulate the space environment of a spacecraft at low orbit was designed with the intent of studying the effect on the properties of optical elements of oxygen atoms impacting at orbital velocity. The four-stage differentially pumped molecular beam facility includes a variety of oxygen atom beam sources which cover a wide range of velocities (1 km/sec to approximately 8 km/sec), in addition to the ultra-clean experimental environmental of an ultra-high vacuum chamber and an optical diagnostic set-up. The primary oxygen atom beam source used to obtain the 8 km/sec O atoms is an arc heated source. It consists of a modified commercially available plasma torch. The modifications include attachments which provide for a nozzle which is used to expand the atomic beam into the vacuum system, and exhaust channels to dispose of excess torch gas. The torch operates in the 'nontransferred' mode of operation, that is the electric arc is confined within the torch. A plasma is formed in helium by a dc arc. A small amount of O2 is injected downstream from the arc where it is thermally dissociated by the hot He into oxygen atoms. The high temperature and isentropic expansion give the oxygen atoms their velocity. Using seeded beam techniques, oxygen atom beams of approximately 3.5 and approximately 1.5 km/sec, respectively, are obtained.

Description: A preliminary assessment of the space shuttle contamination environment was made using data from the first two orbital flight tests, STS-1 and STS-2. Data sources consisted of crew observations during flight, postflight vehicle inspection, and the induced environment contamination monitor which was used on STS-2 and consists of 10 instruments. These instruments are used to measure gas phase contaminants, particle population, humidity, and molecular deposition in the orbiter payload bay during ascent and descent and particle population, molecular deposition, and gas cloud during orbital flight. Results of the measurements described are presented in summary form and indicate low molecular deposition rates for both pressurized and orbital flight.

Description: The structure of the European Space Agency (ESA) is described. The major test facilities used for ESA programs are described. Facility characteristics and special test methods are described.

Description: The training program and procedures developed and implemented at the space simulation laboratory at Martin Marietta Aerospace in Denver are discussed. The training of technicians and professionals as well as preparation for instructors is covered. Training manuals and their compilation are reported as applicable to the specific needs of the laboratory. The development of a space simulation course as part of the Martin Marietta Continuing Education Night School approaching space simulation from an academic viewpoint is presented. Finally, public relations tours of the facility as an informational/educational tool are discussed.

Description: A test involving a flow of hot hydrazine decomposition products at a rate of 13.6 g/s (0.03 lb/s) established the requirement to maintain pressure in an 11.9-m-(39 ft) diameter space chamber below 200 microns. The flow, 2/3 hydrogen and 1/3 nitrogen by volume, continued for several periods ranging from 3 to 15 min. The pressure requirement was necessary to minimize thermal effects of the gas on the test vehicle but was well beyond the capability of the existing facility pumps. Various methods of obtaining additional temporary jump capacity were considered. From these, the slugged-charcoal approach was selected as the quickest and least expensive method to implement.

Description: A thermal balance test (controlled flux intensity) on a simple black dummy spacecraft using IR lamps was performed and evaluated, the latter being aimed specifically at thermal mathematical model (TMM) verification. For reference purposes the model was also subjected to a solar simulation test (SST). The results show that the temperature distributions measured during IR testing for two different model attitudes under steady state conditions are reproducible with a TMM. The TMM test data correlation is not as accurate for IRT as for SST. Using the standard deviation of the temperature difference distribution (analysis minus test) the SST data correlation is better by a factor of 1.8 to 2.5. The lower figure applies to the measured and the higher to the computer-generated IR flux intensity distribution. Techniques of lamp power control are presented. A continuing work program is described which is aimed at quantifying the differences between solar simulation and infrared techniques for a model representing the thermal radiating surfaces of a large communications spacecraft.

Description: The Jet Propulsion Laboratory 25-foot space simulator with its 5.8-m (19-ft) diameter simulated solar beam provides an excellent facility for measuring the optical characteristics of parabolic solar concentrator panels and gores. The virtual source position and size were determined by using a single lamp of the 37 xenon 30-kW source array with only the center lens in the 19-channel optical mixer. This data was used to define the optical test geometry, and it allowed accurate measurement of focal length and surface deviations of the mirror under test. A flux distribution of a typical solar concentrator placed directly on the solar beam gives measurements of performance at the focal point of the parabolic surface.