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: 3C-SiC (beta-SiC) and 6H-SiC p-n junction diodes have been fabricated in regions of both 3C-SiC and 6H-SiC epitaxial layers which were grown side-by-side on low-tilt-angle 6H-SiC substrates via a chemical vapor deposition (CVD) process. Several runs of diodes exhibiting state-of-the-art electrical characteristics were produced, and performance characteristics were measured and compared as a function of doping, temperature, and polytype. The first 3C-SiC diodes which rectify to reverse voltages in excess of 300 V were characterized, representing a six-fold blocking voltage improvement over experimental 3C-SiC diodes produced by previous techniques. When placed under sufficient forward bias, the 3C-SiC diodes emit significantly bright green-yellow light while the 6H-SiC diodes emit in the blue-violet. The 6H-SiC p-n junction diodes represent the first reported high-quality 6H-SiC devices to be grown by CVD on very low-tilt-angle (less than 0.5 deg off the (0001) silicon face) 6H substrates. The reverse leakage current of a 200 micron diameter circular device at 1100 V reverse bias was less than 20 nA at room temperature, and excellent rectification characteristics were demonstrated at the peak characterization temperature of 400 C.

Description: An important aspect of an ASIC (Application Specific Integrated Circuit) design process is verification. The design must not only be functionally accurate, but it must also maintain the correct timing. After a circuit has been laid out, one can utilize the Back Annotation (BA) method to simulate the design and obtain an accurate estimate of performance. However, this can lead to major design changes. It is therefore preferable to eliminate potential problems early in this process. IFA, the Intelligent Front Annotation program, assists in verifying the timing of the ASIC early in the design process. Many difficulties can arise during ASIC design. In a synchronous design, both long path and short path problems can be present. In modern ASIC technologies, the delay through a gate is very dependent on loading. This loading has two main components, the capacitance of the gates being driven and the capacitance of the metal tracks (wires). When using GaAs gate arrays, the metal line capacitance is often the dominating factor. Additionally, the RC delay through the wire itself is significant in sub-micron technologies. Since the wire lengths are unknown before place and route of the entire chip, this would seem to postpone any realistic timing verification until towards the end of the design process, obviously an undesirable situation. The IFA program estimates the delays in an ASIC before layout. Currently the program is designed for Vitesse GaAs gate arrays and, for input, requires the expansion file which is output by the program GED; however, the algorithm is appropriate for many different ASIC types and CAE platforms. IFA is especially useful for devices whose delay is extremely dependent on the interconnection wiring. It estimates the length of the interconnects using information supplied by the user and information in the netlist. The resulting wire lengths are also used to constrain the Place and Route program, ensuring reasonable results. IFA takes locality into account to give a better estimate of wire length, as well as known factors such as fanout and drive. Although the exact location of a cell is not known, an estimate of the wire length can be calculated from the location of the net in the ASIC design structure hierarchy. The length of each net is estimated using the IFA program. This length is then used to run timing analysis or simulation on the design using IFA estimated delay values and to define constraints for Place and Route. Place and Route will use the constraints as limiting values, along with the floor-plan information, to assist the placement. The IFA program has been successfully used in the design of three 350K gate GaAs chips. IFA is written in C language for Sun series computers running SunOS. It is designed to accept input files which are generated by the program GED (CADENCE Design Systems, Inc.; San Jose, CA; 408-943-1234). Sample executables for Sun4 series computers are provided with the distribution medium. IFA requires 32M of RAM for execution. The standard distribution medium is a .25 inch streaming magnetic tape cartridge in UNIX tar format. Documentation is included in the price of the program. IFA was developed in 1992 and is a copyrighted work with all copyright vested in NASA.

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: A computerized interactive harness engineering program has been developed to provide an inexpensive, interactive system which is designed for learning and using an engineering approach to interconnection systems. PCACE is basically a database system that stores information as files of individual connectors and handles wiring information in circuit groups stored as records. This directly emulates the typical manual engineering methods of data handling, thus making the user interface to the program very natural. Data files can be created, viewed, manipulated, or printed in real time. The printed ouput is in a form ready for use by fabrication and engineering personnel. PCACE also contains a wide variety of error-checking routines including connector contact checks during hardcopy generation. The user may edit existing harness data files or create new files. In creating a new file, the user is given the opportunity to insert all the connector and harness boiler plate data which would be part of a normal connector wiring diagram. This data includes the following: 1) connector reference designator, 2) connector part number, 3) backshell part number, 4) cable reference designator, 5) cable part number, 6) drawing revision, 7) relevant notes, 8) standard wire gauge, and 9) maximum circuit count. Any item except the maximum circuit count may be left blank, and any item may be changed at a later time. Once a file is created and organized, the user is directed to the main menu and has access to the file boiler plate, the circuit wiring records, and the wiring records index list. The organization of a file is such that record zero contains the connector/cable boiler plate, and all other records contain circuit wiring data. Each wiring record will handle a circuit with as many as nine wires in the interface. The record stores the circuit name and wire count and the following data for each wire: 1) wire identifier, 2) contact, 3) splice, 4) wire gauge if different from standard, 5) wire/group type, 6) wire destination, and 7) note number. The PCACE record structure allows for a wide variety of wiring forms using splices and shields, yet retains sufficient structure to maintain ease of use. PCACE is written in TURBO Pascal 3.0 and has been implemented on IBM PC, XT, and AT systems under DOS 3.1 with a memory of 512K of 8 bit bytes, two floppy disk drives, an RGB monitor, and a printer with ASCII control characters. PCACE was originally developed in 1983, and the IBM version was released in 1986.

Description: The LOOP computer program was written to simulate the Automatic Frequency Control (AFC) subsystem of a Differential Minimum Shift Keying (DMSK) receiver with a bit rate of 2400 baud. The AFC simulated by LOOP is a first order loop configuration with a first order R-C filter. NASA has been investigating the concept of mobile communications based on low-cost, low-power terminals linked via geostationary satellites. Studies have indicated that low bit rate transmission is suitable for this application, particularly from the frequency and power conservation point of view. A bit rate of 2400 BPS is attractive due to its applicability to the linear predictive coding of speech. Input to LOOP includes the following: 1) the initial frequency error; 2) the double-sided loop noise bandwidth; 3) the filter time constants; 4) the amount of intersymbol interference; and 5) the bit energy to noise spectral density. LOOP output includes: 1) the bit number and the frequency error of that bit; 2) the computed mean of the frequency error; and 3) the standard deviation of the frequency error. LOOP is written in MS SuperSoft FORTRAN 77 for interactive execution and has been implemented on an IBM PC operating under PC DOS with a memory requirement of approximately 40K of 8 bit bytes. This program was developed in 1986.

Description: Accurate computer modeling of passive circular or rectangular waveguide components is often required during the design phase for optimizing frequency response and/or determining the tolerance required on components in order to meet radio frequency specifications. RWGSCAT/CWGSCAT is capable of modeling both types of waveguide components. The Scattering Matrix Program for Circular Waveguide Junctions, CWGSCAT, computes the scattering matrix for a circular waveguide. This includes a dual mode horn and certain types of corrugated horns. RWGSCAT, Rectangular WaveGuide junction SCATtering program, solves for the scattering properties of a rectangular waveguide device, such as a smooth or corrugated rectangular horn, step transformer, or filter. RWGSCAT and CWGSCAT are also available separately as NPO-19091 and NPO-18708, respectively. Many circular waveguide devices can be represented either exactly or approximately as a series of circular waveguide sections which have a common center. In addition, smooth tapers and horns of arbitrary profile may be approximated by a series of small steps. Devices that may be analyzed in this fashion include a simple waveguide step discontinuity, such as that used in a dual mode horn, a stepped matching section, or a corrugated waveguide section with constant varying slot depth. CWGSCAT will accurately predict the reflection and transmission characteristics of such devices, taking into account higher order mode excitation if it occurs as well as multiple reflections and stored energy at each discontinuity. For large devices, with respect to a wavelength where many modes may propagate, the reflection and transmission properties may be required for a higher order mode or series of modes exciting the device. Such interactions are represented best by defining a scattering matrix for the device. The matrix can be determined by using mode matching at each discontinuity present. The results for individual discontinuities are then cascaded to get the matrix for the entire device. Frequently, rectangular waveguide components may be represented either exactly or approximately as a number of different size rectangular waveguides which are connected in series. RWGSCAT will model such devices and accurately predict the reflection and transmission characteristics, taking into account higher order (other than dominant TE 10) mode excitation if it occurs, as well as multiple reflections and stored energy at each discontinuity. For devices which are large with respect to the wavelength of operation, the characteristics of the device may be required for computing a higher order mode or a number of higher order modes exciting the device. Such interactions can be represented by defining a scattering matrix for each discontinuity in the device, and then cascading the individual scattering matrices in order to determine the scattering matrix for the overall device. The individual matrices are obtained using the mode matching method. RWGSCAT and CWGSCAT are written in FORTRAN 77 for IBM PC series and compatible computers running MS-DOS. They have been successfully compiled and implemented using Lahey FORTRAN 77 under MS-DOS. Sample MS-DOS executables and sample input data files are provided on the distribution media. RWGSCAT requires 377K of RAM for execution. CWGSCAT requires 355K of RAM for execution. The standard distribution medium for this program is a set of two 5.25 inch 360K MS-DOS format diskettes. The contents of the diskettes are compressed using the PKWARE archiving tools. The utility to unarchive the files, PKUNZIP.EXE, is included. An electronic copy of the documentation is included on the distribution medium in LaTEX format. RWGSCAT was developed in 1993. CWGSCAT was developed in 1987, and this version was released in 1991. RWGSCAT and CWGSCAT are copyrighted works with all copyrights vested in NASA.

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: Resistivity of a Thin Film Deposited on a Conductive Substrate is a computer program developed to aid in the solution of the class of problems where resistivity measurements are needed for a substance deposited on a substrate of higher resistivity than the deposited layer. One of the ways in which a semiconductor material is characterized is by measurement of its resistivity. In the development of silicon carbide (SiC) for use as a semiconductor material for high temperature applications, it became necessary to measure the resistivity of the thin SiC film while it was still attached to the silicon upon which it had been grown epitaxially. The problem is that the presence of the silicon substrate will introduce error in the measured resistivity of the SiC. This program assumes that the resistivity of a thin film of conducting material deposited on another layer of conducting material is measured using the four-point probe. Using the four-point probe measurements, this program calculates the "true" resistivity of the deposited layer on a substrate of finite and different resistivity. Starting from basic principles, an expression for the ratio of measured voltage difference to injected current is developed. This expression involves the probe spacing, relative thicknesses of the layers, and the substrate resistivity as parameters, as well as the unknown resistivity of the deposited layer. The unknown resistivity can be found by iteratively evaluating the theoretical expression. This must be done numerically. The program is written in FORTRAN 77 and targeted for use on an IBM PC or compatible. It can be modified for use on any machine with a FORTRAN 77 compiler. It requires 46K of memory and has been implemented under MS-DOS 3.2.1. The program was developed in 1986.

Description: Mutual Coupling Program for Circular Waveguide-fed Aperture Array (CWG) was developed to calculate the electromagnetic interaction between elements of an antenna array of circular apertures with specified aperture field distributions. The field distributions were assumed to be a superposition of the modes which could exist in a circular waveguide. Various external media were included to provide flexibility of use, for example, the flexibility to determine the effects of dielectric covers (i.e., thermal protection system tiles) upon the impedance of aperture type antennas. The impedance and radiation characteristics of planar array antennas depend upon the mutual interaction between all the elements of the array. These interactions are influenced by several parameters (e.g., the array grid geometry, the geometry and excitation of each array element, the medium outside the array, and the internal network feeding the array.) For the class of array antenna whose radiating elements consist of small holes in a flat conducting plate, the electromagnetic problem can be divided into two parts, the internal and the external. In solving the external problem for an array of circular apertures, CWG will compute the mutual interaction between various combinations of circular modal distributions and apertures. CWG computes the mutual coupling between various modes assumed to exist in circular apertures that are located in a flat conducting plane of infinite dimensions. The apertures can radiate into free space, a homogeneous medium, a multilayered region or a reflecting surface. These apertures are assumed to be excited by one or more modes corresponding to the modal distributions in circular waveguides of the same cross sections as the apertures. The apertures may be of different sizes and also of different polarizations. However, the program assumes that each aperture field contains the same modal distributions, and calculates the complex scattering matrix between all mode and aperture combinations. The scattering matrix can then be used to determine the complex modal field amplitudes for each aperture with a specified array excitation. CWG is written in VAX FORTRAN for DEC VAX series computers running VMS (LAR-15236) and IBM PC series and compatible computers running MS-DOS (LAR-15226). It requires 360K of RAM for execution. To compile the source code for the PC version, the NDP Fortran compiler and linker will be required; however, the distribution medium for the PC version of CWG includes a sample MS-DOS executable which was created using NDP Fortran with the -vms compiler option. The standard distribution medium for the PC version of CWG is a 3.5 inch 1.44Mb MS-DOS format diskette. The standard distribution medium for the VAX version of CWG is a 1600 BPI 9~track magnetic tape in DEC VAX BACKUP format. The VAX version is also available on a TK50 tape cartridge in DEC VAX BACKUP format. Both machine versions of CWG include an electronic version of the documentation in Microsoft Word for Windows format. CWG was developed in 1993 and is a copyrighted work with all copyright vested in NASA.

Description: Mutual Coupling Program for Circular Waveguide-fed Aperture Array (CWG) was developed to calculate the electromagnetic interaction between elements of an antenna array of circular apertures with specified aperture field distributions. The field distributions were assumed to be a superposition of the modes which could exist in a circular waveguide. Various external media were included to provide flexibility of use, for example, the flexibility to determine the effects of dielectric covers (i.e., thermal protection system tiles) upon the impedance of aperture type antennas. The impedance and radiation characteristics of planar array antennas depend upon the mutual interaction between all the elements of the array. These interactions are influenced by several parameters (e.g., the array grid geometry, the geometry and excitation of each array element, the medium outside the array, and the internal network feeding the array.) For the class of array antenna whose radiating elements consist of small holes in a flat conducting plate, the electromagnetic problem can be divided into two parts, the internal and the external. In solving the external problem for an array of circular apertures, CWG will compute the mutual interaction between various combinations of circular modal distributions and apertures. CWG computes the mutual coupling between various modes assumed to exist in circular apertures that are located in a flat conducting plane of infinite dimensions. The apertures can radiate into free space, a homogeneous medium, a multilayered region or a reflecting surface. These apertures are assumed to be excited by one or more modes corresponding to the modal distributions in circular waveguides of the same cross sections as the apertures. The apertures may be of different sizes and also of different polarizations. However, the program assumes that each aperture field contains the same modal distributions, and calculates the complex scattering matrix between all mode and aperture combinations. The scattering matrix can then be used to determine the complex modal field amplitudes for each aperture with a specified array excitation. CWG is written in VAX FORTRAN for DEC VAX series computers running VMS (LAR-15236) and IBM PC series and compatible computers running MS-DOS (LAR-15226). It requires 360K of RAM for execution. To compile the source code for the PC version, the NDP Fortran compiler and linker will be required; however, the distribution medium for the PC version of CWG includes a sample MS-DOS executable which was created using NDP Fortran with the -vms compiler option. The standard distribution medium for the PC version of CWG is a 3.5 inch 1.44Mb MS-DOS format diskette. The standard distribution medium for the VAX version of CWG is a 1600 BPI 9~track magnetic tape in DEC VAX BACKUP format. The VAX version is also available on a TK50 tape cartridge in DEC VAX BACKUP format. Both machine versions of CWG include an electronic version of the documentation in Microsoft Word for Windows format. CWG was developed in 1993 and is a copyrighted work with all copyright vested in NASA.

Description: The SMall EXplorer (SMEX) Fast Auroral SnapshoT (FAST) spacecraft was developed to investigate plasma physics of auroral phenomena at high orbital altitude. The FAST satellite comprises a variety of deployable booms with sensors on the ends, and instruments that protrude from the main body of the spacecraft to obtain the plasma and electromagnetic fields data. This required the plasma disturbance around the satellite to be kept to a minimum. A non deployable, body mounted solar array was implemented. This led to the design of a light weight solar array substrate with a high degree of structural integrity.

Description: This paper describes how expansions in leaky (or improper) modes may be used to represent the continuous spectrum in an open radiating waveguide. The technique requires a thorough knowledge of the life history of the improper modes as they migrate from improper to proper Riemann surfaces. The method is illustrated by finding the electric field resulting from an impulsively forced current located in the free space above a grounded dielectric slab.

Description: ROTRAN1 is a computer program to calculate the impedance and current gain of a simple transformer. Inputs to the program are primary resistance, primary inductance, secondary (load) resistance, secondary inductance, and mutual inductance. ROTRAN1 was written in BASICA for execution on the IBM PC personal computer. It was written in 1986.

Description: Design and analysis of Environmental Control and Life Support Systems (ECLSS) and Active Thermal Control Systems (ATCS) for spacecraft missions requires powerful software that is flexible and responsive to the demands of particular projects. CASE/A is an interactive trade study and analysis tool designed to increase productivity during all phases of systems engineering. The graphics-based command-driven package provides a user-friendly environment in which the engineer can analyze the performance and interface characteristics of an ECLS/ATC system. The package is useful during all phases of a spacecraft design program, from initial conceptual design trade studies to the actual flight, including pre-flight prediction and in-flight anomaly analysis. The CASE/A program consists of three fundamental parts: 1) the schematic management system, 2) the database management system, and 3) the simulation control and execution system. The schematic management system allows the user to graphically construct a system model by arranging icons representing system components and connecting the components with physical fluid streams. Version 4.1 contains 51 fully coded and documented default component routines. New components can be added by the user through the "blackbox" component option. The database management system supports the storage and manipulation of component data, output data, and solution control data through interactive edit screens. The simulation control and execution system initiates and controls the iterative solution process, displaying time status and any necessary diagnostic messages. In addition to these primary functions, the program provides three other important functional areas: 1) model output management, 2) system utility commands, and 3) user operations logic capacity. The model output management system provides tabular and graphical output capability. Complete fluid constituent mass fraction and properties data (mass flow, pressure, temperature, specific heat, density, and viscosity) is generated at user-selected output intervals and stored for reference. The Integrated Plot Utility (IPU) provides plotting capability for all data output. System utility commands are provided to enable the user to operate more efficiently in the CASE/A environment. The user is able to customize a simulation through optional operations FORTRAN logic. This user-developed code is compiled and linked with a CASE/A model and enables the user to control and timeline component operating parameters during various phases of the iterative solution process. CASE/A provides for transient tracking of the flow stream constituents and determination of their thermodynamic state throughout an ECLSS/ATCS simulation, performing heat transfer, chemical reaction, mass/energy balance, and system pressure drop analysis based on user-specified operating conditions. The program tracks each constituent through all combination and decomposition states while maintaining a mass and energy balance on the overall system. This allows rapid assessment of ECLSS designs, the impact of alternate technologies, and impacts due to changes in metabolic forcing functions, consumables usage, and system control considerations. CASE/A is written in FORTRAN 77 for the DEC VAX/VMS computer series, and requires 12Mb of disk storage and a minimum paging file quota of 20,000 pages. The program operates on the Tektronix 4014 graphics standard and VT100 text standard. The program requires a Tektronix 4014 or later graphics terminal, third party composite graphics/text terminal, or personal computer loaded with appropriate VT100/TEK 4014 emulator software. The use of composite terminals or personal computers with popular emulation software is recommended for enhanced CASE/A operations and general ease of use. The program is available on an unlabeled 9-track 6250 BPI DEC VAX BACKUP format magnetic tape. CASE/A development began in 1985 under contract to NASA/Marshall Space Flight Center. The latest version (4.1) was released in 1990. Tektronix and TEK 4014 are trademarks of Tektronix, Inc. VT100 is a trademark of Digital Equipment Corporation.

Description: BUMPERII is a modular program package employing a numerical solution technique to calculate a spacecraft's probability of no penetration (PNP) from man-made orbital debris or meteoroid impacts. The solution equation used to calculate the PNP is based on the Poisson distribution model for similar analysis of smaller craft, but reflects the more rigorous mathematical modeling of spacecraft geometry, orientation, and impact characteristics necessary for treatment of larger structures such as space station components. The technique considers the spacecraft surface in terms of a series of flat plate elements. It divides the threat environment into a number of finite cases, then evaluates each element of each threat. The code allows for impact shielding (shadowing) of one element by another in various configurations over the spacecraft exterior, and also allows for the effects of changing spacecraft flight orientation and attitude. Four main modules comprise the overall BUMPERII package: GEOMETRY, RESPONSE, SHIELD, and CONTOUR. The GEOMETRY module accepts user-generated finite element model (FEM) representations of the spacecraft geometry and creates geometry databases for both meteoroid and debris analysis. The GEOMETRY module expects input to be in either SUPERTAB Universal File Format or PATRAN Neutral File Format. The RESPONSE module creates wall penetration response databases, one for meteoroid analysis and one for debris analysis, for up to 100 unique wall configurations. This module also creates a file containing critical diameter as a function of impact velocity and impact angle for each wall configuration. The SHIELD module calculates the PNP for the modeled structure given exposure time, operating altitude, element ID ranges, and the data from the RESPONSE and GEOMETRY databases. The results appear in a summary file. SHIELD will also determine the effective area of the components and the overall model, and it can produce a data file containing the probability of penetration values per surface area for each element in the model. The SHIELD module writes this data file in either SUPERTAB Universal File Format or PATRAN Neutral File Format so threat contour plots can be generated as a post-processing feature of the FEM programs SUPERTAB and PATRAN. The CONTOUR module combines the functions of the RESPONSE module and most of the SHIELD module functions allowing determination of ranges of PNP's by looping over ranges of shield and/or wall thicknesses. A data file containing the PNP's for the corresponding shield and vessel wall thickness is produced. Users may perform sensitivity studies of two kinds. The effects of simple variations in orbital time, surface area, and flux may be analyzed by making changes to the terms in the equation representing the average number of penetrating particles per unit time in the PNP solution equation. The package analyzes other changes, including model environment, surface area, and configuration, by re-running the solution sequence with new GEOMETRY and RESPONSE data. BUMPERII can be run with no interactive output to the screen during execution. This can be particularly useful during batch runs. BUMPERII is written in FORTRAN 77 for DEC VAX series computers running under VMS, and was written for use with the finite-element model code SUPERTAB or PATRAN as both a pre-processor and a post-processor. Use of an alternate FEM code will require either development of a translator to change data format or modification of the GEOMETRY subroutine in BUMPERII. This program is available in DEC VAX BACKUP format on a 9-track 1600 BPI magnetic tape (standard distribution media) or on TK50 tape cartridge. The original BUMPER code was developed in 1988 with the BUMPERII revisions following in 1991 and 1992. SUPERTAB is a former name for I-DEAS. I-DEAS Finite Element Modeling is a trademark of Structural Dynamics Research Corporation. DEC, VAX, VMS and TK50 are trademarks of Digital Equipment Corporation.

Description: One of the most important factors in the development of nuclear rocket engine designs is to be able to accurately predict temperatures and pressures throughout a fission nuclear reactor core with axial hydrogen flow through circular coolant passages. CAC is an analytical prediction program to study the heat transfer and fluid flow characteristics of a circular coolant passage. CAC predicts as a function of time axial and radial fluid conditions, passage wall temperatures, flow rates in each coolant passage, and approximate maximum material temperatures. CAC incorporates the hydrogen properties model STATE to provide fluid-state relations, thermodynamic properties, and transport properties of molecular hydrogen in any fixed ortho-para combination. The program requires the general core geometry, the core material properties as a function of temperature, the core power profile, and the core inlet conditions as function of time. Although CAC was originally developed in FORTRAN IV for use on an IBM 7094, this version is written in ANSI standard FORTRAN 77 and is designed to be machine independent. It has been successfully compiled on IBM PC series and compatible computers running MS-DOS with Lahey F77L, a Sun4 series computer running SunOS 4.1.1, and a VAX series computer running VMS 5.4-3. CAC requires 300K of RAM under MS-DOS, 422K of RAM under SunOS, and 220K of RAM under VMS. No sample executable is provided on the distribution medium. Sample input and output data are included. The standard distribution medium for this program is a 5.25 inch 360K MS-DOS format diskette. CAC was developed in 1966, and this machine independent version was released in 1992. IBM-PC and IBM are registered trademarks of International Business Machines. Lahey F77L is a registered trademark of Lahey Computer Systems, Inc. SunOS is a trademark of Sun Microsystems, Inc. VMS is a trademark of Digital Equipment Corporation. MS-DOS is a registered trademark of Microsoft Corporation.

Description: Accurate computer modeling of passive circular waveguide components is often required during the design phase for optimizing frequency response and/or determining the tolerance required on components in order to meet radio frequency specifications. Many circular waveguide devices can be represented either exactly or approximately as a series of circular waveguide sections which have a common center. In addition, smooth tapers and horns of arbitrary profile may be approximated by a series of small steps. The Scattering Matrix Program for Circular Waveguide Junctions, CWGSCAT, computes the scattering matrix for a series of circular waveguide sections. These sections must possess the same center, but the radius and length of each section is completely arbitrary. Devices that may be analyzed include a simple waveguide step discontinuity, such as that used in a dual mode horn, a stepped matching section, or a corrugated waveguide section with constant varying slot depth. Certain types of corrugated horns may also be analyzed with this program. The model used will accurately predict the reflection and transmission characteristics of such devices, taking into account higher order mode excitation if it occurs as well as multiple reflections and stored energy at each discontinuity. For large devices, with respect to a wavelength where many modes may propagate, the reflection and transmission properties may be required for a higher order mode or series of modes exciting the device. Such interactions are represented best by defining a scattering matrix for the device. The matrix can be determined by using mode matching at each discontinuity present. The results for individual discontinuities are then cascaded to get the matrix for the entire device. CWGSCAT is written in FORTRAN to run on IBM PC series computers and compatibles running MS-DOS. It requires 355K of RAM. The standard distribution medium is a 5.25 inch 360K MS-DOS format diskette. CWGSCAT was developed in 1987, and this version was released in 1991. This program is a copyrighted work with all copyright vested in NASA. IBM is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation.

Description: Although extensive knowledge of space station design exists, the information is widely dispersed. The Space Station Freedom Program (SSFP) needs policies and procedures that ensure the use of consistent design objectives throughout its organizational hierarchy. The System Design Tradeoff Model (SDTM) produces information that can be used for this purpose. SDTM is a mathematical model of a set of possible designs for Space Station Freedom. Using the SDTM program, one can find the particular design which provides specified amounts of resources to Freedom's users at the lowest total (or life cycle) cost. One can also compare alternative design concepts by changing the set of possible designs, while holding the specified user services constant, and then comparing costs. Finally, both costs and user services can be varied simultaneously when comparing different designs. SDTM selects its solution from a set of feasible designs. Feasibility constraints include safety considerations, minimum levels of resources required for station users, budget allocation requirements, time limitations, and Congressional mandates. The total, or life cycle, cost includes all of the U.S. costs of the station: design and development, purchase of hardware and software, assembly, and operations throughout its lifetime. The SDTM development team has identified, for a variety of possible space station designs, the subsystems that produce the resources to be modeled. The team has also developed formulas for the cross consumption of resources by other resources, as functions of the amounts of resources produced. SDTM can find the values of station resources, so that subsystem designers can choose new design concepts that further reduce the station's life cycle cost. The fundamental input to SDTM is a set of formulas that describe the subsystems which make up a reference design. Most of the formulas identify how the resources required by each subsystem depend upon the size of the subsystem. Some of the formulas describe how the subsystem costs depend on size. The formulas can be complicated and nonlinear (if nonlinearity is needed to describe how designs change with size). SDTM's outputs are amounts of resources, life-cycle costs, and marginal costs. SDTM will run on IBM PC/XTs, ATs, and 100% compatibles with 640K of RAM and at least 3Mb of fixed-disk storage. A printer which can print in 132-column mode is also required, and a mathematics co-processor chip is highly recommended. This code is written in Turbo C 2.0. However, since the developers used a modified version of the proprietary Vitamin C source code library, the complete source code is not available. The executable is provided, along with all non-proprietary source code. This program was developed in 1989.

Description: The Autonomous Frequency Domain Identification program, AU-FREDI, is a system of methods, algorithms and software that was developed for the identification of structural dynamic parameters and system transfer function characterization for control of large space platforms and flexible spacecraft. It was validated in the CALTECH/Jet Propulsion Laboratory's Large Spacecraft Control Laboratory. Due to the unique characteristics of this laboratory environment, and the environment-specific nature of many of the software's routines, AU-FREDI should be considered to be a collection of routines which can be modified and reassembled to suit system identification and control experiments on large flexible structures. The AU-FREDI software was originally designed to command plant excitation and handle subsequent input/output data transfer, and to conduct system identification based on the I/O data. Key features of the AU-FREDI methodology are as follows: 1. AU-FREDI has on-line digital filter design to support on-orbit optimal input design and data composition. 2. Data composition of experimental data in overlapping frequency bands overcomes finite actuator power constraints. 3. Recursive least squares sine-dwell estimation accurately handles digitized sinusoids and low frequency modes. 4. The system also includes automated estimation of model order using a product moment matrix. 5. A sample-data transfer function parametrization supports digital control design. 6. Minimum variance estimation is assured with a curve fitting algorithm with iterative reweighting. 7. Robust root solvers accurately factorize high order polynomials to determine frequency and damping estimates. 8. Output error characterization of model additive uncertainty supports robustness analysis. The research objectives associated with AU-FREDI were particularly useful in focusing the identification methodology for realistic on-orbit testing conditions. Rather than estimating the entire structure, as is typically done in ground structural testing, AU-FREDI identifies only the key transfer function parameters and uncertainty bounds that are necessary for on-line design and tuning of robust controllers. AU-FREDI's system identification algorithms are independent of the JPL-LSCL environment, and can easily be extracted and modified for use with input/output data files. The basic approach of AU-FREDI's system identification algorithms is to non-parametrically identify the sampled data in the frequency domain using either stochastic or sine-dwell input, and then to obtain a parametric model of the transfer function by curve-fitting techniques. A cross-spectral analysis of the output error is used to determine the additive uncertainty in the estimated transfer function. The nominal transfer function estimate and the estimate of the associated additive uncertainty can be used for robust control analysis and design. AU-FREDI's I/O data transfer routines are tailored to the environment of the CALTECH/ JPL-LSCL which included a special operating system to interface with the testbed. Input commands for a particular experiment (wideband, narrowband, or sine-dwell) were computed on-line and then issued to respective actuators by the operating system. The operating system also took measurements through displacement sensors and passed them back to the software for storage and off-line processing. In order to make use of AU-FREDI's I/O data transfer routines, a user would need to provide an operating system capable of overseeing such functions between the software and the experimental setup at hand. The program documentation contains information designed to support users in either providing such an operating system or modifying the system identification algorithms for use with input/output data files. It provides a history of the theoretical, algorithmic and software development efforts including operating system requirements and listings of some of the various special purpose subroutines which were developed and optimized for Lahey FORTRAN compilers on IBM PC-AT computers before the subroutines were integrated into the system software. Potential purchasers are encouraged to purchase and review the documentation before purchasing the AU-FREDI software. AU-FREDI is distributed in DEC VAX BACKUP format on a 1600 BPI 9-track magnetic tape (standard media) or a TK50 tape cartridge. AU-FREDI was developed in 1989 and is a copyrighted work with all copyright vested in NASA.

Description: In order to optimize frequency response and determine the tolerances required to meet RF specifications, accurate computer modeling of passive rectangular waveguide components is often required. Many rectangular waveguide components may be represented either exactly or approximately as a number of different size rectangular waveguides which are connected in series. RWGSCAT, Rectangular WaveGuide junction SCATtering program, solves for the scattering properties of a waveguide device. This device must consist of a number of rectangular waveguide sections of different cross sectional area which are connected in series. Devices which fall into this category include step transformers, filters, and smooth or corrugated rectangular horns. RWGSCAT will model such devices and accurately predict the reflection and transmission characteristics, taking into account higher order (other than dominant TE 10) mode excitation if it occurs, as well as multiple reflections and stored energy at each discontinuity. For devices which are large with respect to the wavelength of operation, the characteristics of the device may be required for computing a higher order mode or a number of higher order modes exciting the device. Such interactions can be represented by defining a scattering matrix for each discontinuity in the device, and then cascading the individual scattering matrices in order to determine the scattering matrix for the overall device. The individual matrices are obtained using the mode matching method. RWGSCAT is written in FORTRAN 77 for IBM PC series and compatible computers running MS-DOS. It has been successfully compiled and implemented using Lahey FORTRAN 77 under MS-DOS. A sample MS-DOS executable is provided on the distribution medium. It requires 377K of RAM for execution. Sample input data is also provided on the distribution medium. The standard distribution medium for this program is one 5.25 inch 360K MS-DOS format diskette. The contents of the diskette are compressed using the PKWARE archiving tools. The utility to unarchive the files, PKUNZIP.EXE, is included. An electronic copy of the documentation is included on the distribution medium in LaTEX format. RWGSCAT is also offered as a bundle with a related program, CWGSCAT (Scattering Matrix Program for Circular WaveGuide Junctions). Please see the abstract for RWGSCAT/CWGSCAT (COS-10045) for information about the bundled package. RWGSCAT was developed in 1993 and is a copyrighted work with all copyright vested in NASA.

Description: DFACS is an interactive multi-user computer-aided engineering tool for system level electrical integration and cabling engineering. The purpose of the program is to provide the engineering community with a centralized database for entering and accessing system functional definitions, subsystem and instrument-end circuit pinout details, and harnessing data. The primary objective is to provide an instantaneous single point of information interchange, thus avoiding error-prone, time-consuming, and costly multiple-path data shuttling. The DFACS program, which is centered around a single database, has built-in menus that provide easy data input and access for all involved system, subsystem, and cabling personnel. The DFACS program allows parallel design of circuit data sheets and harness drawings. It also recombines raw information to automatically generate various project documents and drawings including the Circuit Data Sheet Index, the Electrical Interface Circuits List, Assembly and Equipment Lists, Electrical Ground Tree, Connector List, Cable Tree, Cabling Electrical Interface and Harness Drawings, Circuit Data Sheets, and ECR List of Affected Interfaces/Assemblies. Real time automatic production of harness drawings and circuit data sheets from the same data reservoir ensures instant system and cabling engineering design harmony. DFACS also contains automatic wire routing procedures and extensive error checking routines designed to minimize the possibility of engineering error. DFACS is designed to run on DEC VAX series computers under VMS using Version 6.3/01 of INGRES QUEL/OSL, a relational database system which is available through Relational Technology, Inc. The program is available in VAX BACKUP format on a 1600 BPI 9-track magnetic tape (standard media) or a TK50 tape cartridge. DFACS was developed in 1987 and last updated in 1990. DFACS is a copyrighted work with all copyright vested in NASA. DEC, VAX and VMS are trademarks of Digital Equipment Corporation. INGRES QUEL/OSL is a trademark of Relational Technology, Inc.

Description: This paper presents the capabilities implemented in the SAX system for an efficient operations management during its in-flight mission. SAX is an Italian scientific satellite for x-ray astronomy whose major mission objectives impose quite tight constraints on the implementation of both the space and ground segment. The most relevant mission characteristics require an operative lifetime of two years, performing scientific observations both in contact and in noncontact periods, with a low equatorial orbit supported by one ground station, so that only a few minutes of communications are available each orbit. This operational scenario determines the need to have a satellite capable of performing the scheduled mission automatically and reacting autonomously to contingency situations. The implementation approach of the on-board operations management, through which the necessary automation and autonomy are achieved, follows a hierarchical structure. This has been achieved adopting a distributed avionic architecture. Nine different on-board computers, in fact, constitute the on-board data management system. Each of them performs the local control and monitors its own functions while the system level control is performed at a higher level by the data handling applications software. The SAX on-board architecture provides the ground operators with different options of intervention by three classes of telecommands. The management of the scientific operations will be scheduled by the operation control center via dedicated operating plans. The SAX satellite flight mode is presently being integrated at Alenia Spazio premises in Turin for a launch scheduled for the end of 1995. Once in orbit, the SAX satellite will be subject to intensive check-out activities in order to verify the required mission performances. An overview of the envisaged procedure and of the necessary on-ground activities is therefore depicted as well.

Description: Galileo sequence design and integration are supported by a suite of formal software tools. Sequence review, however, is largely a manual process with reviewers scanning hundreds of pages of cryptic computer printouts to verify sequence correctness. Beginning in 1990, a series of small, PC based sequence review tools evolved. Each tool performs a specific task but all have a common 'look and feel'. The narrow focus of each tool means simpler operation, and easier creation, testing, and maintenance. Benefits from these tools are (1) decreased review time by factors of 5 to 20 or more with a concomitant reduction in staffing, (2) increased review accuracy, and (3) excellent returns on time invested.

Description: A feasibility study is presented concerning an application of a superconducting linear synchronous motor (LSM) to a large-scale rocket launcher, whose acceleration guide tube of LSM armature windings is constructed 1,500 meters under the ground. The rocket is released from the linear launcher just after it gets to a peak speed of about 900 kilometers per hour, and it flies out of the guide tube to obtain the speed of 700 kilometers per hour at the height of 100 meters above ground. The linear launcher is brought to a stop at the ground surface for a very short time of 5 seconds by a quick control of deceleration. Very large current variations in the single-layer windings of the LSM armature, which are produced at the higher speed region of 600 to 900 kilometers per hour, are controlled successfully by adopting the double-layer windings. The proposed control method makes the rocket launcher ascend stably in the superconducting LSM system, controlling the Coriolis force.

Description: It is clear that computer software is needed to assist in the generation of the equations of motion for complex, flexible spacecraft. Daniel Poelaert of ESTEC has developed the software DISTEL with which he has modeled the structural dynamics for different satellites. He is interested in expanding the capabilities of DISTEL to include structural damping and control systems. Unfortunately, the software has not been released. The author has developed similar software, PDEMOD, which has been used to model the Spacecraft control Laboratory Experiment (SCOLE), the Solar Array Flight Experiment (SAFE), the Mini-MAST truss, and the LACE satellite. PDEMOD has been used also for optimal parameter estimation and integrated control-structures design. PDEMOD is also being extended to include structural damping and control systems which are imbedded into the same equations for the structural dynamics. This paper will address the formulation of the equations for the structural dynamics of spacecraft structures which are constructed of a 3-dimensional arrangement of rigid bodies and flexible beam elements. Control system dynamics are imbedded into the same equations so that model order reduction approximations are not necessary. The input data consists of the physical data of the elements and the topological information describing how the elements are connected. PDEMOD accomplishes the following: (1) automatically assembles the equations of motion for the entire structural model; (2) calculates the modal frequencies; (3) calculates the mode shapes; (4) generates perspective views of the mode shapes; and (5) forms selected transfer functions. The software PDEMOD continues to be developed to provide additional features to assist in analyzing and synthesizing control and structural systems for flexible spacecraft.

Description: Sliding mode control became very popular recently because it makes the closed loop system highly insensitive to external disturbances and parameter variations. Sliding algorithms for flexible structures have been used previously, but these were based on finite-dimensional models. An extension of this approach for differential-difference systems is obtained. That makes if possible to apply sliding-mode control algorithms to the variety of nondispersive flexible structures which can be described as differential-difference systems. The main idea of using this technique for dispersive structures is to reduce the order of the controlled part of the system by applying an integral transformation. We can say that transformation 'absorbs' the dispersive properties of the flexible structure as the controlled part becomes dispersive.

Description: This paper presents an overview of the recent advances in system identification for modal testing and control of large flexible structures. Several techniques are discussed including the Observer/Kalman Filter Identification, the Observer/Controller Identification, and the State-Space System Identification in the Frequency Domain. The System/Observer/Controller Toolbox developed at NASA Langley Research Center is used to show the applications of these techniques to real aerospace structures such as the Hubble spacecraft telescope and the active flexible aircraft wing.

Description: The following topics are discussed: (1) modeling of articulated spacecraft as multi-flex-body systems; (2) nonlinear attitude control by adaptive partial feedback linearizing (PFL) control; (3) attitude dynamics and control for SSF/MRMS; and (4) performance analysis results for attitude control of SSF/MRMS.

Description: The work presented is motivated by the need for a national satellite rescue policy, not the ad hoc policy now in place. In studying different approaches for a national policy, the issue of capture and stabilization of a tumbling spacecraft must be addressed. For a rescue mission involving a tumbling spacecraft, it may be advantageous to have a rescue vehicle which is compact and 'rigid' during the rendezvous/capture phase. After capture, passive stabilization techniques could be utilized as an efficient means of detumbling the resulting system (i.e., both the rescue vehicle and captures spacecraft). Since the rescue vehicle is initially compact and 'rigid,' significant passive stabilization through energy dissipation can only be achieved through the deployment of flexible appendages. Once stabilization is accomplished, retraction of the appendages before maneuvering the system to its final destination may also prove advantageous. It is therefore of paramount interest that we study the effect of appendage deployment/retraction on the attitude stability of a spacecraft. Particular interest should be paid to appendage retraction, since if this process is destabilizing, passive stabilization as proposed may not be useful. Over the past three decades, it has been an 'on-again-off-again affair' with the problem of spacecraft appendage deployment. In most instances, these studies have been numerical simulations of specific spacecraft configurations for which there were specific concerns. The primary focus of these studies was the behavior of the appendage during deployment; the effects of appendage retraction was considered only in one of these studies. What is missing in the literature is a thorough study of the effects of appendage deployment/retraction on the attitude stability of a spacecraft. This paper presents a rigorous analysis of the stability of a spinning spacecraft during the deployment or the retraction of an appendage. The analysis is simplified such that meaningful insights into the problem can be inferred; it is not overly simplified such that critical dynamical behavior is neglected. The system is analyzed assuming that the spacecraft hub is rigid. The appendage deployment mechanism is modeled as a point mass on a massless rod whose length undergoes prescribed changes. Simplified flexibility effects of the appendage are included. The system is examined for stability by linearizing the equations in terms of small deviations from steady, noninterfering coning motion. Routh's procedure for analyzing small deviations from steady motion in dynamical systems is utilized in the analysis. The system of equations are nondimensionalized to facilitate parametric studies. The results are presented in terms of a reduced number of nondimensional parameters so that some general conclusions may be drawn. Verification of the linear analysis is presented through numerical simulations of the complete nonlinear, nonautonomous, coupled equations.

Description: The energy absorber that was developed for the CETA (Crew Equipment and Translation Aid) on Space Station Freedom is a metal on metal frictional type and has a load regulating feature that prevents excessive stroking loads from occurring while in operation. This paper highlights some of the design and operating aspects and the testing of this energy absorber.

Description: A useful adjunct to the manned space station would be a self-contained free-flying laboratory (RoboLab). This laboratory would have a robot operated under telepresence from the space station or ground. Long duration experiments aboard RoboLab could be performed by astronauts or scientists using telepresence to operate equipment and perform experiments. Operating the lab by telepresence would eliminate the need for life support such as food, water and air. The robot would be capable of motion in three dimensions, have binocular vision TV cameras, and two arms with manipulators to simulate hands. The robot would move along a two-dimensional grid and have a rotating, telescoping periscope section for extension in the third dimension. The remote operator would wear a virtual reality type headset to allow the superposition of computer displays over the real-time video of the lab. The operators would wear exoskeleton type arms to facilitate the movement of objects and equipment operation. The combination of video displays, motion, and the exoskeleton arms would provide a high degree of telepresence, especially for novice users such as scientists doing short-term experiments. The RoboLab could be resupplied and samples removed on other space shuttle flights. A self-contained RoboLab module would be designed to fit within the cargo bay of the space shuttle. Different modules could be designed for specific applications, i.e., crystal-growing, medicine, life sciences, chemistry, etc. This paper describes a RoboLab simulation using virtual reality (VR). VR provides an ideal simulation of telepresence before the actual robot and laboratory modules are constructed. The easy simulation of different telepresence designs will produce a highly optimum design before construction rather than the more expensive and time consuming hardware changes afterwards.

Description: A state-of-the-art instrumentation amplifier capable of being used with most types of transducers has been developed at the Kennedy Space Center. This Universal Signal Conditioning Amplifier (USCA) can eliminate costly measurement setup item and troubleshooting, improve system reliability and provide more accurate data than conventional amplifiers. The USCA can configure itself for maximum resolution and accuracy based on information read from a RAM chip attached to each transducer. Excitation voltages or current are also automatically configured. The amplifier uses both analog and digital state-of-the-art technology with analog-to-digital conversion performed in the early stages in order to minimize errors introduced by offset and gain drifts in the analog components. A dynamic temperature compensation scheme has been designed to achieve and maintain 12-bit accuracy of the amplifier from 0 to 70 C. The digital signal processing section allows the implementation of digital filters up to 511th order. The amplifier can also perform real-time linearizations up to fourth order while processing data at a rate of 23.438 kS/s. Both digital and analog outputs are available from the amplifier.

Description: The insulation testing and analysis consists of: identifying and prioritizing NASA wiring requirements; selecting candidate wiring constructions; developing test matrix and formulating test program; managing, coordinating, and conducting tests; and analyzing and documenting data, establishing guidelines and recommendations.

Description: The topics are presented in viewgraph form and include the following: wiring responsibilities; purpose of the program; measurement of program effectiveness; results; and summary.

Description: The topics are presented in viewgraph form and include the following: function of the Space Station Freedom (SSF) Flat Collector Circuit (FCC); requirements of the FCC which affect the selection of the insulation material; data to support the selection of the FCC insulation material; development history; modified design; coverlay testing; effects on modified design on FCC; arc tracking tests performed on FCC; and arc tracking test results.

Description: The objective of this program is to identify, develop, and demonstrate an optimum wire insulation system that is capable of continuous operation at 300 C. The system is to possess a combination of superior electrical (AC or DC), mechanical, and physical properties over the KAPTON (trademark) derived insulations described in MIL-W-81381 and those hybrid constructions identified in Air Force contract F33615-89-C-5606, commonly known as TKT constructions.

Description: Development of a new test method suitable for the assessment of the resistance of aerospace cables to arc tracking for different specific environmental and network conditions of spacecraft is given in view-graph format. The equipment can be easily adapted for tests at different realistic electrical network conditions incorporating circuit protection and the test system works equally well whatever the test atmosphere. Test results confirm that pure Kapton insulated wire has bad arcing characteristics and ETFE insulated wire is considerably better in air. For certain wires, arc tracking effects are increased at higher oxygen concentrations and significantly increased under vacuum. All tests on different cable insulation materials and in different environments, including enriched oxygen atmospheres, resulted in a more or less rapid extinguishing of all high temperature effects at the beginning of the post-test phase. In no case was a self-maintained fire initiated by the arc.

Description: Autonomy is needed for future spacecraft to solve the problems of human operator overload and transmission delay. This paper describes the autonomous spacecraft executive for rendezvous and docking. It is an onboard expert system and has decision making capability for mission planning of nominal and contingency cases. The executive has been developed and verified using a hardware motion based simulator.

Description: Results of a trade study addressing the issues and benefits in using carbon fiber reinforced composites for the Magnetosphere Imager (MI) spacecraft are presented. The MI mission is now part of the Sun/Earth Connection Program. To qualify for this category, new technology and innovative methods to reduce the cost and size have to be considered. Topics addressed cover: (1) what is a composite, including advantages and disadvantages of composites and carbon/graphite fibers; and (2) structural design for MI, including composite design configuration, material selection, and analysis of composite structures.

Description: FEMOT is a finite element program for solving the nonlinear magnetostatic problem. This version uses nonlinear, Newton first order elements. The code can be used for electric motor design and analysis. FEMOT can be embedded within an optimization code that will vary nodal coordinates to optimize the motor design. The output from FEMOT can be used to determine motor back EMF, torque, cogging, and magnet saturation. It will run on a PC and will be available to anyone who wants to use it.

Description: Electrochemical Impedance Spectroscopy (EIS) is a valuable tool for investigating the chemical and physical processes occurring at electrode surfaces. It offers information about electron transfer at interfaces, kinetics of reactions, and diffusion characteristics of the bulk phase between the electrodes. For battery cells, this technique offers another advantage in that it can be done without taking the battery apart. This non-destructive analysis technique can thus be used to gain a better understanding of the processes occurring within a battery cell. This also raises the possibility of improvements in battery design and identification or prediction of battery characteristics useful in industry and aerospace applications. EIS as a technique is powerful and capable of yielding significant information about the cell, but it also requires that the many parameters under investigation can be resolved. This implies an understanding of the processes occurring in a battery cell. Many battery types were surveyed in this work, but the main emphasis was on nickel/metal hydride batteries.

Description: There is a great need to develop a system that can measure accurately atmospheric wind profiles because an accurate data of wind profiles in the atmosphere constitutes single most input for reliable simulations of global climate numerical methods. Also such data helps us understand atmospheric circulation and climate dynamics better. Because of this need for accurate wind measurements, a space-based Laser Atmospheric Winds Sounder (LAWS) is being designed at MSFC to measure wind profiles in the lower atmosphere of the earth with an accuracy of 1 m/s at lower altitudes to 5m/s at higher altitudes. This system uses an orbiting spacecraft with a pulsed laser source and measures the Doppler shift between the transmitted and received frequencies to estimate the atmospheric wind velocities. If a significant return from the ground (sea) is possible, the spacecraft speed and height are estimated from it and these results and the Doppler shift are then used to estimate the wind velocities in the atmosphere. It is expected that at the proposed wavelengths, there will be enough backscatter from the aerosols but there may no be significant return from the ground. So a coherent (heterodyne) detection system is being proposed for signal processing because it can provide high signal to noise ratio and sensitivity and thus make the best use of low ground return. However, for a heterodyne detection scheme to provide the best results, it is important that the receiving aperture be aligned properly for the proposed wind sounder, this amounts to only a few microradians tolerance in alignment. It is suspected that the satellite motion relative to the ground may introduce errors in the order of a few microradians because of special relativity. Hence, the problem of laser scattering off a moving fixed target when the source and receiver are moving, which was not treated in the past in the literature, was analyzed in the following, using relativistic electrodynamics and applied to the case of the space-based coherent lidar, assuming flat ground. Here an interest in developing analytical expression for the location of the receiving point for the return with respect to the satellite, receiving angle and Doppler shift in frequency and amount of tip, all as measured in the satellite moving coordinate system and the diffuse scattering angle at the ground which does not require any compensation. All the three cases of retro-reflection, specular reflection and diffuse scattering by the ground should be treated though retro-reflection and diffuse scattering are more important.

Description: The NASDA office of R&D is studying an automatic technique to capture and berth free-floating satellites using a robot arm on another satellite. A demonstration experiment plan with the Japanese engineering test satellite ETS-7 is being developed based on the basic research on the ground. The overview and key technologies of this experiment plan are presented, and future applications of the automatic capture technique are also reviewed.

Description: ETS-7 (Engineering Test Satellite #7) is an experimental satellite for the in-orbit experiment of the Rendezvous Docking (RVD) and the space robot (RBT) technologies. ETS-7 is a set of two satellites, a chaser satellite and a target satellite. Both satellites will be launched together by NASDA's H-2 rocket into a low earth orbit. Development of ETS-7 started in 1990. Basic design and EM (Engineering Model) development are in progress now in 1994. The satellite will be launched in mid 1997 and the above in-orbit experiments will be conducted for 1.5 years. Design of ETS-7 RBT experiment system and development status are described in this paper.

Description: A robot experiment concept of space truss telemanipulation by National Aerospace Laboratory (NAL) is described in its flight model development. The experiment will be carried out on the Engineering Test Satellite No. 7 (ETS-7) using its robot arm. The satellite is scheduled to be launched in 1997 by National Space Development Agency of Japan (NASDA). The truss flight model is composed of deployable truss system and assemble truss joint. Those truss components will be manipulated by the ETS-7 robot arm using its small grapple fixture type-N (GPF-N), and the experimental task operation will be executed from the ground control station.

Description: The Communications Research Laboratory plans to test an antenna-assembling mechanism on the Engineering Test Satellite 7. The test is one of the application missions for the space robotics experiments that will be conducted mainly by the National Space Development Agency of Japan (NASDA). The purpose of the test is to verify the ability of the antenna assembling mechanism to function in space and to experiment on the teleoperation of a space robot to develop antenna-assembling technology. We present the test experiment plans and the outline of the onboard assembling mechanism.

Description: Photovoltaic electric-powered flight is receiving a great deal of attention in the context of the United States' Unmanned Aerial Vehicle (UAV) program. This paper addresses some of the enabling technical areas and their potential solutions. Of particular interest are the long-duration, high-altitude class of UAV's whose mission it is to achieve altitudes between 60,000 and 100,000 feet, and to remain at those altitudes for prolonged periods performing various mapping and surveillance activities. Addressed herein are studies which reveal the need for extremely light-weight and efficient solar cells, high-efficiency electric motor-driven propeller modules, and power management and distribution control elements. Since the potential payloads vary dramatically in their power consumption and duty cycles, a typical load profile has been selected to provide commonality for the propulsion power comparisons. Since missions vary widely with respect to ground coverage requirements, from repeated orbiting over a localized target to long-distance routes over irregular terrain, we have also averaged the power requirements for on-board guidance and control power, as well as ground control and communication link utilization. In the context of the national technology reinvestment program, wherever possible we modeled components and materials which have been qualified for space and defense applications, yet are compatible with civilian UAV activities. These include, but are not limited to, solar cell developments, electric storage technology for diurnal operation, local and ground communications, power management and distribution, and control servo design. And finally, the results of tests conducted by Wright Laboratory on ultralight, highly efficient MOCVD GaAs solar cells purchased from EPI Materials Ltd. (EML) of the UK are presented. These cells were also used for modeling the flight characteristics of UAV aircraft.

Description: Proliferation of power electronic devices has brought in its wake both deterioration in and demand for quality power supply from the utilities. The power quality problems become apparent when the user's equipment or systems maloperate or fail. Since power quality concerns arise from a wide variety of sources and the problem fixes are better achieved from the expertise of field engineers, development of an expert system for power quality advisement seems to be a very attractive and cost-effective solution for utility applications. An expert system thus developed gives an understanding of the adverse effects of power quality related problems on the system and could help in finding remedial solutions. The paper reports the design of a power quality advisement expert system being developed using CLIPS 6.0. A brief outline of the power quality concerns is first presented. A description of the knowledge base is next given and details of actual implementation include screen output from the program.

Description: The recent discovery of High Temperature Superconductors (HTS) with superconducting transition temperature, T(sub c), above the boiling point of liquid nitrogen has opened the door for using these materials in new and practical applications. These materials have zero resistance to electric current, have the capability of carrying large currents and as such have the potential to be used in high magnetic field applications. One of the space applications that can use superconductors is electromagnetic launch of payloads to low-earth-orbit. An electromagnetic gun-type launcher can be used in small payload systems that are launched at very high velocity, while sled-type magnetically levitated launcher can be used to launch larger payloads at smaller velocities. Both types of launchers are being studied by NASA and the aerospace industry. The use of superconductors will be essential in any of these types of launchers in order to produce the large magnetic fields required to obtain large thrust forces. Low Temperature Superconductor (LTS) technology is mature enough and can be easily integrated in such systems. As for the HTS, many leading companies are currently producing HTS coils and magnets that potentially can be mass-produced for these launchers. It seems that designing and building a small-scale electromagnetic launcher is the next logical step toward seriously considering this method for launching payloads into low-earth-orbit. A second potential application is the use of HTS to build sensitive portable devices for the use in Non Destructive Evaluation (NDE). Superconducting Quantum Interference Devices (SQUID's) are the most sensitive instruments for measuring changes in magnetic flux. By using HTS in SQUID's, one will be able to design a portable unit that uses liquid nitrogen or a cryocooler pump to explore the use of gradiometers or magnetometers to detect deep cracks or corrosion in structures. A third use is the replacement of Infra-Red (IR) sensor leads on Earth Orbit Systems (EOS) with HTS leads. IR detectors on these EOS missions are cooled to a 4.2K to improve their signal to noise ratio. They are connected to data acquisitions systems using manganin wires (low thermal conductors) to reduce the heat load on the cryogen. Replacing these wires with HTS leads will increase the lifetime of these missions by about 50 percent. This is a promising application that is ready for actual implementation on such systems. The analysis also show that an the number of IR detectors increase in larger EOS systems, substantial increase in the lifetime of each mission will be realized by using HTS leads instead of the manganin ones.

Description: This paper describes the Starpicker expert system, a tool for spacecraft operations planning. Both programmatic and technical aspects are discussed.

Description: Cost savings opportunities over the life cycle of a product are highest in the early exploratory phase when different design alternatives are evaluated not only for their performance characteristics but also their methods of fabrication which really control the ultimate manufacturing costs of the product. In the past, Design-To-Cost methodologies for spacecraft design concentrated on the sizing and weight issues more than anything else at the early so-called 'Vehicle Level' (Ref: DOD/NASA Advanced Composites Design Guide). Given the impact of manufacturing cost, the objective of this study is to identify the principal cost drivers for each materials technology and propose a quantitative approach to incorporating these cost drivers into the family of optimization tools used by the Vehicle Analysis Branch of NASA LaRC to assess various conceptual vehicle designs. The advanced materials being considered include aluminum-lithium alloys, thermoplastic graphite-polyether etherketone composites, graphite-bismaleimide composites, graphite- polyimide composites, and carbon-carbon composites. Two conventional materials are added to the study to serve as baseline materials against which the other materials are compared. These two conventional materials are aircraft aluminum alloys series 2000 and series 7000, and graphite-epoxy composites T-300/934. The following information is available in the database. For each material type, the mechanical, physical, thermal, and environmental properties are first listed. Next the principal manufacturing processes are described. Whenever possible, guidelines for optimum processing conditions for specific applications are provided. Finally, six categories of cost drivers are discussed. They include, design features affecting processing, tooling, materials, fabrication, joining/assembly, and quality assurance issues. It should be emphasized that this database is not an exhaustive database. Its primary use is to make the vehicle designer aware of some of the most important aspects of manufacturing associated with his/her choice of the structural materials. The other objective of this study is to propose a quantitative method to determine a Manufacturing Complexity Factor (MCF) for each material being contemplated. This MCF is derived on the basis of the six cost drivers mentioned above plus a Technology Readiness Factor which is very closely related to the Technology Readiness Level (TRL) as defined in the Access To Space final report. Short of any manufacturing information, our MCF is equivalent to the inverse of TRL. As more manufacturing information is available, our MCF is a better representation (than TRL) of the fabrication processes involved. The most likely application for MCF is in cost modeling for trade studies. On-going work is being pursued to expand the potential applications of MCF.

Description: Space station requirements for power have resulted in a need for photovoltaic solar arrays possessing large blanket surface area. However, due to the limited shuttle payload volume solar array designers have been driven to a deployable concept that by nature is extremely flexible. The principal support for this array system is the Folding Articulating Square Truss Mast (FASTMast). In order to accomodate service loads the FASTMast is expected to exhibit nonlinear behavior which could possibly result in structural instability. Presented herein are the results of the Lewis Research Center test and analysis efforts performed in an effort to characterize the FASTMast structural behavior in terms of stability. Results include those obtained from recent nonlinear testing and analysis involving a 1/10 segment of the FASTMast flight article. Implications of these results as they relate to expected behavior of the flight unit will also be discussed.

Description: Flexible, deployable arrays are an attractive alternative to conventional solar arrays for near-term and future space power applications, particularly due to their potential for high specific power and low storage volume. Combined with low-cost flexible thin-film photovoltaics, these arrays have the potential to become an enabling or an enhancing technology for many missions. In order to expedite the acceptance of thin-film photovoltaics for space applications, however, parallel development of flexible photovoltaics and the corresponding deployable structure is essential. Many innovative technologies must be incorporated in these arrays to ensure a significant performance increase over conventional technologies. For example, innovative mechanisms which employ shape memory alloys for storage latches, deployment mechanisms, and array positioning gimbals can be incorporated into flexible array design with significant improvement in the areas of cost, weight, and reliability. This paper discusses recent activities at Martin Marietta regarding the development of flexible, deployable solar array technology. Particular emphasis is placed on the novel use of shape memory alloys for lightweight deployment elements to improve the overall specific power of the array. Array performance projections with flexible thin-film copper-indium-diselenide (CIS) are presented, and government-sponsored solar array programs recently initiated at Martin Marietta through NASA and Air Force Phillips Laboratory are discussed.

Description: The objective was to design a method to measure weight and center of gravity (C.G.) location for Space Station Modules by adding sensors to the existing Rack Insertion End Effector (RIEE). Accomplishments included alternative sensor placement schemes organized into categories. Vendors were queried for suitable sensor equipment recommendations. Inverse mathematical models for each category determine expected maximum sensor loads. Sensors are selected using these computations, yielding cost and accuracy data. Accuracy data for individual sensors are inserted into forward mathematical models to estimate the accuracy of an overall sensor scheme. Cost of the schemes can be estimated. Ease of implementation and operation are discussed.

Description: It has often been proposed that a vehicle returning from Mars will use aerobraking in the Earth's atmosphere to dissipate hyperbolic excess velocity to capture into Earth orbit. Here a different system for dissipating excess velocity without expenditure of reaction mass, magnetobraking, is proposed. Magnetobraking uses the force on an electrodynamic tether in the Earth's magnetic field to produce thrust. An electrodynamic tether is deployed from the spacecraft as it approaches the Earth. The Earth's magnetic field produces a force on electrical current in the tether. If the tether is oriented perpendicularly to the Earth's magnetic field and to the direction of motion of the spacecraft, force produced by the Earth's magnetic field can be used to either brake or accelerate the spacecraft without expenditure of reaction mass. The peak acceleration on the Mars return is 0.007 m/sq sec, and the amount of braking possible is dependent on the density and current-carrying capacity of the tether, but is independent of length. A superconducting tether is required. The required critical current is shown to be within the range of superconducting technology now available in the laboratory.

Description: In this age of shrinking military, civil, and commercial space budgets, an off-the-shelf solution is needed to provide a multimission approach to spacecraft control. A standard operational interface which can be applied to multiple spacecraft allows a common approach to ground and space operations. A trend for many space programs has been to reduce operational staff by applying autonomy to the spacecraft and to the ground stations. The Spacecraft Command Language (SCL) system developed by Interface and Control Systems, Inc. (ICS) provides an off-the-shelf solution for spacecraft operations. The SCL system is designed to provide a hyper-scripting interface which remains standard from program to program. The spacecraft and ground station hardware specifics are isolated to provide the maximum amount of portability from system to system. Uplink and downlink interfaces are also isolated to allow the system to perform independent of the communications protocols chosen. The SCL system can be used for both the ground stations and the spacecraft, or as a value added package for existing ground station environments. The SCL system provides an expanded stored commanding capability as well as a rule-based expert system on-board. The expert system allows reactive control on-board the spacecraft for functions such as electrical power systems (EPS), thermal control, etc. which have traditionally been performed on the ground. The SCL rule and scripting capability share a common syntax allowing control of scripts from rules and rules from scripts. Rather than telemeter over sampled data to the ground, the SCL system maintains a database on-board which is available for interrogation by the scripts and rules. The SCL knowledge base is constructed on the ground and uploaded to the spacecraft. The SCL system follows an open-systems approach allowing other tasks to communicate with SCL on the ground and in space. The SCL system was used on the Clementine program (launched January 25, 1994) and is required to have bidirectional communications with the guidance, navigation, and control (GNC) algorithms which were written as another task. Sequencing of the spacecraft maneuvers are handled by SCL, but the low-level thruster pulse commands are handled by the GNC software. Attitude information is reported back as telemetry, allowing the SCL expert system to inference on the changing data. The Clementine SCL flight software was largely reused from another Naval Center for Space Technology (NCST) satellite program. This paper details the SCL architecture and how an off-the-shelf solution makes sense for multimission spacecraft programs. The Clementine mission will be used as a case study in the application of the SCL to a 'fast track' program. The benefits of such a system in a 'better, cheaper, faster' climate will be discussed.

Description: The agenda of this presentation includes the Orbiter wire selection requirements, the Orbiter wire usage, fabrication and test requirements, typical wiring installations, Kapton wire experience, NASA Kapton wire testing, summary, and backup data.

Description: Aerothermal environments as encountered during the reentry of spaceplanes or during the cruise of hypersonic aircrafts represent complex loading conditions for the external structures of those vehicles. In order to shield against the aerodynamic heating a special Thermal Protection System (TPS) is required which is designed as a light weight structure to reduce the weight penalty. TPS is therefore vulnerable to vibroacoustic fatigue caused by the pressure fluctuations of the environment. Because of the complex interactions between the loading forces and the resulting structural response which make an analytical treatment difficult and in order to provide means for fatigue testing IABG has designed and built a thermoacoustic facility which recently became operational. The facility is capable to produce surface temperatures up to 1.300 C at sound pressure levels up to 160 dB. This paper describes the design of the facility, some operational test work it also deals with problems associated with the facility instrumentation.

Description: A software tool, the Relative Collision Matrix (RCM), has been developed to provide a quick-look representation of the shortterm collision hazard to space systems from a fragmentation event in Earth orbit. The software performs multiple fragmentation simulations of space objects to quantify the probability of collision for a satellite or a constellation of satellites nearby. Previously, the results were displayed in a color matrix format which showed the relative hazard of each constellation. The RCM can be used for scientific research and operational assessments even though it was designed for test and evaluation applications. Because of its successful use as an analytical tool, the capabilities of RCM are being extended by enhancing the orbital hazard analysis routines, developing ballistic trajectory hazard analysis routines, and expanding the breakup modeling. Improvements are also being made to the RCM's usability and presentation quality by developing a graphical user interface and by providing graphical animated and nonanimated output.

Description: In this paper the approximation problem for a class of optimal compensators for flexible structures is considered. The particular case of a simply supported truss with an offset antenna is dealt with. The nonrational positive real optimal compensator transfer function is determined, and it is proposed that an approximation scheme based on a continued fraction expansion method be used. Comparison with the more popular modal expansion technique is performed in terms of stability margin and parameters sensitivity of the relative approximated closed loop transfer functions.

Description: The electrical utility power system requirements were determined for a Maglev line from San Diego to San Francisco and Sacramento with a maximum capacity of 12,000 passengers an hour in each direction at a speed of 300 miles per hour, or one train every 30 seconds in each direction. Basically the Maglev line requires one 50-MVA substation every 12.5 miles. The need for new power lines to serve these substations and their voltage levels are based not only on equipment loading criteria but also on limitations due to voltage flicker and harmonics created by the Maglev system. The resulting power system requirements and their costs depend mostly on the geographical area, urban or suburban with 'strong' power systems, or mountains and rural areas with 'weak' power systems. A reliability evaluation indicated that emergency power sources, such as a 10-MW battery at each substation, were not justified if sufficient redundancy is provided in the design of the substations and the power lines serving them. With a cost of $5.6 M per mile, the power system requirements, including the 12-kV DC cables and the inverters along the Maglev line, were found to be the second largest cost component of the Maglev system, after the cost of the guideway system ($9.1 M per mile), out of a total cost of $23 M per mile.

Description: Three-axis attitude determination is equivalent to finding a coordinate transformation matrix which transforms a set of reference vectors fixed in inertial space to a set of measurement vectors fixed in the spacecraft. The attitude determination problem can be expressed as a constrained optimization problem. The constraint is that a coordinate transformation matrix must be proper, real, and orthogonal. A transformation matrix can be thought of as optimal in the least-squares sense if it maps the measurement vectors to the reference vectors with minimal 2-norm errors and meets the above constraint. This constrained optimization problem is known as Wahba's problem. Several algorithms which solve Wahba's problem exactly have been developed and used. These algorithms, while steadily improving, are all rather complicated. Furthermore, they involve such numerically unstable or sensitive operations as matrix determinant, matrix adjoint, and Newton-Raphson iterations. This paper describes an algorithm which minimizes Wahba's loss function, but without the constraint. When the constraint is ignored, the problem can be solved by a straightforward, numerically stable least-squares algorithm such as QR decomposition. Even though the algorithm does not explicitly take the constraint into account, it still yields a nearly orthogonal matrix for most practical cases; orthogonality only becomes corrupted when the sensor measurements are very noisy, on the same order of magnitude as the attitude rotations. The algorithm can be simplified if the attitude rotations are small enough so that the approximation sin(theta) approximately equals theta holds. We then compare the computational requirements for several well-known algorithms. For the general large-angle case, the QR least-squares algorithm is competitive with all other know algorithms and faster than most. If attitude rotations are small, the least-squares algorithm can be modified to run faster, and this modified algorithm is faster than all but a similarly specialized version of the QUEST algorithm. We also introduce a novel measurement averaging technique which reduces the n-measurement case to the two measurement case for our particular application, a star tracker and earth sensor mounted on an earth-pointed geosynchronous communications satellite. Using this technique, many n-measurement problems reduce to less than or equal to 3 measurements; this reduces the amount of required calculation without significant degradation in accuracy. Finally, we present the results of some tests which compare the least-squares algorithm with the QUEST and FOAM algorithms in the two-measurement case. For our example case, all three algorithms performed with similar accuracy.

Description: A relatively general formulation for studying the dynamics and control of an arbitrary spacecraft with interconnected flexible bodies has been developed. This self-contained and comprehensive numerical algorithm using system modes is applicable to a large class of spacecraft configurations of contemporary and future interest. Here, versatility of the approach is demonstrated through the dynamics and control studies aimed at the evolving Space Station Freedom.

Description: The SpaceHab 1 flight on STS-57 served as a test platform for evaluation of two space station payloads. The first payload evaluated a space station maintenance concept using a sweep signal generator and a 48-channel logic analyzer to perform fault detection and isolation. Crew procedures files, test setup diagram files, and software to configure the test equipment were created on the ground and uplinked on the astronauts' voice communication circuit to perform tests in flight. In order to use these files, the portable computer was operated in a multi-window configuration. The test data transmitted to the ground allowing the ground staff to identify the cause of the fault and provide the crew with the repair procedures and diagrams. The crew successfully repaired the system under test. The second payload investigated hand soldering and de-soldering of standard components on printed circuit (PC) boards in zero gravity. It also used a new type of intra-vehicular foot restraints which uses the neutral body posture in zero-g to provide retention of the crew without their conscious attention.

Description: Very often servicing of satellites is necessary to replace components which are responsible for anomalous behavior of satellite operations due to adverse interactions with the natural space environment. A major difficulty with this diagnosis is that those responsible for diagnosing these anomalies do not have the tools to assess the role of the space environment causing the anomaly. To address this issue, we have under development a new rule-based, expert system for diagnosing spacecraft anomalies. The knowledge base consists of over two-hundred rules and provides links to historical and environmental databases. Environmental causes considered are bulk charging, single event upsets (SEU), surface charging, and total radiation dose. The system's driver translates forward chaining rules into a backward chaining sequence, prompting the user for information pertinent to the causes considered. When the user selects the novice mode, the system automatically gives detailed explanations and descriptions of terms and reasoning as the session progresses, in a sense teaching the user. As such it is an effective tutoring tool. The use of heuristics frees the user from searching through large amounts of irrelevant information and allows the user to input partial information (varying degrees of confidence in an answer) or 'unknown' to any question. The system is available on-line and uses C Language Integrated Production System (CLIPS), an expert shell developed by the NASA Johnson Space Center AI Laboratory in Houston.

Description: An alert was issued because of the arc-tracking possibilities of installed polyimide wire harnesses. MSFC undertook a program to try to enhance the safety and reliability of these harnesses. Photographs are presented showing the need for inspections of installed wiring harnesses.

Description: Electrical wires are considered as EEE parts and are covered within the ESA SCC specification series (ESA SCC 3901/XXX). This specification defines the principal properties of the wires including insulation/lay-up and electrical properties. Some additional space related materials requirements are also included, requirements such as outgassing and silver plating thickness. If a project has additional materials requirements over and above those covered by the relevant SCC specification, then additional testing is required. This is especially true for crewed spacecraft. The following topics are discussed in this context: additional requirements for manned spacecraft; flammability; arc tracking; thermal decomposition; microbial surface growth; and ageing.

Description: We ran some tests on the effect of dimming of metal halide (MH) lamps upon the stability and the spectral quality of the light output. Lamps used were a new Philips lamp HPI-T 250W, a similar Philips lamp with a few thousand burning hours and a new Osram lamp HQI-T 250W/D. The ballast was a BBC type DJ 250/2KS, the starter a BAS TORGI type MZN 250 SE and the dimmer an Elstrom Control System type ERHQ-T 250. Power was derived from a Philips stabilizer, type PE 1602. Lamp output was monitored with a PAR meter. Spectra were taken at 100% and at 50% output as measured with the PAR meter. Lamps were allowed to stabilize at any setting for 30 minutes before measurements were made. Lamp manufacturers advise against dimming for fear of poor stability and intolerable changes of the spectrum. However, none of the lamps showed a decrease in stability, no flicker or wandering of the discharge, and the changes of the spectrum were not negligible, but certainly not dramatic. Lamps of either manufacture retain their white color, relative peak heights of spectral lines did shift, but no gaps in the spectrum occurred. Spectra taken at 50% with 30 minutes intervals coincided. Differences between the new and the older Philips lamp were noticeable, but not really significant.

Description: This talk is an overview of discharge lamp technology commonly employed in general lighting, with emphasis on issues pertinent to lighting for plant growth. Since the audience is primarily from the plant growth community, and this begins the light source part of the program, we will start with a brief description of the discharge lamps. Challenges of economics and of thermal management make lamp efficiency a prime concern in controlled environment agriculture, so we will emphasize science considerations relating to discharge lamp efficiency. We will then look at the spectra and ratings of some representative lighting products, and conclude with a discussion of technological advances.

Description: The Galileo Jupiter orbital mission using the Low Gain Antenna (LGA) requires a higher degree of spacecraft state knowledge than was originally anticipated. Key elements of the revised design include onboard buffering of science and engineering data and extensive processing of data prior to downlink. In order to prevent loss of data resulting from overflow of the buffers and to allow efficient use of the spacecraft resources, ground based models of the spacecraft processes will be implemented. These models will be integral tools in the development of satellite encounter sequences and the cruise/playback sequences where recorded data is retrieved.

Description: The Earth Radiation Budget Satellite (ERBS), Compton Gamma Ray Observatory (CGRO), Upper Atmosphere Research Satellite (UARS), and Extreme Ultraviolet Explorer (EUVE) spacecraft are operated from NASA's Goddard Space Flight Center (GSFC) in Greenbelt, Maryland. On-board power subsystems for each satellite employ NASA Standard 50 Ampere-hour (Ah) nickel-cadmium batteries in a parallel configuration. To date, these batteries have exhibited degradation over periods from several months (anomalous behavior, UARS and CGRO (MPS-1); to little if any, EUVE) to several years (old age, normal behavior, ERBS). Since the onset of degraded performance, each mission's Flight Operations Team (FOT), under the direction of their cognizant GSFC Project Personnel and Space Power Application Branch's Engineers has closely monitored the battery performance and implemented several charge control schemes in an effort to extend battery life. Various software and hardware solutions have been developed to minimize battery overcharge. Each of the four sections of this paper covers a brief overview of each mission's operational battery management and its associated spacecraft battery performance. Also included are new operational procedures developed on-orbit that may be of special interest to future mission definition and development.

Description: Outlined in this presentation is the background to insulation constructions for aerospace wiring applications, the Air Force wiring policy, the purpose and contract requirements of new insulation constructions, the test plan, and the test results.

Description: The sources for electrical power on a lunar base are said to include solar/chemical, nuclear (static conversion), and nuclear (dynamic conversion). The transmission of power via transmission lines is more practical than power beaming or superconducting because of its low cost and reliable, proven technology. Transmission lines must have minimum mass, maximum efficiency, and the ability to operate reliably in the lunar environment. The transmission line design includes conductor material, insulator material, conductor geometry, conductor configuration, line location, waveform, phase selection, and frequency. This presentation oulines the design. Liquid and gaseous dielectrics are undesirable for long term use in the lunar vacuum due to a high probability of loss. Thus, insulation for high voltage transmission line will most likely be solid dielectric or vacuum insulation.

Description: The mission of polymer composites materials technology is to develop materials and processing technology to meet DoD and commercial needs. The following are outlined in this presentation: high performance capacitors, high temperature aerospace insulation, rationale for choosing Foster-Miller (the reporting industry), the approach to the development and evaluation of high temperature insulation materials, and the requirements/evaluation parameters. Supporting tables and diagrams are included.

Description: New Low Earth Orbit (LEO) requirements of space environment wiring are compared with traditional requirements. The pyrolysis of Kapton is reviewed for the LeRc vacuum chamber and the 1989 SSF. SEEB modeling of Kapton pyrolysis is also presented.

Description: The objectives of these projects include the following: validate method used to screen wire insulation with arc tracking characteristics; determine damage resistance to arc as a function of source voltage and insulation thickness; investigate propagation characteristics of Kapton at low voltages; and investigate pyrolytic properties of polyimide insulated (Kapton) wire for low voltage (less than 35 VDC) applications. Supporting diagrams and tables are presented.

Description: The objectives of this research were to investigate possible events that could cause the Kapton to pyrolyze, and to investigate the degree of damage when the Kapton pyrolyzes. Supporting diagrams and tables are presented.

Description: The Navy is experiencing a severe aircraft electrical wiring maintenance problem as a result of the extensive use of an aromatic polyimide insulation that is deteriorating at a rate that was unexpected when this wire was initially selected. This problem has significantly affected readiness, reliability, and safety and has greatly increased the cost of ownership of Naval aircraft. Failures in wire harnesses have exhibited arcing and burning that will propagate drastically, to the interruption of many electrical circuits from a fault initiated by the failure of deteriorating wires. There is an urgent need for a capability to schedule aircraft rewiring in an orderly manner with a logically derived determination of which aircraft have aged to the point of absolute necessity. Excessive maintenance was demonstrated to result from the accelerated aging due to the parameters of moisture, temperature, and strain that exist in the Naval Aircraft environment. Laboratory studies have demonstrated that MIL-W-81381 wire insulation when aged at high humidities followed the classical Arrhenius thermal aging relationship. In an extension of the project a multifactor formula was developed that is now capable of predicting life under varying conditions of these service parameters. An automated test system has also been developed to analyze the degree of deterioration that has occurred in wires taken from an aircraft in order to obtain an assessment of remaining life. Since it is both physically and financially impossible to replace the wiring in all the Navy's aircraft at once, this system will permit expedient scheduling so that those aircraft that are most probable to have wiring failure problems can be overhauled first.

Description: This presentation outlines the background to the concern of using Kapton wire for aerospace vehicles and proposes it should not be utilized in new builds for spacecraft power applications. A NASA HQ investigation concluded that the risk of Kapton arc-tracking/flashover is a credible threat to the shuttle orbiter, but rationale is presented for continued flight for the time being. Recommendations for the protection of the shuttle and the build of the space station are given.

Description: Electrohydrodynamic sample distortion during continuous flow electrophoresis is an experiment to be conducted during the second International Microgravity Laboratory (IML-2) in July 1994. The specific objective of this experiment is the distortion caused by the difference in dielectric constant between the sample and surrounding buffer. Although the role of sample conductivity in electrohydrodynamic has been the subject of both flight and ground experiments, the separate role of dielectric constant, independent of sample conductivity, has not been measured. This paper describes some of the laboratory research and model development that will support the flight experiment on IML-2.

Description: The Space Shuttle acceleration environment is characterized. The acceleration environment is composed of a residual or quasi-steady component and higher frequency components induced by vehicle structural modes and the operation of onboard machinery. Quasi-steady accelerations are generally due to atmospheric drag, gravity gradient effects, and rotational forces. These accelerations tend to vary with the orbital frequency (approx. 10(exp -4) Hz) and have magnitudes less than or equal to 10(exp -6) g(sub 0) (where 1 g(sub 0) is terrestrial gravity). Higher frequency g-jitter is characterized by oscillatory disturbances in the 1-100 Hz range and transient components. Oscillatory accelerations are related to the response of large flexible structures like antennae, the Spacelab module, and the Orbiter itself, and to the operation of rotating machinery. The Orbiter structural modes in the 1-10 Hz range, are excited by oscillatory and transient disturbances and tend to dominate the energy spectrum of the acceleration environment. A comparison of the acceleration measurements from different Space Shuttle missions reveals the characteristic signature of the structural modes of the Orbiter overlaid with mission specific hardware induced disturbances and their harmonics. Transient accelerations are usually attributed to crew activity and Orbiter thruster operations. During crew sleep periods, the acceleration levels are typically on the order of 10(exp -6) g(sub 0) (1 micro-g). Crew work and exercise tend to raise the accelerations to the 10(exp -3) g(sub 0) (1 milli-g) level. Vernier reaction control system firings tend to cause accelerations of 10(exp -4) g(sub 0), while primary reaction control system and Orbiter maneuvering system firings cause accelerations as large as 10(exp -2) g(sub 0). Vibration isolation techniques (both active and passive systems) used during crew exercise have been shown to significantly reduce the acceleration magnitudes.

Description: Anticipating the construction of the international space station, on-orbit modal identification of space platforms through optimally placed accelerometers is an area of recent activity. Unwanted vibrations in the platform could affect the results of experiments which are planned. Therefore, it is important that sensors (accelerometers) be strategically placed to identify the amount and extent of these unwanted vibrations, and to validate the mathematical models used to predict the loads and dynamic response. Due to cost, installation, and data management issues, only a limited number of sensors will be available for placement. This work evaluates and compares four representative sensor placement algorithms for modal identification. Most of the sensor placement work to date has employed only numerical simulations for comparison. This work uses experimental data from a fully-instrumented truss structure which was one of a series of structures designed for research in dynamic scale model ground testing of large space structures at NASA Langley Research Center. Results from this comparison show that for this cantilevered structure, the algorithm based on Guyan reduction is rated slightly better than that based on Effective Independence.

Description: A new algorithm is developed for inflight magnetometer bias determination without knowledge of the attitude. This algorithm combines the fast convergence of a heuristic algorithm currently in use with the correct treatment of the statistics and without discarding data. The algorithm performance is examined using simulated data and compared with previous algorithms.

Description: The initial assembly of Space Station Freedom involves the Space Shuttle, its Remote Manipulation System (RMS) and the evolving Space Station Freedom. The dynamics of this coupled system involves both the structural and the control system dynamics of each of these components. The modeling and analysis of such an assembly is made even more formidable by kinematic and joint nonlinearities. The current practice of modeling such flexible structures is to use finite element modeling in which the mass and interior dynamics is ignored between thousands of nodes, for each major component. The model characteristics of only tens of modes are kept out of thousands which are calculated. The components are then connected by approximating the boundary conditions and inserting the control system dynamics. In this paper continuum models are used instead of finite element models because of the improved accuracy, reduced number of model parameters, the avoidance of model order reduction, and the ability to represent the structural and control system dynamics in the same system of equations. Dynamic analysis of linear versions of the model is performed and compared with finite element model results. Additionally, the transfer matrix to continuum modeling is presented.

Description: The COMET attitude determination and control system, using inverse dynamics and a novel torque distribution/momentum management technique, has shown great flexibility, performance, and robustness. Three-axis control with two wheels is an inherent consequence of inverse dynamics control which allows for reduction in spacecraft weight and cost, or alternatively, provides a simple means of failure-redundancy for three-wheel spacecraft. The control system, without modification, has continued to perform well in spite of large changes in spacecraft mass properties and mission orbit altitude that have occurred during development. This flexibility has obviated imposition of early stringent ADACS design constraints and has greatly reduced commonly incurred ADACS modification costs and delay associated with program maturation.