ALBERT

Description: This paper presents the results of a transient computer simulation that was developed to study phase change energy storage techniques for Space Station Freedom (SSF) solar dynamic (SD) power systems. Such SD systems may be used in future growth SSF configurations. Two solar dynamic options are considered in this paper: Brayton and Rankine. Model elements consist of a single node receiver and concentrator, and takes into account overall heat engine efficiency and power distribution characteristics. The simulation not only computes the energy stored in the receiver phase change material (PCM), but also the amount of the PCM required for various combinations of load demands and power system mission constraints. For a solar dynamic power system in low earth orbit, the amount of stored PCM energy is calculated by balancing the solar energy input and the energy consumed by the loads corrected by an overall system efficiency. The model assumes an average 75 kW SD power system load profile which is connected to user loads via dedicated power distribution channels. The model then calculates the stored energy in the receiver and subsequently estimates the quantity of PCM necessary to meet peaking and contingency requirements. The model can also be used to conduct trade studies on the performance of SD power systems using different storage materials.

Description: The Solar Space Power Analysis Code, SOSPAC, was developed to examine the solar thermal and photovoltaic power generation options available for a satellite or spacecraft in low earth orbit. SOSPAC is a preliminary systems analysis tool and enables the engineer to compare the areas, weights, and costs of several candidate electric and thermal power systems. The configurations studied include photovoltaic arrays and parabolic dish systems to produce electricity only, and in various combinations to provide both thermal and electric power. SOSPAC has been used for comparison and parametric studies of proposed power systems for the NASA Space Station. The initial requirements are projected to be about 40 kW of electrical power, and a similar amount of thermal power with temperatures above 1000 degrees Centigrade. For objects in low earth orbit, the aerodynamic drag caused by suitably large photovoltaic arrays is very substantial. Smaller parabolic dishes can provide thermal energy at a collection efficiency of about 80%, but at increased cost. SOSPAC allows an analysis of cost and performance factors of five hybrid power generating systems. Input includes electrical and thermal power requirements, sun and shade durations for the satellite, and unit weight and cost for subsystems and components. Performance equations of the five configurations are derived, and the output tabulates total weights of the power plant assemblies, area of the arrays, efficiencies, and costs. SOSPAC is written in FORTRAN IV for batch execution and has been implemented on an IBM PC computer operating under DOS with a central memory requirement of approximately 60K of 8 bit bytes. This program was developed in 1985.

Description: Slush hydrogen, a mixture of the solid and liquid phases of hydrogen, is a possible source of fuel for the National Aerospace Plane (NASP) Project. Advantages of slush hydrogen over liquid hydrogen include greater heat capacity and greater density. However, practical use of slush hydrogen as a fuel requires systems of lines, valves, etc. which are designed to deliver the fuel in slush form with minimal solid loss as a result of pipe heating or flow friction. Engineers involved with the NASP Project developed FLUSH to calculate the pressure drop and slush hydrogen solid fraction loss for steady-state, one-dimensional flow. FLUSH solves the steady-state, one-dimensional energy equation and the Bernoulli equation for pipe flow. The program performs these calculations for each two-node element--straight pipe length, elbow, valve, fitting, or other part of the piping system--specified by the user. The user provides flow rate, upstream pressure, initial solid hydrogen fraction, element heat leak, and element parameters such as length and diameter. For each element, FLUSH first calculates the pressure drop, then figures the slush solid fraction exiting the element. The code employs GASPLUS routines to calculate thermodynamic properties for the slush hydrogen. FLUSH is written in FORTRAN IV for DEC VAX series computers running VMS. An executable is provided on the tape. The GASPLUS physical properties routines which are required for building the executable are included as one object library on the program media (full source code for GASPLUS is available separately as COSMIC Program Number LEW-15091). FLUSH is available in DEC VAX BACKUP format on a 9-track 1600 BPI magnetic tape (standard media) or on a TK50 tape cartridge. FLUSH was developed in 1989.

Description: Accurate simulation of nuclear thermal propulsion systems using computational methods will permit reductions in testing and, thus, the time and cost of achieving a flight ready status for systems utilizing this advanced technology. An accurate simulation must maintain a "balance-of-plant" where the required pump work equals the supplied turbine work. This turbopump assembly balancing must be integrated into the overall system analysis models. TPA was developed to balance turbine and pump work using performance maps. It requires the inlet properties, performance maps, and shaft speed. TPA then computes the exit conditions and work terms. The work terms can then be balanced by varying the input shaft speed. The objective of the pump analysis is to determine the propellant state properties at the pump exit and the pump work. The pump analysis algorithm for liquid flow assumes that the shaft speed, the propellant state properties at the pump entrance, the propellant flow rate, the pump entrance and exit areas, as well as performance curves, are all known. The analysis of both the pump pressure rise and pump efficiency curves is required. The objective of the turbine analysis is to determine the propellant state properties at the turbine exit and the turbine work. The turbine analysis algorithm assumes that the shaft speed, the propellant state properties at the turbine entrance, the propellant flow rate, the turbine root mean square blade diameter, the turbine entrance and exit areas, as well as performance curves, are all known. The analysis also requires the turbine flow parameter curve and the turbine total efficiency curve. TPA is written in FORTRAN 77 to be machine independent. The TPA package includes the NBS+_PH2 code, which is also available separately (LEW-15505). TPA has been successfully implemented on a DEC VAX series computer running VMS, a Sun4 series computer running SunOS, and an IBM PC compatible computer running MS-DOS. Lahey F77L3 EM/32 v. 5.01 or higher is required for compilation on an IBM PC compatible computer; however, a PC executable is included on the distribution diskette. The standard distribution medium for this program is one 5.25 inch 360K MS-DOS format diskette. The diskette's contents have been compressed using PKWARE's archiving tools. The utility to unarchive the file, PKUNZIP.EXE, is included. Alternate distribution media and formats are available upon request. TPA was developed in 1993.

Description: Experimental results are reported for submerged injection pressurization and expulsion tests of a 4.89 cu m liquid hydrogen tank. The pressurant injector was positioned near the bottom of the test vessel to simulate liquid engulfment of the pressurant gas inlet, a condition that may occur in low-gravity conditions. Results indicate a substantial reduction in pressurization efficiency with pressurant gas requirements approximately five times greater than ideal amounts. Consequently, submerged vapor injection should be avoided as a low-gravity autogenous pressurization method whenever possible. The work presented herein validates that pressurant requirements are accurately predicted by a homogeneous thermodynamic model when the submerged injection technique is employed.

Description: Presented here are results of a test program undertaken to further define the response of the solar dynamic radiator to hypervelocity impact (HVI). Tests were conducted on representative radiator panels (under ambient, nonoperating conditions) over a range of velocity. Target parameters are also varied. Data indicate that analytical penetration predictions are conservative (i.e., pessimistic) for the specific configuration of the solar dynamic radiator. Test results are used to define the solar dynamic radiator reliability with respect to HVI more rigorously than previous studies. Test data, reliability, and survivability results are presented.

Description: Attention is given to a space-borne engine plume experiment study to fly an experiment which will both verify and quantify the reduced contamination from advanced rhenium-iridium earth-storable bipropellant rockets (hot rockets) and provide a correlation between high-fidelity, in-space measurements and theoretical plume and surface contamination models. The experiment conceptual design is based on survey results from plume and contamination technologists throughout the U.S. With respect to shuttle use, cursory investigations validate Hitchhiker availability and adaptability, adequate remote manipulator system (RMS) articulation and dynamic capability, acceptable RMS attachment capability, adequate power and telemetry capability, and adequate flight altitude and attitude/orbital capability.

Description: A new procedure, dubbed the Munich Method, has been proposed recently for the modeling of rocket engine performance. The author of the Munich Method claims it to be an extension and improvement of the thermodynamic procedures used to model rocket engines in the NASA-Lewis chemical equilibrium program. An examination of the Munich Method shows that it contains several flaws. If these defects are corrected then the Munich Method will produce results identical to those generated by the NASA-Lewis Code.

Description: Mars Environmental Survey Mission (MESUR) is intended to be a low cost, near term mission to study the chemical composition of Martian surface, structure and circulation of Mars' atmosphere, and the structure and dynamics of Mars' interior. To meet the science objectives of the mission, a network of 16 landers will be placed at various latitudes and longitudes of Mars. The paper identifies various options for powering the landers and lists the attributes of each option. The current strategy is to precede the lander launch and deployment with an engineering demo mission called the Pathfinder which would demonstrate the landing strategy and the engineering subsystems. The paper briefly describes the challenges involved in the power system development for the Pathfinder mission.

Description: The Photovoltaic Array Space Power Plus Diagnostics flight experiment (PASP+) subsumes twelve solar array modules which represent the state of the art in the space photovoltaic array industry. Each of the twelve modules individually feature specific photovoltaic technologies such as advanced semiconductor materials, multi-bandgap structures, lightweight array designs, advanced interconnect technologies, or concentrator array designs. This paper will describe each module in detail including the configuration, components, materials, anticipated on orbit performance, and some of the aspects of each array technology. The layout of each module and the photovoltaic cells or array cross section will be presented graphically. A discussion on the environmental constraints and materials selection will be included as well as a delineation of the differences between the modules and the baseline array configuration in its intended application.

Description: Utilizing commercial or military satellites as testbeds for subsystems is a potential platform for small devices. Electric propulsion is a viable and upcoming subsystem that is of high interest to planetary mission engineers as well as commercial satellite developers. It is proposed that by incorporating small lightweight electric propulsion devices onto small satellites as external or 'bolt-on' experiments, an increase in the number of flight opportunities can occur. Specific problems are spacecraft body interaction, contamination effects, thermal interface problems, power conditioning control electronics, and propulsion feed system interfaces.

Description: Experience with many spacecraft configurations boosted by a variety of launch vehicles indicates that the maximum loads experienced throughout most of the structure are inertial in origin. These loads arise from the dynamic elastic response of the flight vehicle to the transient disturbances of launch and flight, and are highly dependent on the dynamic characteristics of both the spacecraft and the launch vehicle. It has proved to be most advantageous, in the analysis of this critical dependency of loads upon vehicle dynamic properties, to establish a mathematical model in terms of normal mode characteristics. In this way, the vibration behavior of an elastomechanical structure (or substructure) can be described by means of the so-called modal or natural degrees of freedom. The conduct of a mode survey test and the use of a suitably test-verified model in loads analyses is essential to the flight worthiness certification process of space systems. The desirability of such tests is confirmed by the fact that, almost invariably, significant deficiencies in the analytical models are revealed by the results. Therefore, this experimental program was undertaken to determine those properties of a solid-propellant rocket motor (SRM) which are required to characterize a dynamic model. Random ambient-excited accelerations were measured at a series of stations along the motor for the purpose of identifying the motor beam-like stiffnesses in bending, shear, and torsion. From a system identification point of view, it is significant that stiffness properties of a subsystem (the motor) are determined from modes of the full system (motor/stand configuration) using mode shape data of the subsystem only. This contrasts with traditional system identification approaches which rely upon complete system mode shapes.

Description: An account is given of the significance for U.S. spacecraft development of a nuclear thermal rocket (NTR) reactor concept that has been developed in the (formerly Soviet) Commonwealth of Independent States (CIS). The CIS NTR reactor employs a hydrogen-cooled zirconium hydride moderator and ternary carbide fuels; the comparatively cool operating temperatures associated with this design promise overall robustness.

Description: Experimental results of no-vent fill testing with liquid hydrogen in a 34 liter stainless steel tank are presented. More than 40 tests were performed with various liquid inlet temperatures, inlet flowrates, initial tank wall temperatures, and liquid injection techniques. Maximum pressure within the receiver tank was limited to 0.207 MPa (30 psia), and fill levels equal to or exceeding 90 percent by volume were achieved in 40 percent of the tests. Three liquid injection techniques were employed; top spray, upward pipe discharge, and bottom diffuser. Effects of each of the various parameters on the tank pressure history and final fill level are evaluated. The final fill level is found to be indirectly proportional to the initial wall and inlet liquid temperatures and directly proportional to the inlet liquid flowrate. Furthermore, the top spray is the most efficient no-vent fill method of the three configurations examined. The success of this injection method is primarily due to condensation of the ullage vapor onto the incoming liquid droplets. Ullage condensation counteracts the tank pressure rise resulting from energy exchange between the fluid and the warmer tank walls and from ullage compression.

Description: A great deal of experimentation and analysis was performed to quantify penetration thresholds of components which will experience orbital debris impacts. Penetration was found to depend upon mission specific parameters such as orbital altitude, inclination, and orientation of the component; and upon component specific parameters such as material, density and the geometry particular to its shielding. Experimental results are highly dependent upon shield configuration and cannot be extrapolated with confidence to alternate shield configurations. Also, current experimental capabilities are limited to velocities which only approach the lower limit of predicted orbital debris velocities. Therefore, prediction of the penetrating particle size for a particular component having a complex geometry remains highly uncertain. An approach is described which was developed to assess on-orbit survivability of the solar dynamic radiator due to micrometeoroid and space debris impacts. Preliminary analyses are presented to quantify the solar dynamic radiator survivability, and include the type of particle and particle population expected to defeat the radiator bumpering (i.e., penetrate a fluid flow tube). Results of preliminary hypervelocity impact testing performed on radiator panel samples (in the 6 to 7 km/sec velocity range) are also presented.

Description: All solar arrays have biased surfaces that can be exposed to the space environment. It has been observed that when the array bias is less than a few hundred volts negative, then the exposed conductive surfaces may undergo arcing in the space plasma. A theory for arcing is developed on these high voltage solar arrays that ascribes the arcing to electric field runaway at the interface of the plasma, conductor, and solar cell dielectric. Experiments were conducted in the laboratory for the High Voltage Solar Array experiment that will fly on the Japanese Space Flyer Unit (SFU) in 1994. The theory was compared in detail with the experiment and shown to give a reasonable explanation for the data. The combined theory and ground experiments were then used to develop predictions for the SFU flight.

Description: Issues affecting the long-term operational performance of the Advanced Photovoltaic Solar Array (APSA) are discussed, with particular attention given to circuit electrical integrity from shadowed and cracked cell modules. The successful integration of individual advanced array components provides a doubling of array specific performance from the previous NASA-developed advanced array (SAFE). Flight test modules both recently fabricated and under fabrication are described. The development of advanced high-performance blanket technology for future APSA enhancement is presented.

Description: Results are presented from a study of the coupling of an SP-100 nuclear reactor with either a Stirling or Brayton power system, at the 100 kWe level, for a power generating system suitable for operation in the lunar and Martian surface environments. In the lunar environment, the reactor and primary coolant loop would be contained in a guard vessel to protect from a loss of primary loop containment. For Mars, all refractory components, including the reactor, coolant, and power conversion components will be contained in a vacuum vessel for protection against the CO2 environment.

Description: The solar dynamic power system chosen by NASA for Space Station Freedom SSF requires validation in the vacuum/microgravity environment. The planned ground tests of a proof-of-concept 2 kWe solar dynamic system with address subsystem integration issues involving such major system components as the behavior of thermal energy storage materials. This test scheme will be scalable to 25 kWe, and will be flight-configured to incorporate features of the SSF's power module design.

Description: NASA-Marshall conducted a study of electrical power system faults with a view to the development of AI control systems for a spacecraft power system breadboard. The results of this study have been applied to a multichannel high voltage dc spacecraft power system, the Large Autonomous Spacecraft Electrical Power System (LASEPS) breadboard. Some of the faults encountered in testing LASEPS included the shorting of a bus an a falloff in battery cell capacity.

Description: The solar array module plasma interactions experiment (SAMPIE) is an approved NASA flight experiment manifested for Shuttle deployment in early 1994. The SAMPIE experiment is designed to investigate the interaction of high voltage space power systems with ionospheric plasma. To study the behavior of solar cells, a number of solar cell coupons (representing design technologies of current interest) will be biased to high voltages to measure both arcing and current collection. Various theories of arc suppression will be tested by including several specially modified cell coupons. Finally, SAMPIE will include experiments to study the basic nature of arcing and current collection. This paper describes the rationale for a space flight experiment, the measurements to be made, and the significance of the expected results. A future paper will present a detailed discussion of the engineering design.

Description: A new approach for development of a 50-kW directly solar-pumped iodine laser (DSPIL) system as a space-based power station was made using a confocal unstable resonator (CUR). The CUR-based DSPIL has advantages, such as performance enhancement, reduction of total mass, and simplicity which alleviates the complexities inherent in the previous system, master oscillator/power amplifier (MOPA) configurations. In this design, a single CUR-based DSPIL with 50-kW output power was defined and compared to the MOPA-based DSPIL. Integration of multiple modules for power requirements more than 50-kW is physically and structurally a sound approach as compared to building a single large system. An integrated system of multiple modules can respond to various mission power requirements by combining and aiming the coherent beams at the user's receiver.

Description: The NASA-sponsored 2kW(e) Solar Dynamic Space Power System Ground Test Demonstration Program requires the physical and the thermodynamic integration of the Brayton Isotope Power System (BIPS) Turboalternator-Compressor (TAC) and recuperator with a heat receiver, solar concentrator, and radiator based on Space Station Freedom designs. The aim of the designs is to provide a cost-effective, minimal-risk, viable ground test system. This paper describes the BIPS TAC configuration and performance characteristics along with the cycle analysis of BIPS TAC.

Description: As the principal contractor to NASA-Lewis Research Center, Mechanical Technology Incorporated is under contract to develop free-piston Stirling power converters in the context of the competitive multiyear Space Stirling Technology Program. The first generation Stirling power converter, the component test power converter (CTPC) initiated cold end testing in 1991, with hot testing scheduled for summer of 1992. This paper reviews the test progress of the CTPC and discusses the potential of Stirling technology for various potential missions at given point designs of 250 watts, 2500 watts, and 25,000 watts.

Description: The Hubble Space Telescope (HST) solar array consists of two identical double roll-out wings designed after the Hughes flexible roll-up solar array (FRUSA) and was developed by the European Space Agency (ESA) to meet specified HST power output requirements at the end of 2 years, with a functional lifetime of 5 years. The requirement that the HST solar array remain functional both mechanically and electrically during its 5-year lifetime meant that the array must withstand 30,000 low Earth orbit (LEO) thermal cycles between approximately +100 and -100 C. In order to evaluate the ability of the array to meet this requirement, an accelerated thermal cycle test in vacuum was conducted at NASA's Marshall Space Flight Center (MSFC), using two 128-cell solar array modules which duplicated the flight HST solar array. Several other tests were performed on the modules. The thermal cycle test was interrupted after 2,577 cycles, and a 'cold-roll' test was performed on one of the modules in order to evaluate the ability of the flight array to survive an emergency deployment during the dark (cold) portion of an orbit. A posttest static shadow test was performed on one of the modules in order to analyze temperature gradients across the module. Finally, current in-flight electrical performance data from the actual HST flight solar array will be tested.

Description: NASA's HST uses Ni-H2 batteries. NASA-Marshall has been conducting developmental tests of such batteries in both six-battery and 22-cell single battery arrays. Tests have recently been conducted on such batteries with a view to the possible need to free additional memory in the HST onboard computer; the electrical power system could contribute to this end by eliminating its software control charge mode capability, which requires significant computer memory capacity.

Description: The orbiting shadowing analysis computer program was developed by NASA in order to assess the shadowing effects on various power systems. The algorithms, the inputs and outputs are discussed. Examples of typical shadowing analysis, performed for the International Space Station Freedom, the International Space Station Alpha and the joint United States/International Mir Solar Dynamic Flight Experimental Project are presented.

Description: An evaluation is made of the potential benefits of C60 molecules as a basis for ion propulsion. Because C60 is storable, its use may result in a larger usable propellant fraction than previous methods of cluster ion propulsion. C60 may also relax such engineering constraints as grid spacing, which restrict the performance of noble gas ion propulsion. The behavior of C60 in a plasma discharge environment, as well as various electron impact cross sections of the molecule, will greatly afftect the feasibility of the concept.

Description: An Isotope-Powered Thermal Storage Unit (ITSU), that would store and utilize heat energy in a 'pulsed' fashion in space operations, is described. Properties of various radioisotopes are considered in conjunction with characteristics of thermal energy storage materials, to evaluate possible implementation of such a device. The utility of the unit is discussed in light of various space applications, including rocket propulsion, power generation, and spacecraft thermal management.

Description: Nuclear propulsion technology offers substantial benefits to the ambitious piloted and robotic solar system exploration missions of the Space Exploration Initiative (SEI). This paper summarizes a workshop jointly sponsored by NASA, DoE, and DoD to assess candidate nuclear electric propulsion technologies. Twenty-one power and propulsion concepts are reviewed. Nuclear power concepts include solid and gaseous fuel concepts, with static and dynamic power conversion. Propulsion concepts include steady state and pulsed electromagnetic engines, a pulsed electrothermal engine, and a steady state electrostatic engine. The technologies vary widely in maturity. The workshop review panels concluded that compelling benefits would accrue from the development of nuclear electric propulsion systems, and that a focused, well-funded program is required to prepare the technologies for SEI missions.

Description: Three types of electric thrusters currently under development at JPL have potential to support future missions which utilize multimegawatt nuclear electric propulsion. These electric thrusters are the electron bombardment ion thruster, the magnetoplasmadynamic (MPD) thruster, and the electron-cyclotron-resonance (ECR) thruster. The electron bombardment ion thruster is a relatively mature technology which has been developed for operation at kilowatt power levels but will require new development for application in the multimegawatt regime. The MPD engine represents a technology which may be very well suited to steady-state multimegawatt applications but which has been limited to sub-scale (100's of kW) and pulsed (MW) testing thus far. The ECR plasma engine represents a class of very promising new concepts which are still in the basic research phase of development, but which may possess important fundamental advantages over other electric thruster technologies. Models of these thrusters are described and used to make projections of thrusters specific mass, efficiency, and power handling capacity for operation in the multimegawatt regime.

Description: MgF2 was investigated as a phase-change energy-storage material for LEO power systems using solar heat to run thermal cycles. It provides a high heat of fusion per unit mass at a high melting point (1536 K). Theoretical evaluation showed the basic chemical compatibility of liquid MgF2 with refractory metals at 1600 K, though transient high pressures of H2 can occur in a closed container due to reaction with residual moisture. The compatibility was tested in two refractory metal containers for over 2000 h. Some showed no deterioration, while there was evidence that the fluoride reacted with hafnium in others. Corollary tests showed that the MgF2 supercooled by 10-30 K and 50-90 K.

Description: The development of a high-efficiency CW YLF laser doped with Er,Tm,Ho: and featuring a strongly focusing resonator that collects a high density of pump power on the active crystal is described. The emission is investigated at 2.06 microns and a tuning range both at liquid-nitrogen (77 K) and at dry-ice (210 K) temperature. The noise characteristics and the long-term power stability of the laser is studied with an eye to employing this source for high-resolution spectroscopy in the 2-micron wavelength region. The detection of several absorption lines of NH3 at low pressure is described. The output power of the laser as a function of the power impinging on the crystal for different transmission of the output mirror is illustrated. The best result obtained is 1.46 W output for 3.2 W of argon pump. The minimum threshold achieved is 3.5 mW with a 1-percent transmission mirror. It is concluded that it is possible to develop a highly efficient Ho:YLF laser featuring low noise and sufficient tunability for high-resolution spectroscopy in the 2-micron region.

Description: Electric-field (EF) and radio-frequency (RF) emissions generated in the exhaust plumes of the diagnostic testbed facility thruster (DTFT) and the SSME are examined briefly for potential applications to plume diagnostics and engine health monitoring. Hypothetically, anomalous engine conditions could produce measurable changes in any characteristic EF and RF spectral signatures identifiable with a 'healthly' plumes. Tests to determine the presence of EF and RF emissions in the DTFT and SSME exhaust plumes were conducted. EF and RF emissions were detected using state-of-the-art sensors. Analysis of limited data sets show some apparent consistencies in spectral signatures. Significant emissions increases were detected during controlled tests using dopants injected into the DTFT.

Description: Hybrid rockets employing liquid oxidizer and solid fuel grain answers to cost, safety, reliability, and environmental impact concerns that have become as prominent as performance in recent years. The oxidizer-free grain has limited sensitivity to grain anomalies, such as bond-line separations, which can cause catastrophic failures in solid rocket motors. An account is presently given of the development effort associated with the AMROC commercial hybrid booster and component testing efforts at NASA-Marshall. These hybrid rockets can be fired, terminated, inspected, evaluated, and restarted for additional testing.

Description: An evaluation is conducted of the use of novel, metallized-propellant propulsion systems to reduce launch masses for manned Mars missions on the basis of density and I(sp) enhancements. The use of metallized propellants leads to a 3.3 percent LEO mass saving which represents 38,000 kg less than O2/H2 propulsion. Attention is given to the possibility of using space-storable propellants for the Mars excursion vehicle, as an alternative to the storing of cryogenic H2 on Mars, albeit at the cost of lower I(sp).

Description: The Mars Rover and Sample Return mission's radioisotope thermoelectric generator (RTG) is presently subjected to a structural and mass analysis in view of a reference mission scenario, an illustrative Rover design and Martian activities agenda, and RTG power system requirements and environmental constraints. The modular heat-source stack in the Rover RTG can be held together by axial load springs. The RTGs should be mounted on the Rover with a vertical orientation, in order to avoid the buildup of windborne Martian sand on its heat-rejection surfaces.

Description: The expert system described in this paper analyzes spectral emissions of rocket engine exhaust plumes and shows major promise for use in engine health diagnostics. Plume emission spectroscopy is an important tool for diagnosing engine anomalies, but it is time-consuming and requires highly skilled personnel. The expert system was created to alleviate such problems. The system accepts a spectral plot in the form of wavelength vs intensity pairs and finds the emission peaks in the spectrum, lists the elemental emitters present in the data and deduces the emitter that produced each peak. The system consists of a conventional language component and a commercially available inference engine that runs on an Apple Macintosh computer. The expert system has undergone limited preliminary testing. It detects elements well and significantly decreases analysis time.

Description: The aerospace industry, like many other industries, regularly applies optimization techniques to develop designs which reduce cost, maximize performance, and minimize weight. The desire to minimize weight is of particular importance in space-related products since the costs of launch are directly related to payload weight, and launch vehicle capabilities often limit the allowable weight of a component or system. With these concerns in mind, this paper presents the optimization of a space-based power generation system for minimum mass. The goal of this work is to demonstrate the use of optimization techniques on a realistic and practical engineering system. The power system described uses thermoelectric devices to convert heat into electricity. The heat source for the system is a nuclear reactor. Waste heat is rejected from the system to space by a radiator.

Description: The topics addressed are: (1) phobos power plant; (2) fusion power/propulsion system; (3) surface power from an orbiting spacecraft; (4) RTG replacement; (5) MHD-thermoelectric burst reactor; (6) TAU Voyage power/propulsion device; (7) ESCAPE to ODYSSEY.

Description: The Texas A&M Nuclear and Aerospace engineering departments have worked on five different projects for the NASA/USRA Advanced Design Program during the 1987/88 year. The aerospace department worked on two types of lunar tunnelers that would create habitable space. The first design used a heated cone to melt the lunar regolith, and the second used a conventional drill to bore its way through the crust. Both used a dump truck to get rid of waste heat from the reactor as well as excess regolith from the tunneling operation. The nuclear engineering department worked on three separate projects. The NEPTUNE system is a manned, outer-planetary explorer designed with Jupiter exploration as the baseline mission. The lifetime requirement for both reactor and power-conversion systems was twenty years. The second project undertaken for the power supply was a Mars Sample Return Mission power supply. This was designed to produce 2 kW of electrical power for seven years. The design consisted of a General Purpose Heat Source (GPHS) utilizing a Stirling engine as the power conversion unit. A mass optimization was performed to aid in overall design. The last design was a reactor to provide power for propulsion to Mars and power on the surface. The requirements of 300 kW of electrical power output and a mass of less than 10,000 Rg were set. This allowed the reactor and power conversion unit to fit within the Space Shuttle cargo bay.

Description: Direct characterization procedures were used to determine the relaxation modulus as a function of time, temperature, and state of strain. Using the quasi-elastic method of linearviscoelasticity, these properties were employed in a finite element computer code to analyze a thick-walled, nonlinear viscoelastic cylinder in the state of plane strain bonded to a thin (but stiff) elastic casing and subjected to slow thermal cooling. The viscoelastic solution is then expressed as a sequence of elastic finite element solutions. The strain-dependent character of the relaxation modulus is included by replacing the single relaxation curve used in the linear viscoelastic theory by a family of relaxation functions obtained at various strain levels. These functions may be regarded as a collection of stress histories or responses to specific loads (in this case, step strains) with which the cooldown solution is made to agree by iterations on the modulus and strain level.

Description: The author briefly describes why he thinks air-breathing propulsion merits serious consideration as an alternative or supplement to rocket propulsion for space shuttle missions. Several aspects of hypersonic ramjet technology are discussed which are indicative of the current state of development and of the compromises which are made in arriving at effective engine configuration concepts. Points of interest in the current NASA Hypersonic Research Engine Project are cited as to exemplify the actual development of a hydrogen-fueled, regeneratively cooled, flight-weight, dual-combustion mode hypersonic ramjet.

Description: The damage potentials to the space shuttle orbiter caused by the high velocity particles contained in the exhaust plume of an upper stage is discussed. In particular, the particle size distribution and composition, the velocity of the particles and the expected contribution from shuttle launched upper stages are addressed. Particle size estimates based on historical data are compapred with those derived from upper stage motor performance testing. The particle velocities as determined by the best available plume computational technique are presented. The shuttle is scheduled to launch approximately 135 upper-stages over its lifetime looking at the currently scheduled flights and averaging over a yearly basis yields the contribution of particulates from the uper-stages. On the average, 91,645 llbs of Al2O3 will be ejected on each launch. The analysis to determine how much of this 91,645 lbs will remain in orbit or the decay rate is yet to be accomplished.

Description: The objective of this experiment is to determine the effects of long-term orbital exposure on the materials used in solid-rocket space motors. Specifically, structural materials and propellants from the STAR/PAM-D series motors and the PAM DII/IPSM-II motors will be tested, as well as advanced composite case and nozzle materials planned for future use. The experiment approach is to expose samples of solid-rocket propellant, liner, insulation, case, and nozzle specimens to the space environment and to compare preflight and postflight measurements of various mechanical, chemical, and ballistic properties.

Description: Proposed experiments for analyzing rocket plumes are reported. Two groups of experiments were studied: (1) those that would help define some of the parameters that characterize the plume and (2) those that would enable evaluation of some of the contamination effects of the plume environment on various items of interest. The items investigated, the purpose of the investigation, are given in tabular form.

Description: The relative effects of several candidate SSRM propellant formulations and their plume impingement effects on HRSI and RCC materials were evaluated. Nine solid propellant formulations were tested. The selected propellant matrix allowed an evaluation of propellants with and without metal additives, with and without burning rate catalyst, and low (approximately 1927 C) and high (approximately 2649 C) combustion temperatures. Motors were fired at a simulated SRB staging altitude of 3.96 km (130,000 ft) (nominal). The altitude pressure was predicted to drop approximately 0.6 km (20,000 ft) during a motor firing. All motors were loaded with 1.8 to 2.3 kg (4 to 5 lb) of propellant and burned for approximately 2 s.

Description: Efforts made to determine the vulnerability of Orbiter and ET materials located at various positions within exhaust plumes from test SSRM's using four different propellant formulations are discussed. Data also cover the effect on TPS materials from a single SSRM plume and dual SSRM plumes, and definitions of test SSRM plume environment at material specimen locations.

Description: The hydraulic actuation system of the space shuttle main engine is discussed. The system consists of five electrohydraulic actuators and a single engine filter used to control the five different propellant valves, which in turn control thrust and mixture ratio of the space shuttle main engine. The hydraulic actuation system provides this control with a precision of 98.7 percent or an error in position no greater than 1.3 percent of full scale rotational travel for critical positions.

Description: The system components and operation of the space shuttle solid rocket booster (SRB) dewatering set are described. The SRB dewatering set consists of a nozzle plug, control console, remote control unit, power distribution unit, umbilical cable, interconnect cables, and various handling and storage items. The nozzle plug (NP) is a remotely controlled, tethered underwater vehicle that is launched from the retrieval vessel (RV) by a crane, descends down the side of the SRB, and is positioned below the SRB nozzle. A TV camera mounted at the top of the NP central core is used by the control console operator to visually guide the NP during descent and docking. The NP is then driven up and locked into the nozzle. Compressed air is passed through the umbilical from the RV, through the NP and into the SRB motor. The water inside the SRB is expelled causing the SRB to rotate to a near horizontal attitude on the surface of the water.

Description: Zero gravity testing in the KC-135 aircraft of flat fold flexible solar array test specimens sufficiently demonstrated the adequacy of the panel design. The aircraft flight crew provided invaluable assistance and significantly contributed to the design and development of the flexible solar array, and ultimately to the potential success of the solar electric propulsion solar array shuttle flight experiment program.

Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a system to be used for measuring propellant flow in a liquid monopropellant rocket engine. Design guidelines rather than detailed specifications are provided for the critical components of each portion of the system. System elemental uncertainties are presented in an appendix.

Description: Basic periods in the history of the development of ramjet engine theory are cited. The periods include the first experimental tests as well as the development of basic ideas and theoretical development of the cosmic ramjet engine.

Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a pressure measuring system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) pressure measurement system elemental uncertainties; (2) short- and long-term pressure measurement system uncertainty; (3) shunt calibration of pressure transducers; (4) special considerations for vacuum measurement; and (5) methods of determining the dynamic characteristics of pressure transducers. Design guidelines are provided for the critical components of each portion of the system to provide a pressure measurement system which meets the performance criteria specified.

Description: The design, installation, checkout, calibration, and operation of a temperature measuring system to be used during tests of a liquid monopropellant rocket engine are discussed. Appendixes include: (1) temperature measurement system elemental uncertainties, and (2) tables and equations for use with thermocouples and resistance thermometers. Design guidelines are given for the critical components of each portion of the system to provide an optimum temperature measurement system which meets the performance criteria specified.

Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a thrust measurement system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) thrust measurement system elemental uncertainties; (2) short- and long-term thrust measurement system uncertainty; and (3) shunt calibration of force transducers.

Description: Definitions, algorithms, and procedures for the reduction of monopropellant thruster measurements to performance parameters are provided. A brief discussion of acquisition and recording systems is also included. Emphasis is placed upon monopropellant hydrazine engines, and some parameters relate specifically to the catalytic decomposition of hydrazine (e.g., percent ammonia dissociation). Performance of other types of monopropellant thrusters may, however, be determined by using procedures similar to those discussed. Two appendixes are included: (1) theoretical performance of monopropellant hydrazine; and (2) calculation of rotational performance.

Description: A space based orbital transfer vehicle (SBOTV) and ground based OTV's (GBOTV) are compared for debris protection, space based OTV maintenance provisions, flight performance, onorbit refueling, and launch and return operations. Debris protection has a severe impact on the SBOTV, while the penalty for the GBOTV is much less severe. A key technology issue is the protection capability of composite materials. Onorbit maintenance is critical for SBOTV. Reduction of losses during the various transfers is th maine problem with refueling a SBOTV. Zero-g propellant storage and transfer is an important technology area for SBOTV. A reusable shroud must be developed to return GBOTV's if a Shuttle derivative vehicle is used. The advantage of space basing lies in more efficient use of the launch vehicle. Since most of the mass going to LEO is OTV propellant, and the launches to deliver the SBOTV propellant are generally mass limited, substantially fewer launches are required to support the SBOTV.

Description: Results of the liquid rocket booster study initiated by NASA to define an alternative to solid rocket boosters, are presented. The prime study contractors, Martin Marietta Corporation and General Dynamics, have identified liquid rocket booster configurations that can increase shuttle performance to 70 klb. These boosters will provide improved reliability, hold down, verification prior to vehicle release, engine-out and abort capabilities. Phasing of these boosters into Space Transportation System (STS) operations without adversely affecting flight rate is described.

Description: The redesigned solid rocket motor of the Space Shuttle is described. Improvements over the model that led to the loss of the Space Shuttle Challenger are outlined. Scale and full-size tests carried out to verify the quality of the redesign are described. A unique feature of the test program is the introduction of deliberate flaws into some test articles. Post-flight evaluation of the redesigned boosters show excellent results.

Description: A contactless method for the determination of the free-carrier density and the composition distribution across the thickness of 3-5 multi-layer solar cell structures, using the Raman scattering method, is developed. The method includes a step analysis of Raman spectra from optical phonons and phonon-plasmon modes of different layers. The method provides simultaneous measurements of the element composition and the thickness of the structure's layers together with the free-carrier density. The results of measurements of the free-carrier density composition distributions of the liquid phase epitaxy grown AlGaAs/GaAs and GaSb solar cell structures are presented and discussed.

Description: The photovoltaic array space power (PASP)-Plus solar cell flight experiment is described, and the observed performances of different solar cell types during the first six months of their operation, are summarized. The solar cell types include single crystal and amorphous silicon, GaAs, several multijunction cell types, indium phosphide and GaAs/GaSb concentrator cells. The radiation degradation experienced by some of the solar cell types agrees with theoretical predictions. Other samples, including silicon, are degraded less than predicted. Effects, including the increase in temperature of all the experiments and the effect of sun glint on cell measurement, are discussed.

Description: The conceptual design and experimental results for two types of space application concentrator photovoltaic modules, employing reflective optical elements, are presented. The first type is based on the use of compound parabolic concentrators, the second type is based on the use of line-focus parabolic troughs. Lightweight concentrators are formed with nickel foil coated silver with a diamond-like carbon layer protection. Secondary optical elements, including lenses and cones, are introduced for a better matching of concentrators and solar cells. Both types of modules are characterized by concentration ratios in the range 20x to 30x, depending on the chosen range of misorientation angles. The estimated specific parameters of these modules operating with single junction AlGaAs/GaAs solar cells are 240 W/sq m and 3 kg/sq m.

Description: Two contact systems for use on shallow junction InP solar cells are described. The feature shared by these two contact systems is the absence of the metallurgical intermixing that normally takes place between the semiconductor and the contact metallization during the sintering process. The n(+)pp(+) cell contact system, consisting of a combination of Au and Ge, not only exhibits very low resistance in the as-fabricated state, but also yields post-sinter resistivity values of 1(exp -7) ohms-sq cm, with effectively no metal-InP interdiffusion. The n(+)pp(+)cell contact system, consisting of a combination of Ag and Zn, permits low resistance ohmic contact to be made directly to a shallow junction p/n InP device without harming the device itself during the contacting process.

Description: The effects of irradiation of In(0.53)Ga(0.47)As/InP (InGaAs/InP) solar cells illuminated through a transparent InP substrate with 1 MeV electrons were measured. These solar cells were developed for bottom cells in tandem solar photovoltaic cell structures. Some InGaAs/InP heterostructures with four layers were grown by liquid phase epitaxy. The structure of the solar cells allowed lightly doped materials in n and p photoactive layers to be used. The base dopant levels ranged from 1.10(exp 17) to 5.10(exp 17) cm(exp -3). The open circuit voltage and the short circuit current were moderately degraded after irradiation with 10(exp 16) cm(exp-2) 1 MeV electrons. This behavior is explained in terms of the device structure and the n and p layer thicknesses.

Description: To provide the transportation of lunar base elements to the moon, large high-energy propulsion systems will be required. Advanced propulsion systems for lunar missions can provide significant launch mass reductions and payload increases. These mass reductions and added payload masses can be translated into significant launch cost savings for the lunar base missions. The masses in low Earth orbit (LEO) were compared for several propulsion systems: nitrogen tetroxide/monomethyl hydrazine (NTO/MMH), oxygen/methane (O2/CH4), oxygen/hydrogen (O2/H2), and metallized O2/H2/Al propellants. Also, the payload mass increases enabled with O2/H2 and O2/H2/Al systems were addressed. In addition, many system design issues involving the engine thrust levels, engine commonality between the transfer vehicle and the excursion vehicle, and the number of launches to place the lunar mission vehicles into LEO will be discussed. Analyses of small lunar missions launched from a single STS-C flight are also presented.

Description: The conceptual design of a single-stage-to-orbit (SSTO) vehicle using a wide variety of evolutionary technologies has recently been completed as a part of NASA's Advanced Manned Launch System (AMLS) study. The employment of new propulsion system technologies is critical to the design of a reasonably sized, operationally efficient SSTO vehicle. This paper presents the propulsion system requirements identified for this near-term AMLS SSTO vehicle. Sensitivities of the vehicle to changes in specific impulse and sea-level thrust-to-weight ratio are examined. The results of a variety of vehicle/propulsion system trades performed on the near-term AMLS SSTO vehicle are also presented.

Description: A Nuclear Thermal Propulsion Workshop, co-sponsored by NASA, DOE and DOD, was held in Cleveland, Ohio on July 10-12, 1990. The workshop was to provide a database of nuclear propulsion concepts and technologies to assist in planning a nuclear propulsion project, identify high priority activities to be initiated early, and to provide cost and schedule estimates for development of concepts to technology readiness level 6 - full system verification in a simulated environment. Sixteen concepts were presented to Technology Review Panels (TRP), and discussed. Each concept was compared to a baseline manned Mars mission. A preliminary comparison of ratings made by the TRP's is presented herein for mission benefit, safety, technical risk, and development cost.

Description: The Space Exploration Initiative (SEI) calls for lunar and Martian exploration missions for which solid-core nuclear thermal rockets (NTRs), in virtue of their single-stage, fully-reusable nature, are ideally suited. NTRs promise double the specific impulse of chemical propulsion. A lunar mission employing a reusable NTR is currently being conducted by NASA. The NTR would be assembled in LEO in such a way that it remained 'radioactively cold' during earth-to-orbit deployment by a heavy-lift chemical booster, and therefore presented no radioactive hazard. Also under consideration is a particle-bed reactor in which the hydrogen propulsive fluid directly cools coated-particle fuel spheres.

Description: The Advanced Solid Rocket Motor (ASRM), currently in its design and development phase, will become the higher reliability/performance next-generation booster for the Space Shuttle. The ASRM design improves safety through the elimination of 229 potential leak paths, thereby eliminating a total of 312 causes of failure. The ASRM will also allow an additional 12,000 lb of payload to be carried by the Shuttle Orbiter. Construction of the ASRM will promote competition through the use of a government owned/contractor operated manufacturing facility. Continuous-mix propellant grain manufacture will be used to enhance safety, quality, and efficiency.

Description: The key to the success of the NASA Space Transportation Main Engine (STME) is its enhanced reliability and reduced costs, rather than enhanced performance and reduced weight. Reliability and cost improvements derive from the design of an engine with fewer components, wider operating margins, conventional materials, and innovative manufacturing techniques; this is exemplified by the STME's combustion chamber, in which the structural jacket and all minifolds are cast as a single piece. 'Platelet' technology is under active consideration as a basis for the nozzle design. The use of hydrostatic bearings in both the turbine and pump ends of the STME turbopump will eliminate bearing-life limitations.

Description: A conceptual design of a model-based fault detection and diagnosis system is developed for the Space Shuttle main engine. The design approach consists of process modeling, residual generation, and fault detection and diagnosis. The engine is modeled using a discrete time, quasilinear state-space representation. Model parameters are determined by identification. Residuals generated from the model are used by a neural network to detect and diagnose engine component faults. Fault diagnosis is accomplished by training the neural network to recognize the pattern of the respective fault signatures. Preliminary results for a failed valve, generated using a full, nonlinear simulation of the engine, are presented. These results indicate that the developed approach can be used for fault detection and diagnosis. The results also show that the developed model is an accurate and reliable predictor of the highly nonlinear and very complex engine.

Description: A majority of the liquid-fueled rocket vehicles developed in the past have been plagued by an instability known as POGO. The POGO phenomenon involves dynamics of the vehicle structure, dynamics of the propellant in the feedline, and the engine dynamic transfer function. Each of these three items must be accurately known in order to determine stability. Metallic bellows are commonly used as segments of propellant feedlines for rocket-propelled vehicles to accommodate temperature-induced length variations, manufacturing tolerances, and gimbaling of the engines. These bellows sections deform radially and change volume when internal pressure varies, and the magnitude of such deformation is much higher than that for the straight, cylindrical segments of the line. The greater flexibility of the bellows decreases the frequency of acoustic oscillations in the line. Calculating elastic stiffness is difficult due to the radial deformation of a bellows section. SHELL was developed specifically to calculate changes in volume of a bellows due to changes in internal pressure. Input to the program consists of tables describing the material, the geometry of the convolutions and loading. The output gives displacements and volume change that can be used for POGO or waterhammer analysis. SHELL is written in standard FORTRAN 77. This program was originally developed on a Univac 1100 series computer and has been successfully implemented on IBM 370 series computers running MVS and DEC VAX series computers running VMS. The main memory requirement for running SHELL under VMS is 116K. The program source code, IBM JCL for compiling and running SHELL, and sample input are provided with the program. SHELL is available on a 9-track 1600 BPI ASCII CARD IMAGE magnetic tape. This program was developed in 1989. IBM is a trademark of International Business Machines Corporation. DEC, VAX and VMS are registered trademarks of Digital Equipment Corporation. Univac 1100 is a trademark of Unisys Corporation.

Description: NASA-Lewis has undertaken the planning and coordination of a joint NASA/DOE/DOD Nuclear Propulsion Project which will investigate both nuclear electric and nuclear thermal concepts. The three-agency team has been tasked with the development of an Interagency Agreement and Memorandum of Understanding, as well as the drafting of a statement as to astronaut crew guidelines and values, the assessment of human-rating requirements, the development of an interagency safety and environmental assessment plan, and the development of test facility requirements. Attention is to be given to the role of SP-100 for nuclear-electric propulsion applications.

Description: NASA-Lewis has undertaken the conceptual development of spacecraft nuclear propulsion systems with DOE support, in order to establish the bases for Space Exploration Initiative lunar and Mars missions. This conceptual evolution project encompasses nuclear thermal propulsion (NTP) and nuclear electric propulsion (NEP) systems. A technology base exists for NTP in the NERVA program files; more fundamental development efforts are entailed in the case of NEP, but this option is noted to offer greater advantages in the long term.

Description: An expert system which diagnoses various seal leakage faults in the High Pressure Oxidizer Turbopump of the SSME was developed using G2 real-time expert system. Three major functions of the software were implemented: model-based data generation, real-time expert system reasoning, and real-time input/output communication. This system is proposed as one module of a complete diagnostic system for the SSME. Diagnosis of a fault is defined as the determination of its type, severity, and likelihood. Since fault diagnosis is often accomplished through the use of heuristic human knowledge, an expert system based approach has been adopted as a paradigm to develop this diagnostic system. To implement this approach, a software shell which can be easily programmed to emulate the human decision process, the G2 Real-Time Expert System, was selected. Lessons learned from this implementation are discussed.

Description: An overview is presented of hybrid rocket propulsion systems whereby combining solids and liquids for launch vehicles could produce a safe, reliable, and low-cost product. The primary subsystems of a hybrid system consist of the oxidizer tank and feed system, an injector system, a solid fuel grain enclosed in a pressure vessel case, a mixing chamber, and a nozzle. The hybrid rocket has an inert grain, which reduces costs of development, transportation, manufacturing, and launch by avoiding many safety measures that must be taken when operating with solids. Other than their use in launch vehicles, hybrids are excellent for simulating the exhaust of solid rocket motors for material development.

Description: The in-flight tests and the operational sequences of the Superfluid Helium On-Orbit Transfer (SHOOT) experiment are outlined. These tests include the transfer of superfluid helium at a variety of rates, the transfer into cold and warm receivers, the operation of an extravehicular activity coupling, and tests of a liquid acquisition device. A variety of different types of instrumentation will be required for these tests. These include pressure sensors and liquid flow meters that must operate in liquid helium, accurate thermometry, two types of quantity gauges, and liquid-vapor sensors.

Description: The technology of the SP-100 space nuclear power system program is compared to that of more familiar solar-power systems. The SP-100 program develops, validates, and demonstrates the technology for space nuclear power systems in the range of 10 to 1000 kilowatts electric for use in future military and civilian space missions. Mission applications, including earth orbiting platforms and lunar/Mars surface power, are enhanced or made possible by SP-100 technology. Attention is given to the SP-100 reference flight system design, the SP-100 nuclear reactor and nuclear-reactor shield, the platform-mounted, tethered, and free-flying reactors, and installation, operation, and disposal options, as well as lunar-Mars surface applications. The SP-100 is presented as one of the nuclear energy sources needed for long-life, compact, lightweight, continuous high power independent of solar orientation, specific orbits, or missions.

Description: To obtain the Advanced Launch System (ALS) primary goals of reduced costs and improved operability, there must be significant reductions in the launch operations and servicing requirements relative to current vehicle designs and practices. One of the primary methods for achieving these goals is by using vehicle electrical power system and controls for all actuation and avionics requirements. A brief status review of the ALS and its associated Advanced Development Program is presented to demonstrate maturation of those technologies that will help meet the overall operability and cost goals. The electric power and actuation systems are highlighted as a specific technology ready not only to meet the stringent ALS goals (cryogenic field valves and thrust vector controls with peak power demands to 75 hp), but also those of other launch vehicles, military and civilian aircraft, lunar/Martian vehicles, and a multitude of commercial applications.

Description: The radio-isotope thermoelectric generator (RTG) for Ulysses' electronic supply is described noting that lack of sufficient sunlight renders usual solar cell power generation ineffective due to increased distance from sun. The history of the RTG in the U.S.A. is reviewed citing the first RTG launch in 1961 with an electrical output of 2.7 W and the improved Ulysses RTG, which provides 285 W at mission beginning and 250 W at mission end. The RTG concept is discussed including the most recent RTG technology developed by the DOE, the General Purpose Heat Source RTG (GPHS-RTG). The system relies upon heat generated by radioactive decay using radioactive plutonium-238, which is converted directly to energy using the Seebeck method.

Description: A radiatively-heated multicouple for use in the next generation of radioisotope thermoelectric generator (RTG) will employ 20 individual couples within a single cell, so that 40 n- and p-semiconductor legs will be interconnected in series. At the hot end of the RTG, the legs will be electrically interconnected using silicon molybdenum; on the cold side, the legs are interconnected by tungsten. The entire cell is then mechanically attached to a radiator, which conducts heat away and radiates it into space. Deep-space applications will use RTGs developed for vacuum operation; thermoelectric converter power systems using a unicouple configuration have flown on such missions as Pioneers 10 and 11, which used lead telluride thermoelectric converters, and Voyagers I and II, which used silicon germanium-based thermoelectrics.