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The Sabatier/Electrolysis (S/E) process is an attractive approach to in situ propellant production (ISPP), and a breadboard demonstration of this process at Lockheed Martin Astronautics funded by JPL performed very well, with high conversion efficiency, and reliable diurnal operation. There is a net usage of hydrogen in the S/E process, and this has been the principal problem for this approach to ISPP.
Our paper describes a technique for generating tunable UV laser radiation between 250-300 nm capable of energies up to 30-5O mJ/pulse. The tunability of this source is attractive for selecting ozone absorption cross sections which are optimal for ozone DIAL detection throughout the troposphere. A Nd:YAG laser is used to pump a pulsed titanium sapphire laser which is then frequency tripled into the UV. Titanium sapphire (TiS) lases robustly between 750-900 nm. In initial experiments we have converted 110 mJ of 867 nm from a TiS laser into 28 mJ at 289 nm. The energy conversion efficiency was 62% for doubling into 433 nm and 25% into 289 nm.
Completely integrated thermophotovoltaic (TPV) power sources in the range of 100 to 500 watts are being developed. The technical approach taken in this project focuses on optimizing the integrated performance of the primary subsystems in order to yield high energy conversion efficiency and cost effectiveness. An important aspect of the approach is the use of a narrow band fibrous emitter radiating to a bandgap matched photovoltaic array to minimize thermal and optical recuperation requirements, as well as the non-recoverable heat losses. For the prototype system, fibrous ytterbia emitters radiating in a narrow band centered at 980 nm are matched with high efficiency silicon photoconverters. The integrated system includes a dielectric stack filter for optical energy recovery and a ceramic recuperator for thermal energy recovery. The prototype TPV system uses a rapid mix distributed fuel delivery system with controlled feeding of the fuel and heated air into a flame at the surface of the emitter. This makes it possible to operate at air preheat temperatures well above the auto-ignition temperature of the fuel thereby substantially increasing the system efficiency. The system has been operated with air preheat temperatures up to 1367 K and has produced a uniform narrow band radiation over the surface of the emitter with this approach. The design of the system is described and test data for the system and some of the key components are presented. The results from a system model, which show the impact of various parameters on system performance, are also discussed.
The objective of this research is to investigate system level performance and design issues associated with magnetic flux compression devices for aerospace power generation and propulsion. The proposed concept incorporates the principles of magnetic flux compression for direct conversion of nuclear/chemical detonation energy into electrical power. Specifically a magnetic field is compressed between an expanding detonation driven diamagnetic plasma and a stator structure formed from a high temperature superconductor (HTSC). The expanding plasma cloud is entirely confined by the compressed magnetic field at the expense of internal kinetic energy. Electrical power is inductively extracted, and the detonation products are collimated and expelled through a magnetic nozzle. The long-term development of this highly integrated generator/propulsion system opens up revolutionary NASA Mission scenarios for future interplanetary and interstellar spacecraft. The unique features of this concept with respect to future space travel opportunities are as follows: ability to implement high energy density chemical detonations or ICF microfusion bursts as the impulsive diamagnetic plasma source; high power density system characteristics constrain the size, weight, and cost of the vehicle architecture; provides inductive storage pulse power with a very short pulse rise time; multimegajoule energy bursts/terawatt power bursts; compact pulse power driver for low-impedance dense plasma devices; utilization of low cost HTSC material and casting technology to increase magnetic flux conservation and inductive energy storage; improvement in chemical/nuclear-to-electric energy conversion efficiency and the ability to generate significant levels of thrust with very high specific impulse; potential for developing a small, lightweight, low cost, self-excited integrated propulsion and power system suitable for space stations, planetary bases, and interplanetary and interstellar space travel; potential for attaining specific impulses approaching 10 (exp 6) seconds, which would enable missions to the outer planets within ten years and missions at interstellar distances within fifty years.
Free-piston Stirling technology efforts in the past typically were intended to address power needs in the multi-kilowatt range. The Stirling power option was attractive primarily because of the high conversion efficiency and potential for long life. In recent years, several technology efforts have focused on the free-piston Stirling power d convertor for space power applications, however the more recent efforts are intended to provide power at levels far below one kilowatt. Through a variety of projects funded by both NASA and DOE, the free-piston Stirling convertor technology has successfully demonstrated high efficiency and long life. Other areas of concern, such as control of multi-convertor systems, and vibration reduction have also been addressed. Efforts are being initiated to address issues such as electro-magnetic interference (EMI), radiation tolerance of organic materials, and the ability to operate through launch loads and survive with integrity. The status of the technology in these areas will be briefly discussed in this paper.
The hot-electron bolometer mixer made from a high-T, superconductor (HTS) was introduced recently as an alternative to a Schottky mixer at THz frequencies. The performance of the mixer depends on the total thermal conductance for heat removal from the phonon sub-system due to either length-dependent phonon diffusion or phonon escape to the substrate. We have measured both the length and temperature dependencies of the IF bandwidth of the mixers fabricated from 25-35 mn thick YBCO films on MgO and sapphire substrates. The films were grown by a laser deposition technique and electron-beam lithography was used to define bridge lengths down to 50 nm. Mixer measurements were done using signal frequencies in the range of 1-100 GHz. For 50 nm and 400 nm long devices on MgO, the 3-dB bandwidth was about 100 MHz. At temperatures below 60 K, the hot-electron plateau was clearly seen starting around 2-3 GHz. At temperatures above 70 K, the flux-flow effects begin to dominate and the IF bandwidth increases to 1-8 GHz, while the conversion efficiency drops by several dB. This temperature dependence of the IF bandwidth can account for previously reported unexpectedly high bandwidth of HTS mixers.
This report summarizes the participation of the University of Denver in an airborne measurement program, SULFUR 6, which was undertaken in late September and early October of 1998 by the Deutsches Zentrum fur Luft und Raumfahrt (DLR). Scientific findings from two papers that have been published or accepted and from one manuscript that is in preparation are presented. The SULFUR 6 experiment was designed to investigate the emissions from subsonic aircraft to constrain calculations of possible atmospheric chemical and climatic effects. The University of Denver effort contributed toward the following SULFUR 6 goals: (1) To investigate the relationship between fuel sulfur content (FSC--mass of sulfur per mass of fuel) and particle number and mass emission index (El--quantity emitted per kg of fuel burned); (2) To provide upper and lower limits for the mass conversion efficiency (nu) of fuel sulfur to gaseous and particulate sulfuric acid; (3) To constrain models of volatile particle nucleation and growth by measuring the particle size distribution between 3 and 100 nm at aircraft plume ages ranging from 10(exp -1) to 10(exp 3) s; (4) To determine microphysical and optical properties and bulk chemical composition of soot particles in aircraft exhaust; and (5) To investigate the differences in particle properties between aircraft plumes in contrail and non-contrail situations. The experiment focused on emissions from the ATTAS research aircraft (a well characterized, but older technology turbojet) and from an in-service Boeing 737-300 aircraft provided by Lufthansa, with modem, high-bypass turbofan engines. Measurements were made from the DLR Dassault Falcon 900 aircraft, a modified business jet. The Atmospheric Effects of Aviation Program (AEAP) provided funding to operate an instrument, the nucleation-mode aerosol size spectrometer (N-MASS), during the SULFUR 6 campaign and to analyze the data. The N-MASS was developed at the University of Denver with the support of NOAA's Office of Global Programs and NASA's AEAP and measures particle size distributions in the 4-100 nm range.
Deep space missions have a strong need for compact, high power density, reliable and long life electrical power generation and storage under extreme temperature conditions. Conventional power generating devices become inefficient at very low temperatures (temperatures lower than 200 K encountered during Mars missions for example) and rechargeable energy storage devices cannot be operated thereby limiting mission duration. At elevated temperatures (for example for planned solar probe or Venus lander missions), thin film interdiffusion destroys electronic devices used for generating and storing power. Solar power generation strongly depends upon the light intensity, which falls rapidly in deep interplanetary missions (beyond 5 AU), and in planetary missions in the sun shadow or in dusty environments (Mars, for example). Radioisotope thermoelectric generators (RTGs) have been successfully used for a number of deep space missions RTGs. However, their energy conversion efficiency and specific power characteristics are quite low, and this technology has been limited to relatively large systems (more than 100 W). The National Aeronautics and Space Administration (NASA) and the Jet Propulsion Laboratory (JPL) have been planning the use of much smaller spacecrafts that will incorporate a variety of microdevices and miniature vehicles such as microdetectors, microsensors, and microrovers. Except for electrochemical batteries and solar cells, there are currently no available miniaturized power sources. Novel technologies that will function reliably over a long duration mission (ten years and over), in harsh environments (temperature, pressure, and atmosphere) must be developed to enable the success of future space missions. It is also expected that such micropower sources could have a wide range of terrestrial applications, in particular when the limited lifetime and environmental limitations of batteries are key factors. Additional information is contained in the original extended abstract.
A compact and high-pulse-energy Ti:Sapphire laser with its Third Harmonic Generation (THG) has been developed for an airborne ozone differential absorption lidar (DIAL) to study the distributions and concentrations of the ozone throughout the troposphere. The Ti:Sapphire laser, pumped by a frequency-doubled Nd:YAG laser and seeded by a single mode diode laser, is operated either at 867 nm or at 900 nm with a pulse repetition frequency of 20 Hz. High energy laser pulses (more than 110 mJ/pulse) at 867 nm or 900 nm with a desired beam quality have been achieved and utilized to generate its third harmonic at 289nm or 300nm, which are on-line and off-line wavelengths of an airborne ozone DIAL. After being experimentally compared with Beta-Barium Borate (beta - BaB2O4 or BBO) nonlinear crystals, two Lithium Triborate (LBO) crystals (5 x 5 x 20 cu mm) are selected for the Third Harmonic Generation (THG). In this paper, we report the Ti:Sapphire laser at 900 nm and its third harmonic at 300 nm. The desired high ultraviolet (UV) output pulse energy is more than 30 mJ at 300 nm and the energy conversion efficiency from 900 nm to 300 nm is 30%.
Considerable attention has been recently received on the impact of aircraft-produced aerosols upon the global climate. Sampling particles directly from jet engines has been performed by different research groups in the U.S. and Europe. However, a large variation has been observed among published data on the conversion efficiency and emission indexes of jet engines. The variation results surely from the differences in test engine types, engine operation conditions, and environmental conditions. The other factor that could result in the observed variation is the performance of sampling probes used. Unfortunately, it is often neglected in the jet engine community. Particle losses during the sampling, transport, and dilution processes are often not discussed/considered in literatures. To address this issue, we evaluated the performance of one sampling probe by challenging it with monodisperse particles. A significant performance difference was observed on the sampling probe evaluated under different temperature conditions. Thermophoretic effect, nonisokinetic sampling and turbulence loss contribute to the loss of particles in sampling probes. The results of this study show that particle loss can be dramatic if the sampling probe is not well designed. Further, the result allows ones to recover the actual size distributions emitted from jet engines.
Using the exact solution of the axisymmetric pulsar magnetosphere derived in a previous publication and the conservation laws of the associated MHD flow, we show that the Lorentz factor of the outflowing plasma increases linearly with distance from the light cylinder. Therefore, the ratio of the Poynting to particle energy flux, generically referred to as sigma, decreases inversely proportional to distance, from a large value (typically approx. greater than 10(exp 4)) near the light cylinder to sigma approx. = 1 at a transition distance R(sub trans). Beyond this distance the inertial effects of the outflowing plasma become important and the magnetic field geometry must deviate from the almost monopolar form it attains between R(sub lc), and R(sub trans). We anticipate that this is achieved by collimation of the poloidal field lines toward the rotation axis, ensuring that the magnetic field pressure in the equatorial region will fall-off faster than 1/R(sup 2) (R being the cylindrical radius). This leads both to a value sigma = a(sub s) much less than 1 at the nebular reverse shock at distance R(sub s) (R(sub s) much greater than R(sub trans)) and to a component of the flow perpendicular to the equatorial component, as required by observation. The presence of the strong shock at R = R(sub s) allows for the efficient conversion of kinetic energy into radiation. We speculate that the Crab pulsar is unique in requiring sigma(sub s) approx. = 3 x 10(exp -3) because of its small translational velocity, which allowed for the shock distance R(sub s) to grow to values much greater than R(sub trans).
Fission technology can enable rapid, affordable access to any point in the solar system. If fission propulsion systems are to be developed to their full potential; however, near-term customers must be identified and initial fission systems successfully developed, launched, and operated. Studies conducted in fiscal year 2001 (IISTP, 2001) show that fission electric propulsion (FEP) systems operating at 80 kWe or above could enhance or enable numerous robotic outer solar system missions of interest. At these power levels it is possible to develop safe, affordable systems that meet mission performance requirements. In selecting the system design to pursue, seven evaluation criteria were identified: safety, reliability, testability, specific mass, cost, schedule, and programmatic risk. A top-level comparison of three potential concepts was performed: an SP-100 based pumped liquid lithium system, a direct gas cooled system, and a heatpipe cooled system. For power levels up to at least 500 kWt (enabling electric power levels of 125-175 kWe, given 25-35% power conversion efficiency) the heatpipe system has advantages related to several criteria and is competitive with respect to all. Hardware-based research and development has further increased confidence in the heatpipe approach. Successful development and utilization of a "Phase 1" fission electric propulsion system will enable advanced Phase 2 and Phase 3 systems capable of providing rapid, affordable access to any point in the solar system.
A combination of CN counts, Ames wire impactor size analyses and optical particle counter data in aircraft exhaust results in a continuous particle size distribution between 0.01 micrometer and 1 micrometer particle radius sampled in the exhaust of a Boeing 757 research aircraft. The two orders of magnitude size range covered by the measurements correspond to 6-7 orders of magnitude particle concentration. CN counts and small particle wire impactor data determine a nucleation mode, composed of aircraft-emitted sulfuric acid aerosol, that contributes between 62% and 85% to the total aerosol surface area and between 31% and 34% to its volume. Soot aerosol comprises 0.5% of the surface area of the sulfuric acid aerosol. Emission indices are: EIH2SO4 = 0.05 g/kgFUEL and (0.2-0.5) g/kgFUEL (for 75 ppmm and 675 ppmm fuel-S, respectively), 2.5E4<EISOOT<1.3E-3 g/kgFUEL, and EICN 8E14 and 1.3E16 particles/kgFUEL (for 75 and 675 ppmm fuel-S). The sulfur (gas) to H2SO4 (particle) conversion efficiency is between 10% and 25%.
Fission technology can enable rapid, affordable access to any point in the solar system. If fission propulsion systems are to be developed to their full potential; however, near-term customers must be identified and initial fission systems successfully developed, launched, and operated. Studies conducted in fiscal year 2001 (IISTP, 2001) show that fission electric propulsion (FEP) systems with a specific mass at or below 50 kg/kWjet could enhance or enable numerous robotic outer solar system missions of interest. At the required specific mass, it is possible to develop safe, affordable systems that meet mission requirements. To help select the system design to pursue, eight evaluation criteria were identified: system integration, safety, reliability, testability, specific mass, cost, schedule, and programmatic risk. A top-level comparison of four potential concepts was performed: a Testable, Passive, Redundant Reactor (TPRR), a Testable Multi-Cell In-Core Thermionic Reactor (TMCT), a Direct Gas Cooled Reactor (DGCR), and a Pumped Liquid Metal Reactor.(PLMR). Development of any of the four systems appears feasible. However, for power levels up to at least 500 kWt (enabling electric power levels of 125-175 kWe, given 25-35% power conversion efficiency) the TPRR has advantages related to several criteria and is competitive with respect to all. Hardware-based research and development has further increased confidence in the TPRR approach. Successful development and utilization of a "Phase I" fission electric propulsion system will enable advanced Phase 2 and Phase 3 systems capable of providing rapid, affordable access to any point in the solar system.
The Vanguard I dish-Stirling module program, initiated in 1982, produced the Vanguard I module, a commercial prototype erected by the Advanco Corporation. The module, which automatically tracks the sun, combines JPL mirrored concentrator technology, an advanced Stirling Solar II engine/generator, a low cost microprocessor-controlled parabolic dish. Vanguard I has a 28% sunlight to electricity conversion efficiency. If tests continue to prove the system effective, Advanco will construct a generating plant to sell electricity to local utilities. An agreement has also been signed with McDonnell Douglas to manufacture a similar module.
Recent advances in crystalline solar cell technology are reviewed. Dual-junction and triple-junction solar cells are presently available from several U. S. vendors. Commercially available triple-junction cells consisting of GaInP, GaAs, and Ge layers can produce up to 27% conversion efficiency in production lots. Technology status and performance figures of merit for currently available photovoltaic arrays are discussed. Three specific NASA mission applications are discussed in detail: Mars surface applications, high temperature solar cell applications, and integrated microelectronic power supplies for nanosatellites.
This viewgraph presentation provides an overview of material science and technology in the area of propulsion energetics. The authors note that conventional propulsion systems are near peak performance and further refinements in manufacturing, engineering design and materials will only provide incremental increases in performance. Energetic propulsion technologies could potential solve the problems of energy storage density and energy-to-thrust conversion efficiency. Topics considered include: the limits of thermal propulsion systems, the need for energetic propulsion research, emerging energetic propulsion technologies, materials research needed for advanced propulsion, and potential research opportunities.
This viewgraph representation provides an overview of research being conducted at NASA's Marshall Space Flight Center. Conventional propulsion systems are at near peak performance levels but will not enable the science and exploration deep space missions NASA envisions. Energetic propulsion technologies can make these missions possible but only if the fundamental problems of energy storage density and energy to energy thrust conversion efficiency are solved. Topics covered include: research rationale, limits of thermal propulsion systems, need for propulsion energetics research, emerging energetic propulsion technologies, and potential research opportunities.