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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2014

Mean Flow Augmented Acoustics in Rocket Systems

Oscillatory motion in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. The customary approach to modeling acoustic waves inside a rocket chamber is to apply the classical inhomogeneous wave equation to the combustion gas. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while the acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The converging section of a rocket nozzle, where gradients in pressure, density, and velocity become large, is a notable region where this approach is not applicable. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. An accurate model of the acoustic behavior within this region where acoustic modes are influenced by the presence of a steady mean flow is required for reliable stability predictions. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, (del squared)(psi) − (lambda/c)(exp 2)(psi) − M(dot)[M(dot)(del)(del(psi))] − 2(lambda(M/c) + (M(dot)del(M))(dot)del(psi)−2(lambda)(psi)[M(dot)del(1/c)]=0 (1) with M as the Mach vector, c as the speed of sound, and lambda as the complex eigenvalue. French apply the finite volume method to solve the steady flow field within the combustion chamber and nozzle with inviscid walls. The complex eigenvalues and eigenvector are determined with the use of the ARPACK eigensolver. The present study employs the COMSOL Multphysics framework to solve the coupled eigenvalue problem using the finite element approach. The study requires one way coupling of the CFD High Mach Number Flow (HMNF) and mathematics module. The HMNF module evaluated the gas flow inside of a solid rocket motor using St. Robert's law modeling solid propellant burn rate, slip boundary conditions, and the supersonic outflow condition. Results from the HMNF model are used by the coefficient form of the mathematics module to determine the eigenvalues of the AVPE. The mathematics model is truncated at the nozzle sonic line, where a zero flux boundary condition is self-satisfying. The remaining boundaries are modeled with a zero flux boundary condition, assuming zero acoustic absorption on all surfaces. Pertinent results from these analyses are the complex valued eigenvalue and eigenvectors. Comparisons are made to the French results to evaluate the modeling approach. A comparison of the French results with that of the present analysis is displayed in figures 1 and 2, respectively. The graphic shows the first tangential eigenvector's real (a) and imaginary (b) values.

Fischbach, Sean R.↗
NASA NTRS2008

Intelligent Memory Module Overcomes Harsh Environments

Solar cells, integrated circuits, and sensors are essential to manned and unmanned space flight and exploration, but such systems are highly susceptible to damage from radiation. Especially problematic, the Van Allen radiation belts encircle Earth in concentric radioactive tori at distances from about 6,300 to 38,000 km, though the inner radiation belt can dip as low as 700 km, posing a severe hazard to craft and humans leaving Earth s atmosphere. To avoid this radiation, the International Space Station and space shuttles orbit at altitudes between 275 and 460 km, below the belts range, and Apollo astronauts skirted the edge of the belts to minimize exposure, passing swiftly through thinner sections of the belts and thereby avoiding significant side effects. This radiation can, however, prove detrimental to improperly protected electronics on satellites that spend the majority of their service life in the harsh environment of the belts. Compact, high-performance electronics that can withstand extreme environmental and radiation stress are thus critical to future space missions. Increasing miniaturization of electronics addresses the need for lighter weight in launch payloads, as launch costs put weight at a premium. Likewise, improved memory technologies have reduced size, cost, mass, power demand, and system complexity, and improved high-bandwidth communication to meet the data volume needs of the next-generation high-resolution sensors. This very miniaturization, however, has exacerbated system susceptibility to radiation, as the charge of ions may meet or exceed that of circuitry, overwhelming the circuit and disrupting operation of a satellite. The Hubble Space Telescope, for example, must turn off its sensors when passing through intense radiation to maintain reliable operation. To address the need for improved data quality, additional capacity for raw and processed data, ever-increasing resolution, and radiation tolerance, NASA spurred the development of the Radiation Tolerant Intelligent Memory Stack (RTIMS).

Source record↗
NASA NTRS1992

A feasibility assessment of magnetic bearings for free-piston Stirling space power converters

This report describes a design and analysis study performed by Mechanical Technology Incorporated (MTI) under NASA Contract NAS3-26061. The objective of the study was to assess the feasibility and efficacy of applying magnetic bearings to free-piston Stirling-cycle power conversion machinery of the type currently being evaluated for possible use in long-term space missions. The study was performed for a 50-kWe Reference Stirling Space Power Converter (RSSPC) system consisting of two 25-kWe free-piston Stirling engine modules. Two different versions of the RSSPC engine modules have been defined under NASA Contract NAS3-25463. These modules currently use hydrostatic gas bearings to support the reciprocating displacer and power piston assemblies. Results of this study show that active magnetic bearings of the attractive electromagnetic type are technically feasible for RSSPC application provided that wire insulation with 60,000-hr life capability at 300 C can be developed for the bearing coils. From a design integration standpoint, both versions of the RSSPC were found to be conceptually amenable to magnetic support of the power piston assembly. However, only one version of the RSSPC was found to be amendable to magnetic support of the displacer assembly. Unacceptable changes to the basic engine design would be required to incorporate magnetic displacer bearings into the second version. Complete magnetic suspension of the RSSPC can potentially increase overall efficiency of the Stirling cycle power converter by 0.53 to 1.4 percent (0.15 to 0.4 efficiency points). Magnetic bearings will also overcome several operational concerns associated with hydrostatic gas bearing systems. However, these advantages are accompanied by a 5 to 8 percent increase in specific mass of the RSSPC, depending on the RSSPC version employed. Additionally, magnetic bearings are much more complex, both mechanically and particularly electronically, than hydrostatic bearings. Accordingly, long-term stability and reliability represent areas of uncertainty for magnetic bearings. Considerable development effort will be required to establish the long-term suitability of these bearings for Stirling space power applications.

Curwen, Peter W.↗
NASA NTRS2015

Advanced Stirling Convertor (ASC) Technology Maturation

The Advanced Stirling Convertor (ASC) development effort was initiated by NASA Glenn Research Center (GRC) with contractor Sunpower Inc. to develop high efficiency thermal-to-electric power conversion technology for NASA Radioisotope Power Systems. Early successful performance demonstrations led to the expansion of the project as well as adoption of the technology by the Department of Energy (DOE) and system integration contractor Lockheed Martin Space Systems Company as part of the Advanced Stirling Radioisotope Generator (ASRG) flight project. The ASRG integrates a pair of ASCs to convert the heat from a pair of General Purpose Heat Source (GPHS) modules into electrical power. The expanded NASA ASC effort included development of several generations of ASC prototypes or Engineering Units to help prepare the ASC technology and Sunpower for flight implementation. Sunpower later had two parallel contracts allowing the last of the NASA Engineering Units called ASC-E3 to serve as pathfinders for the ASC-F flight convertors being built for DOE. The ASC-E3 convertors utilized the ASC-F flight specifications and were built using the ASC-F design and process documentation. Shortly after the first ASC-F Pair achieved initial operation, due to budget constraints, the DOE ASRG flight development contract was terminated. NASA continues to invest in the development of Stirling RPS technology including continued production of the ASC-E3 convertors, seven of which have been delivered with one additional unit in production. Starting in FY2015, Stirling Convertor Technology Maturation has been reorganized as an element of the RPS Stirling Cycle Technology Development (SCTD) Project and long-term plans for continued Stirling technology advancement are in reformulation. This paper provides a status on the ASC project, an overview of advancements made in the design and production of the ASC at Sunpower, and a summary of acceptance tests, reliability tests, and tactical tests at NASA GRC that demonstrate the capabilities of the ASC.

radioisotope↗
NASA NTRSDate not supplied

Capability Gaps Assessment and Identification of Critical Technology Elements for Mars Transit Habitat

The Habitation Systems Development Office (HP40) at NASA Marshall Space Flight Center supports systems engineering, integration, and project management for next generation space habitats. For in space operations and eventual transport of humans to Mars, NASA will rely on a Mars Transit Habitat (TH). The TH will be designed for an up to 1,200-day Mars mission and will carry all food and supplies needed to support four crew for this duration. In the current concept of operations, Mars TH transfers to near rectilinear halo orbit (NRHO) following launch and docks at Gateway as a visiting vehicle. While there, the TH will complete system shakedown testing and a series of analog missions which will grow from 3 to 6+ months in duration TH also augments Gateway’s habitation capabilities beyond 60-days. Proposed Gateway-TH missions will far exceed the longest duration cislunar human missions to date. These shakedown missions will also be the first operational readiness tests of Mars TH’s long-duration deep space systems, and of the split crew (two crew on the surface, two crew in space) operations that are vital to the approach for the first human Mars mission. Once shakedown missions are complete, Mars TH departs Gateway to aggregate with the Mars propulsion system in NRHO before onboarding the crew and final supplies in Earth orbit via a co-manifested Orion-logistics module. Orion and the LM return to Earth prior to the now aggregated Deep Space Transport vehicle’s journey to Mars. Development of the Mars TH requires significant technology development and maturation. Each year the agency performs a capability gaps assessment, where gaps developed by subject matter experts (SMEs) in various engineering/science disciplines are linked to architectural elements in formulation and prioritized. A gap captures the difference between the current state-of-the-art and the maturity of the capability that is needed to enable or enhance a mission as it is currently envisioned in the government reference architecture. HP40 conducted a gap analysis for Mars TH which will be summarized in this poster. Gaps classified as enabling (which means the mission cannot achieve success without gap closure) were subsequently used to identify critical technology elements (CTEs) for Mars TH. This identification of CTEs was also informed by an examination of the product breakdown structure for Mars TH and focused conversations with SMEs in specific technology areas. CTEs identified for Mars TH to date include the following (note this is not a comprehensive list – CTEs listed represent those in MSFC’s capability areas): inflatable softgoods for habitation; enhanced CO2 recovery; life support systems with greater levels of reliability and maintainability; autonomous guidance, navigation, command and control; and radiators for the Mars TH application. The habitation systems development team is currently delving deeper into each CTE to assess technology approaches being pursued, their maturity, and the degree of difficulty in maturation to meet projected Mars TH timelines. This poster will summarize work to date on the identification of enabling capability gaps linked to Mars TH and provide insight into the associated CTEs and technology maturation efforts.

technology development↗
NASA NTRS1975

Interim Report of the Astronomy Spacelab Payloads Study: Ultraviolet and Optical Astronomy - Volume 2

The Space Shuttle will comprise NASA's primary transportation system into near-earth orbit during the 1980s. The Shuttle will provide the astronomical community with a major new capability to send a wide variety of instrumentation into orbit, to utilize it there under manned or automatic control for periods from seven to thirty days, and to return it to the ground. To this end the European Space Research Organization (ESRO) is developing Spacelab, an array of interchangeable components -pressurized manned modules, unpressurized pallets and related support systems - to be mounted in the Shuttle payload bay. Spacelab will offer important opportunities to carry out astronomical research with instruments optimized for specific objectives. With a high flight frequency and with the ability to modify or interchange telescopes and instruments between flights, one will not need to make rigid long-term commitments to specific and compromised telescope/instrument/ detector combinations as is the case for automated satellites. Observational techniques demanding the physical return of data and equipment - in particular the use of photographic film, instruments requiring tight calibration controls, cryogens, high-risk detectors and degradeable optical coatings -will open research areas not readily addressed by automated satellites. Although Shuttle flight duration will be limited to periods from seven to thirty days, substantial data can be obtained with a single instrument on short missions, if targets are carefully selected and prioritized, and a large number of instruments can be accommodated on a single flight. Important astronomical data are regularly obtained on sounding rocket flights of five minutes duration. Spacelab will provide far longer observing periods for large and small telescopes, with resources greatly exceeding those of sounding rockets, while retaining much of the sounding rocket philosophy in terms of instrument flexibility, simplicity, reliability assurance and development costs. The Ultraviolet and Optical Astronomy Program on Spacelab is being planned to provide optical astronomers with relatively simple and regular access to the extended wavelength coverage, the superior image quality and the darkness of the night sky available above the earth's atmosphere. In a rapidly evolving science one cannot predict what problems will be timely in the 1980s. Extrapolating from the current epoch, however, one can envision a continuing interest in such areas as the structure, composition and phenomenology of planetary surfaces and atmospheres; the composition and physical nature of the interstellar medium; the composition, structure and life history of stars, especially those in advanced stages of evolution; the stellar population of other galaxies; such enigmatic phenomena as X-ray binary black holes, pulsars, active galactic nuclei and quasars; large scale interactions between galaxies and the nature of the intergalactic medium; precise calibration of the Hubble law for the expansion of the universe; the average density of the universe; and the existence of extraterrestrial life.

Source record↗
NASA NTRS2003

Enough Is Enough

The spacecraft was nearly integrated and had passed some of its early mechanical and electrical testing. One of its instruments, the Proportional Counter Array (PCA), had a gas leak in one of the five proportional counter modules that made up the array. The science division where the instrument was being developed wanted a gas replenishment system added to assure the PCA would last for the entire mission. Adding a gas replenishment system would mean interrupting spacecraft integration and testing; developing a new subsystem and integrating it onto the spacecraft; modifying all the PCA modules; including a complex integration of the instrument onto the spacecraft; and implementing a more complex performance and environmental test process. It was the wrong answer because it made a simple design more complex and added little value to the mission at a major cost in time and dollars. Our mission couldn't afford the additional budget and schedule risks. XTE was the latest of a long line of projects being managed by my Explorer Program Office, but it was unique in being the first project we had agreed to do for a fixed price. NASA HQ agreed, in return, to provide us with the funding profile we needed to make it happen. We were both trying to break the unhealthy spiral in the Explorer program that saw current missions overrunning and pushing subsequent missions downstream to the point where their science was becoming marginal. The science community was upset and wanted better performance from NASA. I summarized my arguments to the director. The Engineering Directorate had taken responsibility for the spacecraft development when we established XTE as an in-house project at Goddard Space Flight Center, and also was supporting the PCA development. "It adds complexity," I reiterated. "It's a significant cost impact for only a marginal reliability increase". His response was music to my ears, "Jim, I won't stand in your way, but you'll have to convince the scientists and engineers."

Barrowman, James↗
NASA NTRS2016

Advanced Stirling Convertor (ASC) Technology Maturation

The Advanced Stirling Convertor (ASC) development effort was initiated by NASA Glenn Research Center with contractor Sunpower, Inc., to develop high-efficiency thermal-to-electric power conversion technology for NASA Radioisotope Power Systems (RPSs). Early successful performance demonstrations led to the expansion of the project as well as adoption of the technology by the Department of Energy (DOE) and system integration contractor Lockheed Martin Space Systems Company as part of the Advanced Stirling Radioisotope Generator (ASRG) flight project. The ASRG integrates a pair of ASCs to convert the heat from a pair of General Purpose Heat Source (GPHS) modules into electrical power. The expanded NASA ASC effort included development of several generations of ASC prototypes or engineering units to help prepare the ASC technology and Sunpower for flight implementation. Sunpower later had two parallel contracts allowing the last of the NASA engineering units called ASC-E3 to serve as pathfinders for the ASC-F flight convertors being built for DOE. The ASC-E3 convertors utilized the ASC-F flight specifications and were built using the ASC-F design and process documentation. Shortly after the first ASC-F pair achieved initial operation, due to budget constraints, the DOE ASRG flight development contract was terminated. NASA continues to invest in the development of Stirling RPS technology including continued production of the ASC-E3 convertors, seven of which have been delivered with one additional unit in production. Starting in fiscal year 2015, Stirling Convertor Technology Maturation has been reorganized as an element of the RPS Stirling Cycle Technology Development (SCTD) Project and long-term plans for continued Stirling technology advancement are in reformulation. This paper provides a status on the ASC project, an overview of advancements made in the design and production of the ASC at Sunpower, and a summary of acceptance tests, reliability tests, and tactical tests at NASA Glenn that demonstrate the capabilities of the ASC.

Stirling cycle↗
NASA NTRS2004

Epoxy Adhesives for Stator Magnet Assembly in Stirling Radioisotope Generators (SRG)

As NASA seeks to fulfill its goals of exploration and understanding through missions planned to visit the moons of Saturn and beyond, a number of challenges arise from the idea of deep space flight. One of the first problems associated with deep space travel is electrical power production for systems on the spacecraft. Conventional methods such as solar power are not practical because efficiency decreases substantially as the craft moves away from the Sun. The criterion for power generation during deep space missions are very specific, the main points requiring high reliability, low mass, minimal vibration and a long lifespan. A Stirling generator, although fairly old in concept, is considered to be a potential solution for electrical power generation for deep space flight. A Stirling generator works on the same electromagnetic principles of a standard generator, using the linear motion of the alternator through the stationary stator which produces electric induction. The motion of the alternator, however, is produced by the heating and cooling dynamics of pressurized gases. Essentially heating one end and cooling another of a contained gas will cause a periodic expansion and compression of the gas from one side to the other, which a displacer translates into linear mechanical motion. NASA needs to confirm that the materials used in the generator will be able to withstand the rigors of space and the life expectancy of the mission. I am working on the verification of the epoxy adhesives used to bond magnets to the steel lamination stack to complete the stator; in terms of in-service performance and durability under various space environments. Understanding the proper curing conditions, high temperature properties, and degassing problems as well as production difficulties are crucial to the long term success of the generator. system and steel substrate used in the stator. To optimize the curing conditions of the epoxies, modulated differential scanning calorimetry analysis was done as a function of cure time and temperatures. Adhesion bond strength was tested at various temperatures with lap shear samples using Hiperco 50 substrate to ensure that the proper adhesive is being used. To try and solve the problem of bondline thickness, micro glass beads of 0.0017" in diameter were investigated to see if any other physical properties of the epoxy were affected. Efforts will be made to develop a standard, optimized, fabrication process/procedure of sub-scale magnet-stator assemblies for various adhesive performance evaluation studies under simulated generator conditions. Also, accelerated aging testing will be done in a pressurized canister with stator assembly samples for three years to verify if any degassing or thermal degradation of the epoxy occurs. The necessity of verifying the correct epoxy adhesive system for the stator magnet in the SRG is crucial because failure of the stator assembly would jeopardize the electrical system, and thereby the entire mission itself. My work involves specimen fabrications, testing, and data analyses of the epoxy adhesive system for the stator magnet in the SRG is crucial because failure of the stator assembly would jeopardize the electrical system, and thereby the entire mission itself.

Cater, George M.↗
NASA NTRS2022

Flow Visualization, Heat Transfer, and Critical Heat Flux of Flow Boiling in Earth Gravity with Saturated Liquid-Vapor Mixture Inlet Conditions – In Preparation for Experiments Onboard the International Space Station

This study investigates flow boiling of n-Perfluorohexane with saturated two-phase mixture inlet in a rectangular channel of dimensions 114.6-mm heated length, 2.5-mm width, and 5 mm-height. The experiments were performed as part of the Mission Sequence Testing of the Flow Boiling and Condensation Experiment’s (FBCE) Flow Boiling Module (FBM) in the vertical upflow configuration in Earth gravity using the same experimental system that was launched to the International Space Station (ISS) in August 2021. The operating parameters varied are heating configuration (single- and double-sided), mass velocity (380 – 2400 kg/m2s), inlet quality (0.011 – 0.519), and inlet pressure (120 – 179 kPa). High-speed video photographs are presented to explain the two-phase flow patterns within the channel’s heated length. Flow patterns are constituted by low-density and high-density fronts moving along the channel, with the high-density fronts gradually reducing in length due to evaporation. Heat transfer results in terms of flow boiling curves, streamwise wall temperature profiles, streamwise heat transfer coefficient profiles, and average heat transfer coefficients are presented and trends discussed. CHF data from the present experiments are combined with prior databases to compile a consolidated FBCE-CHF database for saturated inlet to expand the ranges of operating conditions and include other flow orientations in Earth gravity. Experimental CHF trends are also discussed. The interfacial lift-off model shows a good CHF predictive accuracy evidenced by a mean absolute error of 11.97% for this consolidated database after constraining it to mass velocities greater than or equal to 500 kg/m2s. Finally, this study confirmed reliability of the upcoming ISS experiments for saturated inlet conditions and the collected Earth-gravity data will be compared to ISS microgravity data.

Flow boiling↗
NASA NTRS2022

A Novel High-Performance Mission-Enabling Multi-Purpose Radioisotope Heat Source

Recent studies indicate science mission concepts targeting access to the sub-surface oceans of icy moons require ice-penetrating cryobots powered by advanced Radioisotope Power Systems (RPS). These systems would deliver waste heat for ice-melting in the range of 10 kW. Minimizing the transit time through the kilometers-thick ice shells to just a few years requires these RPS to utilize heat sources having a higher thermal energy volumetric density than the existing flight-qualified General-Purpose Heat Source (GPHS). A Compact Heat Source (CPHS)has been conceptualized in which the graphite impact shells(GIS) of the existing GPHS are rearranged in a hexagonal aeroshell containing seven GIS per module, as opposed to the standard two per module; offering a thermal energy density of 0.57 W/cm3versus 0.29 W/cm3 offered by the GPHS simply from the repackaging of Technology Readiness Level (TRL)9 subassemblies. Preliminary thermal modeling of the CPHS integrated into a notional radioisotope thermoelectric generator(RTG) structure further suggests that centerline temperatures are well within allowable limits during nominal operation. Given the need for the CPHS for a subset of missions, it is worth exploring the applicability of the CPHS for more general RTG purposes. We discuss herein how the CPHS may be implemented with either heritage or in-development thermoelectric converter technologies into a Next-Generation RTG concept. Due to a higher energy density, the legacy heat rejection fin arrangement must be modified to permit a sufficiently low cold-side temperature. Preliminary finite element analysis suggests fin-root temperatures can be kept as low as 520 K while allowing the generator to fit within the usable dimensions of currently available United States Department of Energy shipping containers. Such temperatures would certainly be compatible with the use of high temperature thermoelectric converter technologies.. A prime candidate is the heritage silicon-germanium (SiGe) unicouple, whose design could be adapted by approximately halving the leg-length, but without changes in hot and cold junction interfaces, which are features critical to the proven performance and reliability of these devices. The estimated Beginning of Life power for a SiGe-based CPHS-RTG using 12 CPHS for a thermal inventory of 10.5 kW is greater than 600 W under deep space operating conditions. Using higher performance segmented couples currently in development that are based on skutterudite,La3−xTe4and 14-1-11 Zintl thermoelectric materials in lieu of the SiGe unicouples would increase the power level to more than1 kW. The high specific power (We/kg) attribute of CPHS-RTGs found in this study could potentially enable Radioisotope Electric Propulsion (REP) mission concepts. Past NASA REP mission concept studies identified specific power needs in excess of 6to 8 We/kg. Based on a GPHS-RTG-like system configuration, we show that at fin root temperatures between 530 K and 570K (deep space environment), specific powers exceeding 10 We/kg are achievable using high performance segmented thermoelectric converters. The compact sizing and power density of the CPHS-RTG would constitute a significant step upgrade in specific power when compared to heritage GPHS-RTG (approximately5.1 We/kg) and off-the-shelf Multi-Mission RTG (approximately 2.6 We/kg).

Nesmith, Bill J.↗
NASA NTRSDate not supplied

In-Situ Reaction Monitoring for Lunar Applications Utilizing a Single Quadrupole Residual Gas Analyzer

With a revived focus to create a permeant presence on the moon and in preparation for future Mars exploration, it is imperative that all re-sources are utilized to their fullest potential. In-Situ Resource Utilization (ISRU) will be critical for future mission success as it would enable independent operation of missions while minimally relying on the complex supply chain created between the Earth, moon and Mars. One of the most critical resources that have been identified for ISRU is the creation of liquid oxygen (O2) for not only breathable air, but also for rocket fuel. On the moon and potentially Mars, one of the most O2 rich resources is in the presence of regolith. Although there are numerous minerals within regolith and various processes to extract O2, this is not the scope of this paper. Since the O2 extraction method can vary, a critical standardized analytical method is needed to verify O2 (or oxygenated precursor compounds) for extraction efficiency and purity. Our group utilizes a modified consumer off-the-shelf (COTS), Residual Gas Analyzer (RGA) Single Quadrupole Mass Spectrometer (QMS). Due to the current field mission requirements, the gas analysis module will need to rely only on mass-to-charge ratios (m/z) and peak intensities to differentiate and quantitate targeted gas-phase reactants or products, i.e. carbon monoxide (CO), carbon dioxide (CO2), methane (CH4) and hydrogen (H2). However, due to solar wind implanted volatiles and various other compounds present in lunar regolith, such as nitrogen (N2), helium (He), and deuterium (D2), deconvolution and quantification of isobaric compounds becomes quite difficult. Specifically, the presence of N2 makes quantification of CO, an oxygenated precursor, difficult due to both compounds having a peak at m/z 28. Approaches to differentiate isobaric components: Initially, the solution to overcome isobaric overlap was to use isotopic ratios, particularly be-tween carbon-13 and nitrogen-15. Although limited literature discuss isotopic composition and attempt to quantify the ratios, variation is particularly high and limited to a few lunar regolith samples. The following were identified as potential approaches to quantify isobaric species, N2 and CO, using our RGA. Threshold Ionization. CO and N2 both have peaks at m/z 28, but only N2 is capable of producing a peak at m/z 7. While a N2 peak at m/z 7 is generally rare, instrument parameters can be optimized in order to maximize the peak intensity. Typically a standard ionization energy of 70 eV is used to monitor a batch process and would produce reliable results for a majority of the reactants or products that we are interested in quantifying. However, to analyze the rare molecular fragment of N2 (N2+ at m/z 7), a different ionization energy may be beneficial. Threshold Ionization Mass Spectrometry (TIMS) will be explored to optimize the N2+ peak at m/z 7 by varying parameters such as ionization energy and pressure. Therefore, a quantitative value at m/z 7 will be used to determine the product gas concentration of N2, which will be subtracted from the total concentration determined at m/z 28 enabling us to approximate the amount of CO. Secondary Detector (infrared gas detector, IR). Another approach to quantitate and differentiate be-tween isobaric components is to leverage alternative modes of detection. Unlike mass spectrometry, which generally uses a separation step prior to analysis (i.e. gas chromatography), IR produces energy that is specifically absorbed depending on the bond and configuration of the molecule. This technique could be leveraged since it is generally only active on non-symmetric molecules, i.e. CO. Utilizing concentrations of CO determined from another detector could help correct concentrations of CO obtained from the QMS. However, the inclusion of another detector adds additional mission requirements, such as power, software changes and avionics updates that add additional cost and schedule requirements.

mass spectrometry↗
NASA NTRS2016

Challenges with Electrical, Electronics, and Electromechanical Parts for James Webb Space Telescope

James Webb Space Telescope (JWST) is the space-based observatory that will extend the knowledge gained by the Hubble Space Telescope (HST). Hubble focuses on optical and ultraviolet wavelengths while JWST focuses on the infrared portion of the electromagnetic spectrum, to see the earliest stars and galaxies that formed in the Universe and to look deep into nearby dust clouds to study the formation of stars and planets. JWST, which commenced creation in 1996, is scheduled to launch in 2018. It includes a suite of four instruments, the spacecraft bus, optical telescope element, Integrated Science Instrument Module (ISIM, the platform to hold the instruments), and a sunshield. The mass of JWST is approximately 6200 kg, including observatory, on-orbit consumables and launch vehicle adaptor. Many challenges were overcome while providing the electrical and electronic components for the Goddard Space Flight Center hardware builds. Other difficulties encountered included developing components to work at cryogenic temperatures, failures of electronic components during development and flight builds, Integration and Test electronic parts problems, and managing technical issues with international partners. This paper will present the context of JWST from a EEE (electrical, electronic, and electromechanical) perspective with examples of challenges and lessons learned throughout the design, development, and fabrication of JWST in cooperation with our associated partners including the Canadian Space Agency (CSA), the European Space Agency (ESA), Lockheed Martin and their respective associated partners. Technical challenges and lessons learned will be discussed.

EEE Parts↗
NASA NTRS2002

Mountain-Top-to-Mountain-Top Optical Link Demonstration

A mountain-top-to-mountain-top optical communications demonstration was conducted between the JPL Table Mountain Facility (TMF), Wrightwood, California, and Strawberry Peak (SP), Lake Arrowhead, California, during the months of August and September of 2000. The link was nearly horizontal at an altitude of 2 km and spanned a range of 46.8 km. A 780-nm multibeam beacon broadcast from TMF was received by the JPL Optical Communications Demonstrator (OCD) located at SP. The received beacon was utilized as a pointing reference to retransmit an 852-nm communications laser beam, modulated at 400 Mb/s by a PN7 pseudo-random bit stream (PRBS) sequence. The long atmospheric path resulted in atmospheric-turbulence-induced angle-of-arrival fluctuations of the beacon at the OCD aperture. A .ne-steering control loop was used to track the resulting beacon centroid motion and update the pointing of the communications laser beam transmitted from SP to TMF. Fried parameters, or r0, inferred from focal spot sizes received at SP were 4 to 5 cm whereas, using the spot sizes received at TMF, they were 2 to 3 cm. In both cases, theory predicts larger r0 values. The predicted angle-of-arrival fluctuations compare well with measured rms displacements exhibited by the focal spots at either end of the link. An uncompensated error of ~1.1 rad in the x-axis and ~2 rad in the y-axis was obtained using centroid data logged by the OCD. Average bit-error rates of 10-5 were recorded for extended periods of time. An atmospheric coherence length r0 of 3 to 5 cm was inferred using the focal-plane spot size measured on the CCD tracking sensor and compared to a predicted value of 5 to 7 cm using reasonable atmospheric models. The irradiance bounds required for the CCD tracking sensor to perform centroiding was found to range from 2000 to 3000 integrated pixel counts, although a more reliable range was 600 to 3000, indicating a dynamic range of 6 to 11 dB. The motion of the spot on the focal plane was also recorded and yielded rms angle-of-arrival-induced focal-plane displacement of 9 to 11 m, compared to a predicted value of 7.8 to 11 m. The irradiance measurements made with the OCD clearly indicate that best tracking performance is obtained when the mean received signal is 2000 to 2200 counts. The best tracking performance resulted in an rms uncompensated error of 1.1 rad. The uncompensated error appeared to increase with either an increase or decrease in mean beacon-signal level. The rms uncompensated error determined by deriving the beacon displacement power spectral density from the beacon centroid-versustime data and the rejection function of the control loop yielded a higher value of 3.4 rad.

Biswas, A.↗
NASA NTRS1994

Advanced technology for a satellite multichannel demultiplexer/demodulator

Satellite on-board processing is needed to efficiently service multiple users while at the same time minimizing earth station complexity. The processing satellite receives a wideband uplink at 30 GHz and down-converts it to a suitable intermediate frequency. A multichannel demultiplexer then separates the composite signal into discrete channels. Each channel is then demodulated by bulk demodulators, with the baseband signals routed to the downlink processor for retransmission to the receiving earth stations. This type of processing circumvents many of the difficulties associated with traditional bent-pipe repeater satellites. Uplink signal distortion and interference are not retransmitted on the downlink. Downlink power can be allocated in accordance with user needs, independent of uplink transmissions. This allows the uplink users to employ different data rates as well as different modulation and coding schemes. In addition, all downlink users have a common frequency standard and symbol clock on the satellite, which is useful for network synchronization in time division multiple access schemes. The purpose of this program is to demonstrate the concept of an optically implemented multichannel demultiplexer (MCD). A proof-of-concept (POC) model has been developed which has the ability to receive a 40 MHz wide composite signal consisting of up to 1000 40 kHz QPSK modulated channels and perform the demultiplexing process. In addition a set of special test equipment (STE) has been configured to evaluate the performance of the POC model. The optical MCD is realized as an acousto-optic spectrum analyzer utilizing the capability of Bragg cells to perform the required channelization. These Bragg cells receive an optical input from a laser source and an RF input (the signal). The Bragg interaction causes optical output diffractions at angles proportional to the RF input frequency. These discrete diffractions are optically detected and output to individual demodulators for baseband conversion. Optimization of the MCD design was conducted in order to achieve a compromise between two opposing sources of signal degradation: adjacent channel interference and intersymbol interference. The system was also optimized to allow simple, inexpensive ground stations communications with the MCD. These design goals led to the realization of a POC MCD which demonstrates the demultiplexing function with minimal signal degradation. Performance evaluation results using the STE equipment indicate that the dynamic range of the demultiplexer in the presence of adjacent and multiple channel loading is 40 - 50 dB. Measured bit error rate (BER) probabilities varied from the predicted theoretical results by one dB or less. The performance of the proof-of-concept model indicate that the development of a space qualified optically implemented MCD are feasible. The advantages to such an implementation include reduced size, weight and power and increased reliability when compared with electronic approaches. All of these factors are critical to on-board satellite processors. Further optimization can be conducted which trade ground station complexity and MCD performance to achieve desired system results.

Abramovitz, Irwin J.↗
NASA NTRS2009

Design Evolution and Verification of the A-3 Chemical Steam Generator

Following is an overview of the Chemical Steam Generator system selected to provide vacuum conditions for a new altitude test facility, the A-3 Test Stand at Stennis Space Center (SSC) in Bay St. Louis, MS. A-3 will serve as NASA s primary facility for altitude testing of the J-2X rocket engine, to be used as the primary propulsion device for the upper stages of the Ares launch vehicles. The Chemical Steam Generators (CSGs) will produce vacuum conditions in the test cell through the production and subsequent supersonic ejection of steam into a diffuser downstream of the J-2X engine nozzle exit. The Chemical Steam Generators chosen have a rich heritage of operation at rocket engine altitude test facilities since the days of the Apollo program and are still in use at NASA White Sands Test Facility (WSTF) in New Mexico. The generators at WSTF have been modified to a degree, but are still very close to the heritage design. The intent for the A-3 implementation is to maintain this heritage design as much as possible, making minimal updates only where necessary to substitute for obsolete parts and to increase reliability. Reliability improvements are especially desired because the proposed system will require 27 generators, which is nine times the largest system installed in the 1960s. Improvements were suggested by the original design firm, Reaction Motors, by NASA SSC and NASA WSTF engineers, and by the A-3 test stand design contractor, Jacobs Technology, Inc. (JTI). This paper describes the range of improvements made to the design to date, starting with the heritage generator and the minor modifications made over time at WSTF, to the modernized configuration which will be used at A-3. The paper will discuss NASA s investment in modifications to SSC s E-2 test facility fire a full-scale Chemical Steam Generator in advance of the larger steam system installation at A-3. Risk mitigation testing will be performed in early 2009 at this test facility to verify that the CSGs operate as expected. The generator which will undergo this testing is of the most recent A-3 configuration, and will be instrumented far in excess of what is normally required for operation. The extra data will allow for easier troubleshooting and more complete knowledge of expected generator performance. In addition, the early testing will give SSC personnel experience in operating the CSG systems, which will expedite the process of installation and activation at A-3. Each Chemical Steam Generator is supported by a complement of valves, instruments, and flow control devices, with the entire assembly called a "module." The generators will be installed in groups of three, historically called "units". A module is so called because of its modular ability to be replaced or serviced without disturbing the other two modules installed on the same unit. A module is pictured in Figure 1, shown with its generator secured by white bands in its shipping (vs. installed) configuration. The heritage system at WSTF is composed of a single unit (three generator modules), pictured in Figure 2 as it was installed in 1965. In contrast, A-3 will have nine units operating in parallel to achieve vacuum conditions appropriate for testing the J-2X engine. Each of the combustors operates in two modes and achieves the so-called "full-steam" mode after all three of its stages ignite. Ignition of the first stage is achieved by exciting a spark plug; the second stage and main stage are lit by the flame front of the previous stage. The main stage burns approximately 97% of the total propellant flow and uses the heat energy to vaporize water into superheated steam. While the main stage remains unlit, the combustor is in so-called "idle" mode. In the WSTF system, this idle mode is not optimized for water usage, and does not need to be, as the water is pumped from a large reservoir. The water supply at A-3 will be contained in tanks with finite volume, so water optimization is preferred for the modnized configuration. Multiple solutions for this issue have been proposed, with the leading concept being a change to the operational definition of "idle mode," with the generator running in a lower heat flux condition.

Kirchner, Casey K.↗
NASA NTRS2013

Hyperspectral Microwave Atmospheric Sounder (HyMAS) Architecture and Design Accommodations

The Hyperspectral Microwave Atmospheric Sounder (HyMAS) is being developed at Lincoln Laboratories and accommodated by the Goddard Space Flight Center for a flight opportunity on a NASA research aircraft. The term "hyperspectral microwave" is used to indicate an all-weather sounding that performs equivalent to hyperspectral infrared sounders in clear air with vertical resolution of approximately 1 km. Deploying the HyMAS equipped scanhead with the existing Conical Scanning Microwave Imaging Radiometer (CoSMIR) shortens the path to a flight demonstration. Hyperspectral microwave is achieved through the use of independent RF antennas that sample the volume of the Earth s atmosphere through various levels of frequencies, thereby producing a set of dense, spaced vertical weighting functions. The simulations proposed for HyMAS 118/183-GHz system should yield surface precipitation rate and water path retrievals for small hail, soft hail, or snow pellets, snow, rainwater, etc. with accuracies comparable to those of the Advanced Technology Microwave Sounder. Further improvements in retrieval methodology (for example, polarization exploitation) are expected. The CoSMIR instrument is a packaging concept re-used on HyMAS to ease the integration features of the scanhead. The HyMAS scanhead will include an ultra-compact Intermediate Frequency Processor (IFP) module that is mounted inside the door to improve thermal management. The IFP is fabricated with materials made of Low-Temperature Co-fired Ceramic (LTCC) technology integrated with detectors, amplifiers, A/D conversion and data aggregation. The IFP will put out 52 channels of 16 bit data comprised of 4-9 channel data streams for temperature profiles and 2-8 channel streams for water vapor. With the limited volume of the existing CoSMIR scanhead and new HyMAS front end components, the HyMAS team at Goddard began preliminary layout work inside the new drum. Importing and re-using models of the shell, the scan head computer, and the slip rings developed for CoSMIR was the starting point. The next step was to modify the antenna faceplate to accommodate the dimensions of the three dual polarization Gaussian Optics Antenna (GOA) assemblies. Two mechanical concepts for the core technology, the hyperspectral IFP, were captured in a design tradeoff. Connector models considered minimum bend radii for the IFP analog connectors. Hyperspectral imaging is accomplished by strategically using a short wavelength intermediate frequency of 18-29 GHz, and thus reducing the size of components in the connection of the front end to the IFP. The SMK (2.92mm) Series connector will lay near the hinge line to minimize its flexing. The digital output of the IFP will use a Serial Peripheral Interface (SPI) that must be accommodated by the scan head computer. To make that computer more reliable, maintainable, and forward compatible with the 52 HyMAS channels, a testbed of the scan head, calibration, and archive computers and the PIC24 microprocessor that resides on the IFP is in development. The computers will be programmed using a new framework application called Interoperable Remote Component (IRC). This software allows flexibility to program computers that communicate with each other and can adapt easily to the emerging HyMAS requirements for data format, algorithms, and graphical user interface (GUI). It is expected that the CoSMIR instrument will cut over to the IRC after it is adapted on an updated CoSMIR testbed.

Hilliard, Lawrence↗
NASA NTRS1997

Quantitative Measurement of Oxygen in Microgravity Combustion

A low-gravity environment, in space or in ground-based facilities such as drop towers, provides a unique setting for studying combustion mechanisms. Understanding the physical phenomena controlling the ignition and spread of flames in microgravity has importance for space safety as well as for better characterization of dynamical and chemical combustion processes which are normally masked by buoyancy and other gravity-related effects. Due to restrictions associated with performing measurements in reduced gravity, diagnostic methods which have been applied to microgravity combustion studies have generally been limited to capture of flame emissions on film or video, laser Schlieren imaging and (intrusive) temperature measurements using thermocouples. Given the development of detailed theoretical models, more sophisticated diagnostic methods are needed to provide the kind of quantitative data necessary to characterize the properties of microgravity combustion processes as well as provide accurate feedback to improve the predictive capabilities of the models. When the demands of space flight are considered, the need for improved diagnostic systems which are rugged, compact, reliable, and operate at low power becomes apparent. The objective of this research is twofold. First, we want to develop a better understanding of the relative roles of diffusion and reaction of oxygen in microgravity combustion. As the primary oxidizer species, oxygen plays a major role in controlling the observed properties of flames, including flame front speed (in solid or liquid flames), extinguishment characteristics, flame size and flame temperature. The second objective is to develop better diagnostics based on diode laser absorption which can be of real value in both microgravity combustion research and as a sensor on-board Spacelab as either an air quality monitor or as part of a fire detection system. In our prior microgravity work, an eight line-of-sight fiber optic system measured water vapor mole fractions in the NASA Lewis 2.2-sec Drop Tower. In that system, the laser and all electronics resided at the top of the drop tower and was connected via a fiber optic cable to the rig, on which a 'pitch and catch' set of fiber collimating lenses were used to transmit the laser beam across a jet diffusion flame. This system required eight independent detection/demodulation units and had poor spatial resolution. This research builds on this earlier work, resulting in an improved capability for quantitative, nonintrusive measurement of major combustion species. A vertical cavity surface-emitting diode laser (VCSEL) and a continuous spatial scanning method permit the measurement of temporal and spatial profiles of the concentrations and temperatures of molecular oxygen. High detection sensitivity is achieved with wavelength modulation spectroscopy (WMS). One-g experiments are performed using a slot diffusion flame. Microgravity measurements on a solid fuel (cellulose sheet) system are planned for the NASA Lewis 2.2-second Drop Tower Facility.

Silver, Joel A.↗