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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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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp −6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe the design and testing of the engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re-deposit within the feed system. Heating is accomplished through a number of individual zones to control the overall power expended on heating the system and insulation is employed to minimize the amount of power used to heat the system prior to thruster operation.

Polzin, Kurt A.↗
NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗
NASA NTRS1981

Head for high speed spinner having a vacuum chuck

The head for a high-speed spinner is characterized by a substantially cylindrical body adapted to be mounted at the distal end of a vertically oriented drive shaft. A vacuum chuck with an upwardly facing chamber is circumscribed by an annular surface for receiving in supported relation a silicon chip. An ordered array of low-pressure cavities is defined about the periphery of the body and connected in communication with the chamber via radially extended bores in order to translate low pressures to the chamber as the head is angularly displaced. A pressure differential is thereby established across the chip for securing it to the head.

Lombardi, F.↗
DOE OSTI2025

Readout electronics for low occupancy High-Pressure Gas TPCs

High-Pressure Gas Time Projection Chambers (HPgTPCs) have benefits such as low energy thresholds, magnetisability, and 4π acceptance, making them ideal for neutrino experiments such as DUNE. We present the design of an FPGA-based solution optimised for Gaseous Argon Near Detector (ND-GAr), which is part of the Phase-II more capable near detector for DUNE. These electronics reduce the cost significantly compared to using collider readout electronics, which are typically designed for much higher occupancy and therefore, for example, need much larger numbers of FPGAs and power per channel. We demonstrate the performance of our electronics with the Teststand for an Overpressurised Argon Detector (TOAD) at Fermilab in the US at a range of pressures and gas mixtures up to 4.5 barA, reading out ∼10 000 channels from a Multi-Wire Proportional Chamber (MWPC). The operation took place between April and July of 2024. We measure the noise characteristics of the system to be sufficiently low, and we identify sources of noise that can be further mitigated in the next iteration. We also note that the cooling scheme used in the test requires improvement before full-scale deployment. Despite these necessary improvements, we show that the system can fulfil the needs of a HPgTPC for a fraction of the price of collider readout electronics.

Data acquisition concepts↗
NASA NTRS1990

Optimization of the propulsion cycles for advanced shuttles. II - Performance model methodology

In 1988 computer codes from NASA and DLR were combined to analyze and optimize rocket engines for advance launch vehicles. Further effort has addressed performance propulsion parameters such as chamber pressure, cycle selection (including mixed and tripropellant cycles), nozzle extension, and new technologies. Results of the study indicate that lower dry masses can be reached with the staged combustion cycle than with the gas-generator cycle, using oxygen/hydrogen at relatively high mixture ratio (7:1) and modest chamber pressure (220 bar). For low dry masses the staged combustion cycle is only excelled by the dual-expander cycle using hydrogen and propane as fuels. The emphasis here is to explain the procedure and boundary conditions used for the analysis and the variety of configurations treated, as well as to demonstrate the methods used by presenting some of the major results.

Manski, Detlef↗
NASA NTRS2010

Rocket Engine Altitude Simulation Technologies

John C. Stennis Space Center is embarking on a very ambitious era in its rocket engine propulsion test history. The first new large rocket engine test stand to be built at Stennis Space Center in over 40 years is under construction. The new A3 Test Stand is designed to test very large (294,000 Ibf thrust) cryogenic propellant rocket engines at a simulated altitude of 100,000 feet. A3 Test Stand will have an engine testing chamber where the engine will be fired after the air in the chamber has been evacuated to a pressure at the simulated altitude of less than 0.16 PSIA. This will result in a very unique environment with extremely low pressures inside a very large chamber and ambient pressures outside this chamber. The test chamber is evacuated of air using a 2-stage diffuser / ejector system powered by 5000 lb/sec of steam produced by 27 chemical steam generators. This large amount of power and flow during an engine test will result in a significant acoustic and vibrational environment in and around A3 Test Stand.

Woods, Jody L.↗
NASA NTRS1988

The simulation of the alternate turbopump development high pressure oxygen and fuel turbopumps for the space shuttle main engine using the Shaberth computer program

The Space Shuttle Main Engine (SSME) is basically comprised of a combustion chamber and nozzle, high and low pressure oxygen turbopumps and high and low pressure fuel turbopumps. In the current configuration, the high pressure fuel (HPTFP) and high pressure oxygen turbopumps (HPOTP) have experienced a history of ball bearing wear. The wear problem can be attributed to numerous factors including the hydrodynamic axial and radial loads caused by the flow of liquid oxygen and liquid hydrogen through the turbopump impellers and turbine. Also, friction effects between the rolling elements, races, and cage can create thermally induced bearing geometry changes. To alleviate some of the current configuration problems, an alternate turbopump development (ATD) was proposed. However, the ATD HPOTP and HPTFP are constrained to operate interchangeably with the current turbopumps, thus, the operation conditions must be similar. The ATD configuration features a major change in bearings used to support the integrated shaft, impeller, and turbine system. A single ball and single roller will replace the pump-end and turbine and duplex ball bearings. The Shaft-Bearing-Thermal (SHABERTH) computer code was used to model the ATD HPOTP and ATD HPFTP configurations. A two bearing model was used to simulate the HPOTP and HPFTP bearings and shaft geometry. From SHABERTH, a comparison of bearing reaction loads, frictional heat generation rates, and Hertz contact stresses will be attempted with analysis at the 109 percent and 65 percent power levels.

Mcdonald, Gary H.↗
NASA NTRS2002

Toward Martian agriculture: responses of plants to hypobaria

The recent surge of interest in human missions to Mars has also generated considerable interest in the responses of plants to hypobaria (reduced atmospheric pressure), particularly among those in the advanced life support community. Potential for in situ resource utilization, challenges in meeting engineering constraints for mass and energy, the prospect of using lightweight plant growth structures on Mars, and the minimal literature on plant responses to low pressure all suggest much needed research in this area. However, the limited literature on hypobaria combined with previous findings on plant responses to atmospheric composition and established principles of mass transfer of gases suggest that some plants will be capable of tolerating and growing at pressures below 20 kPa; and for other species, perhaps as low as 5-10 kPa. In addition, normal and perhaps enhanced growth of many plants will likely occur at reduced partial pressures of oxygen (e.g., 5 kPa). Growth of plants at such low and partial pressures indicates the feasibility of cultivating plants in lightweight, transparent "greenhouses" on the surface of Mars or in other extraterrestrial or extreme environment locations. There are numerous, accessible terrestrial analogs for moderately low pressure ranges, but not for very low and extremely low atmospheric pressures. Research pertaining to very low pressures has been historically restricted to the use of vacuum chambers. Future research prospects, approaches, and priorities for plant growth experiments at low pressure are considered and discussed as they apply to prospects for Martian agriculture.

NASA Discipline Life Support Systems↗
NASA NTRS1990

Test plan pressure fed thrust chamber technology

Aerojet is developing the technology for the design of a reliable, low cost, efficient, and lightweight LOX/RP-1 pressure fed engine. This technology program is a direct result of Aerojet's liquid rocket booster (LRB) study and previous NASA studies that identified liquid engines using high bulk density hydrocarbon fuels as very attractive for a space transportation system (STS). Previous large thrust LOX/RP-1 engine development programs were characterized by costly development problems due to combustion instability damage. The combustion stability solution was typically obtained through trial and error methods of minimizing instability damage by degrading engine performance. The approach to this program was to utilize existing and newly developed combustion analysis models and design methodology to create a thrust chamber design with features having the potential of producing reliable and efficient operation. This process resulted in an engine design with a unique high thrust-per-element OFO triplet injector utilizing a low cost modular approach. Cost efficient ablative materials are baselined for the injector face and chamber. Technology demonstration will be accomplished through a hot fire test program using appropriately sized subscale hardware. This subscale testing will provide a data base to supplement the current industry data bank and to anchor and validate the applied analysis models and design methodology. Once anchored and validated, these analysis models and design methodology can be applied with greatly increased confidence to design and characterize a large scale pressure fed LOX/RP-1 thrust chamber. The objective of this test program is to generate a data base that can be used to anchor and validate existing analysis models and design methodologies and to provide early concept demonstration of a low cost, efficient LOX/RP-1 thrust chamber. Test conditions and hardware instrumentation were defined to provide data sufficient to characterize combustion stability, performance, and thermal operation over a wide thrust chamber throttling range.

Dunn, Glenn↗
NASA NTRS1980

Low-thrust chemical propulsion

Results from investigations leading to the definition of low thrust chemical engine concepts are described. From the thrust chamber cooling analyses, regenerative/radiation-cooled LO2/H2 thrust chambers offered the largest thrust and chamber pressure operational envelope primarily due to the superior cooling capability of hydrogen and its low critical pressure. Regenerative/radiation-cooled LO2/CH4 offered the next largest operational envelope. The maximum chamber pressure for film/radiation-cooling was significantly lower than for regenerative/radiation-cooling. As in regeneration-cooling, LO2/H2 thrust chambers achieved the highest maximum chamber pressure, LO2/CH4 film/radiation-cooling was found not feasible and LO2/RP-1 film/radiation-cooling was extremely limited. In the engine cycle/configuration evaluation, the engine cycle matrix was defined through the incorporation of the heat transfer results. Engine cycle limits were established with the fuel-cell power cycle achieving the highest chamber pressure; however, the fuel cell system weights were excessive. The staged combustion cycle achieved the next highest chamber pressure but the preburner operational feasibility was in question.

Shoji, J. M.↗
DOE OSTI2020

Time Projection Chamber (TPC) detectors for nuclear astrophysics studies with gamma beams

Gamma-Beams at the HIS facility in the USA and anticipated at the ELI-NP facility, now constructed in Romania, present unique new opportunities to advance research in nuclear astrophysics; not the least of which is resolving open questions in oxygen formation during stellar helium burning via a precise measurement of the 12C() reaction. Time projection chamber (TPC) detectors operating with low pressure gas (as an active target) are ideally suited for such studies. We look over the progress of the current research program and plans for the future at the HIS facility with the optical readout TPC (O-TPC) and the development of an electronic readout TPC for the ELI-NP facility (ELITPC).

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS2004

Laboratory Apparatus Generates Dual-Species Cold Atomic Beam

A laser cooling apparatus that generates a cold beam of rubidium and cesium atoms at low pressure has been constructed as one of several intermediate products of a continuing program of research on laser cooling and atomic physics. Laser-cooled atomic beams, which can have temperatures as low as a microkelvin, have been used in diverse applications that include measurements of fundamental constants, atomic clocks that realize the international standard unit of time, atom-wave interferometers, and experiments on Bose-Einstein condensation. The present apparatus is a prototype of one being evaluated for use in a proposed microgravitational experiment called the Quantum Interferometric Test of Equivalence (QuITE). In this experiment, interferometric measurements of cesium and rubidium atoms in free fall would be part of a test of Einstein s equivalence principle. The present apparatus and its anticipated successors may also be useful in other experiments, in both microgravity and normal Earth gravity, in which there are requirements for dual-species atomic beams, low temperatures, and low pressures. The apparatus includes a pyramidal magneto-optical trap in which the illumination is provided by multiple lasers tuned to frequencies characteristic of the two atomic species. The inlet to the apparatus is located in a vacuum chamber that contains rubidium and cesium atoms at a low pressure; the beam leaving through the outlet of the apparatus is used to transfer the atoms to a higher-vacuum (lower-pressure) chamber in which measurements are performed. The pyramidal magneto-optical trap is designed so that the laser cooling forces in one direction are unbalanced, resulting in a continuous cold beam of atoms that leak out of the trap (see figure). The radiant intensity (number of atoms per unit time per unit solid angle) of the apparatus is the greatest of any other source of the same type reported to date. In addition, this is the first such apparatus capable of producing a slow, collimated beam that contains two atomic species at the same time.

Thompson, Robert↗
DOE OSTI2023

The MIGDAL experiment: Measuring a rare atomic process to aid the search for dark matter

We present the Migdal In Galactic Dark mAtter expLoration (MIGDAL) experiment aiming at the unambiguous observation and study of the so-called Migdal effect induced by fast-neutron scattering. It is hoped that this elusive atomic process can be exploited to enhance the reach of direct dark matter search experiments to lower masses, but it is still lacking experimental confirmation. Our goal is to detect the predicted atomic electron emission which is thought to accompany nuclear scattering with low, but calculable, probability, by deploying an Optical Time Projection Chamber filled with a low-pressure gas based on CF 4 . Initially, pure CF 4 will be used, and then in mixtures containing other elements employed by leading dark matter search technologies — including noble species, plus Si and Ge. High resolution track images generated by a Gas Electron Multiplier stack, together with timing information from scintillation and ionisation readout, will be used for 3D reconstruction of the characteristic event topology expected for this process — an arrangement of two tracks sharing a common vertex, with one belonging to a Migdal electron and the other to a nuclear recoil. Different energy-loss rate distributions along both tracks will be used as a powerful discrimination tool against background events. In this article we present the design of the experiment, informed by extensive particle and track simulations and detailed estimations of signal and background rates. In pure CF 4 we expect to observe 8.9 (29.3) Migdal events per calendar day of exposure to an intense D–D (D–T) neutron generator beam at the NILE facility located at the Rutherford Appleton Laboratory (UK). With our nominal assumptions, 5σ median discovery significance can be achieved in under one day with either generator.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS1989

Saltation thresholds and entrainment of fine particles at Earth and Martian pressures

An open circuit wind tunnel designed to operate in a large vacuum chamber was built at NASA-Ames to investigate saltation threshold, flux, deflation rates, and other aeolian phenomena on the planet Mars. The vacuum chamber will operate at pressures as low as 4 mbar, and the tunnel operates at windspeeds as high as 150 m/sec. Either air or CO2 can be used as a working fluid. It was found that, to a first order approximation, the same dynamic pressure was required at Martian pressure to entrain or saltate particles as was required on Earth, although wind and particle speed are considerably higher at Martian pressure. A 2nd wind tunnel, designed to operate aboard the NASA KC-135 0-g aircraft to obtain information on the effect of gravity on saltation threshold and the interparticle force at 0-g, is also described and test data presented. Some of the experiments are summarized and various aspects of low pressure aeolian entrainment for particles 12 to 100 micron in diameter are discussed, some of them unique to low pressure testing and some common in Earth pressure particle transport testing. The facility, the modes of operation, and the materials used are described.

Leach, Rodman↗
DOE OSTI2023

Secondary scintillation properties of multi-layer THGEMs operated in low-pressure CF 4 and Ar/5%Xe

We present a measurement of the secondary scintillation yield produced by two-layer Thick Gas Electron Multipliers (M-THGEMs) in pure Tetrafluoromethane (CF 4 ) gas and in Ar mixed with 5% Xe in low-pressures down to 20 Torr. The detector was irradiated with 5.49 MeV alpha particles from a low-rate 241-Am source. The secondary scintillation light generated during the gas avalanche process was read out by a Hamamatsu photomultiplier tube (model R8520-406), sensitive to a broad wavelength range (160–650 nm). The avalanche charge was collected on the bottom electrode of M-THGEM and correlated to the scintillation light on an event-by-event basis. We observed that, for both gas types, the value of the photon to electron production ratio (0.4 ph/el in CF 4 and 0.1 ph/el in Ar/5%Xe) increases with the thickness of the M-THGEM electrodes and varies significantly with the pressure, being higher at lower values. The decrease in electroluminescence yield at higher pressures is much more pronounced in the Ar/Xe mixture. In addition, because of a larger gas avalanche volume, the electroluminescence light yield is larger in thicker M-THGEM structures. Presented results are particularly useful for designing the next generation of Optical-readout Time Projection Chambers (O-TPCs) operated at low-pressure CF 4 ; applications include experimental nuclear physics with rare isotope beams, dark matter detection with directional sensitivity and observation of the Migdal effect in a low-pressure Optical TPC.

47 OTHER INSTRUMENTATION↗
DOE OSTI2021

Numerical investigation of a fueled pre-chamber spark-ignition natural gas engine

Pre-chamber spark-ignition (PCSI) is a leading advanced ignition concept for internal combustion engines with the potential to enable diesel-like efficiency in medium-duty/heavy-duty (MD/HD) natural gas (NG) engines. By leveraging distributed ignition sources from multiple turbulent jets, the PCSI technology can deliver extremely short combustion duration in ultra-lean mixtures and significantly improve the engine thermal efficiency. However, in the automotive industry there is a lack of adequate science base and predictive simulation tools required for commercial development of PCSI engines. Here, in this study, Reynolds-Average Navier-Stokes simulations are carried out to describe the combustion process in lean-burn NG engines, focusing on the combustion modeling approach. Two combustion models, multi-zone well-stirred reactor (MZ-WSR) and G-equation, are used to simulate the combustion process in an MD NG engine equipped with a fueled-PCSI system for four operating conditions close to the lean operating limit. A skeletal chemical mechanism and a laminar flame speed tabulation are used to compute the combustion accurately. Simulation results are compared with experimental data regarding measured cylinder pressure, heat release rate, and combustion duration. By dividing the PCSI combustion process into four distinct phases, the difference between the two models’ results for each phase is analyzed in detail. The MZ-WSR model overestimates the combustion duration for early flame kernel growth in the pre-chamber due to the lack of a specific formulation to take turbulence-chemistry interaction into account. Despite the prolonged combustion duration and low pressure built-up inside the pre-chamber, the model matches the combustion rate in the main-chamber. In contrast, the G-equation model delivers good agreements for the pre-chamber combustion and turbulent jet-driven combustion processes. However, the model starts to underestimate the combustion rate in the main-chamber, especially under ultra-lean mixture conditions. Finally, improvements are needed for both models to simulate the later combustion stage that occurred in the near-wall regions.

33 ADVANCED PROPULSION SYSTEMS↗