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At least 415 records · Page 23

All sources NASA NTRS DOE OSTI

NASA NTRS1982

Testing of fuel/oxidizer-rich, high-pressure preburners

Results of an evaluation of high pressure combustion of fuel rich and oxidizer rich LOX/RP-1 propellants using 4.0 inch diameter prototype preburner injectors and chambers are presented. Testing covered a pressure range from 8.9 to 17.5 MN/square meters (1292 to 2540 psia). Fuel rich mixture ratios ranged from 0.238 to 0.367; oxidizer rich mixture ratios ranged from 27.2 to 47.5. Performance, gas temperature uniformity, and stability data for two fuel rich and two ozidizer rich preburner injectors are presented for a conventional like-on-like (LOL) design and a platelet design injector. Kinetically limited combustion is shown by the excellent agreement of measured fuel rich gas composition and C performance data with kinetic model predictions. The oxidizer rich test results support previous equilibrium combustion predictions.

Lawver, B. R.↗

NASA NTRS1996

Design of the Heat Receiver for the U.S./Russia Solar Dynamic Power Joint Flight Demonstration

A joint U.S./Russia program is being conducted to develop, fabricate, launch, and operate a solar dynamic demonstration system on Space Station Mir. The goal of the program is to demonstrate and confirm that solar dynamic power systems are viable for future space applications such as the International Space Station Alpha The major components of the system include a heat receiver, a closed Brayton cycle power conversion unit, a power conditioning and control unit, a concentrator, a radiator, a thermal control system, and a Space Shuttle Carrier. This paper discusses the design of the heat receiver component. The receiver comprises a cylindrical cavity, the walls of which are lined with a series of tubes running the length of the cavity. The engine working fluid, a mixture of xenon and helium, is heated by the concentrated sunlight incident on these tubes. The receiver incorporates integral thermal storage, using a eutectic mixture of lithium fluoride and calcium difluoride as the thermal storage solid-to-liquid phase change materiaL This thermal storage is required to enable power production during eclipse. The phase change material is contained in a series of individual containment canisters.

Strumpf, Hal J.↗

NASA NTRS2004

Assessing Jet-Induced Spatial Mixing in a Rich, Reacting Crossflow

In many advanced low NOx gas turbine combustion techniques, such as rich-burn/quick-mix/lean-burn (RQL), jet mixing in a reacting, hot, fuel-rich crossflow plays an important role in minimizing all pollutant emissions and maximizing combustion efficiency. Assessing the degree of mixing and predicting jet penetration is critical to the optimization of the jet injection design strategy. Different passive scalar quantities, including carbon, oxygen, and helium are compared to quantify mixing in an atmospheric RQL combustion rig under reacting conditions. The results show that the O2-based jet mixture fraction underpredicts the C-based mixture fraction due to jet dilution and combustion, whereas the He tracer overpredicts it possibly due to differences in density and diffusivity. The He-method also exhibits significant scatter in the mixture fraction data that can most likely be attributed to differences in gas density and turbulent diffusivity. The jet mixture fraction data were used to evaluate planar spatial unmixedness, which showed good agreement for all three scalars. This investigation suggests that, with further technique refinement, O2 or a He tracer could be used instead of C to determine the extent of reaction and mixing in an RQL combustor.

Demayo, T. N.↗

NASA NTRS2010

Granular Flow and Dynamics of Lunar Simulants in Excavating Implements

The exploration of the lunar surface will rely on properly designed excavation equipment for surface preparations and for collection of lunar regolith in In-Situ Resource Utilization (ISRU) processes. Performance efficiency, i.e minimizing loading forces while maximizing material collection, and mass and volume reductions are major design goals. The NASA Glenn Research Center has embarked on an experimental program to determine the flow characteristics and dynamic forces produced by excavation operations using various excavator bucket designs. A new large scale soil bin facility, 2.27 m x 5.94 m x 0.76 m (nominally 8 ft. x 20 ft. x 27 in.) in size, capable of accommodating moderately large test implements was used for the simulations of lunar operations. The soil bin is filled with GRC-3simulant (a mixture of industrial sands and silt with a particle size distribution and the bulk mechanical (shear) strength representative of an average of lunar regolith from different regions) and uses motorized horizontal rails and a vertical actuator to drive the implement through the lunar simulant soil. A six-axis load cell and encoders provide well resolved measurements of the three dimensional forces and torques and motion of the bucket. In addition, simultaneous video allows for the analysis of the flow behavior and structure formation of the regolith during excavation. The data may be useful in anchoring soil mechanic models and to provide engineering data for design consideration.

Agui, Juan H.↗

NASA NTRS1993

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Niiya, Karen E.↗

NASA NTRS1991

Definition of two-phase flow behaviors for spacecraft design

Two-phase flow, thermal management systems are currently being considered as an alternative to conventional, single phase systems for future space missions because of their potential to reduce overall system mass, size, and pumping power requirements. Knowledge of flow regime transitions, heat transfer characteristics, and pressure drop correlations is necessary to design and develop two-phase systems. A boiling and condensing experiment was built in which R-12 was used as the working fluid. A two-phase pump was used to circulate a freon mixture and allow separate measurements of the vapor and liquid flow streams. The experimental package was flown five times aboard the NASA KC-135 aircraft which simulates zero-g conditions by its parabolic flight trajectory. Test conditions included stratified and annual flow regimes in 1-g which became bubbly, slug, or annular flow regimes on 0-g. A portion of this work is the analysis of adiabatic flow regimes. The superficial velocities of liquid and vapor have been obtained from the measured flow rates and are presented along with the observed flow regimes.

Reinarts, Thomas R.↗

NASA NTRS1995

Alternate Propulsion Subsystem Concepts Tripropellant Comparison Study

A study was conducted under MSFC contract NAS8-39210 to compare tripropellant and bipropellant engine configurations for the SSTO mission. The objective was to produce an 'apples-to-apples' comparison to isolate the effects of design implementation, designing company, year of design, or technologies included from the basic tripropellant/bipropellant comparison. Consequently, identical technologies were included (e.g., jet pumps) and the same design groundrules and practices were used. Engine power cycles were examined as were turbomachinery/preburner arrangements for each cycle. The bipropellant approach and two tripropellant approaches were separately optimized in terms of operating parameters: exit pressures, mixture ratios, thrust splits, etc. This briefing presents the results of the study including engine weights for both tripropellant and bipropellant engines; dry vehicle weight performance for a range of engine chamber pressures; discusses the basis for the results; examines vehicle performance due to engine cycles and the margin characteristics of various cycles; and identifies technologies with significant payoffs for this application.

Levack, Daniel↗

NASA NTRS2015

The THS Experiment: Simulating Titans Atmospheric Chemistry at Low Temperature (200K)

In Titan's atmosphere, composed mainly of N2 (95-98%) and CH4 (2-5%), a complex chemistry occurs at low temperature, and leads to the production of heavy organic molecules and subsequently solid aerosols. Here, we used the Titan Haze Simulation (THS) experiment, an experimental setup developed at the NASA Ames COSmIC simulation facility to study Titan's atmospheric chemistry at low temperature. In the THS, the chemistry is simulated by plasma in the stream of a supersonic expansion. With this unique design, the gas is cooled to Titan-like temperature ( approximately 150K) before inducing the chemistry by plasma, and remains at low temperature in the plasma discharge (approximately 200K). Different N2-CH4-based gas mixtures can be injected in the plasma, with or without the addition of heavier precursors present as trace elements on Titan, in order to monitor the evolution of the chemical growth. Both the gas- and solid phase products resulting from the plasma-induced chemistry can be monitored and analyzed using a combination of complementary in situ and ex situ diagnostics. A recent mass spectrometry[1] study of the gas phase has demonstrated that the THS is a unique tool to probe the first and intermediate steps of Titan's atmospheric chemistry at Titan-like temperature. In particular, the mass spectra obtained in a N2-CH4-C2H2-C6H6 mixture are relevant for comparison to Cassini's CAPS-IBS instrument. The results of a complementary study of the solid phase are consistent with the chemical growth evolution observed in the gas phase. Grains and aggregates form in the gas phase and can be jet deposited on various substrates for ex situ analysis. Scanning Electron Microscopy images show that more complex mixtures produce larger aggregates. A mass spectrometry analysis of the solid phase has detected the presence of aminoacetonitrile, a precursor of glycine, in the THS aerosols. X-ray Absorption Near Edge Structure (XANES) measurements also show the presence of imine and nitrile functional groups, showing evidence of nitrogen chemistry. These complementary studies show the high potential of THS to better understand Titan's chemistry and the origin of aerosol formation.

Laboratory↗

NASA NTRS2015

Pyrotechnically Actuated Gas Generator Utilizing Aqueous Methanol

A gas-generating device was developed to supplement the ram-air inflation of a supersonic ballute. The device is designed to initially pressurize the ballute following deployment, exposing and orienting its ram-air inlets to free-stream air for complete inflation. The supplemental pressurization decreases the total inflation time, and increases the likelihood of a successful inflation. The device contains a reservoir filled with an aqueous mixture of methanol that, when released in to the interior of the ballute, rapidly vaporizes due to the low ambient pressure. Upon activation of the device, a pair of redundant ring mechanisms initiate pyrotechnic charges that pressurize and rupture the reservoir, resulting in ejection of the methanol in to the ballute. In addition to its role in inflation, the device serves as the structural connection to the ballute. Analytical models were developed for the inflation capability of the device, which were verified using vacuum chamber testing of developmental hardware. Static, deployment, and environmental testing demonstrated the functionality of the ring mechanism and reservoir under several temperature and pressure conditions. Finally, the device was successfully operated during the first Supersonic Flight Dynamics Test (SFDT) of NASA's Low Density Supersonic Decelerator (LDSD) project. The design architecture is scalable to accommodate different quantities of gas generation, can be adjusted to operate in a variety of temperature and atmospheric pressure regimes, and provides a robust device that may be installed with minimal risk to personnel or hardware.

IA↗

NASA NTRS1956

Turbojet-engine Starting and Acceleration

From considerations of safety and reliability in performance of gas-turbine aircraft, it is clear that engine starting and acceleration are of utmost importance. For this reason extensive efforts have been devoted to the investigation of the factors involved in the starting and acceleration of engines. In chapter III it is shown that certain basic combustion requirements must be met before ignition can occur; consequently, the design and operation of an engine must be tailored to provide these basic requirements in the combustion zone of the engine, particularly in the vicinity of the ignition source. It is pointed out in chapter III that ignition by electrical discharges is aided by high pressure, high temperature, low gas velocity and turbulence, gaseous fuel-air mixture, proper mixture strength, and-an optimum spark. duration. The simultaneous achievement of all these requirements in an actual turbojet-engine combustor is obviously impossible, yet any attempt to satisfy as many requirements as possible will result in lower ignition energies, lower-weight ignition systems, and greater reliability. These factors together with size and cost considerations determine the acceptability of the final ignition system. It is further shown in chapter III that the problem of wall quenching affects engine starting. For example, the dimensions of the volume to be burned must be larger than the quenching distance at the lowest pressure and the most adverse fuel-air ratio encountered. This fact affects the design of cross-fire tubes between adjacent combustion chambers in a tubular-combustor turbojet engine. Only two chambers in these engines contain spark plugs; therefore, the flame must propagate through small connecting tubes between the chambers. The quenching studies indicate that if the cross-fire tubes are too narrow the flame will not propagate from one chamber to another. In order to better understand the role of the basic factors in actual engine operation, many investigations have been conducted in single combustors from gas-turbine engines and in full-scale engines in altitude tanks and in flight. The purpose of the present chapter is to discuss the results of such studies and, where possible, to interpret these results qualitatively in terms of the basic requirements reported in chapter III. The discussion parallels the three phases of turbojet engine starting: (1) Ignition of the fuel-air mixture (2) Propagation of flame throughout the combustion zone (3) Acceleration of the engine to operating speed.

Mc Cafferty, R. J.↗

NASA NTRS1988

Design of a ram accelerator mass launch system

The ram accelerator mass launch system has been proposed to greatly reduce the costs of placing acceleration-insensitive payloads into low earth orbit. The ram accelerator is a chemically propelled, impulsive mass launch system capable of efficiently accelerating relatively large masses from velocities of 0.7 km/sec to 10 km/sec. The principles of propulsion are based on those of a conventional supersonic air-breathing ramjet; however the device operates in a somewhat different manner. The payload carrying vehicle resembles the center-body of the ramjet and accelerates through a stationary tube which acts as the outer cowling. The tube is filled with premixed gaseous fuel and oxidizer mixtures that burn in the vicinity of the vehicle's base, producing a thrust which accelerates the vehicle through the tube. This study examines the requirement for placing a 2000 kg vehicle into a 500 km circular orbit with a minimum amount of on-board rocket propellant for orbital maneuvers. The goal is to achieve a 50 pct payload mass fraction. The proposed design requirements have several self-imposed constraints that define the vehicle and tube configurations. Structural considerations on the vehicle and tube wall dictate an upper acceleration limit of 1000 g's and a tube inside diameter of 1.0 m. In-tube propulsive requirements and vehicle structural constraints result in a vehicle diameter of 0.76 m, a total length of 7.5 m and a nose-cone half angle of 7 degrees. An ablating nose-cone constructed from carbon-carbon composite serves as the thermal protection mechanism for atmospheric transit.

Aarnio, Michael↗

NASA NTRS1979

Description and orbit data of variable-conductance heat-pipe system for the communications technology satellite

A variable-conductance heat-pipe system (VCHPS) with methanol as the working fluid and a nitrogen and helium mixture as the control gas was used for the thermal control of a 200 W RF traveling wave tube of the Communication Technology Satellite. Three stainless steel heat pipes (one redundant) and an aluminum radiator were designed to transfer 196 watts for an evaporator temperature of 50 C. The system has operated for three years with no noticeable change in performance. On four occasions the heat pipes apparently deprimed. A short time after reducing the tube power, the heat pipes reprimed and the system continued to operate normally. The description, qualification testing, and orbit data of the VCHPS are presented.

Gedeon, L.↗

NASA NTRS1983

Observations of odd-half cyclotron harmonic emissions in a shell-Maxwellian laboratory plasma

During the last 14 years, the subject of odd-half electron cyclotron harmonic emission from plasma has been given considerable attention. It was apparently first reported to occur in a beam plasma machine. The existence of the emission in space has been well documented through observations made with satellites. Because wavelengths are difficult to observe in space, no wave number spectrum has ever been obtained for the half-odd harmonic emission. Such a spectrum together with a frequency spectrum might provide the basis for a successful modeling of the instability. The present investigation is concerned with the design of a laboratory experiment in which a plasma with an anisotropic velocity distribution is produced and measured. In addition, the dispersion relation is directly measured, and a noise analysis is conducted. The obtained plasma, a mixture of shell and Maxwellian distributions, is found to emit waves within the Bernstein wave branches. By correlation measurements, a mode is found which is essentially an absolute instability in a narrow frequency band.

Urrutia, J. M.↗

NASA NTRS1983

Electrophoretic purification of cells in space - Evaluation of results from STS-3

The procedure and results of Electrophoresis Equipment Verification Test, designed to examine electrophoretic behavior of animal cells is suspension more concentrated than possible on earth and flown on the Shuttle flight STS-3, were discussed. Ground-based laboratory values of electrophoretic mobilities of a mixture of human and rabbit aldehyde-fixed red blood cells (RBC) were compared with those recorded at 11 minute intervals on the Shuttle STS-3. RBC migration and separation observed through photographic records were not as expected. However, cell mobilities and migrating band profiles were consistent with the results of laboratory simulation experiments. It was concluded that zero G electrophoresis of very high concentrations (1 x 10 to the 9th) is possible and similar to electrophoresis of normal cell concentrations on earth.

Sarnoff, B. E.↗

NASA NTRS2000

Low Complexity Turbo-like Codes

This paper addresses the iterative devoding analysis, design, and performance of low complexity turbo like codes.

turbo-like codes iterative decoding analysis conca↗

NASA NTRS2015

Electrical Pressurization Concept for the Orion MPCV European Service Module Propulsion System

The paper presents the design of the pressurization system of the European Service Module (ESM) of the Orion Multi-Purpose Crew Vehicle (MPCV). Being part of the propulsion subsystem, an electrical pressurization concept is implemented to condition propellants according to the engine needs via a bang-bang regulation system. Separate pressurization for the oxidizer and the fuel tank permits mixture ratio adjustments and prevents vapor mixing of the two hypergolic propellants during nominal operation. In case of loss of pressurization capability of a single side, the system can be converted into a common pressurization system. The regulation concept is based on evaluation of a set of tank pressure sensors and according activation of regulation valves, based on a single-failure tolerant weighting of three pressure signals. While regulation is performed on ESM level, commanding of regulation parameters as well as failure detection, isolation and recovery is performed from within the Crew Module, developed by Lockheed Martin Space System Company. The overall design and development maturity presented is post Preliminary Design Review (PDR) and reflects the current status of the MPCV ESM pressurization system.

Helium↗

NASA NTRS1983

Improved 02/H2 Gas Mixture Sensor

Monitor of mixture concentrations uses catalyzed and uncatalyzed temperature probe. Sensor includes Pt-catalyzed temperature probe mounted in line with similar uncatalyzed temperature probe. Use of common temperature probes and standard, flareless, high-pressure tubefittings resulted in design conductive to installation in almost any system. Suitable for use in regenerative fuel cells, life-support systems, and other closed systems.

Moulthrop, L. C.↗

NASA NTRS1984

Experiments on water/melt explosions, nature of products, and models of dispersal

Experiments were carried out in a steel pressure device using controlled amounts of water and thermite melt to examine the mechanical energy released on explosive mixing following the initial contact of the two materials. An experimental design was used to allow the direct calculation of the mechanical energy by the dynamic lift of the device as recorded both optically and physically. A large number of experiments were run to accurately determine the optimum mixture of water and melt for the conversion of thermal to mechanical energy. The maximum efficiency observed was about 12% at a water/thermite mass ratio of 0.50. These experiments are the basis for the development of models of hydroexplosions and melt fragmentation. Particles collected from the experimental products are similar in size and shape to pyroclasts produced by much larger hydrovolcanic explosions. Melt rupture at optimum ratios produces very fine particles whereas rupture at high or low water/melt ratios produces large melt fragments. Grain surface textures in the experimental products are also related to the water/melt ratio and the mechanism of explosive mixing. It is thus possible to have qualitative information about the nature of the explosion from the sizes and shapes of the fragments produced.

Sheridan, M. F.↗