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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Investigation of throat film coolant for advanced LOX/RP-1 thrust chambers

The use of throat film coolant as a method to reduce peak combustion heat flux was investigated. Hot-fire tests were performed using a subscale LOX/RP-1 combustor at a nominal chamber pressure of 2000 psia. A special injection manifold was fabricated to introduce the film coolant into the chamber. Four tests were conducted with RP-1 film coolant injected 4.4 inches upstream of the throat. In order to determine the effect of chamber length on the film coolant, four additional tests were performed with the injection point moved 16.4 inches upstream of the throat. Results from the test program are discussed. Methods for incorporating throat film cooling into a flightweight combustion chamber are also discussed.

Volkmann, Jon C.↗
NASA NTRS1991

Factors controlling the structures of magma chambers in basaltic volcanoes

The depths, vertical extents, and lateral extents of magma chambers and their formation are discussed. The depth to the center of a magma chamber is most probably determined by the density structure of the lithosphere; this process is explained. It is commonly assumed that magma chambers grow until the stress on the roof, floor, and side-wall boundaries exceed the strength of the wall rocks. Attempts to grow further lead to dike propagation events which reduce the stresses below the critical values of rock failure. The tensile or compressive failure of the walls is discussed with respect to magma migration. The later growth of magma chambers is accomplished by lateral dike injection into the country rocks. The factors controlling the patterns of growth and cooling of such dikes are briefly mentioned.

Wilson, L.↗
NASA NTRS1992

Hot fire fatigue testing results for the compliant combustion chamber

A hydrogen-oxygen subscale rocket combustion chamber was designed incorporating an advanced design concept to reduce strain and increase life. The design permits unrestrained thermal expansion of a circumferential direction and, thereby, provides structural compliance during the thermal cycling of hot-fire testing. The chamber was built and test fired at a chamber pressure of 4137 kN/sq m (600 psia) and a hydrogen-oxygen mixture ratio of 6.0. Compared with a conventional milled-channel configuration, the new structurally compliant chamber had a 134 or 287 percent increase in fatigue life, depending on the life predicted for the conventional configuration.

Pavli, Albert J.↗
NASA NTRS1992

Engineering verification of the biomass production chamber

The requirements for life support systems, both biological and physical-chemical, for long-term human attended space missions are under serious study throughout NASA. The KSC 'breadboard' project has focused on biomass production using higher plants for atmospheric regeneration and food production in a special biomass production chamber. This chamber is designed to provide information on food crop growth rate, contaminants in the chamber that alter plant growth requirements for atmospheric regeneration, carbon dioxide consumption, oxygen production, and water utilization. The shape and size, mass, and energy requirements in relation to the overall integrity of the biomass production chamber are under constant study.

Prince, R. P.↗
NASA NTRS1994

Iridium-Coated Rhenium Combustion Chamber

Iridium-coated rhenium combustion chamber withstands operating temperatures up to 2,200 degrees C. Chamber designed to replace older silicide-coated combustion chamber in small rocket engine. Modified versions of newer chamber could be designed for use on Earth in gas turbines, ramjets, and scramjets.

Schneider, Steven J.↗
NASA NTRS1994

In situ detection of tropospheric OH and HO2 by laser-induced fluorescence in a detection chamber at low pressure

Just as in the method of Hard and O'Brien, ambient air is pulled through an approximately 1 mm diameter inlet into a detection chamber that is maintained at a pressure of 2.4 torr. The ambient air stream travels through the detection chamber with a velocity of greater than 100 m sec(exp -1) in a narrow stream, constrained by the addition of an inert gas flow (0.4 torr). The OH molecule is both excited and detected in the A(exp 2)Sigma (v' = 0) yields X(exp 2)II (v'' = 0) transition at 308 nm. Light from a copper vapor-pumped dye laser (rep. rate = 10kHz; pulse length = 20 ns; linewidth = .1 cm(exp -1), and average power = 15 mW), resonant with the Q(sub 1)(3) transition, is multipassed through a White cell and intercepts the air stream as 24 non-overlapping 2mm by 5mm beams. A fast microchannel plate detector is turned off during the laser pulse to prevent saturation of the detector due to Rayleigh and chamber scattering. It is turned on 30-100 nsec after the end of the laser pulse for 300 ns to collect resonance fluorescence from OH. HO2 is detected by chemical conversion to OH by reaction with reagent NO, followed by OH detection. Both the detection sensitivity and the inlet characteristics must be understood for any in situ instrument. For the calibration of the detection sensitivity, OH is produced quantitatively by the fast H + NO2 yields OH + NO reaction in a low pressure, flowing discharge tube connected to the detection chamber. The inlet transmission of OH inlet is calibrated separately.

Brune, William H.↗
NASA NTRS1994

A hydrogen-oxygen rocket engine coolant passage design program (RECOP) for fluid-cooled thrust chambers and nozzles

The design of coolant passages in regeneratively cooled thrust chambers is critical to the operation and safety of a rocket engine system. Designing a coolant passage is a complex thermal and hydraulic problem requiring an accurate understanding of the heat transfer between the combustion gas and the coolant. Every major rocket engine company has invested in the development of thrust chamber computer design and analysis tools; two examples are Rocketdyne's REGEN code and Aerojet's ELES program. In an effort to augment current design capabilities for government and industry, the NASA Lewis Research Center is developing a computer model to design coolant passages for advanced regeneratively cooled thrust chambers. The RECOP code incorporates state-of-the-art correlations, numerical techniques and design methods, certainly minimum requirements for generating optimum designs of future space chemical engines. A preliminary version of the RECOP model was recently completed and code validation work is in progress. This paper introduces major features of RECOP and compares the analysis to design points for the first test case engine; the Pratt & Whitney RL10A-3-3A thrust chamber.

Tomsik, Thomas M.↗
NASA NTRS1993

CFD Analysis of Spray Combustion and Radiation in OMV Thrust Chamber

The Variable Thrust Engine (VTE), developed by TRW, for the Orbit Maneuvering Vehicle (OMV) uses a hypergolic propellant combination of Monomethyl Hydrazine (MMH) and Nitrogen Tetroxide (NTO) as fuel and oxidizer, respectively. The propellants are pressure fed into the combustion chamber through a single pintle injection element. The performance of this engine is dependent on the pintle geometry and a number of complex physical phenomena and their mutual interactions. The most important among these are (1) atomization of the liquid jets into fine droplets; (2) the motion of these droplets in the gas field; (3) vaporization of the droplets (4) turbulent mixing of the fuel and oxidizer; and (5) hypergolic reaction between MMH and NTO. Each of the above phenomena by itself poses a considerable challenge to the technical community. In a reactive flow field of the kind occurring inside the VTE, the mutual interactions between these physical processes tend to further complicate the analysis. The objective of this work is to develop a comprehensive mathematical modeling methodology to analyze the flow field within the VTE. Using this model, the effect of flow parameters on various physical processes such as atomization, spray dynamics, combustion, and radiation is studied. This information can then be used to optimize design parameters and thus improve the performance of the engine. The REFLEQS CFD Code is used for solving the fluid dynamic equations. The spray dynamics is modeled using the Eulerian-Lagrangian approach. The discrete ordinate method with 12 ordinate directions is used to predict the radiative heat transfer in the OMV combustion chamber, nozzle, and the heat shield. The hypergolic reaction between MMH and NTO is predicted using an equilibrium chemistry model with 13 species. The results indicate that mixing and combustion is very sensitive to the droplet size. Smaller droplets evaporate faster than bigger droplets, leading to a well mixed zone in the combustion chamber. The radiative heat flux at combustion chamber and nozzle walls are an order of negligible less than the conductive heat flux. Simulations performed with the heat shield show that a negligible amount of fluid is entrained into the heat shield region. However, the heat shield is shown to be effective in protecting the OMV structure surrounding the engine from the radiated heat.

Giridharan, M. G.↗
NASA NTRS1997

Development of Large Area Emulsion Chamber Methods with a Super Conducting Magnet for Observation of Cosmic Ray Nuclei from 1 GeV to 1,000 TeV (Emulsion Techniques)

The research developed the fundamental techniques of the emulsion chamber methods that permit measurements of the composition and energy spectra of cosmic rays at energies ranging from 1 GeV/n to over 1,000 TeV/n. The research program consisted of exploring new principles and techniques in measuring very high energy cosmic nuclei with large-area emulsion chambers for high statistics experiments. These tasks have been accomplished and their use was essential in successful analysis of the balloon-borne emulsion chamber experiments up to 10(exp 14) eV. It also provided the fundamental technologies for designing large-area detectors that are aimed at measuring the composition at above 1015 eV region. The latter is now partially succeeded by a NASA Mission Concept, Advanced Cosmic Composition Experiments on the Space Station (ACCESS). The cosmic ray group at the University of Alabama in Huntsville has performed technological R & D as well as contributing to the Japanese-American-Emulsion-Chamber-Experiments (JACEE) Collaboration with the regular data analysis. While primary research support for other institutions' efforts in the JACEE experiments came from NSF and DOE, primary support for the University of Alabama in Huntsville was this contract. Supplemental tasks to standardize the data base and hardware upgrades (automatized microscope) had this institutions cooperation. Investigation of new techniques in this program consisted of development of a fast calorimetry, magnetic/scattering selection of high momentum tracks for a pairmeter, and high statistics momentum measurements for low energy nuclei (E < 1 TeV/n). The highest energy calorimetry and a pairmeter have been considered as strawman instruments by the GOAL (Galactic Origin and Acceleration Limit) proposal of the NASA Cosmic Ray Working Group for long- duration balloon flights. We accomplished the objectives of the GOAL program with three circumpolar, Antarctic JACEE balloon flights during 1992 - 1994.

Takahashi, Yoshiyuki↗
NASA NTRS2000

A Regeneratively-Cooled Thrust Chamber for the Fastrac Engine

This document consists of presentation slides about the development of the regeneratively cooled thrust chamber for the Fastrac engine. The Fastrac engine was originally developed to demonstrate low cost design and fabrication methods. It was intended to be used in an expendable booster. The regen thrust chamber enables a more cost efficient test program. Using the low cost design and fabrication methodology designed for the 12K regeneratively cooled chamber, the contractor designed, developed and fabricated a regeneratively cooled thrust chamber for the Fastrac engine.

Brown, Kendall↗
NASA NTRS2000

Development of an Integrated Rocket Thrust Chamber Assembly Analysis and Design Toolkit (TCAT)

There are many individual computational tools and methods used to analyze and design components of a rocket engine thrust chamber assembly (TCA) 11,21. To analyze and design a thrust chamber assembly these tools are usually used in sequence while communication of information between these tools is usually done manually. Each component of a TCA is usually designed and optimized individually with limited considerations on system optimized design. This approach is often iterative and can be prone to error. Also at present, expert knowledge of each tool is required. The objective of this software development effort is to select, integrate and automate the best tools into an easy-to-use computational package, the Thrust Chamber Analysis Toolkit (TCAT). This tool will provide a seamless process to analyze and design all or individual components of a thrust chamber assembly. This paper describes the current status of the MAT software development. The selected component codes and their capabilities are briefly described. User and module interfaces and integration status are presented. MAT pre- and post-processing added options are presented and overall MAT capabilities are demonstrated with the results of sample cases. Ongoing and planned model validation efforts are also described. Finally, the MAT future plan options are discussed.

Farhangi, S.↗
NASA NTRS1948

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine: Combustion-Chamber Characterisitcs - 5

An investigation to determine the performance and operational characteristics of an axial-flow gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet ram-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and corrected horsepower. For the range of corrected engine speeds investigated, overall total-pressure-loss ratio, cycle efficiency, and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. For the range of corrected horsepowers investigated, the total-pressure-loss ratio and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horsepowers investigated at all corrected engine speeds.

Geisenheyner, Robert M.↗
NASA NTRS2003

Vacuum Plasma Spray Forming of Copper Alloy Liners for Regeneratively Cooled Liquid Rocket Combustion Chambers

Vacuum plasma spray (VPS) has been demonstrated as a method to form combustion chambers from copper alloys NARloy-Z and GRCop-84. Vacuum plasma spray forming is of particular interest in the forming of CuCrNb alloys such as GRCop-84, developed by NASA s Glenn Research Center, because the alloy cannot be formed using conventional casting and forging methods. This limitation is related to the levels of chromium and niobium in the alloy, which exceed the solubility limit in copper. Until recently, the only forming process that maintained the required microstructure of CrNb intermetallics was powder metallurgy formation of a billet from powder stock, followed by extrusion. This severely limits its usefulness in structural applications, particularly the complex shapes required for combustion chamber liners. This paper discusses the techniques used to form combustion chambers from CuCrNb and NARloy-Z, which will be used in regeneratively cooled liquid rocket combustion chambers.

Zimmerman, Frank↗
NASA NTRS1947

Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine: Combustion-Chamber Characteristics - V

An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower. For the range of corrected engine speeds investigated, over-all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or ram-pressure ratio. For the range of corrected horse-powers investigated, the total-pressure-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horse-powers investigated at all corrected engine speeds.

Gensenheyner, Robert M.↗
NASA NTRS2005

NSTAR Extended Life Test Discharge Chamber Flake Analysis

The Extended Life Test (ELT) of the NASA Solar Electric Propulsion Technology Readiness (NSTAR) ion thruster was concluded after 30,352 hours of operation. The ELT was conducted using the Deep Space 1 (DS1) back-up flight engine, a 30 cm diameter xenon ion thruster. Post-test inspection of the ELT engine revealed numerous contaminant flakes distributed over the bottom of the cylindrical section of the anode within the discharge chamber (DC). Extensive analyses were conducted to determine the source of the particles, which is critical to the understanding of degradation mechanisms of long life ion thruster operation. Analyses included: optical microscopy (OM) and particle length histograms, field emission scanning electron microscopy (FESEM) combined with energy dispersive spectroscopy (EDS), and atomic oxygen plasma exposure tests. Analyses of the particles indicate that the majority of the DC flakes consist of a layered structure, typically with either two or three layers. The flakes comprising two layers were typically found to have a molybdenum-rich (Mo-rich) layer on one side and a carbon-rich (C-rich) layer on the other side. The flakes comprising three layers were found to be sandwich-like structures with Mo-rich exterior layers and a C-rich interior layer. The presence of the C-rich layers indicates that these particles were produced by sputter deposition build-up on a surface external to the discharge chamber from ion sputter erosion of the graphite target in the test chamber. This contaminant layer became thick enough that particles spalled off, and then were electro-statically attracted into the ion thruster interior, where they were coated with Mo from internal sputter erosion of the screen grid and cathode components. Atomic oxygen tests provided evidence that the DC chamber flakes are composed of a significant fraction of carbon. Particle size histograms further indicated that the source of the particles was spalling of carbon flakes from downstream surfaces. Analyses of flakes taken from the downstream surface of the accelerator grid provided additional supportive information. The production of the downstream carbon flakes, and hence the potential problems associated with the flake particles in the ELT ion thruster engine is a facility induced effect and would not occur in the space environment.

deGroh, Kim K.↗
NASA NTRS2005

Computational Model Tracking Primary Electrons, Secondary Electrons, and Ions in the Discharge Chamber of an Ion Engine

Computational modeling of the plasma located in the discharge chamber of an ion engine is an important activity so that the development and design of the next generation of ion engines may be enhanced. In this work a computational tool called XOOPIC is used to model the primary electrons, secondary electrons, and ions inside the discharge chamber. The details of this computational tool are discussed in this paper. Preliminary results from XOOPIC are presented. The results presented include particle number density distributions for the primary electrons, the secondary electrons, and the ions. In addition the total number of a particular particle in the discharge chamber as a function of time, electric potential maps and magnetic field maps are presented. A primary electron number density plot from PRIMA is given in this paper so that the results of XOOPIC can be compared to it. PRIMA is a computer code that the present investigators have used in much of their previous work that provides results that compare well to experimental results. PRIMA only models the primary electrons in the discharge chamber. Modeling ions and secondary electrons, as well as the primary electrons, will greatly increase our ability to predict different characteristics of the plasma discharge used in an ion engine.

Mahalingam, Sudhakar↗
NASA NTRS2005

Multiphysics Nuclear Thermal Rocket Thrust Chamber Analysis

The objective of this effort is t o develop an efficient and accurate thermo-fluid computational methodology to predict environments for hypothetical thrust chamber design and analysis. The current task scope is to perform multidimensional, multiphysics analysis of thrust performance and heat transfer analysis for a hypothetical solid-core, nuclear thermal engine including thrust chamber and nozzle. The multiphysics aspects of the model include: real fluid dynamics, chemical reactivity, turbulent flow, and conjugate heat transfer. The model will be designed to identify thermal, fluid, and hydrogen environments in all flow paths and materials. This model would then be used to perform non- nuclear reproduction of the flow element failures demonstrated in the Rover/NERVA testing, investigate performance of specific configurations and assess potential issues and enhancements. A two-pronged approach will be employed in this effort: a detailed analysis of a multi-channel, flow-element, and global modeling of the entire thrust chamber assembly with a porosity modeling technique. It is expected that the detailed analysis of a single flow element would provide detailed fluid, thermal, and hydrogen environments for stress analysis, while the global thrust chamber assembly analysis would promote understanding of the effects of hydrogen dissociation and heat transfer on thrust performance. These modeling activities will be validated as much as possible by testing performed by other related efforts.

Wang, Ten-See↗
NASA NTRS1996

Cyclically controlled welding purge chamber

An arrangement for butt-welding cylindrical sections of large, thin-wall tanks includes a rotatable mandrel with side-by-side sets of radial position adjusters. Each set of adjusters bears on one of the tank sections adjacent the seam, to prevent the sections from sagging out-of-round. The mandrel rotates relative to the welder, so that a continuous seam is formed. A purge chamber is fixed in position behind the seam at the weld head, and is flushed with inert gas. The purge chamber includes a two-sided structure which is contiguous with the cylindrical sections and a circumferential vane to form an open-ended tube-like structure, through which the radial position adjusters pass as the mandrel and cylindrical workpiece sections rotate. The tube-like structure is formed into a chamber by a plurality of movable gates which are controlled to maintain a seal while allowing adjusters to progress through the purge chamber.

Gallagher, Robert L.↗