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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Development of boron epoxy rocket motor chambers.

A 71 cm diameter 74 cm length boron/epoxy composite rocket motor chamber was designed based on the geometric configuration of the JPL Applications Technology Satellite titanium alloy apogee motor chamber. Because analyses showed large stress concentrations in the domes, the configuration was modified using the same basic constraints for openings and attachments. The rocket motor chamber was then fabricated by filament winding with boron/epoxy tape and hydrostatically tested to failure at 264 N/sq cm, 57.2 N/sq cm above the design value. Two more rocket motor chambers were fabricated with the same basic constraints, but shortened to 57.6 cm for a smaller propellant load. The first of these short chambers failed in proof because of filament winding fabrication difficulties. The second chamber was successfully fabricated and passed the hydrostatic proof test.

Jensen, W. M.↗
NASA NTRS1990

Cooling of rocket thrust chambers with liquid oxygen

Rocket engines using high pressure liquid oxygen (LOX) and kerosene (RP-1) as the propellants have been considered for future launch vehicle propulsion. Generaly, in regeneratively cooled engines, thefuel is used to cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability at high temperatures and pressures. Therefore, LOX is being considered as an alternative coolant. However, there has been concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot-gas side wall. To address this concern, an investigation was previously conducted with simulated fatigue cracks upstream of the thrust chamber throat. When these chambers were tested, an unexpected melting in the throat region developed which was not in line with the simulated fatigue cracks. The current experimental program was conducted in order to determine the cause for the failure in the earlier thrust chambers and to further investigate the effects of cracks in the thrust chamber liner upstream of the throat. The thrust chambers were tested at oxygen-to-fuel mixture ratios from 1.5 to 2.86 at a nominal chamber pressure of 8.6 MPa. As a result of the test series, the reason for the failure occurring in the earlier work was determined to be injector anomalies. The LOX leaking through the simulated fatigue cracks did not affect the integrity of the chambers.

Armstrong, Elizabeth S.↗
NASA NTRS1990

Cooling of rocket thrust chambers with liquid oxygen

Rocket engines using high pressure liquid oxygen (LOX) and kerosene (RP-1) as the propellants have been considered for future launch vehicle propulsion. Generally, in regeneratively cooled engines, the fuel is used to cool the combustion chamber. However, hydrocarbons such as RP-1 are limited in their cooling capability at high temperatures and pressures. Therefore, LOX is being considered as an alternative coolant. However, there has been concern as to the effect on the integrity of the chamber liner if oxygen leaks into the combustion zone through fatigue cracks that may develop between the cooling passages and the hot-gas side wall. To address this concern, an investigation was previously conducted with simulated fatigue cracks upstream of the thrust chamber throat. When these chambers were tested, an unexpected melting in the throat region developed which was not in line with the simulated fatigue cracks. The current experimental program was conducted in order to determine the cause for the failure in the earlier thrust chambers and to further investigate the effects of cracks in the thrust chamber liner upstream of the throat. The thrust chambers were tested at oxygen-to-fuel mixture ratios from 1.5 to 2.86 at a nominal chamber pressure of 8.6 MPa. As a result of the test series, the reason for the failure occurring in the earlier work was determined to be injector anomalies. The LOX leaking through the simulated fatigue cracks did not affect the integrity of the chambers.

Armstrong, Elizabeth S.↗
NASA NTRS1993

Heat transfer in rocket combustion chambers

Complexities of liquid rocket engine heat transfer which involve the injector faceplate and film cooled walls are being investigated by computational analysis. A conjugate heat transfer analysis was used to describe localized heating phenomena associated with particular injector configurations and film coolant flows. These components were analyzed, and the analyses verified when appropriate test data were available. The component analyses are being synthesized into an overall flowfield/heat transfer model. A Navier-Stokes flow solver, the FDNS code, was used to make the analyses. Particular attention was given to the representation of the thermodynamic properties of the fluid streams. Unit flow models of specific coaxial injector elements have been developed and are being used to describe the flame structure near the injector faceplate.

Anderson, P.↗
NASA NTRS1972

Fabrication of boron/epoxy rocket motor chambers.

Three 71-cm-diameter boron/epoxy composite rocket motor chambers were filament wound. The first chamber was hydrostatically tested; the second failed in proof test as a result of fabrication deficiencies. The third chamber obtained a successful proof test, but deformed later in processing because of hydrolysis. In spite of the deformation, this chamber was test fired and then burst tested.

Jensen, W. M.↗
NASA NTRS2010

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

Hulka, James R.↗
NASA NTRS1990

Non-grey radiation in a liquid rocket combustion chamber

The effects of radiation on droplets gasification in liquid rocket combustion chambers has been studied. The modeling includes a Legendre pseudo-spectral collocation approximation to solve the Radiative Transfer Equation (RTE). The band model has been utilized to account for non-grey emitting and absorbing gases present in the comustion chamber. The GEMCHIP II code has been utilized to study the fuel and oxidizer droplets combustion and interaction. The submodels within this code are capable of accounting for group combustion and conjugate effects between many droplets. The radiative model has been coupled with the GEMCHIP II code accounting for radial effects only, to provide the tool for studying the combustion-radiation coupling effects in a bipropellant system. The results indicate that the gasification/combustion process will be enhanced upstream of the chamber causing thicker flame sheet and associated higher combustion efficiency.

Kehtarnavaz, H.↗
NASA NTRS1992

Formed platelets for low cost regeneratively cooled rocket combustion chambers

Formed platelet technology can be used to fabricate LO2/LH2-fueled rocket propulsion chambers with higher heat capacity, higher cycle life, and lower pressure drops, in conjunction with lower costs due to the application of high volume production methods. The formed-platelet combustor liner is an assembly of identical platelets, each of which is composed of diffusion-bonded, photoetched laminae with coolant flow passages; the platelets are also joined by diffusion bonding to form the combustor's circumference. Attention is given to the fabrication of a platelet combustion chamber generating 40 Klbf of thrust.

Burkhardt, W. M.↗
NASA NTRS1998

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Smith, Timothy D.↗