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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 NTRS1978

Investigation of the burning configuration of a coaxial injector in a combustion chamber

An analytical investigation was made into the stability of the burning configuration of a single coaxial injector surrounded by similar injectors. The stability criteria was based on an average pressure difference along the boundaries of the adjacent stream tubes as calculated using Spaulding's numerical method. The results indicate qualitatively that there is a tendency for the injectors to have different burning configurations. It is believed that the configuration achieved is random, however once the burning configuration is established, it is believed to persist.

Ohara, J.↗
NASA NTRS1978

Reverse-flow combustor for small gas turbines with pressure-atomizing fuel injectors

A reverse flow combustor suitable for a small gas turbine (2 to 3 kg/s mass flow) was used to evalute the effect of pressure atomizing fuel injectors on combustor performance. In these tests an experimental combustor was designed to operate with 18 simplex pressure atomizing fuel injectors at sea level takeoff conditions. To improve performance at low power conditions, fuel was redistributed so that only every other injector was operational. Combustor performance, emissions, and liner temperature were compared over a range of pressure and inlet air temperatures corresponding to simulated idle, cruise, and takeoff conditions typical of a 16 to 1 pressure ratio turbine engine.

Norgren, C. T.↗
NASA NTRS1979

Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

The feasibility of using a heavy hydrocarbon fuel as a rocket propellant is examined. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. Experiments were done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in) to 55.9 cm (22 in). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by reaming each injector several times to provide test data over a range of injector pressure drop.

Pavli, A. J.↗
NASA NTRS1979

Design and evaluation of high performance rocket engine injectors for use with hydrocarbon fuels

An experimental program to determine the feasibility of using a heavy hydrocarbon fuel as a rocket propellant is reported herein. A method of predicting performance of a heavy hydrocarbon in terms of vaporization effectiveness is described and compared to other fuels and to experimental test results. The work was done at a chamber pressure of 4137 KN/sq M (600 psia) with RP-1, JP-10, and liquefied natural gas as fuels, and liquid oxygen as the oxidizer. Combustion length effects were explored over a range of 21.6 cm (8 1/2 in.) to 55.9 cm (22 in.). Four injector types were tested, each over a range of mixture ratios. Further configuration modifications were obtained by 'reaming' each injector several times to provide test data over a range of injector pressure drop.

Pavli, A. J.↗
NASA NTRS1984

Low loss injector for Space Shuttle main engine. Center director's discretionary fund

An efficient propellant injection method to raise the Space Shuttle Main Engine (SSME) thrust and payload is discussed. Relatively large diameter injector elements with low pressure loss are recommended for the main combustion chamber and the pre-burners. Smaller losses admit more propellant flow which then raises thrust. Payload is not only gained by specific impulse but also by thrust. The chamber pressure is stabilized by selecting the proper cavity size for the injector elements while reducing the injection pressure loss which normally is kept high for stability. The rather large injector element recesses provide acoustic damping which makes baffles and acoustic absorbers unnecessary. A tenfold reduction of flow induced stresses which are rather high in the present design is shown. Relaxed tolerances, fewer elements, and better maintenance are offered. The study was conducted under a center director discretionary fund assignment.

Vonpragenau, G. L.↗
NASA NTRS1987

In-circuit fault injector user's guide

A fault injector system, called an in-circuit injector, was designed and developed to facilitate fault injection experiments performed at NASA-Langley's Avionics Integration Research Lab (AIRLAB). The in-circuit fault injector (ICFI) allows fault injections to be performed on electronic systems without special test features, e.g., sockets. The system supports stuck-at-zero, stuck-at-one, and transient fault models. The ICFI system is interfaced to a VAX-11/750 minicomputer. An interface program has been developed in the VAX. The computer code required to access the interface program is presented. Also presented is the connection procedure to be followed to connect the ICFI system to a circuit under test and the ICFI front panel controls which allow manual control of fault injections.

Padilla, Peter A.↗
NASA NTRS1987

Application of advanced diagnostics to airblast injector flows

This effort is concerned with the application of both conventional laser velocimetry and phase Doppler anemometry to the flow produced by an airblast nozzle. The emphasis is placed on the acquisition of data using actual engine injector/swirler components at (noncombusting) conditions simulating those encountered in the engine. The objective of the effort was to test the applicability of the instrumentation to real injector flows, to develop information on the behavior of injectors at high flow, and to provide data useful in the development of physical models of injector flows.

Mcvey, John B.↗
NASA NTRS1989

The structural dynamics analysis of the main injector LOX inlet tee and its redesign

A ground test of a specially instrumented Space Shuttle Main Engine (SSME) revealed excessively high strains in the high pressure Liquid Oxygen (LOX) inlet line which carries LOX to the Main Injector. The inlet tee acts as a manifold which utilizes two splitter vanes to direct the flow of LOX into the main injector. When the high strains were discovered, it was thought that these splitter vanes, coupled with high energy flow was the source. This resulted not only in high strains in the LOX inlet tee, but excessive vibration at a frequency of 4000 Hz located at the gimbal bearing, and at the main injector. This was later to be known as The 4kHz Phenomenon. The dynamic analysis which was used to predict the response due to the 4kHz phenomenon and the analysis used to aid in the design modification required to the hardware to reduce or eliminate the phenomenon is examined.

Romano, E.↗
NASA NTRS1988

Injector element characterization methodology

Characterization of liquid rocket engine injector elements is an important part of the development process for rocket engine combustion devices. Modern nonintrusive instrumentation for flow velocity and spray droplet size measurement, and automated, computer-controlled test facilities allow rapid, low-cost evaluation of injector element performance and behavior. Application of these methods in rocket engine development, paralleling their use in gas turbine engine development, will reduce rocket engine development cost and risk. The Alternate Turbopump (ATP) Hot Gas Systems (HGS) preburner injector elements were characterized using such methods, and the methodology and some of the results obtained will be shown.

Cox, George B., Jr.↗
NASA NTRS1991

Subscale LOX/hydrogen testing with a modular chamber and a swirl coaxial injector

To support basic injector and chamber technology, the response of a modular calorimeter chamber to an injector with swirl coaxial elements was investigated. The tests supported both the Advanced Launch System (ALS) and the Space Shuttle Main Engine programs. Original test plans included chamber pressures up to 2250 psia, the current ALS engine criteria, but hardware limitations forced conditions to be reduced. With four different chamber configurations, a variety of data was obtained with chamber pressures ranging from 1483 psia to 1679 psia and mixture ratios from 5.24 to 6.90. The swirl coaxial injector showed good chamber wall compatibility, despite misaligned elements, and its high performance was independent of fuel. Heating rates and wall temperatures were acceptable and close to predict profiles, and after scaling to a chamber pressure of 2250 psia, heating rates remained acceptable. In addition, a chamber spool made with the vacuum plasma spray process survived hot-fire testing with no detrimental effects.

Elam, Sandra K.↗
NASA NTRS1991

An investigation of a contoured wall injector for hypervelocity mixing augmentation

An experimental and computational investigation of a contoured wall fuel injector is presented. The injector was aimed at enabling shock-enhanced mixing for the supersonic combustion ramjet engines currently envisioned for applications on hypersonic vehicles. Three-dimensional flow field surveys, and temporally-resolved planar Rayleigh scattering measurements are presented for Mach 1.7 helium injection into Mach 6 air. These experimental data are compared directly with a three-dimensional Navier-Stokes simulation of the flow about the injector array. Two dominant axial vorticity sources are identified and characterized. The axial vorticity produced strong convective mixing of the injectant with the freestream. Shock-impingement was particularly effective as it assured seeding of baroclinic vorticity directly on the helium/air interface. The vorticity coalesced into a counter-rotating vortex pair of a sense which produced migration of the helium away from the wall. The influences of spatial averaging on the representation of the flow field as well as the importance of the fluctuating component of the flow in producing molecularly-mixed fluid are addressed.

Waitz, Ian A.↗
NASA NTRS1992

A systematic experimental and computational investigation of a class of contoured wall fuel injectors

The performance of contoured wall fuel injectors for scramjet engine applications is considered. These fuel injectors were aimed at augmenting mixing through axial vorticity production arising from interaction of the fuel/air interface with an oblique shock. The effects of incoming boundary layer height, injector spacing, and injectant to freestream pressure and velocity ratios are examined. Results from 3D flow field surveys and Navier-Stokes simulations are presented.

Waitz, Ian A.↗
NASA NTRS1992

Performance effects resulting from plugged liquid oxygen posts of the Space Shuttle Main Engine Injector

The paper discusses the performance effects resulting from plugged LOX posts of the Space Shuttle Main Engine Injector. The simulation was performed with the REFLEQS 2-D code. Analysis was performed axisymmetrically and injector surface was divided into several regions to account for the mixture ratio variation on the injector surface. The reduction of vaccum specific impulse was approximately 0.01 second per plugged LOX post. This reduction is an order of magnitude higher than the result of Space Shuttle flight reconstruction data. It is presumed that this overprediction is due to the axisymmetric simulation that smears local effects.

Kim, S.↗
NASA NTRS1992

Exit plane H2O concentration measurements correlated with OH PLIF near-injector mixing measurements for scramjet flows

Mixing and combusting high enthalpy flows, similar to those encountered in scramjet engines, were investigated using a shock tunnel to produce the flow in conjunction with non-intrusive optical diagnostics which monitored the performance of two injector configurations. The shock tunnel is configured to produce Mach 3 flow and stagnation enthalpies corresponding to flight equivalent Mach numbers between 7 and 11. A pulsed hydrogen injection capability and interchangeable injector blocks provide a means of examining high speed, high enthalpy reacting flows. Planar laser induced fluorescence (PLIF) of OH molecules in the near injector region produced images which show the combusting and mixing zones for the reacting flow. Line-of-sight exit plane measurement of water concentration and temperature were used to provide a unique method of monitoring exit plane products. These results demonstrated that a velocity matched axial injection system produced a fuel jet that lifted off the floor of the duct. Mixing was observed to increase for this system as a velocity mismatch was introduced. Comparison of exit plane water concentrations for a wall jet injection system and a velocity matched injection system indicated similar mixing performance but an accurate pressure measurement is necessary to further validate the result. In addition, exit plane measurements indicated an approximate steady-state condition was achieved during the 1 to 2 ms test times.

Parker, T. E.↗
NASA NTRS1993

Experimental investigation of a simulated LOX injector flow field

Current development of liquid oxidizer injectors for liquid rocket engines is highly dependent upon experimental investigations of the two-phase liquid oxidizer/gaseous fuel injection plume characteristics. This paper describes an effort to increase the applicability of optical diagnostic approaches to such flow fields find includes an innovative approach for interpreting laser-induced fluorescence signals from axisymmetric two-phase flows. The usefulness of both laser-induced fluorescence and surface scattering techniques are explored herein. The flow field selected for this investigation is an oxidizer-swirled, coaxial flow produced using an injector design under consideration for use in the Space Transportation Main Engine (STME). The liquid oxygen (LOX) flow through this injector is simulated using water. Photographs of laser-induced fluorescence and particle scattering and measured profiles of the mass distribution in this flow field are presented and discussed herein.

Hartfield, Roy J., Jr.↗
NASA NTRS1999

Main Chamber and Preburner Injector Technology

This document reports the experimental and analytical research carried out at the Penn State Propulsion Engineering Research Center in support of NASA's plan to develop advanced technologies for future single stage to orbit (SSTO) propulsion systems. The focus of the work is on understanding specific technical issues related to bi-propellant and tri-propellant thrusters. The experiments concentrate on both cold flow demonstrations and hot-fire uni-element tests to demonstrate concepts that can be incorporated into hardware design and development. The analysis is CFD-based and is intended to support the design and interpretation of the experiments and to extrapolate findings to full-scale designs. The research is divided into five main categories that impact various SSTO development scenarios. The first category focuses on RP-1/gaseous hydrogen (GH2)/gaseous oxygen (GO2) tri-propellant combustion with specific emphasis on understanding the benefits of hydrogen addition to RP-1/oxygen combustion and in developing innovative injector technology. The second category investigates liquid oxygen (LOX)/GH2 combustion at main chamber near stoichiometric conditions to improve understanding of existing LOX/GH2 rocket systems. The third and fourth categories investigate the technical issues related with oxidizer-rich and fuel-rich propulsive concepts, issues that are necessary for developing the full-flow engine cycle. Here, injector technology issues for both LOX/GH2 and LOX/RP-1 propellants are examined. The last category, also related to the full-flow engine cycle, examines injector technology needs for GO2/GH2 propellant combustion at near-stoichiometric conditions for main chamber application.

Santoro, Robert J.↗
NASA NTRS1999

Design of a Novel Gaseous Hydrogen-Oxygen Rocket Injector Element

NASA and Aerojet are developing a Rocket-Based Combined Cycle (RBCC) engine under the Advanced Reusable Technology program. The rocket application requires that the combustion process be stable, complete, and take place in as short a distance as possible without compromising the structural integrity of the injector itself. A novel gaseous hydrogen-oxygen rocket injector element design was arrived at through an iterative design process making extensive use of CFD simulations, which resulted in a design that is meeting design goals. Sub-scale versions of the injector have been built and tested in a unique test-rig and in a sub-scale RBCC engine. The Aerojet RBCC concept integrates small rocket thrusters into the rearfacing base area of struts placed in the flowpath of a scramjet (Supersonic Combusting Ramjet) engine. In one mode of operation, at vehicle takeoff, the rockets provide the primary thrust with additional thrust coming from an ejector effect as air is drawn into the engine inlet, entrained, and accelerated by the rocket exhaust.

Glenn, Dennis E.↗
NASA NTRS1998

Numerical Simulations of a Lobed Fuel Injector

Numerical modeling of the nonreactive mixing processes associated with a lobed fuel injector in a coflowing air stream is presented. The lobed fuel injector is a device which generates strong streamwise vorticity, producing locally high strain rates which can enhance the molecular mixing of reactants while delaying ignition in a controlled manner. Vortex element modeling is used to simulate flow field evolution and fuel element mixing characteristics for this lobed fuel injector. Quantitative predictions for vorticity generation and qualitative results for streamwise rollup compare well qualitatively with recent experimental investigations of this flow field [Smith et al, Phys. Fluids 9, 667 (1997)]. Parametric studies of the effects of lobe amplitude-to-wavelength ratio, lobe angle, and lobe shape for given flow conditions suggest that geometrical features may be optimized to enhance mixing and control reaction processes.

Strickland, J. H.↗