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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 181 records · Page 10
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NASA NTRS2019

Pilot Injector Redesign to Reduce N+3 Cycle Emissions for a Gas-Turbine Combustor

An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASA’s N+3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flame tube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a ‘medium-power’ cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

nitrogen oxides injectors↗
NASA NTRS2019

Pilot Injector Redesign to Reduce N+3 Cycle Emissions for a Gas-Turbine Combustor

An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASA's N plus 3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flametube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a "medium-power" cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

Nitrogen Oxides↗
NASA NTRS1974

Platelet injectors for Space Shuttle orbit maneuvering engine

The Space Shuttle Orbit Maneuvering Subsystem Rocket Engine employs a platelet element injector concept. This injector has demonstrated 316-sec vacuum specific impulse performance under simulated altitude conditions when tested with a milled slot/electroformed nickel close-out regenerative chamber and a full 71 area ratio nozzle. To date, over 300 altitude engine tests and 300 stability bomb tests have demonstrated stable, erosion free operation with this concept to test durations of 150 seconds. The injector and chamber also meet the reusable requirements of the shuttle with a cycle life capability in excess of 1000 cycles. An extensive altitude restart program has also demonstrated OMS-engine operation over large variations in the burn and coast times with helium saturated propellants.

Kahl, R. C.↗
NASA NTRS1975

Nonreactive mixing study of a scramjet swept-strut fuel injector

The results are presented of a cold-mixing investigation performed to supply combustor design information and to determine optimum normal fuel-injector configurations for a general scramjet swept-strut fuel injector. The experimental investigation was made with two swept struts in a closed duct at a Mach number of 4.4 and a nominal ratio of jet mass flow to air mass flow of 0.0295, with helium used to simulate hydrogen fuel. Four injector patterns were evaluated; they represented the range of hole spacing and the ratio of jet dynamic pressure to free-stream dynamic pressure. Helium concentration, pitot pressure, and static pressure in the downstream mixing region were measured to generate the contour plots needed to define the mixing-region flow field and the mixing parameters. Experimental results show that the fuel penetration from the struts was less than the predicted values based on flat-plate data; but the mixing rate was faster and produced a mixing length less than one-half that predicted.

Mcclinton, C. R.↗
NASA NTRS1978

Photographic characterization of spark-ignition engine fuel injectors

Manifold port fuel injectors suitable for use in general aviation spark-ignition engines were evaluated qualitatively on the basis of fuel spray characteristics. Photographs were taken at various fuel flow rates or pressure levels. Mechanically and electronically operated pintle injectors generally produced the most atomization. The plain-orifice injectors used on most fuel-injected general aviation engines did not atomize the fuel when sprayed into quiescent air.

Evanich, P. L.↗
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.↗