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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

High Speed Data Investigation of an LOX/LNG Combustor Assisted by Non-linear Acoustic Simulations

During testing of a regenerative cooled liquid oxygen/liquid natural gas combustor large amplitude combustion instabilities were observed in half of the series conducted. These instabilities produced an unsteady pressure Root Mean Square (RMS) value greater than the chamber steady state pressure. Quick look investigations of the test data presented some standard indications of non-linear tangential instability such as a non-sinusoidal time series, and larger peaks than trough. Also observed in the data quick look were some atypical findings including a first harmonic amplitude larger than the fundamental, and numerous side bands on the fundamental and subsequent harmonics. A simplified model of the combustor was created using a commercially available finite element software. The first tangential acoustic mode was then simulated using a non-linear wave equation in conjunction with a time explicit discontinuous Gelarkin flow solver. Numerous simulations were evaluated providing valuable insight and reproducing the atypical findings in the data. Side bands were created through non-linear signal modulation like an AM radio. This happens due to the non-linear steepening and relaxing that occurs in large amplitude oscillations. The simulation was also able to reproduce a signal with a first harmonic amplitude larger than the fundamental. Investigating the pressure signal at multiple locations across the chamber head end revealed radial dependent frequency content. A literature review of the topic revealed corroboration of these effect in idealized geometries.

Combustion Stability↗
NASA NTRSDate not supplied

High Speed Data Investigation of an LOX/LNG Combustor Assisted by Non-linear Acoustic Simulations

During testing of a regenerative cooled liquid oxygen/liquid natural gas combustor large amplitude combustion instabilities were observed in half of the series conducted. These instabilities produced an unsteady pressure Root Mean Square (RMS) value greater than the chamber steady state pressure. Quick look investigations of the test data presented some standard indications of non-linear tangential instability such as a non-sinusoidal time series, and larger peaks than trough. Also observed in the data quick look were some atypical findings including a first harmonic amplitude larger than the fundamental, and numerous side bands on the fundamental and subsequent harmonics. A simplified model of the combustor was created using a commercially available finite element software. The first tangential acoustic mode was then simulated using a non-linear wave equation in conjunction with a time explicit discontinuous Gelarkin flow solver. Numerous simulations were evaluated providing valuable insight and reproducing the atypical findings in the data. Side bands were created through non-linear signal modulation like an AM radio. This happens due to the non-linear steepening and relaxing that occurs in large amplitude oscillations. The simulation was also able to reproduce a signal with a first harmonic amplitude larger than the fundamental. Investigating the pressure signal at multiple locations across the chamber head end revealed radial dependent frequency content. A literature review of the topic revealed corroboration of these effect in idealized geometries.

Combustion Stability↗
NASA NTRS1982

Evaluation of advanced combustion concepts for dry NO sub x suppression with coal-derived, gaseous fuels

The emissions performance of a rich lean combustor (developed for liquid fuels) was determined for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf (7.0 to 10.3 MJ/NCM). The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher. NOx generation from NH3 is significant at ammonia concentrations significantly less than 0.5%. These levels may occur depending on fuel gas cleanup system design. CO emissions, combustion efficiency, smoke and other operational performance parameters were satisfactory. A test was completed with a catalytic combustor concept with petroleum distillate fuel. Reactor stage NOx emissions were low (1.4g NOx/kg fuel). CO emissions and combustion efficiency were satisfactory. Airflow split instabilities occurred which eventually led to test termination.

Beebe, K. W.↗
NASA NTRS1982

Low NO sub x heavy fuel combustor concept program phase 1A gas tests

The emissions performance of a rich lean combustor (developed for liquid fuels) for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf were assessed. The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. Although meeting NOx goals for the 167 Btu/scf gas, NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx generation from NH3 was significant at ammonia concentrations significantly less tha 0.5%. These levels occur depending on fuel gas cleanup system design, However, NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher.

Cutrone, M. B.↗
NASA NTRS1950

Combustion efficiency and altitude operational limits of three liquid hydrocarbon fuels having high volumetric energy content in a J33 single combustor

Combustion efficiency and altitude operational limits were determined in a J33 single combustor for AN-F-58 fuel and three liquid hydrocarbon fuels having high volumetric energy content (decalin, tetralin, and monomethylnaphthalene) at simulated altitude and combustor inlet-air conditions. At the conditions investigated, the combustion efficiency for the four fuels generally decreased with an increase in volumetric energy content. The altitude operational limits for decalin and tetralin fuels were higher than for AN-F-58 fuel; monomethylnaphthalene fuel gave the lowest altitude operational limit.

Stricker, Edward G↗
NASA NTRS1974

Combustion instability with partial length acoustic liners

The development of analytical techniques and numerical methods for the prediction of the stability behavior of liquid propellant rocket combustors is discussed. The studies involved nonlinear wave propagation effects and nonlinear combustion zone, liner, and nozzle responses. A second primary extension of the basic theory was directed at the inclusion of the effects of injector face baffles. Attention was restricted to the linear (small amplitude) problem, finite Mach numbers, three dimensional oscillations, and distributed combustion effects.

Mitchell, C. E.↗
NASA NTRS1975

Combustion instability with partial length acoustic liners

The development is discussed of analytical techniques and numerical methods for the prediction of the stability behavior of liquid propellant rocket combustors with partial length acoustic liners, injector face baffles, and nonlinear combustion responses. Three primary objectives included: (1) extension of the Green's Function integral-iteration technique previously developed so that nonlinear three dimensional wave propagation effects could be included in stability predictions; (2) development of an analytical technique and a computer program for the prediction of the damping effects caused by injector face baffles; and (3) inclusion of a nonlinear, wave distortation dependent, combustion response model in the global stability analysis and determination of its effect on combustor stability.

Mitchell, C. E.↗
NASA NTRS1980

Combustion stability with baffles, absorbers and velocity sensitive combustion

Analytical and computational techniques were developed to predict the stability behavior of liquid propellant rocket combustors using damping devices such as acoustic liners, slot absorbers, and injector face baffles. Models were developed to determine the frequency and decay rate of combustor oscillations, the spatial and temporal pressure waveforms, and the stability limits in terms of combustion response model parameters.

Mitchell, C. E.↗
NASA NTRS1985

Atomization and Mixing Study

The primary objective was the obtainment of atomization and mixing performance data for a variety of typical liquid oxygen/hydrocarbon injector element designs. Such data are required to establish injector design criteria and to provide critical inputs to liquid rocket engine combustor performance and stability analysis, and computational codes and methods. Deficiencies and problems with the atomization test equipment were identified, and action initiated to resolve them. Test results of the gas/liquid mixing tests indicated that an assessment of test methods was required. A series of 71 liquid/liquid tests were performed.

Ferrenberg, A.↗
NASA NTRS1993

Prediction of prompt NO(x) in hydrocarbon air flames

The gas turbine industry is directing particular attention to very low NOx combustors, whether for aircraft or land based CCGT systems. These low NOx combustors frequently use liquid fuels or natural gas burning under very lean premixed conditions with air or under rich-lean conditions, although only the first case is studied here. In land based systems, diluted steam or nitrogen are sometimes injected into the combustion chambers to reduce flame temperature. The NOx emissions from such systems are the product of three chemical mechanisms which are interrelated: the hydrocarbon prompt NO, the thermal NO (extended Zeldovich mechanism), and the nitrous oxide route to NO. Formation of NO2 from NO also occurs, as well as emission of carbon monoxide and unburnt hydrocarbons. When the fuel-oxidant proportion decreases towards leaner conditions, flame temperatures are lowered, resulting in the total NO being reduced and the thermal-NO contribution greatly diminished to the benefit of the remaining two mechanisms of NO formation. While knowledge of the elementary reactions and their chemical kinetics concerning methane and simple hydrocarbons combustion has existed for a number of years, its use for computer modeling is limited to simple flow dynamics configurations. Nevertheless, understanding of such combusting flows under a wide range of experimental conditions allows for analogies or speculations with more complex actual systems. Such understanding can be achieved by means of one dimensional laminar premixed flame modeling, with a full chemical mechanism which incorporates the three routes of NO formation. Complementary to this understanding is the modeling of the actual combustion system using a full description of the fluid dynamics coupled with a reduced chemical scheme, which is then compared against the first model. The objective of this investigation is to evaluate the relative importance of the three mechanisms of NO formation in lean premixed methane-air combustion with increasing pressure using the one dimensional plug flow package PREMIX, and to test the validity of a three dimensional model with a global chemical mechanism against the one dimensional model in the atmospheric pressure case. Methane is chosen because it is the only mechanism which is reasonably well known and is a good guide to the behavior of other hydrocarbons. The mixture ratio chosen is richer than that in lean gas turbines, but the combustion of this mixture with the low preheat gives realistic gas turbine final flame temperatures. Conditions of NO2 formation are also analyzed in the one dimensional model and results extrapolated to the case of gas turbines.

Valerie Dupont↗
NASA NTRS2010

NASA Tech Briefs, October 2010

Topics covered include: Hybrid Architecture Active Wavefront Sensing and Control; Carbon-Nanotube-Based Chemical Gas Sensor; Aerogel-Positronium Technology for the Detection of Small Quantities of Organic and/or Toxic Materials; Graphene-Based Reversible Nano-Switch/Sensor Schottky Diode; Inductive Non-Contact Position Sensor; High-Temperature Surface-Acoustic-Wave Transducer; Grid-Sphere Electrodes for Contact with Ionospheric Plasma; Enabling IP Header Compression in COTS Routers via Frame Relay on a Simplex Link; Ka-Band SiGe Receiver Front-End MMIC for Transponder Applications; Robust Optimization Design Algorithm for High-Frequency TWTs; Optimal and Local Connectivity Between Neuron and Synapse Array in the Quantum Dot/Silicon Brain; Method and Circuit for In-Situ Health Monitoring of Solar Cells in Space; BGen: A UML Behavior Network Generator Tool; Platform for Post-Processing Waveform-Based NDE; Electrochemical Hydrogen Peroxide Generator; Fabrication of Single, Vertically Aligned Carbon Nanotubes in 3D Nanoscale Architectures; Process to Create High-Fidelity Lunar Dust Simulants; Lithium-Ion Electrolytes Containing Phosphorous-Based, Flame-Retardant Additives; InGaP Heterojunction Barrier Solar Cells; Straight-Pore Microfilter with Efficient Regeneration; Determining Shear Stress Distribution in a Laminate; Self-Adjusting Liquid Injectors for Combustors; Handling Qualities Prediction of an F-16XL-Based Reduced Sonic Boom Aircraft; Tele-Robotic ATHLETE Controller for Kinematics - TRACK; Three-Wheel Brush-Wheel Sampler; Heterodyne Interferometer Angle Metrology; Aligning Astronomical Telescopes via Identification of Stars; Generation of Optical Combs in a WGM Resonator from a Bichromatic Pump; Large-Format AlGaN PIN Photodiode Arrays for UV Images; Fiber-Coupled Planar Light-Wave Circuit for Seed Laser Control in High Spectral Resolution Lidar Systems; On Calculating the Zero-Gravity Surface Figure of a Mirror; Optical Modification of Casimir Forces for Improved Function of Micro- and Nano-Scale Devices; Analysis, Simulation, and Verification of Knowledge-Based, Rule-Based, and Expert Systems; Core and Off-Core Processes in Systems Engineering; Digital Reconstruction Supporting Investigation of Mishaps; and Template Matching Approach to Signal Prediction.

Source record↗
NASA NTRSDate not supplied

Characterization of an Optically Accessible RP/GOx Rotating Detonation Engine

The accurate characterization of the dynamic response of liquid jets to transient detonation wave passage is crucial for optimizing and modeling liquid-fueled rotating detonation combustors. To address this need, we have developed an optically accessible rocket RDE (RP-RDE), which utilizes liquid RP-1 and gaseous oxygen as propellants and operates at elevated pressures. The experimental setup features optical access from the oxidizer injection plenum to the combustor exit, allowing for detailed characterization of injection, mixing, and combustion processes and their impact on the detonation wave structure. To design and develop the liquid injection scheme for the RP-RDE test article, we employed a combined experimental and numerical approach that included cold-flow simulations. Our aim was to optimize the liquid injection process and ensure that the resulting fuel-oxidizer mixture is homogeneous and stable. To this end, high-speed optical imaging will be employed to provide qualitative visualization of the liquid and gaseous phases of the propellant as well as the local heat release during detonation wave propagation through the non-premixed medium. This work will provide critical insights into the behavior of the transient detonation waves in more relevant rocket conditions, which is essential for the optimization and modeling of rotating detonation combustors.

Propulsion↗
NASA NTRS1953

Effect of Axially Staged Fuel Introduction on Performance of One-quarter Sector of Annular Turbojet Combustor

The design principle of injecting liquid fuel at more than one axial station in an annual turbojet combustor was investigated. Fuel was injected into the combustor as much as 5 inches downstream of the primary fuel injectors. Many fuel-injection configurations were examined and the performance results are presented for 11 configurations that best demonstrate the trends in performance obtained. The performance investigations were made at a constant combustor-inlet pressure of 15 inches of mercury absolute and at air flows up to 70 percent higher than values typical of current design practice. At these higher air flows, staging the fuel introduction improved the combustion efficiency considerably over that obtained in the combustor when no fuel staging was employed. At air flows currently encountered in turbojet engines, fuel staging was of minor value. Radial temperature distribution seemed relatively unaffected by the location of fuel-injection stations.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1993

The effect of incomplete fuel-air mixing on the lean limit and emissions characteristics of a Lean Prevaporized Premixed (LPP) combustor

Results are presented from an experimental study of the effect of incomplete fuel-air mixing on the lean limit and emissions characteristics of a lean, prevaporized, premixed (LPP), coaxial mixing tube combustor. Two-dimensional exciplex fluorescence was used to characterize the degree of fuel vaporization and mixing at the combustor inlet under non-combusting conditions. These tests were conducted at a pressure of 4 atm., a temperature of 400 C, a mixer tube velocity of 100 m/sec and an equivalence ratio of .8, using a mixture of tetradecane, 1 methyl naphthalene and TMPD as a fuel simulant. Fuel-air mixtures with two distinct spatial distributions were studied. The exciplex measurements showed that there was a significant amount of unvaporized fuel at the combustor entrance in both cases. One case, however, exhibited a very non-uniform distribution of fuel liquid and vapor at the combustor entrance, i.e., with most of the fuel in the upper half of the combustor tube, while in the other case, both the fuel liquid and vapor were much more uniformly distributed across the width of the combustor entrance. The lean limit and emissions measurements were all made at a pressure of 4 atm. and a mixer tube velocity of 100 m/sec, using Jet A fuel and both fuel-air mixture distributions. Contrary to what was expected, the better mixed case was found to have a substantially leaner operating limit. The two mixture distributions also unexpectedly resulted in comparable NO(x) emissions, for a given equivalence ratio and inlet temperature, however, lower NO(x) emissions were possible in the better mixed case due to its leaner operating limit.

Santavicca, D. A.↗
NASA NTRS1995

Resolution of the buoyancy in the 8-foot high temperature tunnel combustor

Currently, the 8-Foot High Temperature Tunnel (8-Ft. HTT) combustor produces a good profile at only one point (2000 psia and 3650 R with oxygen enrichment). Air is enriched with oxygen (liquid) so that the combustor product gas will contain the volumetric amount of oxygen normally found in air. The oxygen enriched air has a large fraction that is not reacted and flows through the outer periphery of the fuel injector. This ring of cold air in addition to the relatively cold walls of the combustor set up buoyancy forces that produce a segregation of relatively cool gases at the bottom of the combustor exit. The basic problem is to produce a test gas that has uniform properties at all combustor conditions. The combustor temperature may be as high as 3700 R or as low as 2000 R. Combustor pressures can be as high as 3500 psia (no oxygen enrichment) and as low as 600 psia. The segregation is most severe with oxygen enriched air, since its temperature is lower and its density is high. The combustor is lined with nickel 201 and can be operated at about 1600 R maximum. A global mixing process is desired that produces an acceptable profile of temperature, species, and velocity at the exit of the combustor. The ultimate goal is a temperature profile with about 100 R variance and about 2 percent variance in oxygen. The exit total temperature must not be lowered significantly by the mixing apparatus or mechanisms employed. If immersed bodies are used, they must also be kept very hot. All combustor wall modifications must be able to survive the heat and structural conditions of the varied operating conditions. Our approach to resolving this issue is being conducted in three stages: (1) Consider mixing exclusively, (2) Resolve the heat transfer concerns resulting from the chosen mixing strategy, and (3) Solve the material and structural problems resulting from stages (1) and (2). Since the 8-Ft. HTT is unavailable for experimentation, the study is conducted exclusively with computational fluid dynamic (CPD) codes (Fluent/Uns and Rampant 3.1) using unstructured grid through body fitted coordinates. Both CFD codes are general purpose Navier-Stokes solution packages that can solve integral conservation equations for conservation of mass, momentum and energy. The governing equations are discretized using a control-volume finite-element method on unstructured triangular 2-D grids. In the interest of time, a 3-D tetrahedral grid was used to check the 2-D results on one mixing strategy and the 2-D results were confirmed. Preliminary results indicate that excellent mixing can be achieved with a body placed in the center of the flowing hot fluid with a minor modification to the combustor wall similar to a model positioned in the test section of a wind tunnel. The concept here, is to create longitudinal vortices strong enough to bring both fluids into intimate contact with each other near the wall where the cold fluid resides. However, there may be a trade-off in the choice of mixing strategy, heat transfer and structural requirements. Currently, we are examining the more promising geometries for heat transfer concerns and developing strategies for the material of construction of the center body.

Loney, Norman W.↗
NASA NTRS1992

Effect of model selection on combustor performance and stability predictions using ROCCID

The ROCket Combustor Interactive Design (ROCCID) methodology is an interactive computer program that combines previously developed combustion analysis models to calculate the combustion performance and stability of liquid rocket engines. Test data from 213 kN (48,000 lbf) Liquid Oxygen (LOX)/RP-1 combustor with an O-F-O (oxidizer-fuel-oxidizer) triplet injector were used to characterize the predictive capabilities of the ROCCID analysis models for this injector/propellant configuration. Thirteen combustion performance and stability models were incorporated into ROCCID, and ten of them, which have options for triplet injectors, were examined. Calculations using different combinations of analysis models, with little or no anchoring, were carried out on a test matrix of operating combinations matching those of the test program. Results of the computer analyses were compared to test data, and the ability of the model combinations to correctly predict combustion stability or instability was determined. For the best model combination(s), sensitivity of the calculations to fuel drop size and mixing efficiency was examined. Error in the stability calculations due to uncertainty in the pressure interaction index (N) was examined. The recommended model combinations for this O-F-O triplet LOX/RP-1 configuration are proposed.

Giuliani, James E.↗