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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Characterization and Simulation of the Thermoacoustic Instability Behavior of an Advanced, Low Emissions Combustor Prototype

Extensive research is being done toward the development of ultra-low-emissions combustors for aircraft gas turbine engines. However, these combustors have an increased susceptibility to thermoacoustic instabilities. This type of instability was recently observed in an advanced, low emissions combustor prototype installed in a NASA Glenn Research Center test stand. The instability produces pressure oscillations that grow with increasing fuel/air ratio, preventing full power operation. The instability behavior makes the combustor a potentially useful test bed for research into active control methods for combustion instability suppression. The instability behavior was characterized by operating the combustor at various pressures, temperatures, and fuel and air flows representative of operation within an aircraft gas turbine engine. Trends in instability behavior versus operating condition have been identified and documented, and possible explanations for the trends provided. A simulation developed at NASA Glenn captures the observed instability behavior. The physics-based simulation includes the relevant physical features of the combustor and test rig, employs a Sectored 1-D approach, includes simplified reaction equations, and provides time-accurate results. A computationally efficient method is used for area transitions, which decreases run times and allows the simulation to be used for parametric studies, including control method investigations. Simulation results show that the simulation exhibits a self-starting, self-sustained combustion instability and also replicates the experimentally observed instability trends versus operating condition. Future plans are to use the simulation to investigate active control strategies to suppress combustion instabilities and then to experimentally demonstrate active instability suppression with the low emissions combustor prototype, enabling full power, stable operation.

DeLaat, John C.↗
NASA NTRS1931

Performance of a Compression-ignition Engine with a Precombustion Chamber Having High-Velocity Air Flow

Presented here are the results of performance tests made with a single-cylinder, four stroke cycle, compression-ignition engine. These tests were made on a precombustion chamber type of cylinder head designed to have air velocity and tangential air flow in both the chamber and cylinder. The performance was investigated for variable load and engine speed, type of fuel spray, valve opening pressure, injection period and, for the spherical chamber, position of the injection spray relative to the air flow. The pressure variations between the pear-shaped precombustion chamber and the cylinder for motoring and full load conditions were determined with a Farnboro electric indicator. The combustion chamber designs tested gave good mixing of a single compact fuel spray with the air, but did not control the ensuing combustion sufficiently. Relative to each other, the velocity of air flow was too high, the spray dispersion by injection too great, and the metering effect of the cylinder head passage insufficient. The correct relation of these factors is of the utmost importance for engine performance.

Spanogle, J A↗
NASA NTRS2003

Control of Thermo-Acoustics Instabilities: The Multi-Scale Extended Kalman Approach

"Multi-Scale Extended Kalman" (MSEK) is a novel model-based control approach recently found to be effective for suppressing combustion instabilities in gas turbines. A control law formulated in this approach for fuel modulation demonstrated steady suppression of a high-frequency combustion instability (less than 500Hz) in a liquid-fuel combustion test rig under engine-realistic conditions. To make-up for severe transport-delays on control effect, the MSEK controller combines a wavelet -like Multi-Scale analysis and an Extended Kalman Observer to predict the thermo-acoustic states of combustion pressure perturbations. The commanded fuel modulation is composed of a damper action based on the predicted states, and a tones suppression action based on the Multi-Scale estimation of thermal excitations and other transient disturbances. The controller performs automatic adjustments of the gain and phase of these actions to minimize the Time-Scale Averaged Variances of the pressures inside the combustion zone and upstream of the injector. The successful demonstration of Active Combustion Control with this MSEK controller completed an important NASA milestone for the current research in advanced combustion technologies.

Le, Dzu K.↗
NASA NTRS2010

Method and Apparatus for Thermal Spraying of Metal Coatings Using Pulsejet Resonant Pulsed Combustion

An apparatus and method [or thermal spraying a metal coating on a substrate is accomplished with a modified pulsejet and optionally an ejector to assist in preventing oxidation. Metal such a Aluminum or Magnesium may be used. A pulsejet is first initiated by applying fuel, air. and a spark. Metal is inserted continuously in a high volume of meta1 into a combustion chamber of the pulsejet. The combustion is thereafter. controlled resonantly at high frequency and the metal is heated to a molten state. The metal is then transported from the combustion chamber into a tail pipe of said pulsejet and is expelled therefrom at high velocity and deposited on a target substrate.

Paxson, Daniel E.↗
NASA NTRS2014

Method and Apparatus for Thermal Spraying of Metal Coatings Using Pulsejet Resonant Pulsed Combustion

An apparatus and method for thermal spraying a metal coating on a substrate is accomplished with a modified pulsejet and optionally an ejector to assist in preventing oxidation. Metal such as Aluminum or Magnesium may be used. A pulsejet is first initiated by applying fuel, air, and a spark. Metal is inserted continuously in a high volume of metal into a combustion chamber of the pulsejet. The combustion is thereafter controlled resonantly at high frequency and the metal is heated to a molten state. The metal is then transported from the combustion chamber into a tailpipe of said pulsejet and is expelled therefrom at high velocity and deposited on a target substrate.

Paxson, Daniel E.↗
NASA NTRS1936

Some Effects of Injection Advance Angle, Engine-Jacket Temperature, and Speed on Combustion in a Compression-Ignition Engine

An optical indicator and a high-speed motion-picture camera capable of operating at the rate of 2,000 frames per second were used to record simultaneously the pressure development and the flame formation in the combustion chamber of the NACA combustion apparatus. Tests were made at engine speeds of 570 and 1,500 r.p.m. The engine-jacket temperature was varied from 100 degrees to 300 degrees F. And the injection advance angle from 13 degrees after top center to 120 degrees before top center. The results show that the course of the combustion is largely controlled by the temperature and pressure of the air in the chamber from the time the fuel is injected until the time at which combustion starts and by the ignition lag. The conclusion is presented that in a compression-ignition engine with a quiescent combustion chamber the ignition lag should be the longest that can be used without excessive rates of pressure rise; any further shortening of the ignition lag decreased the effective combustion of the engine.

Rothrock, A M↗
NASA NTRS1999

NASA GRC's High Pressure Burner Rig Facility and Materials Test Capabilities

The High Pressure Burner Rig (HPBR) at NASA Glenn Research Center is a high-velocity. pressurized combustion test rig used for high-temperature environmental durability studies of advanced materials and components. The facility burns jet fuel and air in controlled ratios, simulating combustion gas chemistries and temperatures that are realistic to those in gas turbine engines. In addition, the test section is capable of simulating the pressures and gas velocities representative of today's aircraft. The HPBR provides a relatively inexpensive. yet sophisticated means for researchers to study the high-temperature oxidation of advanced materials. The facility has the unique capability of operating under both fuel-lean and fuel-rich gas mixtures. using a fume incinerator to eliminate any harmful byproduct emissions (CO, H2S) of rich-burn operation. Test samples are easily accessible for ongoing inspection and documentation of weight change, thickness, cracking, and other metrics. Temperature measurement is available in the form of both thermocouples and optical pyrometery. and the facility is equipped with quartz windows for observation and video taping. Operating conditions include: (1) 1.0 kg/sec (2.0 lbm/sec) combustion and secondary cooling airflow capability: (2) Equivalence ratios of 0.5- 1.0 (lean) to 1.5-2.0 (rich), with typically 10% H2O vapor pressure: (3) Gas temperatures ranging 700-1650 C (1300-3000 F): (4) Test pressures ranging 4-12 atmospheres: (5) Gas flow velocities ranging 10-30 m/s (50-100) ft/sec.: and (6) Cyclic and steady-state exposure capabilities. The facility has historically been used to test coupon-size materials. including metals and ceramics. However complex-shaped components have also been tested including cylinders, airfoils, and film-cooled end walls. The facility has also been used to develop thin-film temperature measurement sensors.

Robinson, R. Craig↗
NASA NTRS1950

Effect of Ram-jet Pressure Pulsations on Supersonic-diffuser Performance

Experimental study of effects of combustion roughness and controlled mechanical oscillations on diffuser operation was conducted on an 8-inch ram jet at Mach number 1.87. Within experimental accuracy, the optimum mean combustion-chamber static pressure occurred when the maximum instantaneous static pressure equalled the optimum steady-flow value and the corresponding decrement in optimum mean static pressure was equal to one-half the total amplitude of the pressure pulsations. A marked attenuation of amplitude of cold-buzz pressure fluctuations was effected over limited range of subcritical diffuser operation by means of properly tuned rotating disk located in combustion chamber.

Connors, James F↗
NASA NTRS2004

Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

Sims, J. D.↗
NASA NTRS1972

Design and evaluation of an oxidant-fuel-ratio-zoned rocket injector for high performance and ablative engine compatibility

A method for temperature control of the combustion gases in the peripheral zone of a rocket combustor which would reduce ablative throat erosion, prevent melting of zirconia throat inserts, and maintain high combustion performance is discussed. Included are techniques for analyzing and predicting zoned injector performance, as well as the philosophy and method for accomplishing an optimum compromise between high performance and reduced effective gas temperature. The experimental work was done with a 1000-lbf rocket engine which used as propellants N2O4 and a blend of 50-percent N2H4 and 50-percent UDMH at 100-psia chamber pressure and an overall O/F of 2.0. The method selected to provide temperature control was to use 30 percent of the propellant to form a peripheral zone of combustion gases at an O/F of 1.31 and 2700 K. The remaining 70 percent of the propellant in the core was at an O/F of 2.45 to keep the overall O/F at 2.0.

Winter, J. M.↗
NASA NTRS1982

Techniques for enhancing durability and equivalence ratio control in a rich-lean, three-stage ground power gas turbine combustor

Rig tests of a can-type combustor were performed to demonstrate two advanced ground power engine combustor concepts: steam cooled rich-burn combustor primary zones for enhanced durability; and variable combustor geometry for three stage combustion equivalence ratio control. Both concepts proved to be highly successful in achieving their desired objectives. The steam cooling reduced peak liner temperatures to less than 800 K. This offers the potential of both long life and reduced use of strategic materials for liner fabrication. Three degrees of variable geometry were successfully implemented to control airflow distribution within the combustor. One was a variable blade angle axial flow air swirler to control primary airflow while the other two consisted of rotating bands to control secondary and tertiary or dilution air flow.

Schultz, D. F.↗
NASA NTRS1987

Hypersonic propulsion research

The NASA Langley Research Center has conducted hypersonic propulsion research since the 1960s. A variety of inlet concepts were explored and characterized. The emphasis of the inlet program was the development of the short (light weight), fixed geometry, side-wall-compression inlets that operate efficiently over a wide Mach number range. As hypersonic combustion tunnels were developed, programs to study the parameters controlling fuel mixing and combustion with single and multiple strut models were conducted using direct connect test techniques. These various tests supported the design of subscale engine test hardware that integrated inlet and combustor technology and allowed the study of the effect of heat release on thrust and combustor/inlet interaction. A number of subscale engine tests have demonstrated predicted performance levels at Mach 4 and 7 simulated flight conditions.

Northam, G. Burton↗
NASA NTRS1990

Hypersonic propulsion research

The development of technology for the modular airframe integrated scramjet has been the focus of hypersonic propulsion research for several years. An part of this research, a variety of inlet concepts have been explored and characterized. The emphasis of the inlet program has been the development of the short (light weight), fixed geometry, side wall compression inlets that operate efficiently over a wide Mach number range. As hypersonic combustion tunnels were developed, programs to study the parameters controlling fuel mixing and combustion with single and multiple strut models were conducted using direct connect test techniques. These various tests supported the design of subscale engine test hardware that integrated inlet and combustor technology and allowed the study of the effect of heat release on thrust and combustor/inlet interaction. A number of subscale engine tests have shown predicted performance levels at Mach 4 and 7 simulated flight conditions. A few of the highlights from this research program are summarized.

Northam, G. Burton↗
NASA NTRS1957

Relation of Turbojet and Ramjet Combustion Efficiency to Second-Order Reaction Kinetics and Fundamental Flame Speed

Theoretical studies of the turbojet and ramjet combustion process are summarized and the resulting equations are applied to experimental data obtained from various combustor tests. The theoretical treatment assumes that one step in the over-all chain of processes which constitute jet-engine combustion is sufficiently slow to be the rate-controlling step that determines combustion efficiency.

Childs, J Howard↗
NASA NTRS1984

Research on oxidation by air and tempering of Raney nickel electrocatalysts for the H2 anodes of alkali combustion materials cells

The controlled oxidation in air of Raney nickel electrocatalysts was studied, with special attention paid to the quantitative analysis of nickel hydroxide. The content of the latter was determined through X-ray studies, thermogravimetric measurements, and spectral photometric examinations. The dependence of the content on the drying of activated catalyst is determined. The influence of nickel hydroxide on the electrochemical parameters of the catalyst, such as diffusion polarization, is studied, including a measurement of the exchange current density using the potential drop method. Conservation by oxidation in air with ancillary stabilization of the oxide in an H2 flow at 300 C is explored, including reduction by H2, the influence of tempering time, and structural studies on conserved and stabilized catalyst, long term research on the catalyst, including the influence of aging on the reduced catalyst, and the results of impedance measurements are presented.

Selbach, H. J.↗
NASA NTRS1941

Researches on direct injection in internal-combustion engines

These researches present a solution for reducing the fatigue of the Diesel engine by permitting the preservation of its components and, at the same time, raising its specific horsepower to a par with that of carburetor engines, while maintaining for the Diesel engine its perogative of burning heavy fuel under optimum economical conditions. The feeding of Diesel engines by injection pumps actuated by engine compression achieves the required high speeds of injection readily and permits rigorous control of the combustible charge introduced into each cylinder and of the peak pressure in the resultant cycle.

Tuscher, Jean E↗
NASA NTRS1975

International Symposium on Combustion, 15th, Tokyo, Japan, August 25-31, 1974, Proceedings

Recent theoretical and experimental studies concerned with detonation and pressure wave combustion interaction, fire and explosion research and safety, heterogeneous combustion, flame-flow interactions, kinetics of elementary reactions, pollution control in and by combustion systems, and ignition are presented. Some of the topics covered include critical power density for direct initiation of unconfined gaseous detonations, extinction of laminar diffusion flames for liquid fuels, combustion of bulk titanium in oxygen, flame propagation in small spheres of unconfined and slightly confined flammable mixtures, kinetics of the reaction of nitric oxide with hydrogen, production of chemi-ions and formation of CH and CH2 radicals in methane-oxygen and ethylene-oxygen flames, NOx emission characteristics in two-stage combustion, and spherical ignition of oxyhydrogen behind a reflected shock wave. Individual items are announced in this issue.

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