Search NASA⌕ Search

SEARCH · Search NASA

Results for “COMBUSTION CHAMBER”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 397 records · Page 22

Combustion instability coupling with feed system acoustics

High frequency combustion instability has recently been observed by Rocketdyne in a 40K thrust methane/LOX combustion chamber. The oscillations had frequencies as high as 14,000 Hz with pressure amplitudes in the LOX dome of 500 psi at a chamber pressures of 2,000 psi. At this frequency the wave length associated with a period of oscillation is 2.3 inches in LOX and 1.4 inches in methane. These distances are comparable to the lengths of the injector elements which requires that acoustic waves be considered in the feed systems rather than using lumped parameters as is normally considered for feed system coupled oscillations. To expand the capability of existing models, the Feiler and Heidmann feed system coupled instability model was modified to include acoustic oscillations in the feed system. Similarly the vaporization controlled instability model of Heidmann and Wieber was modified to include flow oscillations that would be produced by feed system coupling. The major elements that control oscillations in a rocket combustion chamber are shown and discussed.

Priem, Richard J.↗

Hydrocarbon-fuel/combustion-chamber-liner materials compatibility

Results of material compatibility experiments using hydrocarbon fuels in contact with copper-based combustion chamber liner materials are presented. Mil-Spec RP-1, n- dodecane, propane, and methane fuels were tested in contact with OFHC, NASA-Z, and ZrCu coppers. Two distinct test methods were employed. Static tests, in which copper coupons were exposed to fuel for long durations at constant temperature and pressure, provided compatibility data in a precisely controlled environment. Dynamic tests, using the Aerojet Carbothermal Test Facility, provided fuel and copper compatibility data under realistic booster engine service conditions. Tests were conducted using very pure grades of each fuel and fuels to which a contaminant, e.g., ethylene or methyl mercaptan, was added to define the role played by fuel impurities. Conclusions are reached as to degradation mechanisms and effects, methods for the elimination of these mechanisms, selection of copper alloy combustion chamber liners, and hydrocarbon fuel purchase specifications.

Gage, Mark L.↗

Coke Deposition and Smoke Formation in Turbojet Engines

In the early development of jet engines, it was occasionally found that excessive amounts of coke or other carbonaceous deposits were formed in the combustion chamber. Sometimes a considerable amount of smoke was noted in the-exhaust gases. Excessive coke deposits may adversely affect jet-engine performance in several ways. The formation of excessive amounts of coke on or just downstream of a fuel nozzle (figs. 116(a) and (b)) changes the fuel-spray pattern and possibly affects combustor life and performance. Similar effects on performance can result from the deposition of coke on primary-air entry ports (fig. 116(c)). Sea-level or altitude starting may be impaired by the deposition of coke on spark-plug electrodes (fig. 116(b)), deposits either grounding the electrodes completely or causing the spark to occur at positions other than the intended gap. For some time it was thought that large deposits of coke in turbojet combustion chambers (fig. 116(a)) might break away and damage turbine blades; however, experience has indicated that for metal blades this problem is insignificant. (Cermet turbine blades may be damaged by loose coke deposits.) Finally, the deposition of coke may cause high-temperature areas, which promote liner warping and cracking (fig. 116(d)) from excessive temperature gradients and variations in thermal-expansion rates. Smoke in the exhaust gases does not generally impair engine performance but may be undesirable from a tactical or a nuisance standpoint. Appendix B of reference 1 and references 2 to 4 present data obtained from full-scale engines operated on test stands and from flight tests that indicate some effects on performance caused by coke deposits and smoke. Some information about the mechanism of coke formation is given in reference 5 and chapter IX. The data indicate that (1) high-boiling fuel residuals and partly polymerized products may be mixed with a large amount of smoke formed in the gas phase to account for the consistency, structure, and chemical composition of the soft coke in the dome and (2) the hard deposits on the liner are similar to petroleum coke and may result from the liquid-phase thermal cracking of the fuel. During the early development period of jet engines, it was noted that the excessive coke deposits and exhaust smoke were generally obtained when fuel-oil-type fuels were used. Engines using gasoline-type fuels were relatively free from the deposits and smoke. These results indicated that some type of quality control would be needed in fuel specifications. Also noted was the effect of engine operating conditions on coke deposition. It is possible that, even with a clean-burning fuel, an excessive amount of coke could be formed at some operating conditions. In this case, combustor redesign could possibly reduce the coke to a tolerable level. This chapter is a summary of the various coke-deposition and exhaust-smoke problems connected- with the turbojet combustor. Included are (1) the effect of coke deposition on combustor life or durability and performance; (2) the effect of combustor design, operating conditions, inlet variables, and fuel characteristics on coke deposition; (3) elimination of coke deposits; (4) the effect of operating conditions and fuel characteristics on formation of exhaust smoke; and (5) various bench test methods proposed for determining and controlling fuel quality.

Hibbard, R. R.↗

Space Shuttle Main Engine (SSME) Options for the Future Shuttle

The main engines for the Future Shuttle will focus on improved safety and operability. Performance enhancements may also be required for vehicle safety purposes to achieve more desirable abort scenarios. This paper discusses the potential improvements that will be considered for implementation into the Future Shuttle. Integrated engine and vehicle health management systems will achieve additional system-level reliability improvements over those currently in development. Advanced instrumentation for detecting leaks, analyzing component wear and degradation, and providing sophisticated operational data will be used for reliable engine control and scheduling maintenance operations. A new nozzle and main combustion chamber (MCC) will reduce failure probability by 50% and allow for higher thrust capability without requiring the entire engine to be redesigned. Turbopump improvements may range from minor component improvements to using 3rd-generation pumps built on the advanced concepts demonstrated by the Integrated Powerhead Development (IPD) program and the Space Launch Initiative (SLI) prototype engines.The main engines for the Future Shuttle will focus on improved safety and operability. Performance enhancements may also be required for vehicle safety purposes to achieve more desirable abort scenarios. This paper discusses the potential improvements that will be considered for implementation into the Future Shuttle. Integrated engine and vehicle health management systems will achieve additional system-level reliability improvements over those currently in development. Advanced instrumentation for detecting leaks, analyzing component wear and degradation, and providing sophisticated operational data will be used for reliable engine control and scheduling maintenance operations. A new nozzle and main combustion chamber (MCC) will reduce failure probability by 50% and allow for higher thrust capability without requiring the entire engine to be redesigned. Turbopump improvements may range from minor component improvements to using 3rd-generation pumps built on the advanced concepts demonstrated by the Integrated Powerhead Development (IPD) program and the Space Launch Initiative (SLI) prototype engines.

Jue, Fred↗

Cu-Cr-Nb-Zr Alloy for Rocket Engines and Other High-Heat- Flux Applications

Rocket-engine main combustion chamber liners are used to contain the burning of fuel and oxidizer and provide a stream of high-velocity gas for propulsion. The liners in engines such as the Space Shuttle Main Engine are regeneratively cooled by flowing fuel, e.g., cryogenic hydrogen, through cooling channels in the back side of the liner. The heat gained by the liner from the flame and compression of the gas in the throat section is transferred to the fuel by the liner. As a result, the liner must either have a very high thermal conductivity or a very high operating temperature. In addition to the large heat flux (>10 MW/sq m), the liners experience a very large thermal gradient, typically more than 500 C over 1 mm. The gradient produces thermally induced stresses and strains that cause low cycle fatigue (LCF). Typically, a liner will experience a strain differential in excess of 1% between the cooling channel and the hot wall. Each time the engine is fired, the liner undergoes an LCF cycle. The number of cycles can be as few as one for an expendable booster engine, to as many as several thousand for a reusable launch vehicle or reaction control system. Finally, the liners undergo creep and a form of mechanical degradation called thermal ratcheting that results in the bowing out of the cooling channel into the combustion chamber, and eventual failure of the liner. GRCop-84, a Cu-Cr-Nb alloy, is generally recognized as the best liner material available at the time of this reporting. The alloy consists of 14% Cr2Nb precipitates in a pure copper matrix. Through experimental work, it has been established that the Zr will not participate in the formation of Laves phase precipitates with Cr and Nb, but will instead react with Cu to form the desired Cu-Zr compounds. It is believed that significant improvements in the mechanical properties of GRCop-84 will be realized by adding Zr. The innovation is a Cu-Cr-Nb-Zr alloy covering the composition range of 0.8 to 8.1 weight percent Cr, 0.7 to 7.2 weight percent Nb, 0.1 to 1.5 weight percent Zr, and balance Cu. The alloy combines two known strengthening mechanisms - dispersion strengthening by Cr2Nb precipitates (GRCop-84), and precipitation strengthening by CuxZr (AMZIRC) - to produce a synergistic increase in the capabilities of the alloy with the goal of achieving properties greater than either of the methods could achieve alone. The anticipated advantages of the alloy are higher strength at temperatures up to 700 C, improved creep strength, and significantly higher LCF lives relative to GRCop-84. The thermal expansion, thermal conductivity, and processing of the alloy are anticipated to remain largely unchanged relative to GRCop-84.

Ellis, David L.↗

Diffusivity in turbulent fluid containing two dominant scales, and compressible shear layer according to a kinetic theory

The solution of the two nonequilibrium-degree kinetic equation was first determined for the effective length scale and turbulence energy for a spatially homogeneous turbulence field with two characteristic length scales, where the source for one family of eddies exists. This solution was applied to the evaluation of the eddy diffusivity in the combustion chamber of an internal combustion engine. The result was compared with another existing solution. This was carried out to demonstrate the feasibility of obtaining an effective length-scale equation within the context of the kinetic theory. A formulation and partial solution of the compressible plane shear layer are also presented.

Chung, P. M.↗

Device for quickly sensing the amount of O2 in a combustion product gas

A sensing device comprising an O2 sensor, a pump, a compressor, and a heater is provided to quickly sense the amount of O2 in a combustion product gas. A sample of the combustion product gas is compressed to a pressure slightly above one atmosphere by the compressor. Next, the heater heats the sample between 800 C and 900 C. Next, the pump causes the sample to be flushed against the electrode located in O2 sensor 6000 to 10,000 times per second. Reference air at approximately one atmosphere is provided to the electrode of O2 sensor. Accordingly, the O2 sensor produces a voltage which is proportional to the amount of oxygen in the combustion product gas. This voltage may be used to control the amount of O2 entering into the combustion chamber which produces the combustion product gas.

Singh, Jag J.↗

Analytical and experimental investigations of the oblique detonation wave engine concept

Wave combustors, which include the oblique detonation wave engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using Computational Fluid Dynamics (CFD) codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being performed in an arc heated hypersonic wind tunnel. Several fuel injection design were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Menees, Gene P.↗

Analytical and experimental investigations of the oblique detonation wave engine concept

Wave combustors, which include the Oblique Detonation Wave Engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which will require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using CFD codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being carried out in an arc heated hypersonic wind tunnel. Several fuel injection designs were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Menees, Gene P.↗

Preliminary Tests on the Vaporization of Fuel Sprays

High-speed motion pictures were taken of fuel sprays injected into the combustion chamber of the N.A.C.A. combustion apparatus. Three fuels, ethyl alcohol, gasoline, and fuel oil, which differed considerably in volatility were tested. By maintaining the engine temperature below that required for ignition the spray could be studied from soon after the start of injection until 130 crank degrees later. The results show that the sprays vaporize appreciably so that it is possible for the ignition in high speed compression-ignition engines to take place from the vapor phase.

Rothrock, A M↗

Liquid Propellant Tracing Impingement Injector

An injector for use with the rocket thruster has a plurality of fuel ports separated from a plurality of oxidizer ports. The oxidizer and fuel ports are paired together directing their respective fluids along a path with radial and tangential components so that the two fluids impinge at a predetermined spaced apart distance from the chamber wall of the combustion chamber at an impingement track. By providing the fuel at a steeper angle relative to the chamber walls than the oxidizer, the fuel can be utilized to provide a fuel rich zone near the chamber walls to assist in cooling the chamber walls during operation.

George D Xenofos↗

Post University On-the-Job Training for Engineers

Our national need for qualified scientists and engineers is greater now than at any other time in our history. Fortunately, we can point with pride to this need as a measure of the impact of science and technology on our way of life. In effect, we have made such rapid strides In advancing established sciences and in opening new technological fields that we have proved the value of the scientist and engineer to society, and, as a-result, have created an expanding demand for their services which we must now attempt to satisfy. This demand we face is also due to the changing skills and high degree of specialization required to perform in these new technological fields. The colleges and universities are doing their part to provide current graduates with a modern technical foundation, but we cannot afford to ignore the thousands of experienced engineers and scientists already employed by private industry and government. As employers, we have an obligation to these men and women to see that they are provided with an understanding of the latest advances that modern technology has to offer; that we develop them in particular specialty areas characteristic of a given field of work; and, equally important, that we assist them in the transition from one field to another as the technological emphasis shifts. Practically all technological industries have experienced and continue to experience rapid changes in their activities. The aerospace business, in particular, has been characterized by extremely rapid, in fact revolutionary, changes during the relatively short period of its existence0 At the National Aeronautics and Space Administration, successor to the National Advisory Committee for Aeronautics, for example, we have encountered the fun impact of a changing science and technology. Indeed, as a research organization, we have undoubtedly contributed, in some measure, to this change. Within the NASAs Lewis Research Center, we have approximately 800 research scientists and engineers who have matured professionally in an environment which is essentially one of continuous learning - an experience which comes close to being a form of post graduate training in itself. This environment, in addition to providing continuous evolutionary changes, has also provided two major revolutions which have made this development picture more complex. We will describe these environmental changes which have occurred at the Lewis Research Center and discuss the various techniques and programs we have employed to provide for the professional development of our staff. The Lewis Research Center has had an Interesting and exciting l8-year history of aerospace propulsion research and development. It began during the early years of World War II as an expansion of the Power Plant Division of the NCA Langley Center with the mission of conducting research required for the development of improved reciprocating engines and to study the associated problems of subsonic propulsion aerodynamics, It was only a few years later, however, that turbojet and ramjet propulsion and supersonic flight research became our main concern. This transition to jet type engines and higher speeds was our first major technological change. The aerodynamics of propellers became the aerodynamics of high speed turbine and compressor blades; the fuel ignition and carbon deposition problems were transferred from a cyclical or Intermittent high compression combustion chamber to a continuous combustion zone within a thin-walled metal shell; aerodynamics problems were thrust into the supersonic range; and high temperature materials began to play an increasingly critical role. Although this transition still required the same basic knowledge and principles as before, the new engine types did involve a different emphasis and variety of consideration not generally familiar to our scientists and engineers.

Manganiello, Eugene J.↗

Rapid Analysis and Manufacturing Propulsion Technology (RAMPT)

The RAMPT project is maturing novel design and manufacturing technologies to increase scale, significantly reduce cost, and improve performance for regeneratively-cooled thrust chamber assemblies, specifically the combustion chamber and nozzle for government and industry programs.

Fikes, John C.↗

Combustion Studies of MMA/GOx for a Hybrid Rocket Motor

Poly(methyl methacrylate) (PMMA) is the synthetic polymer of methyl methacrylate (MMA), used in a multitude of applications, including as a solid fuel in hybrid rockets. PMMA undergoes pyrolysis into gaseous MMA (C5H8O2), which then undergoes combustion into carbon dioxide, water, carbon monoxide and low-molecular-weight compounds, including formaldehyde. Hybrid rocket motors as seen in the schematic in Fig. 1, contains the oxidizer in a pressurized tank, which is released through a port in the solid fuel grain in the combustion chamber, after which the combustion products are exhausted through a converging diverging nozzle.

Dhandapani, Chandru↗