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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 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

SSME thrust chamber modeling with Navier Stokes equations

The capability of predicting two-dimensional, compressible and reacting flow in the combustion chamber and nozzle of the Space Shuttle Main Engine (SSME) is demonstrated. A nonorthogonal body fitted coordinate system has been used to represent the combustor and nozzle geometry. The Navier-Stokes equations are solved for the entire thrust chamber with the k-epsilon turbulence model accounting for compressibility and large pressure gradients effects. Results of the computational test cases reveal all expected features of the transonic nozzle flows including location of sonic line, internal shock and boundary layer build-up. Calculated performance parameters such as thrust, flow rate, and specific impulse are also in reasonble agreement with available data. The results show promising potential of solving full Navier-Stokes equations with heat transfer and two-phase combustion in truly comprehensive modeling of rocket engines.

Przekwas, A. J.↗
DOE OSTI2025

Exploring the Effects of Varying Pre-Chamber Geometry in a Heavy-Duty Natural Gas Optical Engine under Dilution Conditions

Pre-chamber combustion is an advanced ignition strategy that has been shown to enhance spark ignition (SI) combustion stability in natural gas (NG) engines by providing distributed ignition sites from turbulent jets and enhancing main-chamber turbulence. Pre-chamber combustion has been proven especially advantageous compared to SI in ultra-lean and dilute operating conditions. This work involves experimental investigation of the effects of varying passive pre-chamber nozzle configuration on pre-chamber and main chamber combustion under simulated exhaust gas recirculation (EGR) dilution (0 and 20%) conditions in a heavy-duty, single-cylinder, optically accessible NG engine at stoichiometric fuel-air ratio. Pre-chamber nozzle configurations include four pre-chambers with constant nozzle area to pre-chamber volume ratio (A/V) with different nozzle sizes and orientations and one configuration with larger nozzles. The optical engine is operated in a skip-fire sequence consisting of 18 motored cycles followed by two consecutive fired cycles to elucidate the effect of combustion residuals (internal EGR) on combustion evolution. Pressure-based diagnostics are used to monitor pre-chamber and subsequent main chamber combustion, and optical diagnostics include high-speed OH* chemiluminescence to visualize the development of pre-chamber jets and the resulting ignition of the main chamber charge. Heat release analysis of the in-cylinder pressure data indicates that a faster pre-chamber pressure rise does not always translate into faster main-chamber combustion. The pre-chamber with the smallest nozzle diameter produced the highest pre-chamber pressure rise and fastest combustion under non-diluted conditions. However, dilution delays the main chamber ignition for smaller nozzles despite a comparable rise in pre-chamber pressure compared to configurations with larger nozzles. This effect is more pronounced for cycles with in-cylinder combustion residuals in addition to external dilution. Additionally, it was observed that pre-chambers with swirling nozzles have a faster pressure rise in the pre-chamber and main chamber under dilute conditions. Optical diagnostics suggest that the main reason for the delay between the pre-chamber pressure rise and main-chamber combustion lies in jet quenching and delayed re-ignition, which can even lead to misfire if jets emitted from small nozzles combined with dilution fail to re-ignite.

Dhotre, Akash [University Of Minnesota-Twin Cities↗
NASA NTRS1984

Transient flow combustion

Non-steady combustion problems can result from engine sources such as accelerations, decelerations, nozzle adjustments, augmentor ignition, and air perturbations into and out of the compressor. Also non-steady combustion can be generated internally from combustion instability or self-induced oscillations. A premixed-prevaporized combustor would be particularly sensitive to flow transients because of its susceptability to flashback-autoignition and blowout. An experimental program, the Transient Flow Combustion Study is in progress to study the effects of air and fuel flow transients on a premixed-prevaporized combustor. Preliminary tests performed at an inlet air temperature of 600 K, a reference velocity of 30 m/s, and a pressure of 700 kPa. The airflow was reduced to 1/3 of its original value in a 40 ms ramp before flashback occurred. Ramping the airflow up has shown that blowout is more sensitive than flashback to flow transients. Blowout occurred with a 25 percent increase in airflow (at a constant fuel-air ratio) in a 20 ms ramp. Combustion resonance was found at some conditions and may be important in determining the effects of flow transients.

Tacina, R. R.↗
NASA NTRS1984

Combustion overview

The objective of this effort is to develop improved analytical models of the internal combustor flow field and liner heat transfer as a means to shorten combustor development time and increase turbine engine hot section life. A four-element approach was selected to meet this objective. First, existing models were utilized to determine their deficiencies. Supporting research was then commenced to improve the exisiting models. While the research effort is in progress, the models are being refined to improve numerics and numerical diffusion. And lastly, the research results and improved numerics will be integrated into existing models.

Schultz, D. F.↗
NASA NTRS2006

Fuel-Air Mixing and Combustion in Scramjets

At flight speeds, the residence time for atmospheric air ingested into a scramjet inlet and exiting from the engine nozzle is on the order of a millisecond. Therefore, fuel injected into the air must efficiently mix within tens of microseconds and react to release its energy in the combustor. The overall combustion process should be mixing controlled to provide a stable operating environment; in reality, however, combustion in the upstream portion of the combustor, particularly at higher Mach numbers, is kinetically controlled where ignition delay times are on the same order as the fluid scale. Both mixing and combustion time scales must be considered in a detailed study of mixing and reaction in a scramjet to understand the flow processes and to ultimately achieve a successful design. Although the geometric configuration of a scramjet is relatively simple compared to a turbomachinery design, the flow physics associated with the simultaneous injection of fuel from multiple injector configurations, and the mixing and combustion of that fuel downstream of the injectors is still quite complex. For this reason, many researchers have considered the more tractable problem of a spatially developing, primarily supersonic, chemically reacting mixing layer or jet that relaxes only the complexities introduced by engine geometry. All of the difficulties introduced by the fluid mechanics, combustion chemistry, and interactions between these phenomena can be retained in the reacting mixing layer, making it an ideal problem for the detailed study of supersonic reacting flow in a scramjet. With a good understanding of the physics of the scramjet internal flowfield, the designer can then return to the actual scramjet geometry with this knowledge and apply engineering design tools that more properly account for the complex physics. This approach will guide the discussion in the remainder of this section.

Drummond, J. Philip↗
NASA NTRS1974

Hypersonic ramjet experiment project. Phase 1: Computer program description, ramjet and scramjet cycle performance

A computer program was developed to describe the performance of ramjet and scramjet cycles. The program performs one dimensional calculations of the equilibrium, real-gas internal flow properties of the engine. The program can be used for the following: (1) preliminary design calculation and (2) design analysis of internal flow properties corresponding to stipulated flow areas. Only the combustion of hydrogen in air is considered in this case.

Jackson, R. J.↗
NASA NTRS1995

Affect of Brush Seals on Wave Rotor Performance Assessed

The NASA Lewis Research Center's experimental and theoretical research shows that wave rotor topping can significantly enhance gas turbine engine performance levels. Engine-specific fuel consumption and specific power are potentially enhanced by 15 and 20 percent, respectively, in small (e.g., 400 to 700 hp) and intermediate (e.g., 3000 to 5000 hp) turboshaft engines. Furthermore, there is potential for a 3- to 6-percent specific fuel consumption enhancement in large (e.g., 80,000 to 100,000 lbf) turbofan engines. This wave-rotor-enhanced engine performance is accomplished within current material-limited temperature constraints. The completed first phase of experimental testing involved a three-port wave rotor cycle in which medium total pressure inlet air was divided into two outlet streams, one of higher total pressure and one of lower total pressure. The experiment successfully provided the data needed to characterize viscous, partial admission, and leakage loss mechanisms. Statistical analysis indicated that wave rotor product efficiency decreases linearly with the rotor to end-wall gap, the square of the friction factor, and the square of the passage of nondimensional opening time. Brush seals were installed to further minimize rotor passage-to-cavity leakage. The graph shows the effect of brush seals on wave rotor product efficiency. For the second-phase experiment, which involves a four-port wave rotor cycle in which heat is added to the Brayton cycle in an external burner, a one-dimensional design/analysis code is used in conjunction with a wave rotor performance optimization scheme and a two-dimensional Navier-Stokes code. The purpose of the four-port experiment is to demonstrate and validate the numerically predicted four-port pressure ratio versus temperature ratio at pressures and temperatures lower than those that would be encountered in a future wave rotor/demonstrator engine test. Lewis and the Allison Engine Company are collaborating to investigate wave rotor integration in an existing turboshaft engine. Recent theoretical efforts include simulating wave rotor dynamics (e.g., startup and load-change transient analysis), modifying the one-dimensional wave rotor code to simulate combustion internal to the wave rotor, and developing an analytical wave rotor design/analysis tool based on macroscopic balances for parametric wave rotor/engine analysis.

Source record↗
NASA NTRS2008

NASA Glenn High Pressure Low NOx Emissions Research

In collaboration with U.S. aircraft engine companies, NASA Glenn Research Center has contributed to the advancement of low emissions combustion systems. For the High Speed Research Program (HSR), a 90% reduction in nitrogen oxides (NOx) emissions (relative to the then-current state of the art) has been demonstrated in sector rig testing at General Electric Aircraft Engines (GEAE). For the Advanced Subsonic Technology Program (AST), a 50% reduction in NOx emissions relative to the 1996 International Civil Aviation Organization (ICAO) standards has been demonstrated in sector rigs at both GEAE and Pratt & Whitney (P&W). During the Ultra Efficient Engine Technology Program (UEET), a 70% reduction in NOx emissions, relative to the 1996 ICAO standards, was achieved in sector rig testing at Glenn in the world class Advanced Subsonic Combustion Rig (ASCR) and at contractor facilities. Low NOx combustor development continues under the Fundamental Aeronautics Program. To achieve these reductions, experimental and analytical research has been conducted to advance the understanding of emissions formation in combustion processes. Lean direct injection (LDI) concept development uses advanced laser-based non-intrusive diagnostics and analytical work to complement the emissions measurements and to provide guidance for concept improvement. This paper describes emissions results from flametube tests of a 9-injection-point LDI fuel/air mixer tested at inlet pressures up to 5500 kPa. Sample results from CFD and laser diagnostics are also discussed.

Tacina, Kathleen M.↗
NASA NTRS2007

High Pressure Low NOx Emissions Research: Recent Progress at NASA Glenn Research Center

In collaboration with U.S. aircraft engine companies, NASA Glenn Research Center has contributed to the advancement of low emissions combustion systems. For the High Speed Research Program (HSR), a 90% reduction in nitrogen oxides (NOx) emissions (relative to the then-current state of the art) has been demonstrated in sector rig testing at General Electric Aircraft Engines (GEAE). For the Advanced Subsonic Technology Program (AST), a 50% reduction in NOx emissions relative to the 1996 International Civil Aviation Organization (ICAO) standards has been at demonstrated in sector rigs at both GEAE and Pratt & Whitney (P&W). During the Ultra Efficient Engine Technology Program (UEET), a 70% reduction in NOx emissions, relative to the 1996 ICAO standards, was achieved in sector rig testing at Glenn in the world class Advanced Subsonic Combustion Rig (ASCR) and at contractor facilities. Low NOx combustor development continues under the Fundamental Aeronautics Program. To achieve these reductions, experimental and analytical research has been conducted to advance the understanding of emissions formation in combustion processes. Lean direct injection (LDI) concept development uses advanced laser-based non-intrusive diagnostics and analytical work to complement the emissions measurements and to provide guidance for concept improvement. This paper describes emissions results from flametube tests of a 9- injection-point LDI fuel/air mixer tested at inlet pressures up to 5500 kPa. Sample results from CFD and laser diagnostics are also discussed.

Chi-Ming, Lee↗
NASA NTRS1986

Numerical analysis of SSME preburner injector atomization and combustion processes

The coaxial spray injection and combustion flowfields of a Space Shuttle Main Engine preburner injector element have been analyzed using a three-phase numerical code. The processes of atomization, evaporation, secondary droplet breakup, and multispecies chemistry, as well as turbulent diffusion, are included. The model produced realistic pictures of the complex internal flowfield, including liquid jet length, spray shape, flame-zone size and characteristics, and predicted temperatures that seem to be in agreement with test data envelopes. It predicted an external group combustion type of flame. Salient combustion and mixing features are discussed and sources of uncertainty are pointed out for future studies.

Liang, P. Y.↗
NASA NTRS1993

Fire safety practices in the Shuttle and the Space Station Freedom

The Shuttle reinforces its policy of fire-preventive measures with onboard smoke detectors and Halon 1301 fire extinguishers. The forthcoming Space Station Freedom will have expanded fire protection with photoelectric smoke detectors, radiation flame detectors, and both fixed and portable carbon dioxide fire extinguishers. Many design and operational issues remain to be resolved for Freedom. In particular, the fire-suppression designs must consider the problems of gas leakage in toxic concentrations, alternative systems for single-failure redundancy, and commonality with the corresponding systems of the Freedom international partners. While physical and engineering requirements remain the primary driving forces for spacecraft fire-safety technology, there are, nevertheless, needs and opportunities for the application of microgravity combustion knowledge to improve and optimize the fire-protective systems.

Friedman, Robert↗
NASA NTRS1980

Preparing aircraft propulsion for a new era in energy and the environment

Improving fuel efficiency, new sources of jet fuel, and noise and emission control are subjects of NASA's aeronautics program. Projects aimed at attaining a 5% fuel savings for existing engines and a 13-22% savings for the next generation of turbofan engines using advanced components, and establishing a basis for turboprop-powered commercial air transports with 30-40% savings over conventional turbofan aircraft at comparable speeds and altitudes, are discussed. Fuel sources are considered in terms of reduced hydrogen and higher aromatic contents and resultant higher liner temperatures, and attention is given to lean burning, improved fuel atomization, higher freezing-point fuel, and deriving jet fuel from shale oil or coal. Noise sources including the fan, turbine, combustion process, and flow over internal struts, and attenuation using acoustic treatment, are discussed, while near-term reduction of polluting gaseous emissions at both low and high power, and far-term defining of the minimum gaseous-pollutant levels possible from turbine engines are also under study.

Stewart, W. L.↗
NASA NTRS1991

Symposium (International) on Combustion, 23rd, Universite d'Orleans, France, July 22-27, 1990, Proceedings

The present symposium on combustion discusses reaction kinetics, NO(x) kinetics, premixed, diffusion, and nonsteady flames, turbulent combustion, hazardous waste, fluidized bed combustion, coal boilers and furnaces, engines, heterogeneous kinetics, heterogeneous, droplet, and microgravity combustion, and high-temperature synthesis. Attention is given to reactions of biphenyl, methylnaphthalenes, and phenanthrene with atomic oxygen in the gas phase, the oxidation of ortho-xylene, the effects of water on combustion kinetics at high pressure, and the formation and measurement of N2O in combustion systems. Topics addressed include large ions in premixed benzene-oxygen flames, the structure and kinetics of CH4/N2O flames, the propagation of unsteady hydrogen premixed flames near flammability limits, and the basic structure of lean propane flames. Also considered are OH measurements of piloted diffusion flames of nitrogen-diluted methane near extinction, waste combustion, preferential oxidation of carbon surfaces, and reburning mechanisms in a pulverized coal combustor.

Source record↗
NASA NTRS1986

Monitoring Temperatures Indirectly in Cooled Combustors

Noninvasive monitoring of temperatures on inner surface of combustion liner of cooled combustor increases liner life, thereby increasing combustor reliability and performance. Technique used in furnaces, reactors, jet engines, rocket engines, stationary turbines, combustors, and heat exchangers. Growth of internal overheating streaks monitored using output from array of thermocouples on outside wall of combustor. Computer analysis of output indicates temperature pattern on combustor lining. Use of data enables timely liner maintenance or overhaul.

Wagner, W.↗
NASA NTRS2001

Droplet Combustion in a Slow Convective Flow

The objective of the present flight experiment definition study is to investigate the effects of slow forced convective flows on the dynamics of isolated single droplet combustion and is designed to complement the quiescent, microgravity droplet combustion experiments (DCE-1 and DCE-2) of Williams and Dryer. The fuels selected for this study are the same as those of DCE, namely, a sooting alkane fuel (heptane) and a non-sooting alcohol (methanol), and imposed flow rates are chosen between 0 and 20 cm/s with varying ambient oxygen concentrations and pressures. Within this velocity range, both accelerating and decelerating flow effects will also be investigated. Two different approaches to generate the forced flow are currently under development in ground-based facilities; the first is a flow tunnel concept where the forced flow is imposed against a stationary droplet, and in the second a tethered droplet is translated at a specified velocity in a quiescent ambient medium. Depending upon the engineering feasibility a selection will be made between these two approaches so that the experiment can be accommodated in the Multiple Droplet Combustion Apparatus (MDCA) currently being designed for the International Space Station. Recently, we have finished designing and fabricating the experimental rigs using both the above mentioned concepts. The flow tunnel concept is implemented in a 2.2 second drop tower rig. Preliminary experiments have been carried out using heptane and methanol in air at atmospheric pressure. The translating droplet apparatus is scheduled to be tested in the 5 second drop facility in the near future. This report presents some of the experimental results obtained for heptane.

Nayagam, V.↗
DOE OSTI2025

Enabling Production of Low Carbon Emissions Steel through CO 2 Capture from Blast Furnace Gases at Cleveland Cliffs’ 5 mtpa Steel Plant at Burns Harbor, Indiana

Dastur International Inc. has prepared a DOE-funded Pre-front-end engineering design (Pre-FEED) study to capture up to 2.8 mtpa of CO 2 from the available Blast Furnace (BF) gases at the Burns Harbor steel plant operated by Cleveland Cliffs in Burns Harbor, Indiana. The project consists of an amine solvent based pre-combustion carbon capture system, along with a unique BF gas conditioning process which significantly optimizes the project design and architecture in terms of higher volumes (2.8 mtpa vis-à-vis 1.6 mtpa without conditioning) & concentration (32 vol% in conditioned gas vs. 22% in raw BF gas) of CO 2 , which can be captured using a single train, resulting in reduced size and capital cost of the CO 2 capture island, and improved overall economics on a $\$$/tCO 2 captured basis.

42 ENGINEERING↗
NASA NTRS2019

Additive Manufacturing Development and Hot-Fire Testing of Liquid Rocket Channel Wall Nozzles Using Blown Powder Directed Energy Deposition Inconel 625 and JBK-75 Alloys

Additive manufacturing (AM) is being investigated at NASA and across much of the rocket propulsion industry as an alternate fabrication technique to create complex geometries for liquid engine components that offers schedule and cost saving opportunities. The geometries that can be created using AM offer a significant advantage over traditional techniques. Internal complexities, such as internal coolant channels for combustion chambers and nozzles that would typically require several operations to manufacture traditionally can be fabricated in one process. Additionally, the coolant channels are closed out as a part of the AM build process, eliminating the complexities of a traditional process like brazing or plating. The primary additive manufacturing technique that has been evaluated is powder bed fusion (PBF), or selective laser melting (SLM), but there is a scale limitation for this technique. There are several alternate additive manufacturing techniques that are being investigated for large-scale nozzles and chambers including directed energy deposition (DED) processes. A significant advantage of the DED processes is the ability to adapt to a robotic or gantry CNC system with a localized purge or purge chamber, allowing unlimited build volume. This paper will discuss the development and hot-fire testing of channel-cooled nozzles fabricated utilizing one form of DED called blown powder deposition. This initial development work using blown powder DED is being explored to form the entire channel wall nozzle with integral coolant channels within a single AM build. Much of this development is focused on the design and DED-fabrication of complex and thin-walled features and on characterization of the materials properties produced with this techniques in order to evolve this process. Subscale nozzles were fabricated using this DED technique and hot-fire tested in Liquid Oxygen/Hydrogen (LOX/GH2) and LOX/Kerosene (LOX/RP-1) environments accumulating significant development time and cycles. The initial materials that were evaluated during this testing were high-strength nickel-based Inconel 625 and JBK-75. Further process development is being completed to increase the scale of this technology for large-scale nozzles. This paper will summarize the general design considerations for DED, specific channel-cooled nozzle design, manufacturing process development, property development, initial hot-fire testing and future developments to mature this technology for regeneratively-cooled nozzles. An overview of future development at NASA will also be discussed.

Gradl, Paul R.↗
NASA NTRS1975

Noise addendum experimental clean combustor program, phase 1

The development of advanced CTOL aircraft engines with reduced exhaust emissions is discussed. Combustor noise information provided during the basic emissions program and used to advantage in securing reduced levels of combustion noise is included. Results are presented of internal pressure transducer measurements made during the scheduled emissions test program on ten configurations involving variations of three basic combustor designs.

Sofrin, T. G.↗