Search NASA⌕ Search

SEARCH · Search NASA

Results for “advanced combustion engine”

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 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
DOE OSTI2021

Gasoline fuels properties for multi-mode operation – Observations in a GDI and the CFR engine

The combustion behavior of five full boiling range RON98 gasoline blends was evaluated for multi-mode operation in a GDI and the CFR octane rating engine. The GDI engine tests were conducted with stoichiometric air-fuel ratio in spark-ignition (SI), and with air-diluted homogeneous charge compression ignition (HCCI) mode. In the CFR engine, the knocking combustion was analyzed under standard RON testing conditions at both peak knocking lambda and stoichiometric air-fuel ratios, whereas compression ignited operation was characterized by utilizing the HCCI number test protocol. Disparate knock limited SI and HCCI combustion behavior was observed for the test fuels, despite four of the fuels having the same RON and octane sensitivity. It was found that knock-limited combustion phasing in the GDI engine did not align well with the RON. However, a detailed comparison of knock-limited SI operation in the GDI and CFR engine revealed that a more appropriate effective RON based on a common knock intensity metric (MAPO) and stoichiometric air-fuel ratio resulted in comparable knock characterization between the two engine platforms. Furthermore, the critical intake air temperature and the critical compression ratio were proposed to characterize knock-limited SI operation, while the minimum intake air heating and compression ratio were used to define a fuel's autoignition propensity for compression ignition operation in the GDI and CFR engine, respectively. Finally, each fuel's characteristic compression ratio needed to obtain knock-limited SI (KLSI) and HCCI operation was used to calculate an effective multi-mode octane number (MM-ON) based on the primary reference fuel blends.

33 ADVANCED PROPULSION SYSTEMS↗
DOE OSTI2025

Rationally Designed Lithium Ion Batteries Towards Displacing Internal Combustion Engines

Group14 Technologies is leading a word-class team composed of Cabot Corporation, Silatronix, Arkema, Pacific Northwest National Laboratories, and Farasis Energy to research, fabricate, test and demonstrate lithium-ion batteries (LIB) implementing ≥30% silicon-carbon (Si-C) composite anode material achieving aggressive next generation automotive energy targets.

25 ENERGY STORAGE↗
DOE OSTI2021

Comparative Study for Pressure Gain Combustion-Gas Turbine System Performance in NGCC Configurations

The Department of Energy (DOE) has a goal to increase natural gas combined cycle (NGCC) plant efficiency to greater than 70 percent on a lower heating value (LHV) basis. Improvement in gas turbine (GT) efficiency is key achieving this goal. While many of the traditional improvement methods (higher firing temperature, advanced materials, improved cooling methods, etc.) may increase current GT and NGCC efficiencies incrementally, combining these traditional advancements with breakthrough technologies such as pressure gain combustion (PGC) may provide significant system efficiency gains. PGC is a combustion process that results in a net pressure increase across the combustor (up to pressure ratio of 2.2) rather than the traditional (<5%) pressure drop. There are multiple PGC technologies being developed that create the pressure rise phenomena, including resonant pulse combustors (RPCs), pulse detonation engines (PDEs), wave rotor combustor (WRC) engines and rotating detonation engines (RDEs). In this study, scoping analyses of PGC concepts applied to NGCC power plants were performed by combining simplified analytical models for the RPC, PDE, WRC, and RDE types of pressure gain processes with a GT model to evaluate impacts of using each of the PGC technologies on key GT performance parameters. Sensitivity analyses were also performed. Ultimately, the PGC technologies were found to have the potential to increase simple cycle efficiency by 4–6 percent (absolute) and NGCC efficiency by 2–3 percent (absolute) based on the results of the current study.

03 NATURAL GAS↗
NASA NTRS2010

Hot-Fire Testing of 100 LB(sub F) LOX/LCH4 Reaction Control Engine at Altitude Conditions

Liquid oxygen/liquid methane (LO2/LCH4 ) has recently been viewed as a potential green propulsion system for both the Altair ascent main engine (AME) and reaction control system (RCS). The Propulsion and Cryogenic Advanced Development Project (PCAD) has been tasked by NASA to develop these green propellant systems to enable safe and cost effective exploration missions. However, experience with LO2/LCH4 as a propellant combination is limited, so testing of these systems is critical to demonstrating reliable ignition and performance. A test program of a 100 lb f reaction control engine (RCE) is underway at the Altitude Combustion Stand (ACS) of the NASA Glenn Research Center, with a focus on conducting tests at altitude conditions. These tests include a unique propellant conditioning feed system (PCFS) which allows for the inlet conditions of the propellant to be varied to test warm to subcooled liquid propellant temperatures. Engine performance, including thrust, c* and vacuum specific impulse (I(sub sp,vac)) will be presented as a function of propellant temperature conditions. In general, the engine performed as expected, with higher performance at warmer propellant temperatures but better efficiency at lower propellant temperatures. Mixture ratio effects were inconclusive within the uncertainty bands of data, but qualitatively showed higher performance at lower ratios.

Marshall, William M.↗
NASA NTRS2020

Advancing Development of Environmental Barrier Coatings Resistant to Attack by Molten Calcium-Magnesium-Aluminosilicate (CMAS)

Ceramic matrix composites (CMCs) are a leading material system to replace metal-based parts in the hot-section of air-breathing turbine engines to improve fuel efficiency in aircraft engines. CMCs have higher temperature capabilities and lower density compared with traditional metallic structural materials. However, silicon-based CMCs are susceptible to oxidation in the harsh combustion environment encountered in turbine engines. Consequently, environmental barrier coatings (EBCs) are being developed to protect CMC components to improve durability and extend service life of CMCs. Sand, volcanic ash and other particulate debris, which are generally comprised of calcium-magnesium-aluminosilicate (CMAS) and other trace oxides, are routinely ingested by aircraft engines. At temperatures above 1200°C, CMAS particulates melt. Near target operating temperatures (~1500°C) of future CMC-based aircraft engines, molten CMAS behaves like a viscous melt that can infiltrate and chemically interact with protective coatings. These interactions can cause premature failure of the EBC system and ultimately the overall CMC engine component. Degradation of candidate EBC materials by molten CMAS will be presented with a focus on recent work, as well as methods of evaluating the complex high-temperature materials interactions, underway at NASA Glenn Research Center.

Wiesner, Valerie L.↗
NASA NTRS1994

Theoretical determination of chemical rate constants using novel time-dependent methods

The work completed within the grant period 10/1/91 through 12/31/93 falls primarily in the area of reaction dynamics using both quantum and classical mechanical methodologies. Essentially four projects have been completed and have been or are in preparation of being published. The majority of time was spent in the determination of reaction rate coefficients in the area of hydrocarbon fuel combustion reactions which are relevant to NASA's High Speed Research Program (HSRP). These reaction coefficients are important in the design of novel jet engines with low NOx emissions, which through a series of catalytic reactions contribute to the deterioration of the earth's ozone layer. A second area of research studied concerned the control of chemical reactivity using ultrashort (femtosecond) laser pulses. Recent advances in pulsed-laser technologies have opened up a vast new field to be investigated both experimentally and theoretically. The photodissociation of molecules adsorbed on surfaces using novel time-independent quantum mechanical methods was a third project. And finally, using state-of-the-art, high level ab initio electronic structure methods in conjunction with accurate quantum dynamical methods, the rovibrational energy levels of a triatomic molecule with two nonhydrogen atoms (HCN) were calculated to unprecedented levels of agreement between theory and experiment.

Dateo, Christopher E.↗
NASA NTRS1998

Parametric Study Conducted of Rocket- Based, Combined-Cycle Nozzles

Having reached the end of the 20th century, our society is quite familiar with the many benefits of recycling and reusing the products of civilization. The high-technology world of aerospace vehicle design is no exception. Because of the many potential economic benefits of reusable launch vehicles, NASA is aggressively pursuing this technology on several fronts. One of the most promising technologies receiving renewed attention is Rocket-Based, Combined-Cycle (RBCC) propulsion. This propulsion method combines many of the efficiencies of high-performance jet aircraft with the power and high-altitude capability of rocket engines. The goal of the present work at the NASA Lewis Research Center is to further understand the complex fluid physics within RBCC engines that govern system performance. This work is being performed in support of NASA's Advanced Reusable Technologies program. A robust RBCC engine design optimization demands further investigation of the subsystem performance of the engine's complex propulsion cycles. The RBCC propulsion system under consideration at Lewis is defined by four modes of operation in a singlestage- to-orbit configuration. In the first mode, the engine functions as a rocket-driven ejector. When the rocket engine is switched off, subsonic combustion (mode 2) is present in the ramjet mode. As the vehicle continues to accelerate, supersonic combustion (mode 3) occurs in the ramjet mode. Finally, as the edge of the atmosphere is approached and the engine inlet is closed off, the rocket is reignited and the final accent to orbit is undertaken in an all-rocket mode (mode 4). The performance of this fourth and final mode is the subject of this present study. Performance is being monitored in terms of the amount of thrust generated from a given amount of propellant.

Steffen, Christopher J., Jr.↗
NASA NTRS2023

Passive Rocket Diffuser Testing: Reacting Flow Performance of Advanced Configurations

Upper stage rocket engines are optimized for operation at the edge of space. Ground testing often requires simulation of high-altitude environments to prevent damage from off-design operation. Passive supersonic diffusers are an effective means of establishing the partial vacuum needed around the nozzle exit. The low cost of diffusers relative to active vacuum sources (ejectors, pumps, etc.) provides a perpetual incentive to improve their performance and expand the envelope of passively testable engines. Second-throat diffusers have been the de facto standard for rocket testing since the 1950s, and numerous studies have focused on parametric optimization of the topology. Yet, even the best second-throat designs may require more driving pressure than an upper stage engine can provide. To overcome conventional limits, engineers at NASA’s Stennis Space Center (SSC) adapted an obscure cold-flow diffuser topology to the extremes of immersion in supersonic combustion products. Aerodynamic experimentation was conducted at SSC’s E-3 test stand to quantify the performance of prototypical hot-fire spike diffuser hardware and enable direct comparison to prior states of the art. A LO X /GH 2 thruster was contoured to approximate the regeneratively-cooled nozzle of an upper-stage engine and fired into 28 downstream diffuser configurations: 16 second-throat, 2 centerbody, and 10 spike. Temperature and pressure were recorded along the test article walls for a range of chamber pressures spanning 3.5 - 5.0 MPa. The second-throat configurations with the lowest start and unstart pressure ratios were shown to outperform equivalent systems in literature and taken as baselines for evaluation of advanced diffuser performance. Compared to these exemplars, centerbody diffusers decreased start pressure ratio by 11% but increased unstart pressure ratio by 15%. Spike diffusers reduced start and unstart pressure ratios by 24% and 9%, respectively.

Rocket↗
NASA NTRS1984

Ceramic composite liner material for gas turbine combustors

Advanced commercial and military gas turbine engines may operate at combustor outlet temperatures in excess of 1920 K (3000 F). At these temperatures combustors liners experience extreme convective and radiative heat fluxes. The ability of a plasma sprayed ceramic coating to reduce liner metal temperature has been recognized. However, the brittleness of the ceramic layer and the difference in thermal expansion with the metal substrate has caused cracking, spalling and some separation of the ceramic coating. Research directed at turbine tip seals (or shrouds) has shown the advantage of applying the ceramic to a compliant metal pad. This paper discusses recent studies of applying ceramics to combustor liners in which yttria stabilized zirconia plasma sprayed on compliant metal substrates which were exposed to near stoichiometric combustion, presents performance and durability results, and describes a conceptual design for an advanced, small gas turbine combustor. Test specimens were convectively cooled or convective-transpiration cooled and were evaluated in a 10 cm square flame tube combustor at inlet air temperatures of 533 K (500 F) and at a pressure of 0.5 MPa (75 psia). The ceramics were exposed to flame temperatures in excess of 2000 K (3320 F). Results appear very promising with all 30 specimens surviving a screening test and one of two specimens surviving a cyclic durability test.

Ercegovic, D. B.↗
NASA NTRS2014

Designing Liquid Rocket Engine Injectors for Performance, Stability, and Cost

NASA is using computational fluid dynamics (CFD) simulations to lower the cost and increase the performance and stability of fuel injector designs for next-generation liquid rocket engines (LREs). The Loci-STREAM CFD code is used to simulate the complex combustion processes inside the LRE. These analyses enable efficient evaluation of the performance and stability characteristics of injector design concepts, while decreasing reliance on the costly test-fail-fix cycle of traditional design approaches. These injector simulations were recently employed as a key part of the design process for an Advanced Booster concept for NASA's heavy-lift Space Launch System (SLS).

Westra, Douglas G.↗
DOE OSTI2020

Operation of internal combustion engine with improved fuel efficiency

A process of operating a spark-ignited internal combustion engine (SI-ICE) with improved fuel efficiency and reduced emissions including under steady state and under lean-operating conditions at high overall air to fuel (AFR) ratios. A first supply of high octane hydrocarbon fuel, such as gasoline or natural gas, and a first supply of oxidant are fed to a fuel reformer to produce a gaseous reformate with a reforming efficiency of greater than 75 percent relative to equilibrium. The gaseous reformate is mixed with a second supply of oxidant, after which the resulting reformate blended oxidant is fed with a second supply of high octane hydrocarbon fuel to the SI-ICE for combustion. Steady state fuel efficiency is improved by more than 3 percent, when the reformate comprises from greater than about 1 to less than about 18 percent of the total volume of reformate blended oxidant fed to the engine.

33 ADVANCED PROPULSION SYSTEMS↗
DOE OSTI2022

Editorial: Particulate Matter Emissions From Conventional and Reformulated Fuel Combustion: Advances in Experiments and Simulations

Particulate matter (PM or soot) emissions generated by combustion of conventional and reformulated fuels adversely impact the environment and human health (Bond et al., 2013; Landrigan et al., 2018; World Health Organization (WHO), 2016), generating public awareness and motivating efforts towards the mitigation of their harmful effects. Soot formation is one of the most complex phenomena in combustion, involving interactions between combustion chemistry, fluid mechanics, mass/heat transport, and particle dynamics, spanning different spatial and temporal scales. A fundamental understanding of soot formation process is thus necessary to achieve a strong reduction of PM emissions and design cleaner and more efficient combustion systems. This has motivated a longstanding and ongoing research activity aimed at improving our understanding of the physical and chemical processes involved in soot formation, well reviewed in these recent works (Wang and Chung, 2019; Michelsen et al., 2020; Martin et al., 2022). Despite the wide interest, the transition from gas-phase molecules to incipient soot particles is still elusive and the successive particle growth and oxidation processes are far from being fully understood, especially in conditions relevant to real-world applications. Recent advances in combustion PM emission diagnostic and computational capabilities helped in improving the predictability of fundamental chemical and aerosol models for practical applications, thus tackling some of the above-mentioned challenges. The aim of this research topic is to display the ongoing research efforts in addressing the existing gaps on particulate formation from various fuel sources and in conditions typical of practical combustion applications (e.g. flames, engines, pool fires), through both experimental and numerical approaches.

42 ENGINEERING↗
DOE OSTI2020

Recent Progress in the Engineering of Polymeric Membranes for CO2 Capture from Flue Gas

CO2 capture from coal- or natural gas-derived flue gas has been widely considered as the next opportunity for the large-scale deployment of gas separation membranes. Despite the tremendous progress made in the synthesis of polymeric membranes with high CO2/N2 separation performance, only a few membrane technologies were advanced to the bench-scale study or above from a highly idealized laboratory setting. Therefore, the recent progress in polymeric membranes is reviewed in the perspectives of capture system energetics, process synthesis, membrane scale-up, modular fabrication, and field tests. These engineering considerations can provide a holistic approach to better guide membrane research and accelerate the commercialization of gas separation membranes for post-combustion carbon capture.

20 FOSSIL-FUELED POWER PLANTS↗
NASA NTRS1988

Assessment, development and application of combustor aerothermal models

The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between: combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamics (CFD) codes have led to their application to complex three-dimensional flows such as those in the gas turbine combustor. A number of US. Government and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aerothermal Modeling and supporting programs are described in this paper.

Aerothermal modeling↗
NASA NTRS1988

Assessment, development, and application of combustor aerothermal models

The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamcis codes have led to their application to complex 3-D flows such as those in the gas turbine combustor. A number of U.S. Government and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aerothermal Modeling and supporting programs are described.

Holdeman, J. D.↗
NASA NTRS1989

Assessment, development, and application of combustor aerothermal models

The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between: combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamics (CFD) codes have led to their application to complex three-dimensional flows such as those in the gas turbine combustor. A number of U.S. Govern- ment and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aero- thermal Modeling and supporting programs are described in this paper

J D Holdeman↗
NASA NTRS1989

High performance alloy electroforming

Electroformed copper and nickel are used in structural applications for advanced propellant combustion chambers. An improved process has been developed by Bell Aerospace Textron, Inc. wherein electroformed nickel-manganese alloy has demonstrated superior mechanical and thermal stability when compared to previously reported deposits from known nickel plating processes. Solution chemistry and parametric operating procedures are now established and material property data is established for deposition of thick, large complex shapes such as the Space Shuttle Main Engine. The critical operating variables are those governing the ratio of codeposited nickel and manganese. The deposition uniformity which in turn affects the manganese concentration distribution is affected by solution resistance and geometric effects as well as solution agitation. The manganese concentration in the deposit must be between 2000 and 3000 ppm for optimum physical properties to be realized. The study also includes data regarding deposition procedures for achieving excellent bond strength at an interface with copper, nickel-manganese or INCONEL 718. Applications for this electroformed material include fabrication of complex or re-entry shapes which would be difficult or impossible to form from high strength alloys such as INCONEL 718.

Malone, G. A.↗
NASA NTRSDate not supplied

Component Applications using Metal Additive Manufacturing Techniques and Materials for Rocket Propulsion

The NASA Marshall Space Flight Center (MSFC) has been involved with various forms of metallic additive manufacturing for use in liquid rocket engine component design, development, and testing since 2010. These AM techniques have been demonstrated to significantly reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing fabrication of designs not feasible by conventional means. The focus at the NASA MSFC for these metal additive manufacturing techniques include laser powder-bed fusion (L-PBF), blown powder directed energy deposition (DED) and arc-based deposition. A variety of components have been evaluated and tested including thrust chamber injectors, injector components such as faceplates, regeneratively-cooled combustion chambers, regeneratively-cooled nozzles, gas generator and preburner hardware, and augmented spark igniters. To support these component applications in harsh environments, NASA has advanced a variety of “standard” additive manufacturing alloys such as those in the superalloy-family and also evolved new alloys including GRCop-84, GRCop-42, NASA HR-1, and JBK-75. The purpose of this presentation is to discuss the various programs at the NASA MSFC using AM to develop, fabricate, and test combustion devices hardware and the evolution of the new additive alloys. Additional information will be provided on the development of multi-metallic additive manufacturing, post-processing of AM techniques including surface enhancements (polishing) techniques, material and process characterization, future development programs, and dissemination of data to industry partners.

Additive Manufacturing↗