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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

"Powdered Magnesium: Carbon Dioxide Combustion for Mars Propulsion"

Powdered magnesium - carbon dioxide combustion is examined as a potential in-situ propellant combination for Mars propulsion. Although this particular combination has relatively low performance in comparison to traditional bi-propellants, it remains attractive as a potential basis for future Martian mobility systems since it could be partially or wholly manufactured from indigenous planetary resources. As a means of achieving high mobility during long-duration Mars exploration missions, the poorer performing in-situ combination can, in fact, become a superior alternative to conventional storable propellants, which would need to be entirely transported from earth. Thus, the engineering aspects of powdered metal combustion devices are discussed including transport/injection of compacted powder, ignition, combustion efficiency, combustion stability, dilution effects, lean burn limits, and slag formation issues. It is suggested that these technological issues could be effectively addressed through a multi-phase research and development effort beginning with basic feasibility tests using an existing dump configured atmospheric pressure burner. Follow-on phases would involve the development and testing of a pressurized research combustor and technology demonstration tests of a prototypical rocket configuration.

Foote, John P.↗
NASA NTRS2007

Powdered Magnesium-Carbon Dioxide Rocket Combustion Technology for In Situ Mars Propulsion

Powdered magnesium (Mg) carbon dioxide (CO2) combustion is examined as a potential in situ propellant combination for Mars propulsion. Although this particular combination has relatively low performance in comparison to traditional bipropellants, it remains attractive as a potential basis for future martian mobility systems, since it could be partially or wholly manufactured from indigenous planetary resources. As a means of achieving high mobility during long-duration Mars exploration missions, the poorer performing in situ combination can, in fact, become a superior alternative to conventional storable propellants, which would need to be entirely transported from Earth. Thus, the engineering aspects of powdered metal combustion devices are discussed including transport/injection of compacted powder, ignition, combustion efficiency, combustion stability, dilution effects, lean burn limits, and slag formation issues. It is suggested that these technological issues could be effectively addressed through a multiphase research and development effort beginning with basic feasibility tests using an existing dump configured atmospheric pressure burner. Follow-on phases would involve the development and testing of a pressurized research combustor and technology demonstration tests of a prototypical rocket configuration.

Foote, J. P.↗
NASA NTRS1958

A Preliminary Investigation of the Performance of a Short-length Turbojet Combustor Using Vaporized Hydrocarbon Fuels

Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Jones, R. E.↗
DOE OSTI2022

Development and Demonstration of a Fuel-Efficient, Class 8 Tractor & Trailer Engine System (SuperTruck II)

Navistar presents the SuperTruck II (ST II) Final Report to the Unites States Department of Energy (US DOE), which covers the five Budget Periods (BPs) from 10-1-2016 through 6-30-2022. For ST II, Navistar built on the achievements of the SuperTruck I (ST I) Program as a catalyst to continue critical research, design and development, testing, and operations to reach the ambitious goals of the ST II project. This approach allowed Navistar to continue contributing to the essential needs of our nation for safe, efficient, and cost-effective delivery of goods and services, as we reduced negative environmental effects and improved operational productivity. This document contains information specified in DOE F 4600.2, Final Scientific/Technical Report DOE F 241.3, B. SCIENTIFIC/TECHNICAL REPORTS, explaining how we met and exceeded program requirements. Throughout this Final Report, Navistar extracted information from documents prepared during the project that represent our management, design and development, building, and testing efforts to meet and exceed SuperTruck II project goals. Navistar followed Plan requirements to achieve / exceed Project Objectives: a) >100% improvement in vehicle freight efficiency (FE) (on ton-MPG basis) relative to 2009 baseline with stretch goal of 140% improvement [actual: 170%); b) >55% engine brake thermal efficiency (BTE) demonstrated in operational engine at a 65-mph cruise point on a dynamometer – ≥31% increase from 2009 baseline [actual: 55.20% of combined BTE) ; and c) development and implementation of commercially cost effective technologies (in terms of a simple payback). Technology selection / development path focused on developing technologies applicable for production within 3-year approach, while ensuring technology readiness and cost of ownership for end users. The Program was organized into five budget periods: Requirements / Technology Assessment and Initial Hardware Testing; Technology Development and Concept Readiness Demonstration; Technology Finalization and Validation Tractor / Trailer Fabrication, Integration and Commissioning Demonstration; and Fuel Economy (FE) and Brake Thermal Efficiency (BTE) and Program Completion. Leadership was provided by DOE, with tasks performed by laboratories (Argonne National Laboratory, Lawrence Livermore National Laboratory); partners at Bosch, TPI, Dana, and J.B. Hunt; , and support from University of Michigan and Clemson University. Navistar lead this team with Principal Investigator / Contracting Officer; Project Manager (PM); Vehicle, Engine, and Aftertreatment Engineers; Finance Manager, Technical Program Leads, and Legal/IP; and other key personnel. Work also included personnel in risk management; funding / budget / finance. Work involved analysis, development, testing, and down selection of individual/system engine, aftertreatment, and vehicle technologies, with integration of selected technologies into a prototype vehicle for demonstration of fuel-efficiency gain. Work also included component/integrated system level development of truck and trailer aerodynamics, base engine efficiency, advanced aftertreatment, combustion efficiency, waste heat recovery, hybrid powertrain, reduced rolling resistance, weight reduction, idle reduction, and driver feedback. As ST II progressed, Navistar performed computer-based modeling / simulations of technologies focused on the primary operational areas: Engine, Aftertreatment, and Vehicle. During the ST II Program, the COVID Virus outbreak unexpectedly challenged by the effects of, which affected staffing, scheduling, design, supplies, availability of materials, production procedures, and testing. The DOE responded by extending the program by three quarters to ensure that project tasks were completed for this vital project. Focus continued on analyzing, developing, testing, and down selecting individual-/system-level engine and vehicle technologies for integration of the final selected technologies into a prototype vehicle that would demonstrate fuel-efficiency gains made possible through these technologies. This included component/integrated system-level development of truck and trailer aerodynamics, base engine efficiency, advanced aftertreatment, combustion efficiency, waste heat recovery, solar power, distributed and intelligent vehicle power, hybrid powertrain, reduced rolling resistance, weight reduction, idle reduction, and driver feedback. Throughout the program, function, reliability, and performance at all levels were ensured through testing. Proof of this approach was demonstrated in multiple, on-road demonstrations: Scenario A (Flatland) Fuel Economy, Scenario B (Hilly) Fuel Economy, and City Cycle Tests. Other benefits derived from ST II included new/improved products, publications, patents, and next-step capabilities related to electric/hydrogen vehicles and autonomous driving.

Zukouski, Russ↗
NASA NTRS1951

Combustion-chamber Performance Characteristics of a Python Turbine-propeller Engine Investigated in Altitude Wind Tunnel

Combustion-chamber performance characteristics of a Python turbine-propeller engine were determined from investigation of a complete engine over a range of engine speeds and shaft horsepowers at simulated altitudes. Results indicated the effect of engine operating conditions and altitude on combustion efficiency and combustion-chamber total pressure losses. Performance of this vaporizing type combustion chamber was also compared with several atomizing type combustion chambers. Over the range of test conditions investigated, combustion efficiency varied from approximately 0.95 to 0.99.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1949

Altitude Performance and Operational Characteristics of 29-inch-diameter Tail-pipe Burner with Several Fuel Systems and Flame Holders on J35 Turbojet Engine

An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning has been conducted in the NACA Lewis altitude wind tunnel. Several fuel systems and flame holders were investigated in a 29-inch-diameter tail-pipe burner to determine the effect of fuel distribution and flame-holder design on tail-pipe-burner performance and operational characteristics over a range of simulated flight conditions. At an altitude of 5000 feet, the type of flame holder used had only a slight effect on the combustion efficiency. As the altitude was increased, the decrease in peak combustion efficiency became more rapid as the blocking area of the flame holder was reduced. At all altitudes investigated, an improvement in the uniformity of the radial distribution of fuel and air slightly increased the peak combustion efficiencies and shifted the peak combustion efficiency to higher tail-pipe fuel-air ratios. The use of an internal cooling liner extending the full length of the tail-pipe combustion chamber provided adequate shell cooling at all flight conditions investigated.

Conrad, E William↗
DOE OSTI2020

SENSITIVITY ANALYSIS OF FUEL PHYSICAL PROPERTY EFFECTS ON SPARK IGNITION ENGINE PERFORMANCE

Alternative fuels are of interest to automakers and regulators due to their potential to reduce net greenhouse gas emissions from transportation sources. Alternative fuels also have fuel properties which may enable advanced combustion modes with higher engine thermal efficiencies. There has been previous work to identify the relationship between various fuel properties and engine performance, but most of this work has been experiments or simulations where the change in properties was obtained through changing the fuel composition, making isolating the effects of individual fuel properties difficult. In this study, numerical simulations have been used to investigate the effects of individual fuel physical properties such as viscosity or heat of vaporization (HoV) on engine performance. Simulations have been performed of two different engine platforms, the first an optical, single-cylinder research engine and the second a multi-cylinder production engine. Both engines are direct-injection spark-ignition engines with pent-roof heads and are designed for automotive applications. Each engine was run at a different operating condition, one stable and one knock-limited. Different base fuels provided a variety of simulated conditions. Up to six different fuel properties were varied as part of Global Sensitivity Analyses performed for each of the engines with multiple performance targets including thermal efficiency, combustion efficiency and combustion phasing. Results show trends that are largely consistent with previous experimental findings using multiple fuels. The engine thermal efficiency was primarily sensitive to the fuel's HoV, with other fuel physical properties having smaller effects. For optical engine results, the magnitude of the effect was greater in this study than expected based on previous experimental results were many fuel physical and chemical properties were varied simultaneously. However, for the multi-cylinder production engine, the relationship between thermal efficiency and HoV was slightly smaller.

engine simulation↗
NASA NTRS1976

Effect of exhaust gas recirculation on emissions from a flame-tube combustor using Liquid Jet A fuel

The effects of uncooled exhaust gas recirculation as an inert diluent on emissions of oxides of nitrogen (NO + NO2) and on combustion efficiency were investigated. Ratios of recirculated combustion products to inlet airflow were varied from 10 to 80 percent by using an inlet air ejector nozzle. Liquid Jet A fuel was used. The flame-tube combustor was 10.2 cm in diameter. It was operated with and without a flameholder present. The combustor pressure was maintained constant at 0.5 MPa. The equivalence ratio was varied from 0.3 to 1.0. The inlet air temperature was varied from 590 to 800 K, and the reference velocity from 10 to 30 m/sec. Increasing the percent recirculation from 10 to 25 had the following effects: (1) the peak NOx emission was decreased by 37 percent, from 8 to 5 g NO2/kg fuel, at an inlet air temperature of 590 K and a reference velocity of 15 m/sec; (2) the combustion efficiency was increased, particularly at the higher equivalence ratios; and (3) for a high combustion efficiency of greater than 99.5 percent, the range of operation of the combustor was nearly doubled in terms of equivalence ratio. Increasing the recirculation from 25 to 50 percent did not change the emissions significantly.

Marek, C. J.↗
NASA NTRS1974

Some aspects of flox-methane rocket engine throttling

Four injector designs and two chamber profiles were experimentally evaluated for structural integrity, combustion efficiency, and resistance to combustion instabilities. Vacuum thrust measurements were used as a primary measure of combustion efficiency. Stability rating to test the sensitivity of the injectors to high frequency combustion was conducted, but not extensively. To map the boundary between stable operation and chugging instability, chamber pressure was throttled downward from 689.5 to 206.9 kN/sq m abs (100 to 30 psia). Best operational results were obtained with an injector configuration having no hydraulic swirlers, a 0.00102-m (0.040-in.) recessed FLOX tube, and a nonflared exit in the methane annulus. This injector design exhibited stable combustion and good integrity of hardware, and it exceeded the design goal efficiency (88 percent) at the 10 to 1 throttled condition.

Sokolowski, D. E.↗
NASA NTRS1948

Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine

A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.

Childs, J. Howard↗
NASA NTRS1995

Cooperative Testing of Rocket Injectors That Use Gaseous Oxygen and Hydrogen

Gaseous oxygen and hydrogen propellants used in a special engine energy cycle called Full-Flow Staged Combustion are believed to significantly increase the lifetime of a rocket engine's pumps. The cycle can also reduce the operating temperatures of the engine. Improving the lifetime of the hardware reduces its overall maintenance and operations costs, and is critical to reducing costs for the joint NASA/industry Reusable Launch Vehicle (RLV). The work in this project will demonstrate the performance and lifetime of one-element and many-element combustors with gaseous O2/H2 injectors. This work supporting the RLV program is a cooperative venture of the NASA Lewis Research Center, the NASA Marshall Space Flight Center, Rocketdyne, and the Pennsylvania State University. Information about gas-gas rocket injector performance with O2/H2 is very limited. Because of this paucity of data, new testing is needed to improve the knowledge base for testing and designing new injectors for the RLV and to improve computer models that predict the combusting gas flows of new injector designs. Therefore, detailed observations and measurements of the combusting flow from many-element injectors in a rocket engine are being sought. These observations and measurements will be done with three different tools: schlieren photography, ultraviolet imaging, and Raman spectroscopy. The schlieren system will take photos of the density differences in combusting flow, the ultraviolet movies will determine the location of the hydroxyl (OH) radical in the combustion flow, and the Raman spectroscopic measurements will provide the combustion temperature and amount of water (H2O), hydrogen (H2), and oxygen (O2) in the combustor. Marshall is providing overall program management, design and computational fluid dynamics (CFD) analyses, as well as funding for the work at Penn State. An existing, windowed combustor and several injectors will be provided by Rocketdyne--two injectors for the initial screening tests and one with an optimized design based on the best design found in the screening tests. Lewis will provide a nozzle and several injectors for the screening test program. The configuration of the injectors will be based on a design chosen by all the participants, and their elements will be based on the coaxial and impinging flow. Lewis also will provide the instrumentation for the flow-field measurements: schlieren, ultraviolet imaging, and Raman spectroscopy. In addition, thermocouples will measure heat flow on the injector face. Other traditional measurements of rocket performance will be made as well: chamber pressure, mass flow of each propellant, purge flow, and the barrier cooling gas flow. Penn State will conduct single-element testing with the injector elements from both the Rocketdyne and the jointly designed injectors. A wide variety of traditional and nontraditional injector designs will be tested in this program. The results will be valuable in computational fluid dynamics code validation and overall rocket combustion efficiency measurements. Correlations between combustion efficiency, laser measurements of species, and ultraviolet and visible light photography will also be made. Thus far, several different single-element injectors have been tested at Penn State and Lewis. The experimental setup of a rocket engine with a viewing window is shown. The combusting flow is shown. The results are helping engineers design the many element injectors.

Source record↗
NASA NTRS1990

Solid-liquid staged combustion space boosters

NASA has begun to evaluate solid-liquid hybrid propulsion for launch vehicle booster. A three-phase program was outlined to identify, acquire, and demonstrate technology needed to approximate solid and liquid propulsion state of the art. Aerojet has completed a Phase 1 study and recommends a solid-liquid staged combustion concept in which turbopump fed LO2 is burned with fuel-rich solid propellant effluent in aft-mounted thrust chambers.These reasonably sized thrust chambers are LO2 regeneratively cooled, supplemented with fuel-rich barrier cooling. Turbopumps are driven by the resulting GO2 coolant in an expander-bleed-burnoff cycle. Turbine exhaust pressurizes the LO2 tankage directly, and the excess is bled into supersonic nozzle splitlines, where it combusts with the fuel rich boundary layer. Thrust vector control is enhanced by supersonic nozzle movement on flexseal mounts. Every hybrid solid-liquid concept examined improves booster energy management and launch propellant safety compared to current solid boosters. Solid-liquid staged combustion improves hybrid performance by improving both combustion efficiency and combustion stability, especially important for large boosters. These improvements result from careful fluid management and use of smaller combustors. The study shows NASA safety, reliability, cost, and performance criteria are best met with this concept, wherein simple hardware relies on several separate emerging technologies, all of which have been demonstrated successfully.

Culver, D. W.↗
NASA NTRS2021

Design Guidelines for Swirl-Venturi Fuel-Air Mixers for Lean Direct Injection Combustors

Design guidelines for aircraft gas turbine combustors are provided based on the swirl-venturi lean direction injection (SVLDI) scheme. The guidelines are developed through a combination of literature review and internal research at NASA GRC that sought to develop an understanding of how LDI geometry affects NOx emissions, minimum stable f/a, and combustion dynamics. Key objectives are to improve low power operability by reducing the fuel/air ratio achievable with the SV-LDI concept and to reduce NOx emissions for both the subsonic landing-takeoff cycle and supersonic cruise conditions. The design of any combustion system will be a set of compromises between flame stability and decreasing combustion emissions. This document is to serve as a guide for selecting the best options given the overall mission. This document is organized as follows. In section 1, we introduce and describe the SV-LDI concept as a way to reduce NOx emissions. Section 2 provides descriptions of the individual components that make up a single swirl-venturi LDI element — the venturi, swirler, and fuel injector and their primary purposes. Considerable focus is applied to how changes on the air side components affect the flow field. Section 3 examines variations within a single SV-LDI element swirler: blade thickness, presence of a diffuser section, position of fuel nozzle, et cetera. We include examples from the literature to emphasize key aspects of these changes. In Section 4 we consider the interaction of a single element with the combustion chamber and with neighboring elements. In Section 5, we look at fuel staging effects on emissions, combustion efficiency, and combustion dynamics. Throughout these sections, we will highlight any findings we consider universal, and in Section 6 we consolidate those findings to provide guidelines for use in SV-LDI designs.

gas turbine combustion↗
DOE OSTI2024

Development of prechamber enabled mixing-controlled combustion strategy for ultra-low methane emissions from lean burn natural gas engines

This numerical study explores the optimization of Prechamber Enabled Mixing-Controlled Combustion (PC-MCC) using natural gas in heavy-duty engines, aiming to enhance combustion efficiency to minimize methane slip and NOx emissions. The approach involves a prechamber ignition system, distinct from conventional spark ignition (SI) systems, to initiate combustion of direct injected natural gas. By leveraging the robust ignition characteristics of the prechamber, the PC-MCC method demonstrates significant potential in achieving efficient combustion akin to diesel engines but with lower greenhouse gas emissions. The research evaluates the effects of various geometric and operational parameters on the combustion process and emissions, including prechamber volume, nozzle diameter, direct injector (DI) geometry, and engine operating strategies. Computational Fluid Dynamics (CFD) simulations are utilized, focusing on a heavy-duty, single-cylinder engine modeled after the Caterpillar C9.3B engine. Key findings indicate that a prechamber volume of 3 cc, coupled with a nozzle diameter of 2.75 mm for two prechamber holes, strikes an optimal balance between combustion efficiency and emissions reduction. This configuration ensures robust combustion across a range of operating conditions while maintaining methane slip within targeted limits. Further investigation into DI geometry shows the significance of the injector umbrella angle and nozzle diameter in shaping the fuel-air mixing and combustion dynamics. An umbrella angle of 130° and a nozzle diameter of 300 microns are identified as optimal, promoting rapid and efficient combustion with minimized methane and NOx emissions. The study also investigates the impact of injection timing and pressure, highlighting their roles in controlling combustion timing and influencing emissions levels. Advanced injection timing is found to be crucial in achieving the desired low methane slip, whereas retarded injection timing assists to reduce NOx emissions while having a slight increase in methane emissions. Operating strategies incorporating various levels of Exhaust Gas Recirculation (EGR) are assessed for their effectiveness in further reducing emissions. The research demonstrates that a judicious combination of internal hot EGR and careful calibration of DI pressure and SOI timing can achieve significant reductions in NOx emissions while keeping methane slip under control. Specifically, an internal EGR level of 15%–25%, combined with DI pressures of 200–300 bar and injection timings at or after top dead center, is recommended. These findings contribute valuable insights into the development of advanced combustion techniques for natural gas engines, offering a viable pathway to reduce methane slip without compromising engine efficiency or performance. The PC-MCC system presents a promising solution for the future of heavy-duty natural gas engine technology to reduce methane emissions.

Nsaif, Osama↗
DOE OSTI2021

Performance of direct injected propane and gasoline in a high stroke-to-bore ratio SI engine: Pathways to diesel efficiency parity with ultra low soot

This work explores pathways to achieve diesel-like, high-efficiency combustion with stoichiometric 3-way catalyst compatible combustion in a single-cylinder spark ignition (SI) research engine. A unique high stroke-to-bore engine design (1.5:1) with cooled exhaust gas recirculation (EGR) and high compression ratio ( r c ) was used to improve engine efficiency by up to 30% compared with a production turbocharged gasoline direct injection spark ignition engine. Engine experiments were conducted with both 91 RON E10 gasoline and liquified petroleum gas (LPG) (i.e. autogas) and were compared to legacy gasoline data on the production engine. Geometric compression ratio ( r c ) of 13.3:1 was used for both fuels with additional experiments at 16.8:1 for LPG only. Measurements of exhaust soot particle size and number concentrations were made with both fuels. Significant reduction in soot particles across the whole particle size range were achieved with LPG due to the elimination of in-cylinder liquid films. The effects of EGR, late intake valve closing (IVC) and fuel characteristics were investigated through their effects on efficiency, combustion stability and soot production. Results of 47% gross thermal efficiency, and 45% net thermal efficiency at stoichiometric engine operation, at up to 17 bar IMEP and 2000 r/min with 16.8:1 r c were achieved with LPG. Estimated brake efficiency values were compared to a contemporary medium duty diesel engine illustrating the benefits of the chosen path for achieving diesel efficiency parity.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1975

Effect of hydrogen injection stability and emissions of an experimental premixed prevaporized propane burner

Hydrogen in quantities up to 5 percent by weight of the total fuel flow was injected into a premixed propane burner. The hydrogen was either premixed with the propane and air upstream of the burner or introduced as a torch at the flameholder. Emissions of total nitrogen oxides, carbon monoxide, and unburned hydrocarbon are reported as are combustion efficiencies and lean blowout limits. To maintain at least 99 percent combustion efficiency at a 700 K inlet mixture temperature with no hydrogen added, it was necessary to burn with a propane equivalence ratio of 0.525. When 4 percent hydrogen was premixed with the propane and air, a combustion efficiency greater than 99 percent was recorded at a propane equivalence ratio of 0.425. The total nitrogen oxides (NOx) emissions corresponding to these two conditions were 0.8 g NO2/kg equivalent propane and 0.44 g NO2/kg equivalent propane, respectively. The hydrogen torch did not reduce NOx emissions.

Anderson, D. N.↗
NASA NTRS1993

STE thrust chamber technology: Main injector technology program and nozzle Advanced Development Program (ADP)

The purpose of the STME Main Injector Program was to enhance the technology base for the large-scale main injector-combustor system of oxygen-hydrogen booster engines in the areas of combustion efficiency, chamber heating rates, and combustion stability. The initial task of the Main Injector Program, focused on analysis and theoretical predictions using existing models, was complemented by the design, fabrication, and test at MSFC of a subscale calorimetric, 40,000-pound thrust class, axisymmetric thrust chamber operating at approximately 2,250 psi and a 7:1 expansion ratio. Test results were used to further define combustion stability bounds, combustion efficiency, and heating rates using a large injector scale similar to the Pratt & Whitney (P&W) STME main injector design configuration including the tangential entry swirl coaxial injection elements. The subscale combustion data was used to verify and refine analytical modeling simulation and extend the database range to guide the design of the large-scale system main injector. The subscale injector design incorporated fuel and oxidizer flow area control features which could be varied; this allowed testing of several design points so that the STME conditions could be bracketed. The subscale injector design also incorporated high-reliability and low-cost fabrication techniques such as a one-piece electrical discharged machined (EDMed) interpropellant plate. Both subscale and large-scale injectors incorporated outer row injector elements with scarfed tip features to allow evaluation of reduced heating rates to the combustion chamber.

Source record↗
NASA NTRS1952

Effect of design changes and operating conditions on combustion and operational performance of a 28-inch diameter Ram-jet engine / T. B. Shillito and Shigeo Nakanishi

The results of an altitude test-chamber investigation of the effects of a number of design changes and operating conditions on altitude peformance of a 28-inch diameter ram jet engine are presented. Most of the investigation was for a simulated flight Mach number of 2.0 above the tropopause. Fuel-air distribution, gutter width, the presence of a pilot flame, cimbustion-chamber-inlet temperature, and exhaust-nozzle throat area were found to have significant effects on limits of combustion. Combustion efficiency increased with increasing combustion-chamber-inlet temperature and was adversely affected by an increase in the exhaust-nozzld area. Similiar lean limits of combustion were obtained for both Diesel fuel and normal heptane, but combustion efficiences obtained with Diesel fuel were lower than those obtained with normal heptane.

NOZZLES, EXHAUST - ENGINES, JET↗