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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 397 records · Page 22
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

NOx Emissions Performance and Correlation Equations for a Multipoint LDI Injector

Lean Direct Injection (LDI) is a combustor concept that reduces nitrogen oxides (NOx) emissions.This paper looks at a 3-zone multipoint LDI concept developed by Parker Hannifin Corporation. The concept was tested in a flame-tube test facility at NASA Glenn Research Center. Due to test facility limitations, such as inlet air temperature and pressure, the flame-tube test was not able to cover the full set of engine operation conditions. Three NOx correlation equations were developed based on assessing NOx emissions dependencies on inlet air pressure (P3), inlet air temperature (T3), and fuel air equivalence ratio(theta) to estimate the NOx emissions at the unreachable high engine power conditions. As the results, the NOx emissions are found to be a strong function of combustion inlet air temperature and fuel air equivalence ratio but a weaker function of inlet air pressure. With these three equations, the NOx emissions performance of this injector concept is calculated as a 66 reduction relative to the ICAO CAEP-6 standard using a 55:1 pressure-ratio engine cycle. Uncertainty in the NOx emissions estimation increases as the extrapolation range departs from the experimental conditions. Since maximum inlet air pressure tested was less than 50 of the full power engine inlet air pressure, a future experiment at higher inlet air pressure conditions is needed to confirm the NOx emissions dependency on inlet air pressure.

NOx correlation equations↗
NASA NTRS2015

NOx Emissions Performance and Correlation Equations for a Multipoint LDI Injector

Lean Direct Injection (LDI) is a combustor concept that reduces nitrogen oxides (NOx) emissions. This paper looks at a 3-zone multipoint LDI concept developed by Parker Hannifin Corporation. The concept was tested in a flame-tube test facility at NASA Glenn Research Center. Due to test facility limitations, such as inlet air temperature and pressure, the flame-tube test was not able to cover the full set of engine operation conditions. Three NOx correlation equations were developed based on assessing NOx emissions dependencies on inlet air pressure (P3), inlet air temperature (T3), and fuel air equivalence ratio (Ф) to estimate the NOx emissions at the unreachable high engine power conditions. As the results, the NOx emissions are found to be a strong function of combustion inlet air temperature and fuel air equivalence ratio but a weaker function of inlet air pressure. With these three equations, the NOx emissions performance of this injector concept is calculated as a 66% reduction relative to the ICAO CAEP-6 standard using a 55:1 pressure-ratio engine cycle. Uncertainty in the NOx emissions estimation increases as the extrapolation range departs from the experimental conditions. Since maximum inlet air pressure tested was less than 50% of the full power engine inlet air pressure, a future experiment at higher inlet air pressure conditions is needed to confirm the NOx emissions dependency on inlet air pressure.

NOx correlation equations↗
NASA NTRS2001

"2001: A Space Odyssey" Revisited: The Feasibility of 24 Hour Commuter Flights to the Moon Using NTR Propulsion with LUNOX Afterburners

The prospects for "24 hour" commuter flights to the Moon. similar to that portrayed in 2001: A Space Odyssey but on a more Spartan scale. are examined using two near term. "high leverage" technologies-liquid oxygen (LOX)-augmented nuclear thermal rocket (NTR) propulsion and "lunar-derived" oxygen (LUNOX) production. Iron-rich volcanic glass. or "orange soil," discovered during the Apollo 17 mission to Taurus-Littrow. has produced a 4% oxygen yield in recent NASA experiments using hydrogen reduction. LUNOX development and utilization would eliminate the need to transport oxygen supplies from Earth and is expected to dramatically reduce the size, cost and complexity of space transportation systems. The LOX-augmented NTR concept (LANTR) exploits the high performance capability of the conventional liquid hydrogen (LH2)-cooled NTR and the mission leverage provided by LUNOX in a unique way. LANTR utilizes the large divergent section of its nozzle as an "afterburner" into which oxygen is injected and supersonically combusted with nuclear preheated hydrogen emerging from the engine's choked sonic throat-essentially "scramjet propulsion in reverse." By varying the oxygen-to-hydrogen mixture ratio, the LANTR engine can operate over a wide range of thrust and specific impulse (Isp) values while the reactor core power level remains relatively constant. The thrust augmentation feature of LANTR means that "big engine" performance can be obtained using smaller. more affordable. easier to test NTR engines. The use of high-density LOX in place of low-density LH2 also reduces hydrogen mass and tank volume resulting in smaller space vehicles. An implementation strategy and evolutionary lunar mission architecture is outlined which requires only Shuttle C or "in-line" Shuttle-derived launch vehicles, and utilizes conventional NTR-powered lunar transfer vehicles (LTVs), operating in an "expendable mode" initially, to maximize delivered surface payload on each mission. The increased payload is dedicated to installing "modular" LUNOX production units with the intent of supplying LUNOX to lunar landing vehicles (LLVs) and then LTVs at the earliest possible opportunity. Once LUNOX becomes available in low lunar orbit (LLO), monopropellant NTRs would be outfitted with an oxygen propellant module, feed system and afterburner nozzle for "bipropellant" operation. Transition to a "reusable" mission architecture now occurs with smaller, LANTR-powered LTVs delivering approximately 400% more payload on each piloted round trip mission than earlier expendable "all LH2" NTR systems. As initial lunar outposts grow to eventual lunar settlements and LUNOX production capacity increases, the LANTR concept can enable a rapid "commuter" shuttle capable of 24 hour "one way" trips to and from the Moon. A vast deposit of "iron-rich" volcanic glass beads identified at just one candidate site - located at the southeastern edge of Mare Serenitatis - could supply sufficient LUNOX to support daily commuter flights to the Moon for the next 9000 years'.

Borowski, Stanley K.↗
NASA NTRS2009

Distributed Control Architecture for Gas Turbine Engine

The transformation of engine control systems from centralized to distributed architecture is both necessary and enabling for future aeropropulsion applications. The continued growth of adaptive control applications and the trend to smaller, light weight cores is a counter influence on the weight and volume of control system hardware. A distributed engine control system using high temperature electronics and open systems communications will reverse the growing trend of control system weight ratio to total engine weight and also be a major factor in decreasing overall cost of ownership for aeropropulsion systems. The implementation of distributed engine control is not without significant challenges. There are the needs for high temperature electronics, development of simple, robust communications, and power supply for the on-board electronics.

Dennis Culley↗
DOE OSTI2022

Passive selective catalytic reduction (SCR) system and method for low-temperature combustion (LTC) engine

A combustion mode module is configured to switch operation of a low-temperature combustion (LTC) engine between a spark ignition (SI) mode, a positive valve overlap (PVO) mode, and a negative valve overlap (NVO) mode. A spark control module is configured to control a spark plug to generate a spark in a cylinder of the LTC engine when the LTC engine is operating in the SI mode. A valve control module is configured to control intake and exhaust valves of the cylinder to yield a PVO and a NVO when the LTC engine is operating in the PVO mode and the NVO mode, respectively. An air/fuel (A/F) control module is configured to adjust a desired A/F ratio of the LTC engine to a rich A/F ratio when operation of the LTC engine is switched to the PVO mode from either one of the SI mode and the NVO mode.

Yun, Hanho↗
NASA NTRS1999

Carbon Nanotube Composites: Strongest Engineering Material Ever?

The primary goal of the carbon nanotube project at Johnson Space Center (JSC) is to fabricate structural materials with a much higher strength-to-weight ratio than any engineered material today, Single-wall nanotubes present extraordinary mechanical properties along with new challenges for materials processing. Our project includes nanotube production, characterization, purification, and incorporation into applications studies. Now is the time to move from studying individual nanotubes to applications work. Current research at JSC focuses on structural polymeric materials to attempt to lower the weight of spacecraft necessary for interplanetary missions. These nanoscale fibers present unique new challenges to composites engineers. Preliminary studies show good nanotube dispersion and wetting by the epoxy materials. Results of tensile strength tests will also be reported. Other applications of nanotubes are also of interest for energy storage, gas storage, nanoelectronics, field emission, and biomedical uses.

Mayeaux, Brian↗
NASA NTRS1953

Effect of Fuel-air Ratio Concentration in Combustion Zone on Combustion Performance of a 16-inch Ram-jet Engine

The results of fuel-air ratio and temperature surveys made in the flame stabilizing zone of a 16-inch ram-jet combustor indicated that the distribution of combustible mixture immediately downstream of a flame holder should be mechanically controlled if a preferred mixture composition is to be maintained in that zone. A sloping-baffle combustor configuration gave combustion efficiencies of 90 percent or greater over a range of fuel-air ratios from 0.010 to 0.045.

COMBUSTION - RAM-JET ENGINES↗
NASA NTRS1973

Advanced space engine preliminary design

A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.

Cuffe, J. P. B.↗
NASA NTRS1975

Preliminary evaluation of a heat pipe heat exchanger on a regenerative turbofan

A preliminary evaluation was made of a regenerative turbofan engine using a heat pipe heat exchanger. The heat exchanger had an effectiveness of 0.70, a pressure drop of 3 percent on each side, and used sodium for the working fluid in the stainless steel heat pipes. The engine was compared to a reference turbofan engine originally designed for service in 1979. Both engines had a bypass ratio of 4.5 and a fan pressure ratio of 2.0. The design thrust of the engines was in the 4000 N range at a cruise condition of Mach 0.98 and 11.6 km. It is shown that heat pipe heat exchangers of this type cause a large weight and size problem for the engine. The penalties were too severe to be overcome by the small uninstalled fuel consumption advantage. The type of heat exchanger should only be considered for small airflow engines in flight applications. Ground applications might prove more suitable and flexible.

Kraft, G. A.↗
NASA NTRS2008

Heat Transfer Measurements for a Film Cooled Turbine Vane Cascade

Experimental heat transfer and pressure measurements were obtained on a large scale film cooled turbine vane cascade. The objective was to investigate heat transfer on a commercial high pressure first stage turbine vane at near engine Mach and Reynolds number conditions. Additionally blowing ratios and coolant density were also matched. Numerical computations were made with the Glenn-HT code of the same geometry and compared with the experimental results. A transient thermochromic liquid crystal technique was used to obtain steady state heat transfer data on the mid-span geometry of an instrumented vane with 12 rows of circular and shaped film cooling holes. A mixture of SF6 and Argon gases was used for film coolant to match the coolant-to-gas density ratio of a real engine. The exit Mach number and Reynolds number were 0.725 and 2.7 million respectively. Trends from the experimental heat transfer data matched well with the computational prediction, particularly for the film cooled case.

Poinsatte, Philip E.↗
DOE OSTI2022

Effects of Port Angle on Scavenging of an Opposed Piston Two-Stroke Engine

Opposed-piston 2-stroke (OP-2S) engines have the potential to achieve higher thermal efficiency than a typical diesel engine. However, the uniflow scavenging process is difficult to control over a wide range of speeds and loads. Scavenging performance is highly sensitive to pressure dynamics, port timings, and port design. This study proposes an analysis of the effects of port vane angle on the scavenging performance of an opposed-piston 2-stroke engine via simulation. A CFD model of a three-cylinder opposed-piston 2-stroke was developed and validated against experimental data collected by Achates Power Inc. One of the three cylinders was then isolated in a new model and simulated using cycle-averaged and cylinder-averaged initial/ boundary conditions. This isolated cylinder model was used to efficiently sweep port angles from 12 degrees to 29 degrees at different pressure ratios. Results indicate that scavenging performance is correlated with the bulk swirl ratio generated by these port angles. Scavenging performance is also sensitive to the pressure ratio across the engine. It was concluded that, for a given pressure ratio, the smallest port angle produces the best scavenging; however, previous work indicates lower bulk swirl ratio during combustion can have an impact on emissions formation.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS1978

Aerodynamic characteristics of a counter-rotating, coaxial, hingeless rotor helicopter model with auxiliary propulsion

A wind-tunnel model test at advance ratios from 0 to 0.3 with and without auxiliary jet engine thrust is reported. At each advance ratio and engine thrust, both the control power and the aircraft stability were measured. The results indicate that there is a cross-coupling for collective pitch and longitudinal cyclic pitch inputs. The control power for these inputs increased with advance ratio. There was also cross-coupling for differential collective pitch inputs. The airframe was longitudinally unstable, but the instability was less at the highest advance ratio tested. The airframe showed both positive effective dihedral and positive directional stability.

Phelps, A. E., III↗
NASA NTRS1992

Orbit transfer rocket engine technology program: Advanced engine study

In Task D.6 of the Advanced Engine Study, three primary subtasks were accomplished: (1) design of parametric data; (2) engine requirement variation studies; and (3) vehicle study/engine study coordination. Parametric data were generated for vacuum thrusts ranging from 7500 lbf to 50,000 lbf, nozzle expansion ratios from 600 to 1200, and engine mixture ratios from 5:1 to 7:1. Failure Modes and Effects Analysis (FMEA) was used as a departure point for these parametric analyses. These data are intended to assist in definition and trade studies. In the Engine Requirements Variation Studies, the individual effects of increasing the throttling ratio from 10:1 to 20:1 and requiring the engine to operate at a maximum mixture ratio of 12:1 were determined. Off design engine balances were generated at these extreme conditions and individual component operating requirements analyzed in detail. Potential problems were identified and possible solutions generated. In the Vehicle Study/Engine Study coordination subtask, vehicle contractor support was provided as needed, addressing a variety of issues uncovered during vehicle trade studies. This support was primarily provided during Technical Interchange Meetings (TIM) in which Space Exploration Initiative (SEI) studies were addressed.

Erickson, C. M.↗
NASA NTRS2003

Space Shuttle Main Engine Implications for the Abort-to-Orbit Off-the-Pad Study

In 2001, the Space Shuttle Main Engine (SSME) project office was contacted by the Space Shuttle Ascent Guidance, Navigation, and Control group to provide the engine perspective for an Abort-to-Orbit (ATO) study. The purpose of the AT0 Off-the-Pad study was to determine the feasibility of eliminating the Return to Launch Site and Transatlantic abort modes by using a five-segment solid rocket booster and throttling the remaining SSMEs to a higher power level. This would enable all abort modes to be Abort-to-Orbit. The SSME project office at Marshall Space Flight Center collaborated with MSFC's Space Transportation Directorate and Rocketdyne Propulsion and Power to provide the AT0 Off-the-Pad study with the analysis required. Power levels at 109%, 1 11% and 1 13% of rated power level were studied as well as mixture ratio decreases down to 5.85. SSME was to evaluate and define the technical and programmatic impacts to certify the SSME to these abort power levels. The SSME systems analysis group performed a steady state analysis using the SSME power balance model to determine if there were any technical issues associated with higher power level, low mixture ratio operation. Based on each power level/mixture ratio combination, an engine certification plan was created and a preliminary probabilistic risk assessment was performed. The results showed favorable results for higher power 1evel/lower mixture ratio SSME operation. In nearly all performance and redline parameters, the traded engine operation was encompassed by nominal engine performance of a prior engine configuration.

Schoffstoll, Dayna L.↗
NASA NTRS1976

Summary report on effects at temperature, humidity, and fuel-air ratio on two air-cooled light aircraft engines

Five different engine models were tested to experimentally characterize emissions and to determine the effects of variation in fuel-air ratio and spark timing on emissions levels and other operating characteristics such as cooling, misfiring, roughness, power acceleration, etc. The results are given of two NASA reports covering the Avco Lycoming 0-320-D engine testing and the recently obtained results on the Teledyne Continental TSIO-360-C engine.

Kempke, E. E., Jr.↗
NASA NTRS2000

Data Oscillation Resolution of Propellant Flowmeter Used in FASTRAC Engine Testing

The Stennis Space Centers' horizontal test facility, Marshall Space Flight Centers' propulsion test article and the X-34 flight vehicle are designed with V-cone flowmeters for measurement of both RP-1 and LOX flow-rates for Fastrac engine testing. Delta pressure transducer data from these flowmeters are used to calibrate the RP-1 and LOX mixture ratio in the Fastrac engine. Data from the V-Cone flowmeter delta pressure transducers have excessive oscillation. The delta pressure oscillations have caused flowrate data fluctuations that interfered with making the accurate readings necessary to calibrate the RP-1 and LOX mixture ratio required for Fastrac engine operation. The objective of this report is to document the flowmeter data oscillation problem and the method used to obtain more reliable flowmeter data.

Heflin, J.↗
NASA NTRS2007

Implementation of In-Situ Impedance Techniques on a Full Scale Aero-Engine System

Determination of acoustic liner impedance for jet engine applications remains a challenge for the designer. Although suitable models have been developed that take account of source amplitude and the local flow environment experienced by the liner, experimental validation of these models has been difficult. This is primarily due to the inability of researchers to faithfully mimic the environment in jet engine nacelles in the laboratory. An in-situ measurement technique, one that can be implemented in an actual engine, is desirable so an accurate impedance can be determined for future modeling and quality control. This paper documents the implementation of such a local acoustic impedance measurement technique that is used under controlled laboratory conditions as well as on full scale turbine engine liner test article. The objective for these series of in-situ measurements is to substantiate treatment design, provide understanding of flow effects on installed liner performance, and provide modeling input for fan noise propagation computations. A series of acoustic liner evaluation tests are performed that includes normal incidence tube, grazing incidence tube, and finally testing on a full scale engine on a static test stand. Lab tests were intended to provide insight and guidance for accurately measuring the impedance of the liner housed in the inlet of a Honeywell Tech7000 turbofan. Results have shown that one can acquire very reasonable liner impedance data for a full scale engine under realistic test conditions. Furthermore, higher fidelity results can be obtained by using a three-microphone coherence technique that can enhance signal-to-noise ratio at high engine power settings. This research has also confirmed the limitations of this particular type of in-situ measurement. This is most evident in the installation of instrumentation and its effect on what is being measured.

Gaeta, R. J.↗
NASA NTRS1997

SSME Key Operations Demonstration

A Space Shuttle Main Engine (SSME) test program was conducted between August 1995 and May 1996 using the Technology Test Bed (TTB) Engine. SSTO vehicle studies have indicated that increases in the propulsion system operating range can save significant weight and cost at the vehicle level. This test program demonstrated the ability of the SSME to accommodate a wide variation in safe operating ranges and therefore its applicability to the SSTO mission. A total of eight tests were completed with four at Marshall Space Flight Center's Advanced Engine Test Facility and four at the Stennis Space Center (SSC) A-2 attitude test stand. Key demonstration objectives were: 1) Mainstage operation at 5.4 to 6.9 mixture ratio; 2) Nominal engine start with significantly reduced engine inlet pressures of 50 psia LOX and 38 psia fuel; and 3) Low power level operation at 17%, 22%, 27%, 40%, 45%, and 50% of Rated Power Level. Use of the highly instrumented TTB engine for this test series has afforded the opportunity to study in great detail engine system operation not possible with a standard SSME and has significantly contributed to a greater understanding of the capabilities of the SSME and liquid rocket engines in general.

Anderson, Brian↗