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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 NTRS1952

Altitude Performance of Annular Combustor Type Turbojet Engine with JFC-2 Fuel /james W. Useller, James L. Harp, Jr. and Zelmar Barson

An investigation was made comparing the performance of JFC-2 fuel and unleaded, clear gasoline in a 3000-pound-thrust turbojet engine. The JFC-2 fuel was a blend of percent diesel fuel and 25 percent aviation gasoline. Engine combustion efficiency was equal to that obtained with gasoline at rated engine speed and altitudes up to 35,000 feet, but at lower engine speeds or at higher altitudes the JFC-2 fuel gave lower combustion efficiency. No discernible difference was obtained in starting or low-speed combustiion blow-out characteristics of the two fuels. Turbine-discharge radial temperature profiles were nearly the same at altitudes up to 35,000 feet.

Useller, James W↗
NASA NTRS2001

Optical Diagnosis of Gas Turbine Combustors Being Conducted

Researchers at the NASA Glenn Research Center, in collaboration with industry, are reducing gas turbine engine emissions by studying visually the air-fuel interactions and combustion processes in combustors. This is especially critical for next generation engines that, in order to be more fuel-efficient, operate at higher temperatures and pressures than the current fleet engines. Optically based experiments were conducted in support of the Ultra-Efficient Engine Technology program in Glenn's unique, world-class, advanced subsonic combustion rig (ASCR) facility. The ASCR can supply air and jet fuel at the flow rates, temperatures, and pressures that simulate the conditions expected in the combustors of high-performance, civilian aircraft engines. In addition, this facility is large enough to support true sectors ("pie" slices of a full annular combustor). Sectors enable one to test true shapes rather than rectangular approximations of the actual hardware. Therefore, there is no compromise to actual engine geometry. A schematic drawing of the sector test stand is shown. The test hardware is mounted just upstream of the instrumentation section. The test stand can accommodate hardware up to 0.76-m diameter by 1.2-m long; thus sectors or small full annular combustors can be examined in this facility. Planar (two-dimensional) imaging using laser-induced fluorescence and Mie scattering, chemiluminescence, and video imagery were obtained for a variety of engine cycle conditions. The hardware tested was a double annular sector (two adjacent fuel injectors aligned radially) representing approximately 15 of a full annular combustor. An example of the two-dimensional data obtained for this configuration is also shown. The fluorescence data show the location of fuel and hydroxyl radical (OH) along the centerline of the fuel injectors. The chemiluminescence data show C2 within the total observable volume. The top row of this figure shows images obtained at an engine low-power condition, and the bottom row shows data from a higher power operating point. The data show distinctly the differences in flame structure between low-power and high-power engine conditions, in both location and amount of species produced (OH, C2) or consumed (fuel). The unique capability of the facility coupled with its optical accessibility helps to eliminate the need for high-pressure performance extrapolations. Tests such as described here have been used successfully to assess the performance of fuel-injection concepts and to modify those designs, if needed.

Hicks, Yolanda R.↗
NASA NTRS1980

Experimental combustor study program

Advanced combustor concepts are evaluated as a means of accommodating possible future broad specification fuels. The three advanced double annular combustor concepts consisted of (1) a concept employing high pressure drop fuel nozzles for improved atomization, (2) a concept with premixing tubes in the main stage, and (3) a concept with the pilot stage on the inside and the main stage on the sideout, which is the reverse of the other two concepts. All of the advanced concepts show promise for reduced sensitivity to fuel hydrogen content. Some hardware problems were encountered, but these problems could be quickly resolved if refinement tests were conducted. The design with the premixing main stage was selected for a parametric test because of its low NOx emissions level, carbon free dome, and very low dome temperatures which were essentially independent of fuel type. The other advanced designs also had low done temperatures. The premixing dome design liner temperatures exhibited less sensitivity to fuel type than did the base-line combustor, although more sensitivity than observed for concept 1. The inner liner hot spot and the observed smoke results for the premixing design suggest that the fuel-air mixture was not as uniform as desired.

Kasper, J. M.↗
NASA NTRSDate not supplied

Advanced Diagnostics for Megahertz Imaging of Mixing, Fuel Spray, and Combustion Processes for Rotating Detonation Combustors

Recent advancements in megahertz rate, high-power, burst-mode laser technology are leveraged to perform and explore imaging measurements that spatially and temporally resolve the mixing, combustion, and detonation flow field in two laboratory-scale rotating detonation combustors (RDCs). In a non-premixed annular RDC, multiple imaging diagnostics are explored to investigate gaseous and liquid injector behavior, the detonation wave structure, and the propellant refill. In one instance, OH planar laser-induced fluorescence (OH-PLIF) imaging is performed up to a 2 MHz repetition rate to track the combustion products and reaction zone locations. In the same annular RDC, a single liquid fuel jet is injected, and laser-based 355-nm imaging of the fuel spray is performed up to a 1 MHz repetition rate. For this configuration, the annular RDC is used as a detonation driver to impose periodic detonation waves to interact with the fuel spray. Moreover, in this annular RDC, a range of tracer-based laser imaging measurements are explored to time-resolve the unsteady oxidizer air recovery and refill process. In a non-premixed linear RDC, planar imaging measurements of the fuel mixing are performed up to a 200 kHz repetition rate using PLIF of a tracer in the fuel supply. The fuel mixing imaging helps explain the origin of the observed pre and post wave burning, detonation structure, and enables quantifying injector recovery timescales. This paper will provide a high-level broad survey of diagnostics applied in these RDCs, lessons learned, and interesting observations.

Combustion↗
NASA NTRSDate not supplied

Advanced Diagnostics For Megahertz Imaging Of Mixing, Fuel Spray, And Combustion Processes For Rotating Detonation Combustors

Recent advancements in megahertz rate, high-power, burst-mode laser technology are leveraged to perform and explore imaging measurements that spatially and temporally resolve the mixing, combustion, and detonation flow field in two laboratory-scale rotating detonation combustors (RDCs). In a non-premixed annular RDC, multiple imaging diagnostics are explored to investigate gaseous and liquid injector behavior, the detonation wave structure, and the propellant refill. In one instance, OH planar laser-induced fluorescence (OH-PLIF) imaging is performed up to a 2 MHz repetition rate to track the combustion products and reaction zone locations. In the same annular RDC, a single liquid fuel jet is injected, and laser-based 355-nm imaging of the fuel spray is performed up to a 1 MHz repetition rate. For this configuration, the annular RDC is used as a detonation driver to impose periodic detonation waves to interact with the fuel spray. Moreover, in this annular RDC, a range of tracer-based laser imaging measurements are explored to time-resolve the unsteady oxidizer air recovery and refill process. In a non-premixed linear RDC, planar imaging measurements of the fuel mixing are performed up to a 200 kHz repetition rate using PLIF of a tracer in the fuel supply. The fuel mixing imaging helps explain the origin of the observed pre and post wave burning, detonation structure, and enables quantifying injector recovery timescales. This paper will provide a high-level broad survey of diagnostics applied in these RDCs, lessons learned, and interesting observations.

Combustion↗
NASA NTRS1979

Analytical evaluation of the impact of broad specification fuels on high bypass turbofan engine combustors

Six conceptual combustor designs for the CF6-50 high bypass turbofan engine and six conceptual combustor designs for the NASA/GE E3 high bypass turbofan engine were analyzed to provide an assessment of the major problems anticipated in using broad specification fuels in these aircraft engine combustion systems. Each of the conceptual combustor designs, which are representative of both state-of-the-art and advanced state-of-the-art combustion systems, was analyzed to estimate combustor performance, durability, and pollutant emissions when using commercial Jet A aviation fuel and when using experimental referee board specification fuel. Results indicate that lean burning, low emissions double annular combustor concepts can accommodate a wide range of fuel properties without a serious deterioration of performance or durability. However, rich burning, single annular concepts would be less tolerant to a relaxation of fuel properties. As the fuel specifications are relaxed, autoignition delay time becomes much smaller which presents a serious design and development problem for premixing-prevaporizing combustion system concepts.

Taylor, J. R.↗
NASA NTRS1973

Nuclear characteristics of a fissioning uranium plasma test reactor with light-water cooling

An analytical study was performed to determine a design configuration for a cavity test reactor. Test section criteria were that an average flux of 10 to the 15th power neutrons/sq cm/sec (E less than or equal to 0.12 eV) be supplied to a 61-cm-diameter spherical cavity at 200-atm pressure. Design objectives were to minimize required driver power, to use existing fuel-element technology, and to obtain fuel-element life of 10 to 100 full-power hours. Parameter calculations were made on moderator region size and material, driver fuel arrangement, control system, and structure in order to determine a feasible configuration. Although not optimized, a configuration was selected which would meet design criteria. The driver fuel region was a cylindrical annular region, one element thick, of 33 MTR-type H2O-cooled elements (Al-U fuel plate configuration), each 101 cm long. The region between the spherical test cavity and the cylindrical driver fuel region was Be (10 vol. % H2O coolant) with a midplane dimension of 8 cm. Exterior to the driver fuel, the 25-cm-thick cylindrical and axial reflectors were also Be with 10 vol. % H2O coolant. The entire reactor was contained in a 10-cm-thick steel pressure vessel, and the 200-atm cavity pressure was equalized throughout the driver reactor. Fuel-element life was 50 hr at the required driver power of 200 MW. Reactor control would be achieved with rotating poison drums located in the cylindrical reflector region. A control range of about 18 percent delta k/k was required for reactor operation.

Whitmarsh, C. L., Jr.↗
NASA NTRS1956

Performance of a short combustor at high altitudes using hydrogen fuel

Performance characteristics of a 16-inch annular-type combustor installed in a full-scale engine using gaseous-hydrogen fuel were obtained at simulated altitudes from 66,000 to 86,000 feet at a flight Mach number of 0.8. Combustion efficiencies of 86 percent were obtained at 86,000 feet (combustor pressure, 420 lb/sq ft abs). Combustor blowout was not encountered during the investigation.

Sivo, Joseph N↗
NASA NTRS1976

Performance and emission characteristics of swirl-can combustors to near-stoichiometric fuel-air ratio

Emissions and performance characteristics were determined for two full annular swirl-can combustors operated to near stoichiometric fuel-air ratio. Test condition variations were as follows: combustor inlet-air temperatures, 589, 756, 839, and 894 K; reference velocities, 24 to 37 meters per second; inlet pressure, 62 newtons per square centimeter; and fuel-air ratios, 0.015 to 0.065. The combustor average exit temperature and combustor efficiency were calculated from the combustor exhaust gas composition. For fuel-air ratios greater than 0.04, the combustion efficiency decreased with increasing fuel-air ratios in a near-linear manner. Increasing the combustor inlet air temperature tended to offset this decrease. Maximum oxides of nitrogen emission indices occurred at intermediate fuel-air ratios and were dependent on combustor design. Carbon monoxide levels were extremely high and were the primary cause of poor combustion efficiency at the higher fuel-air ratios. Unburned hydrocarbons were low for all test conditions. For high fuel-air ratios SAE smoke numbers greater than 25 were produced, except at the highest inlet-air temperatures.

Diehl, L. A.↗
NASA NTRS1978

Correlation of combustor acoustic power levels inferred from internal fluctuating pressure measurements

Combustion chamber acoustic power levels inferred from internal fluctuating pressure measurements are correlated with operating conditions and chamber geometries over a wide range. The variables include considerations of chamber design (can, annular, and reverse-flow annular) and size, number of fuel nozzles, burner staging and fuel split, airflow and heat release rates, and chamber inlet pressure and temperature levels. The correlated data include those obtained with combustion component development rigs as well as engines.

Vonglahn, U. H.↗
NASA NTRS1978

Correlation of combustor acoustic power levels inferred from internal fluctuating pressure measurements

Combustion chamber acoustic power levels inferred from internal fluctuating pressure measurements are correlated with operating conditions and chamber geometries over a wide range. The variables include considerations of chamber design (can, annular, and reverse-flow annular) and size, number of fuel nozzles, burner staging and fuel split, airflow and heat release rates, and chamber inlet pressure and temperature levels. The correlated data include those obtained with combustion component development rigs as well as engines.

Von Glahn, U. H.↗
NASA NTRS1973

Effect of fuel zoning and fuel nozzle design on pollution emissions at ground idle conditions for a double-annular ram-induction combustor

An exhaust emission survey was conducted on a double-annular ram induction combustor at simulated ground idle conditions. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. The emission levels of total hydrocarbon (THC), carbon monoxide, carbon dioxide, and nitric oxide were measured. The effects of fuel zoning, fuel nozzle design, and operating conditions (inlet temperature and reference Mach number) on the level of these emissions were determined. At an overall combustor fuel/air ratio of 0.007, fuel zoning reduced THC emissions by a factor of 5 to 1. The reduction in THC emissions is attributed to the increase in local fuel/air ratio provided by the fuel zoning. An alternative method of increasing fuel/air ratio would be to operate with larger-than-normal compressor overboard bleed; however, analysis on this method indicated an increase in idle fuel consumption of 20 percent. The use of air-atomizing nozzles reduced the THC emissions by 2 to 1.

Clements, T. R.↗
NASA NTRS1969

Development of an Advanced Annular Combustor

The objective of the effort described in this report was to determine the structural durability of a full-scale advanced annular turbojet combustor using ASTM A-1 type fuel and operating at conditions typical of advanced supersonic aircraft. A full-scale annular combustor of the ram-induction type was fabricated and subjected to a 325-hour cyclic endurance test at conditions representative of operation in a Mach 3.0 aircraft. The combustor exhibited extensive cracking and scoop burning at the end of the test program. But these defects had no appreciable effect on combustor performance, as performance remained at a high level throughout the endurance program. Most performance goals were achieved with pressure loss values near 6% and 8%, and temperature rise variation ratio (deltaTVR) values near 1.25 and l.22 at takeoff and cruise conditions, respectively. Combustion efficiencies approached l004 and the exit radial temperature profiles were approximately as desired.

Rusnak, J. P.↗
NASA NTRS1979

Experimental clean combustor program: Diesel no. 2 fuel addendum, phase 3

A CF6-50 engine equipped with an advanced, low emission, double annular combustor was operated 4.8 hours with No. 2 diesel fuel. Fourteen steady-state operating conditions ranging from idle to full power were investigated. Engine/combustor performance and exhaust emissions were obtained and compared to JF-5 fueled test results. With one exception, fuel effects were very small and in agreement with previously obtained combustor test rig results. At high power operating condition, the two fuels produced virtually the same peak metal temperatures and exhaust emission levels. At low power operating conditions, where only the pilot stage was fueled, smoke levels tended to be significantly higher with No. 2 diesel fuel. Additional development of this combustor concept is needed in the areas of exit temperature distribution, engine fuel control, and exhaust emission levels before it can be considered for production engine use.

Gleason, C. C.↗