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Results for “combustion efficiency”

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 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1948

Simulated Altitude Performance of Combustors for the 24C Turbojet Engine. III - Performance of Rectangular-Slot Baskets: Performance of Rectangular-Slot Baskets - 3

The performance of an annular combustion chamber from a 24C turbojet engine was investigated over a range of simulated altitudes from 20,000 to 55,000 feet and corrected engine rotor speeds from 6000 to 13,000 rpm at a simulated ram-pressure ratio of 1.04. The purpose of the investigation was to determine the effects on the altitude operational limits, combustor-outlet gas temperature distribution, combustion efficiencies, and combustor inlet-to-outlet total-pressure drops of two changes in the 24C-4B basket air-passage arrangements that were designed to improve combustor-outlet temperature distribution. These changes were: (a) replacement of the downstream secondary air holes with large rectangular slots further upstream (rectangular-slot basket), and (b) enlargement of anticoking holes in the rectangular-slot basket (modified rectangular-slot basket). The results indicate that improved outlet-gas temperature distribution of each succeeding combustor basket investigated was attained at a sacrifice in the altitude limit of operation. The altitude limits of operation of the combustor with the original basket ranged from 34,000 feet at a corrected engine speed of 6000 rpm to a maximum of 52,000 feet at 12 ' 500 rpm. The altitude limits of the rectangular-slot basket were about 2000 feet lower throughout the engine speed range than those of the original basket. The altitude limits of the combustor with the modified rectangular-slot basket were about equivalent to those of the other baskets in the corrected-engine-speed range from 12,000 to 12,500 rpm but were about 10,000 feet lower than those of the original basket in the corrected-engine-speed range from 6000 to 9000 rpm. For the same inlet-air conditions, the combustion efficiencies were highest for the original basket and progressively lower for each of the other two baskets. The combustor inlet-to-outlet pressure drops of all three combustor baskets at the same operating conditions were within +/- 10 percent of the pressure drop of the original basket.

Tischler, Adelbert O.↗
NASA NTRS1993

Test results from a simple, low-cost, pressure-fed liquid hydrogen/liquid oxygen rocket combustor

A simple, low-cost rocket engine was designed, fabricated, and successfully hot fire tested over a wide range of interface conditions and operating parameters. The engine used low enthalpy hydrogen (45 to 70 R, 200 to 390 psia) and oxygen (139 to 163 R, 210 to 480 psia) propellants pressure-fed directly from facility cryogenic tanks. The engine demonstrated excellent performance, with 97% average combustion efficiency, and absence of combustion instabilities. Engine design chamber pressure was 300 psia, yielding about 16,500 pounds thrust at sea level with a 3:1 expansion ration test nozzle. The engine used a fixed-element injector based on TRW's unique coaxial pintle design, but was operated at 60%, 80%, and 100% thrust levels by throttling facility propellant valves. The engine was tested at propellant mixture ratios (O/F) from 5.8 to 8.4; design O/F was 6.6. To document combustion stability, in five tests RDX explosive pulse guns were detonated in radial and tangential directions across the combustion chamber during steady-state operation. The largest disturbance consisted of simultaneous detonation of a 20-grain radial gun and a 40-grain tangential gun. In no case was an instability, either feed system mode or chamber acoustic mode, excited. High-frequency piezoelectric pressure transducers documented stable recovery from disturbance overpressures within 40 milliseconds of peak pressure. A total of 67 firing tests, accumulating 149 seconds of firing time above 10% P(sub c), were performed. Since parametric testing required run durations of only 2 to 3 seconds, a heat sink combustion chamber was employed for most runs. To evaluate the feasibility of a low-cost ablative system for a flight engine design, one 20-second continuous firing was conducted with a silicone rubber chamber/throat/nozzle liner cast in one piece directly into the engine. The ablative engine operated at the equivalent of 309 seconds sea level specific impulse, when adjusted to a 98% efficient 6:1 expansion ration nozzle, and 430 seconds vacuum specific impulse, when adjusted to a 98% efficient 50:1 expansion ratio nozzle. This engine and test series represent an initial subscale demonstration of a new booster-class engine that eliminates the cost and complexity associated with regenerative cooling and typical engine cycles. This paper presents a description of the engine design and discussion and summary data plots of the performance measured during the parametric testing.

Dressler, G. A.↗
NASA NTRS2016

Alternative Bio-Derived JP-8 Class Fuel and JP-8 Fuel: Flame Tube Combustor Test Results Compared using a GE TAPS Injector Configuration

This paper presents results from tests in a flame tube facility, where a bio-derived alternate fuel was compared with JP-8 for emissions and general combustion performance. A research version of General Electric Aviation (GE) TAPS injector was used for the tests. Results include combustion efficiency from gaseous emission measurements, 2D planar laser-based imaging as well as basic flow visualization of the flame. Four inlet test conditions were selected that simulate various engine power conditions relevant to NASA Fundamental Aeronautics Supersonics Project and Environmentally Responsible Aviation Program. One inlet condition was a pilot-only test point. The other three inlet conditions incorporated fuel staging via a split between the pilot and main circuits of either 10%/90% or 20%/80%. For each engine power condition, three fuel mixes were used: 100% JP-8; 100% alternative; and a blend of the two, containing 75% alternative. Results for the inlet cases that have fuel split between pilot and main, indicate that fuel from the pilot appears to be evaporated by the time it reaches the dome exit. Main circuit liquid evaporates within a downstream distance equal to annulus height, no matter the fuel. Some fuel fluorescence images for a 10%/90% fuel staging case show a distinct difference between JP-8 and bio-derived fuel. OH PLIF results indicate that OH forms in a region more centrally-located for the JP-8 case downstream of the pilot, in its central recirculation region (CRZ). For the bio-derived Hydrotreated Renewable Jet (HRJ) fuel, however, we do not see much OH in the CRZ. The OH image structure near the dome exit is similar for the two fuels, but farther downstream the OH in the CRZ is much more apparent for the JP-8 than for the alternate fuel. For all conditions, there was no discernable difference between fuel types in combustion efficiency or emissions.

gas turbine combustors↗
NASA NTRS1947

Effect of combustor-inlet conditions on performance of an annular turbojet combustor

The combustion performance, and particularly the phenomenon of altitude operational limits, was studied by operating the annular combustor of a turbojet engine over a range of conditions of air flow, inlet pressure, inlet temperature, and fuel flow. Information was obtained on the combustion efficiencies, the effect on combustion of inlet variables, the altitude operational limits with two different fuels, the pressure losses in the combustor, the temperature and velocity profiles at the combustor outlet, the extent of afterburning, the fuel-injection characteristics, and the condition of the combustor basket.

Childs, J Howard↗
NASA NTRS1998

Results of a 24-inch Hybrid Motor Performance Uncertainty Analysis

The subscale (11 and 24-inch) hybrid motors at the Marshall Space Flight Center (MSFC) have been used as versatile and cost effective testbeds for developing new technology. Comparisons between motor configuration, ignition systems, feed systems, fuel formulations, and nozzle materials have been carried out without detailed consideration as to haw "good" the motor performance data were. For the 250,000 lb/thrust motor developed by the Hybrid Propulsion Demonstration Program consortium, this shortcoming is particularly risky because motor performance will likely be used as put of a set of downselect criteria to choose between competing ignition and feed systems under development. This analysis directly addresses that shortcoming by applying uncertainty analysis techniques to the experimental determination of the characteristic velocity, theoretical characteristic velocity, and characteristic velocity efficiency for a 24-inch motor firing. With the adoption of fuel-lined headends, flow restriction, and aft mixing chambers, state of the an 24-inch hybrid motors have become very efficient However, impossibly high combustion efficiencies (some computed as high as 108%) have been measured in some tests with 11-inch motors. This analysis has given new insight into explaining how these efficiencies were measured to be so high, and into which experimental measurements contribute the most to the overall uncertainty.

Sims, Joseph D.↗
NASA NTRS1951

Effects of Altitude on Turbojet Engine Performance

Component and over-all performance characteristics of several turbojet engines investigated in the altitude facilities of the NACA Lewis Laboratory during the last several years are summarized to indicate the effects of altitude on turbojet engine performance. Data presented show that failure of turbojet engine performance to generalize for all altitudes can be traced to reductions in compressor efficiency, corrected air flow, and combustion efficiency at altitude. In addition, it is shown that although engines of different design may have equal thrusts at sea level, the thrusts at altitude may vary widely because of differences in compressor performance characteristics from one engine to another.

COMPRESSORS - EFFICIENCY↗
NASA NTRS1981

Design, fabrication, test, and delivery of a high-pressure oxygen/RP-1 injector

A summary of the design analyses for a liquid rocket injector using oxygen and RP-1 propellants at high chamber pressures of 20,682 kPa (3000 psia) is presented. This analytical investigation includes combustion efficiency versus injector element type, combustion stability, and combustor cooling requirements. The design and fabrication of a subscale injector/acoustic resonantor assembly capable of providing a nominal thrust of 222K N (50,000 lbF) is presented.

Schoenman, L.↗
NASA NTRS1957

Investigation of Combustion in 16-inch Ram Jet Under Simulated Conditions of High Altitude and High Mach Number

Results obtained with three different burner configurations in a connected-pipe investigation of a 16-inch ram jet are presented. The radial position of the fuel injector and the engine-outlet area both affected burner performance. For a given configuration, only slight changes in total-pressure ratio across the combustion chamber were obtained over the complete range of operation. With one burner, combustion efficiencies obtained at a combustion-chamber-inlet static pressure of 800 pounds per square foot absolute were greater than those obtained at 1920 pounds per square foot absolute.

ENGINES, RAM JET↗
NASA NTRS1952

Evaluation of five conical center-body supersonic diffusers at several angles of attack

Five supersonic inlets having conical center bodies were investigated in a 16-inch ram jet engine with and without combustion at Mach numbers of 1.7, 1.9, and 2.0 and angles of attack from 0 degrees to 10 degrees. Four of the inlets were of the low-mass-flow-ratio type whereas the fifth inlet was of the high-mass-flow-ratio type. The stable operating range of the four low-nass-flow-ratio inlets decreased considerably as the angle of attack was increased despite an increase of the cone angle or the use of boundary-layer bleed. The stable range of the high-mass-flow-ratio inlet increased considerably both at 0 degrees and at greater angles of attack as the free-stream Mach number was decreased. Free-stream Mach number, angle of attack, and inlet type had negligible effects on combustion efficiency, temperature ratio, or combustion-chamber inlet Mach number in the supercritical region.

Englert, Gerald W↗
NASA NTRS2017

Comparison of Mixing Characteristics for Several Fuel Injectors on an Open Plate and in a Ducted Flowpath Configuration at Hypervelocity Flow Conditions

In order to reduce the cost and complexity associated with fuel injection and mixing experiments for high-speed flows, and to further enable optical access to the test section for nonintrusive diagnostics, the Enhanced Injection and Mixing Project (EIMP) utilizes an open flat plate configuration to characterize inert mixing properties of various fuel injectors for hypervelocity applications. The experiments also utilize reduced total temperature conditions to alleviate the need for hardware cooling. The use of "cold" flows and non-reacting mixtures for mixing experiments is not new, and has been extensively utilized as a screening technique for scramjet fuel injectors. The impact of reduced facility-air total temperature, and the use of inert fuel simulants, such as helium, on the mixing character of the flow has been assessed in previous numerical studies by the authors. Mixing performance was characterized for three different injectors: a strut, a ramp, and a flushwall. The present study focuses on the impact of using an open plate to approximate mixing in the duct. Toward this end, Reynolds-averaged simulations (RAS) were performed for the three fuel injectors in an open plate configuration and in a duct. The mixing parameters of interest, such as mixing efficiency and total pressure recovery, are then computed and compared for the two configurations. In addition to mixing efficiency and total pressure recovery, the combustion efficiency and thrust potential are also computed for the reacting simulations.

Drozda, Tomasz G.↗
NASA NTRS2018

Comparison of Mixing Characteristics for Several Fuel Injectors at Mach 8, 12, and 15 Hypervelocity Flow Conditions

CFD analysis is presented of the mixing characteristics and performance of three fuel injectors at flight Mach numbers of 8, 12, and 15. The Reynolds-averaged simulations (RAS) were carried out using the VULCAN-CFD solver. The high Mach number flow conditions match those of the experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies relevant to flight Mach numbers greater than 8. The injectors include a fuel placement device, a strut, and a fluidic vortical mixer, a ramp. These fuel injectors accomplish the necessary task of distributing and mixing fuel into the supersonic cross-flow, albeit via different strategies. For comparison, a flush-wall injector is also included. This type of injector generally represents the simplest method of introducing fuel into a scramjet combustor. The three injectors represent the baseline configurations of the EIMP experiments. The mixing parameters of interest, such as mixing efficiency and total pressure recovery, are computed from the RAS and compared for the three flight conditions and injector configurations. In addition to mixing efficiency and total pressure recovery, the combustion efficiency and thrust potential are also computed for the reacting simulations. Plotting the total pressure recovery and thrust potential as a function of mixing efficiency provides added insight into critical aspects of combustor performance as the flight condition and injector type are varied.

Shenoy, Rajiv R.↗