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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Combustor technology for future small gas turbine aircraft

Future engine cycles proposed for advanced small gas turbine engines will increase the severity of the operating conditions of the combustor. These cycles call for increased overall engine pressure ratios which increase combustor inlet pressure and temperature. Further, the temperature rise through the combustor and the corresponding exit temperature also increase. Future combustor technology needs for small gas turbine engines is described. New fuel injectors with large turndown ratios which produce uniform circumferential and radial temperature patterns will be required. Uniform burning will be of greater importance because hot gas temperatures will approach turbine material limits. The higher combustion temperatures and increased radiation at high pressures will put a greater heat load on the combustor liners. At the same time, less cooling air will be available as more of the air will be used for combustion. Thus, improved cooling concepts and/or materials requiring little or no direct cooling will be required. Although presently there are no requirements for emissions levels from small gas turbine engines, regulation is expected in the near future. This will require the development of low emission combustors. In particular, nitrogen oxides will increase substantially if new technologies limiting their formation are not evolved and implemented. For example, staged combustion employing lean, premixed/prevaporized, lean direct injection, or rich burn-quick quench-lean burn concepts could replace conventional single stage combustors.

Lyons, Valerie J.↗
NASA NTRS1991

Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

In order to increase the performance of a high performance, advanced expander-cycle engine combustor, higher chamber pressures are required. In order to increase chamber pressure, more heat energy is required to be transferred to the combustor coolant circuit fluid which drives the turbomachinery. This requirement was fulfilled by increasing the area exposed to the hot-gas by using combustor ribs. A previous technology task conducted 2-d hot air and cold flow tests to determine an optimum rib height and configuration. In task C.5 a combustor calorimeter was fabricated with the optimum rib configuration, 0.040 in. high ribs, in order to determine their enhancing capability. A secondary objective was to determine the effects of mixture ratio changers on the enhancement during hot-fire testing. The program used the Rocketdyne Integrated Component Evaluator (ICE) reconfigured into a thrust chamber only mode. The test results were extrapolated to give a projected enhancement from the ribs for a 16 in. long cylindrical combustor at 15 Klb nominal thrust level. The hot-gas wall ribs resulted in a 58 percent increase in heat transfer. When projected to a full size 15K combustor, it becomes a 46 percent increase. The results of those tests, a comparison with previous 2-d results, the effects of mixture ratio and combustion gas flow on the ribs and the potential ramifications for expander cycle combustors are detailed.

Brown, William S.↗
NASA NTRS1994

The effect of incomplete fuel-air mixing on the lean blowout limit, lean stability limit and NO(x) emissions in lean premixed gas turbine combustors

Gas turbine engines for both land-based and aircraft propulsion applications are facing regulations on NOx emissions which cannot be met with current combustor technology. A number of alternative combustor strategies are being investigated which have the potential capability of achieving ultra-low NOx emissions, including lean premixed combustors, direct injection combustors, rich burn-quick quench-lean burn combustors and catalytic combustors. The research reported in this paper addresses the effect of incomplete fuel-air mixing on the lean limit performance and the NOx emissions characteristics of lean premixed combustors.

Shih, W.-P.↗
NASA NTRS2001

Experimental and Computational Study of Trapped Vortex Combustor Sector Rig with Tri-pass Diffuser

The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL (Rich burn/Quick mix/Lean burn) modes of combustion. The present work describes the operational principles of the TVC, and provides detailed performance data on a configuration featuring a tri-pass diffusion system. Performance data include EINOx (NO(sub x) emission index) results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable in comparison to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

Hendricks, Robert C.↗
NASA NTRS2002

Measurement and Computation of Supersonic Flow in a Lobed Diffuser-Mixer for Trapped Vortex Combustors

The trapped vortex combustor (TVC) pioneered by Air Force Research Laboratories (AFRL) is under consideration as an alternative to conventional gas turbine combustors. The TVC has demonstrated excellent operational characteristics such as high combustion efficiency, low NO(x) emissions, effective flame stabilization, excellent high-altitude relight capability, and operation in the lean-burn or rich burn-quick quench-lean burn (RQL) modes of combustion. It also has excellent potential for lowering the engine combustor weight. This performance at low to moderate combustor mach numbers has stimulated interest in its ability to operate at higher combustion mach number, and for aerospace, this implies potentially higher flight mach numbers. To this end, a lobed diffuser-mixer that enhances the fuel-air mixing in the TVC combustor core was designed and evaluated, with special attention paid to the potential shock system entering the combustor core. For the present investigation, the lobed diffuser-mixer combustor rig is in a full annular configuration featuring sixfold symmetry among the lobes, symmetry within each lobe, and plain parallel, symmetric incident flow. During hardware cold-flow testing, significant discrepancies were found between computed and measured values for the pitot-probe-averaged static pressure profiles at the lobe exit plane. Computational fluid dynamics (CFD) simulations were initiated to determine whether the static pressure probe was causing high local flow-field disturbances in the supersonic flow exiting the diffuser-mixer and whether shock wave impingement on the pitot probe tip, pressure ports, or surface was the cause of the discrepancies. Simulations were performed with and without the pitot probe present in the modeling. A comparison of static pressure profiles without the probe showed that static pressure was off by nearly a factor of 2 over much of the radial profile, even when taking into account potential axial displacement of the probe by up to 0.25 in. (0.64 cm). Including the pitot probe in the CFD modeling and data interpretation lead to good agreement between measurement and prediction. Graphical inspection of the results showed that the shock waves impinging on the probe surface were highly nonuniform, with static pressure varying circumferentially among the pressure ports by over 10 percent in some cases. As part of the measurement methodology, such measurements should be routinely supplemented with CFD analyses that include the pitot probe as part of the flow-path geometry.

Brankovic, Andreja↗
NASA NTRS2004

Experimental and Computational Study of Trapped Vortex Combustor Sector Rig With Tri-Pass Diffuser

The Trapped Vortex Combustor (TVC) potentially offers numerous operational advantages over current production gas turbine engine combustors. These include lower weight, lower pollutant emissions, effective flame stabilization, high combustion efficiency, excellent high altitude relight capability, and operation in the lean burn or RQL modes of combustion. The present work describes the operational principles of the TVC, and extends diffuser velocities toward choked flow and provides system performance data. Performance data include EINOx results for various fuel-air ratios and combustor residence times, combustion efficiency as a function of combustor residence time, and combustor lean blow-out (LBO) performance. Computational fluid dynamics (CFD) simulations using liquid spray droplet evaporation and combustion modeling are performed and related to flow structures observed in photographs of the combustor. The CFD results are used to understand the aerodynamics and combustion features under different fueling conditions. Performance data acquired to date are favorable compared to conventional gas turbine combustors. Further testing over a wider range of fuel-air ratios, fuel flow splits, and pressure ratios is in progress to explore the TVC performance. In addition, alternate configurations for the upstream pressure feed, including bi-pass diffusion schemes, as well as variations on the fuel injection patterns, are currently in test and evaluation phases.

Hendricks, R. C.↗
NASA NTRS2002

Sector Tests of a Low-NO(sub x), Lean, Direct- Injection, Multipoint Integrated Module Combustor Concept Conducted

The low-emissions combustor development described is directed toward advanced high pressure aircraft gas-turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low power conditions. Low-NOx combustors can be classified into rich-burn and lean-burn concepts. Lean-burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) concepts. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibility of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone, and thus, it does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, good atomization is necessary and the fuel must be mixed quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP. The LDI concept described is a multipoint fuel injection/multiburning zone concept. Each of the multiple fuel injectors has an air swirler associated with it to provide quick mixing and a small recirculation zone for burning. The multipoint fuel injection provides quick, uniform mixing and the small multiburning zones provide for reduced burning residence time, resulting in low NOx formation. An integrated-module approach was used for the construction where chemically etched laminates, diffusion bonded together, combine the fuel injectors, air swirlers, and fuel manifold into a single element. The multipoint concept combustor was demonstrated in a 15 sector test. The configuration tested had 36 fuel injectors and fuel-air mixers that replaced two fuel injectors in a conventional dual-annular combustor. During tests, inlet temperatures were up to 870 K and inlet pressures were up to 5400 kPa. A correlation was developed that related the NOx emissions with the inlet temperature, inlet pressure, fuel-air ratio, and pressure drop. At low-power conditions, fuel staging was used so that high combustion efficiency was obtained with only one-fourth of the fuel injectors flowing. The test facility had optical access, and visual images showed the flame to be very short, approximately 25 mm long.

Tacina, Robert R.↗
NASA NTRS2003

Active Control of High-Frequency Combustor Instability Demonstrated

To reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities-high-pressure oscillations much like sound waves that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the combustor and turbine safe operating life. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Propulsion and Power Program, the NASA Glenn Research Center in partnership with Pratt & Whitney, United Technologies Research Center, and Georgia Institute of Technology is developing technologies for the active control of combustion instabilities.

DeLaat, John C.↗
NASA NTRS2003

Hydrogen Fuel Capability Added to Combustor Flametube Rig

Facility capabilities have been expanded at Test Cell 23, Research Combustor Lab (RCL23) at the NASA Glenn Research Center, with a new gaseous hydrogen fuel system. The purpose of this facility is to test a variety of fuel nozzle and flameholder hardware configurations for use in aircraft combustors. Previously, this facility only had jet fuel available to perform these various combustor flametube tests. The new hydrogen fuel system will support the testing and development of aircraft combustors with zero carbon dioxide (CO2) emissions. Research information generated from this test rig includes combustor emissions and performance data via gas sampling probes and emissions measuring equipment. The new gaseous hydrogen system is being supplied from a 70 000-standard-ft3 tube trailer at flow rates up to 0.05 lb/s (maximum). The hydrogen supply pressure is regulated, and the flow is controlled with a -in. remotely operated globe valve. Both a calibrated subsonic venturi and a coriolis mass flowmeter are used to measure flow. Safety concerns required the placement of all hydrogen connections within purge boxes, each of which contains a small nitrogen flow that is vented past a hydrogen detector. If any hydrogen leaks occur, the hydrogen detectors alert the operators and automatically safe the facility. Facility upgrades and modifications were also performed on other fluids systems, including the nitrogen gas, cooling water, and air systems. RCL23 can provide nonvitiated heated air to the research combustor, up to 350 psig at 1200 F and 3.0 lb/s. Significant modernization of the facility control systems and the data acquisition systems was completed. A flexible control architecture was installed that allows quick changes of research configurations. The labor-intensive hardware interface has been removed and changed to a software-based system. In addition, the operation of this facility has been greatly enhanced with new software programming and graphic operator interface stations. Glenn s RCL23 facility systems were successfully checked out in the spring of 2002, and hydrogen combustor research testing began in the summer of 2002.

Frankenfield, Bruce J.↗
NASA NTRS2008

Characterization and Simulation of the Thermoacoustic Instability Behavior of an Advanced, Low Emissions Combustor Prototype

Extensive research is being done toward the development of ultra-low-emissions combustors for aircraft gas turbine engines. However, these combustors have an increased susceptibility to thermoacoustic instabilities. This type of instability was recently observed in an advanced, low emissions combustor prototype installed in a NASA Glenn Research Center test stand. The instability produces pressure oscillations that grow with increasing fuel/air ratio, preventing full power operation. The instability behavior makes the combustor a potentially useful test bed for research into active control methods for combustion instability suppression. The instability behavior was characterized by operating the combustor at various pressures, temperatures, and fuel and air flows representative of operation within an aircraft gas turbine engine. Trends in instability behavior versus operating condition have been identified and documented, and possible explanations for the trends provided. A simulation developed at NASA Glenn captures the observed instability behavior. The physics-based simulation includes the relevant physical features of the combustor and test rig, employs a Sectored 1-D approach, includes simplified reaction equations, and provides time-accurate results. A computationally efficient method is used for area transitions, which decreases run times and allows the simulation to be used for parametric studies, including control method investigations. Simulation results show that the simulation exhibits a self-starting, self-sustained combustion instability and also replicates the experimentally observed instability trends versus operating condition. Future plans are to use the simulation to investigate active control strategies to suppress combustion instabilities and then to experimentally demonstrate active instability suppression with the low emissions combustor prototype, enabling full power, stable operation.

DeLaat, John C.↗
NASA NTRS2008

Characterization and Simulation of Thermoacoustic Instability in a Low Emissions Combustor Prototype

Extensive research is being done toward the development of ultra-low-emissions combustors for aircraft gas turbine engines. However, these combustors have an increased susceptibility to thermoacoustic instabilities. This type of instability was recently observed in an advanced, low emissions combustor prototype installed in a NASA Glenn Research Center test stand. The instability produces pressure oscillations that grow with increasing fuel/air ratio, preventing full power operation. The instability behavior makes the combustor a potentially useful test bed for research into active control methods for combustion instability suppression. The instability behavior was characterized by operating the combustor at various pressures, temperatures, and fuel and air flows representative of operation within an aircraft gas turbine engine. Trends in instability behavior vs. operating condition have been identified and documented. A simulation developed at NASA Glenn captures the observed instability behavior. The physics-based simulation includes the relevant physical features of the combustor and test rig, employs a Sectored 1-D approach, includes simplified reaction equations, and provides time-accurate results. A computationally efficient method is used for area transitions, which decreases run times and allows the simulation to be used for parametric studies, including control method investigations. Simulation results show that the simulation exhibits a self-starting, self-sustained combustion instability and also replicates the experimentally observed instability trends vs. operating condition. Future plans are to use the simulation to investigate active control strategies to suppress combustion instabilities and then to experimentally demonstrate active instability suppression with the low emissions combustor prototype, enabling full power, stable operation.

DeLaat, John C.↗
NASA NTRS2012

A Comparison of Combustor-Noise Models

The present status of combustor-noise prediction in the NASA Aircraft Noise Prediction Program (ANOPP)1 for current-generation (N) turbofan engines is summarized. Several semi-empirical models for turbofan combustor noise are discussed, including best methods for near-term updates to ANOPP. An alternate turbine-transmission factor2 will appear as a user selectable option in the combustor-noise module GECOR in the next release. The three-spectrum model proposed by Stone et al.3 for GE turbofan-engine combustor noise is discussed and compared with ANOPP predictions for several relevant cases. Based on the results presented herein and in their report,3 it is recommended that the application of this fully empirical combustor-noise prediction method be limited to situations involving only General-Electric turbofan engines. Long-term needs and challenges for the N+1 through N+3 time frame are discussed. Because the impact of other propulsion-noise sources continues to be reduced due to turbofan design trends, advances in noise-mitigation techniques, and expected aircraft configuration changes, the relative importance of core noise is expected to greatly increase in the future. The noise-source structure in the combustor, including the indirect one, and the effects of the propagation path through the engine and exhaust nozzle need to be better understood. In particular, the acoustic consequences of the expected trends toward smaller, highly efficient gas-generator cores and low-emission fuel-flexible combustors need to be fully investigated since future designs are quite likely to fall outside of the parameter space of existing (semi-empirical) prediction tools.

Hultgren, Lennart S.↗
NASA NTRS2012

Core Noise: Overview of Upcoming LDI Combustor Test

This presentation is a technical summary of and outlook for NASA-internal and NASA-sponsored external research on core (combustor and turbine) noise funded by the Fundamental Aeronautics Program Fixed Wing Project. The presentation covers: the emerging importance of core noise due to turbofan design trends and its relevance to the NASA N+3 noise-reduction goal; the core noise components and the rationale for the current emphasis on combustor noise; and the current and planned research activities in the combustor-noise area. Two NASA-sponsored research programs, with particular emphasis on indirect combustor noise, "Acoustic Database for Core Noise Sources", Honeywell Aerospace (NNC11TA40T) and "Measurement and Modeling of Entropic Noise Sources in a Single-Stage Low-Pressure Turbine", U. Illinois/U. Notre Dame (NNX11AI74A) are briefly described. Recent progress in the development of CMC-based acoustic liners for broadband noise reduction suitable for turbofan-core application is outlined. Combustor-design trends and the potential impacts on combustor acoustics are discussed. A NASA GRC developed nine-point lean-direct-injection (LDI) fuel injector is briefly described. The modification of an upcoming thermo-acoustic instability evaluation of the GRC injector in a combustor rig to also provide acoustic information relevant to community noise is presented. The NASA Fundamental Aeronautics Program has the principal objective of overcoming today's national challenges in air transportation. The reduction of aircraft noise is critical to enabling the anticipated large increase in future air traffic. The Quiet Performance Research Theme of the Fixed Wing Project aims to develop concepts and technologies to dramatically reduce the perceived community noise attributable to aircraft with minimal impact on weight and performance.

Hultgren, Lennart S.↗
NASA NTRS2020

Emission Characteristics of an Axially Staged Sector Combustor for a Small Core High OPR Subsonic Aircraft Engine

This paper presents the nitrogen oxides, carbon monoxide, and particulate matter emissions of a single sector axially staged combustor sector designed and fabricated by United Technologies Research Center (UTRC) in partnership with NASA under a compact low-emissions combustor contract supported by the NASA Advanced Air Transport Technology (AATT) N+3 project. The test was conducted at NASA Glenn Research Center's CE-5 combustion test facility. The facility provided inlet air temperatures up to 922 K and pressures up to 19.0 bar. The combustor design concept, called Axially Controlled Stoichiometry (ACS), was developed by Pratt & Whitney (P&W) under NASA's Environmentally Responsible Aviation (ERA) program for an N+2 combustor for use in twin-aisle subsonic aircraft engines. Under the N+3 project the ACS combustor was scaled-down for application to small-core N+3 engines for use in single-aisle aircraft. The results show that the NOx and CO emissions characteristics are similar in both the N+2 and N+3 applications. The non-volatile particulate matter (nvPM) emissions trends are similar to CO emissions with an exception at high fuel-air ratio, as inlet air temperature and pressure conditions change from taxi to approach. Three NOx correlation equations are generated to describe theNOx emissions of this combustor. The percentage landing and takeoff (LTO) NOx reduction of the N+3 ACS combustor is between 82% and 89% relative to the ICAO CAEP/6 standard, which meets the NASA N+3 goal of exceeding 80% LTO NOx reduction.

He, Zhuohui J.↗
NASA NTRS2024

Hybrid Thermally Efficient Core (HyTEC) HyTEC Phase 1 – Combustor Final Report

The objective of the HyTEC – Combustor Technology project is to develop technology for a compact, low emissions, rich-burn combustor that maintains a high-level of durability. To accomplish this, the combustor will incorporate CMC liners and a CMC Dome to enable improved air utilization for mixing and NOx reduction. In particular, three technology areas were matured: metallic swirler attachment to the CMC dome, Nickel diffusion in CMC, and EBC advancements with improved durability. The maturation of these technology areas is critical to successfully designing a combustor for an engine intended to demonstrate increased thermal efficiency with integrated high-power density-core engine technologies. As part of the project, multiple swirler attachment architectures were matured through the design process, manufactured, and tested in simulated engine conditions. These architectures utilized two different approaches to attachment: a pressure loaded approach and a clamped approach. Each approach required testing at simulated engine conditions to mature the technology. These tests were developed to ensure that the probable failure mechanisms for each architecture were assessed relative to its long-term durability. After successfully testing both approaches, the pressure loaded approach was chosen based on its simplicity and its similarity to the attachment approach used on lean burn combustors. Finally, this successful testing resulted in the swirler attachment technology achieving TRL 4 status. Another aspect of concern with the metallic swirler attachment is the potential for nickel diffusion into the CMC to cause durability shortfalls. A series of tests were completed to understand both the amount of nickel that is expected to diffuse into the CMC and the effect this diffusion will have on mechanical properties of the CMC. After successfully completing this testing, the reduction in local mechanical properties was found to not have an impact on the overall durability of the CMC Dome. To further improve the overall durability, three different EBC coating architectures were also studied. One of these coatings being the current state of the art for CMC liners with the other two being proposed improvements. Coupons were produced using each coating and subjected to a series of laboratory testing. One coating was found to have an overall superior durability in laboratory testing and survived rig testing without indication of degradation. This successful testing resulted in the EBC technology achieving TRL 5 status. Overall, a design approach was successfully down-selected that achieves both the TPM requirements and the need for TRL 4+ status. This design incorporates a pressure loaded swirler attachment with an improved EBC coating (Architecture 1). This combination allows the continued use of a CMC dome and results in an improved durability for the proposed compact core combustor.

CMC↗
DOE OSTI2025

Characterization of Heat Transfer in a Trapped-Vortex Combustor Designed for High-Temperature Material Testing

Next generation gas turbines will reduce carbon emissions by improving their cycle efficiencies and implementing hydrogen as a low-carbon fuel. The combined-cycle efficiency can be improved by increasing the combustor outlet temperature and decreasing the cooling air required for hot-section parts. Ultra-high temperature ceramic matrix composites (CMC) can survive increasingly extreme environments due to their strong mechanical properties at expected conditions. The introduction of CMCs in high-hydrogen flame environments requires testing at realistic conditions that mimic the thermochemical and fluid mechanic states in gas-turbine combustors. In this work, we describe a new experiment that has been designed to test high-temperature materials in a combustor-relevant environment. Its trapped-vortex combustor chamber design allows for high levels of fuel flexibility and wide flame stability limits. Testing of monolithic silicon carbide (SiC) samples in this environment was done to characterize the heat transfer to material samples over a wide range of operating conditions, including variations in fuel composition and thermal power. Heat flux measurements were verified in non-reacting and reacting environments after which they were benchmarked with previous literature to verify the operation of the facility with heated air. Testing in combustor-relevant conditions shows high levels of heat flux to the material samples in the combustor, particularly downstream of the anchored flame.

Richins, Porter [Pennsylvania State University , ,↗
NASA NTRS1972

Pollution measurements of a swirl-can combustor

Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for an experimental, annular, swirl can combustor. The combustor was 42 inches in diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 600, 900 and 1050 F, inlet pressures of 5 to 6 atmospheres, reference velocities of 69 to 120 feet per second and fuel-air ratios of 0.014 to 0.0695. Tests were also conducted at a simulated engine idle condition. Results demonstrated that swirl can combustors produce oxides of nitrogen levels substantially lower than conventional combustor designs. These reductions are attributed to reduced dwell times resulting from short combustor length, quick mixing of combustion gases with diluent air, and to uniform fuel distributions resulting from the swirl can approach. Radial staging of fuel at idle conditions resulted in increases in combustion efficiencies and corresponding reductions in pollutant levels.

Niedzwiecki, R. W.↗
NASA NTRS1973

Performance of an annual combustor designed for a low-cost turbojet engine

Performance tests were conducted on a combustor designed for use in a low-cost turbojet engine. Low-cost features included the use of very inexpensive simplex fuel nozzles and combustor liners of perforated sheet material. Combustion efficiencies at the altitude-cruise and sea-level design points were approximately 94 and 96 percent, respectively. The combustor isothermal total-pressure loss was 8.8 percent at the altitude-cruise-condition diffuser-inlet Mach number of 0.335. The combustor-exit temperature pattern factor was less than 0.3 at the altitude-cruise, sea-level-cruise, and sea-level-static design conditions. The combustor-exit average radial temperature profiles at all conditions were in very good agreement with the design profile. The intense mixing required because of the very high combustor heat-release rate had an adverse effect on ignition capability at altitude windmilling design conditions.

Fear, J. S.↗