Design and preliminary results of a semitranspiration cooled (Lamilloy) liner for a high-pressure high-temperature combustor

NASA Technical Reports Server (NTRS)

Wear, J. D.; Trout, A. M.; Smith, J. M.; Jones, R. E.

1978-01-01

A Lamilloy combustor liner was designed, fabricated and tested in a combustor at pressures up to 8 atmospheres. The liner was fabricated of a three layer Lamilloy structure and designed to replace a conventional step louver liner. The liner is to be used in a combustor that provides hot gases to a turbine cooling test facility at pressures up to 40 atmospheres. The Lamilloy liner was tested extensively at lower pressures and demonstrated lower metal temperatures than the conventional liner, while at the same time requiring about 40 percent less cooling air flow. Tests conducted at combustor exit temperatures in excess of 2200 K have not indicated any cooling or durability problems with the Lamilloy linear.

Genetic algorithm to optimize the design of main combustor and gas generator in liquid rocket engines

NASA Astrophysics Data System (ADS)

Son, Min; Ko, Sangho; Koo, Jaye

2014-06-01

A genetic algorithm was used to develop optimal design methods for the regenerative cooled combustor and fuel-rich gas generator of a liquid rocket engine. For the combustor design, a chemical equilibrium analysis was applied, and the profile was calculated using Rao's method. One-dimensional heat transfer was assumed along the profile, and cooling channels were designed. For the gas-generator design, non-equilibrium properties were derived from a counterflow analysis, and a vaporization model for the fuel droplet was adopted to calculate residence time. Finally, a genetic algorithm was adopted to optimize the designs. The combustor and gas generator were optimally designed for 30-tonf, 75-tonf, and 150-tonf engines. The optimized combustors demonstrated superior design characteristics when compared with previous non-optimized results. Wall temperatures at the nozzle throat were optimized to satisfy the requirement of 800 K, and specific impulses were maximized. In addition, the target turbine power and a burned-gas temperature of 1000 K were obtained from the optimized gas-generator design.

Numerical Prediction of Non-Reacting and Reacting Flow in a Model Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

Davoudzadeh, Farhad; Liu, Nan-Suey

2005-01-01

The three-dimensional, viscous, turbulent, reacting and non-reacting flow characteristics of a model gas turbine combustor operating on air/methane are simulated via an unstructured and massively parallel Reynolds-Averaged Navier-Stokes (RANS) code. This serves to demonstrate the capabilities of the code for design and analysis of real combustor engines. The effects of some design features of combustors are examined. In addition, the computed results are validated against experimental data.

Experimental clean combustor program, phase 2

NASA Technical Reports Server (NTRS)

Gleason, C. C.; Rogers, D. W.; Bahr, D. W.

1976-01-01

The primary objectives of this three-phase program are to develop technology for the design of advanced combustors with significantly lower pollutant emission levels than those of current combustors, and to demonstrate these pollutant emission reductions in CF6-50C engine tests. The purpose of the Phase 2 Program was to further develop the two most promising concepts identified in the Phase 1 Program, the double annular combustor and the radial/axial staged combustor, and to design a combustor and breadboard fuel splitter control for CF6-50 engine demonstration testing in the Phase 3 Program. Noise measurement and alternate fuels addendums to the basic program were conducted to obtain additional experimental data. Twenty-one full annular and fifty-two sector combustor configurations were evaluated. Both combustor types demonstrated the capability for significantly reducing pollutant emission levels. The most promising results were obtained with the double annular combustor. Rig test results corrected to CF-50C engine conditions produced EPA emission parameters for CO, HC, and NOX of 3.4, 0.4, and 4.5 respectively. These levels represent CO, HC, and NOX reductions of 69, 90, and 42 percent respectively from current combustor emission levels. The combustor also met smoke emission level requirements and development engine performance and installation requirements.

Engine-Scale Combustor Rig Designed, Fabricated, and Tested for Combustion Instability Control Research

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Breisacher, Kevin J.

2000-01-01

Low-emission combustor designs are prone to combustor instabilities. Because active control of these instabilities may allow future combustors to meet both stringent emissions and performance requirements, an experimental combustor rig was developed for investigating methods of actively suppressing combustion instabilities. The experimental rig has features similar to a real engine combustor and exhibits instabilities representative of those in aircraft gas turbine engines. Experimental testing in the spring of 1999 demonstrated that the rig can be tuned to closely represent an instability observed in engine tests. Future plans are to develop and demonstrate combustion instability control using this experimental combustor rig. The NASA Glenn Research Center at Lewis Field is leading the Combustion Instability Control program to investigate methods for actively suppressing combustion instabilities. Under this program, a single-nozzle, liquid-fueled research combustor rig was designed, fabricated, and tested. The rig has many of the complexities of a real engine combustor, including an actual fuel nozzle and swirler, dilution cooling, and an effusion-cooled liner. Prior to designing the experimental rig, a survey of aircraft engine combustion instability experience identified an instability observed in a prototype engine as a suitable candidate for replication. The frequency of the instability was 525 Hz, with an amplitude of approximately 1.5-psi peak-to-peak at a burner pressure of 200 psia. The single-nozzle experimental combustor rig was designed to preserve subcomponent lengths, cross sectional area distribution, flow distribution, pressure-drop distribution, temperature distribution, and other factors previously found to be determinants of burner acoustic frequencies, mode shapes, gain, and damping. Analytical models were used to predict the acoustic resonances of both the engine combustor and proposed experiment. The analysis confirmed that the test rig configuration and engine configuration had similar longitudinal acoustic characteristics, increasing the likelihood that the engine instability would be replicated in the rig. Parametric analytical studies were performed to understand the influence of geometry and condition variations and to establish a combustion test plan. Cold-flow experiments verified that the design values of area and flow distributions were obtained. Combustion test results established the existence of a longitudinal combustion instability in the 500-Hz range with a measured amplitude approximating that observed in the engine. Modifications to the rig configuration during testing also showed the potential for injector independence. The research combustor rig was developed in partnership with Pratt & Whitney of West Palm Beach, Florida, and United Technologies Research Center of East Hartford, Connecticut. Experimental testing of the combustor rig took place at United Technologies Research Center.

Effects of fuel nozzle design on performance of an experimental annular combustor using natural gas fuel

NASA Technical Reports Server (NTRS)

Wear, J. D.; Schultz, D. F.

1972-01-01

Tests of various fuel nozzles were conducted with natural gas fuel in a full-annulus combustor. The nozzles were designed to provide either axial, angled, or radial fuel injection. Each fuel nozzle was evaluated by measuring combustion efficiency at relatively severe combustor operating conditions. Combustor blowout and altitude ignition tests were also used to evaluate nozzle designs. Results indicate that angled injection gave higher combustion efficiency, less tendency toward combustion instability, and altitude relight characteristics equal to or superior to those of the other fuel nozzles that were tested.

Design and preliminary results of a fuel flexible industrial gas turbine combustor

NASA Technical Reports Server (NTRS)

Novick, A. S.; Troth, D. L.; Yacobucci, H. G.

1981-01-01

The design characteristics are presented of a fuel tolerant variable geometry staged air combustor using regenerative/convective cooling. The rich/quench/lean variable geometry combustor is designed to achieve low NO(x) emission from fuels containing fuel bound nitrogen. The physical size of the combustor was calculated for a can-annular combustion system with associated operating conditions for the Allison 570-K engine. Preliminary test results indicate that the concept has the potential to meet emission requirements at maximum continuous power operation. However, airflow sealing and improved fuel/air mixing are necessary to meet Department of Energy program goals.

Combustion Characteristics Analysis of Improved Combustor Structure of Micro Turbine Engine

NASA Astrophysics Data System (ADS)

Chen, Hai

2018-05-01

In order to improve the performance of micro combustor, the 60 slots of the original combustor were modified into 120 slots for the MIT 6-wafer micro-combustor. The performance of the micro combustor with the improved and original design was compared through numerical simulation, and stable operating ranges was studied. It was found that the improved combustor can stabilize the flame under the condition of higher fuel/air mixture mass flow rate.

Pollution measurements of a swirl-can combustor

NASA Technical Reports Server (NTRS)

Niedzwiecki, R. W.; Jones, R. E.

1972-01-01

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.

Phase I Final Scientific Report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lu, Xijia; Fetvedt, Jeremy; Dimmig, Walker

This Final Scientific Report addresses the accomplishments achieved during Phase I of DE- FE0023985, Coal Syngas Combustor Development for Supercritical CO 2 Power Cycles. The primary objective of the project was to develop a coal syngas-fueled combustor design for use with high-pressure, high-temperature, oxy-fuel, supercritical CO 2 power cycles, with particular focus given to the conditions required by the Allam Cycle. The primary goals, from the Statement of Project Objectives, were to develop: (1) a conceptual design of a syngas-fueled combustor-turbine block for a 300MWe high-pressure, oxy-fuel, sCO2 power plant; (2) the preliminary design of a 5MWt test combustor; andmore » (3) the definition of a combustor test program. Accomplishments for each of these goals are discussed in this report.« less

Combustor concepts for aircraft gas turbine low-power emissions reduction

NASA Technical Reports Server (NTRS)

Mularz, E. J.; Gleason, C. C.; Dodds, W. J.

1978-01-01

Several combustor concepts were designed and tested to demonstrate significant reductions in aircraft engine idle pollutant emissions. Each concept used a different approach for pollutant reductions: the hot wall combustor employs a thermal barrier coating and impingement cooled liners; the recuperative cooling combustor preheats the air before entering the combustion chamber; and the catalytic converter combustor is composed of a conventional primary zone followed by a catalytic bed for pollutant cleanup. The designs are discussed in detail and test results are presented for a range of aircraft engine idle conditions. The results indicate that ultralow levels of unburned hydrocarbons and carbon monoxide emissions can be achieved.

Spray combustion experiments and numerical predictions

NASA Technical Reports Server (NTRS)

Mularz, Edward J.; Bulzan, Daniel L.; Chen, Kuo-Huey

1993-01-01

The next generation of commercial aircraft will include turbofan engines with performance significantly better than those in the current fleet. Control of particulate and gaseous emissions will also be an integral part of the engine design criteria. These performance and emission requirements present a technical challenge for the combustor: control of the fuel and air mixing and control of the local stoichiometry will have to be maintained much more rigorously than with combustors in current production. A better understanding of the flow physics of liquid fuel spray combustion is necessary. This paper describes recent experiments on spray combustion where detailed measurements of the spray characteristics were made, including local drop-size distributions and velocities. Also, an advanced combustor CFD code has been under development and predictions from this code are compared with experimental results. Studies such as these will provide information to the advanced combustor designer on fuel spray quality and mixing effectiveness. Validation of new fast, robust, and efficient CFD codes will also enable the combustor designer to use them as additional design tools for optimization of combustor concepts for the next generation of aircraft engines.

Fuel properties effect on the performance of a small high temperature rise combustor

NASA Technical Reports Server (NTRS)

Acosta, Waldo A.; Beckel, Stephen A.

1989-01-01

The performance of an advanced small high temperature rise combustor was experimentally determined at NASA-Lewis. The combustor was designed to meet the requirements of advanced high temperature, high pressure ratio turboshaft engines. The combustor featured an advanced fuel injector and an advanced segmented liner design. The full size combustor was evaluated at power conditions ranging from idle to maximum power. The effect of broad fuel properties was studied by evaluating the combustor with three different fuels. The fuels used were JP-5, a blend of Diesel Fuel Marine/Home Heating Oil, and a blend of Suntec C/Home Heating Oil. The fuel properties effect on the performance of the combustion in terms of pattern factor, liner temperatures, and exhaust emissions are documented.

Assessment, development, and application of combustor aerothermal models

NASA Technical Reports Server (NTRS)

Holdeman, J. D.; Mongia, H. C.; Mularz, E. J.

1989-01-01

The gas turbine combustion system design and development effort is an engineering exercise to obtain an acceptable solution to the conflicting design trade-offs between combustion efficiency, gaseous emissions, smoke, ignition, restart, lean blowout, burner exit temperature quality, structural durability, and life cycle cost. For many years, these combustor design trade-offs have been carried out with the help of fundamental reasoning and extensive component and bench testing, backed by empirical and experience correlations. Recent advances in the capability of computational fluid dynamics codes have led to their application to complex 3-D flows such as those in the gas turbine combustor. A number of U.S. Government and industry sponsored programs have made significant contributions to the formulation, development, and verification of an analytical combustor design methodology which will better define the aerothermal loads in a combustor, and be a valuable tool for design of future combustion systems. The contributions made by NASA Hot Section Technology (HOST) sponsored Aerothermal Modeling and supporting programs are described.

An efficient liner cooling scheme for advanced small gas turbine combustors

NASA Technical Reports Server (NTRS)

Paskin, Marc D.; Mongia, Hukam C.; Acosta, Waldo A.

1993-01-01

A joint Army/NASA program was conducted to design, fabricate, and test an advanced, small gas turbine, reverse-flow combustor utilizing a compliant metal/ceramic (CMC) wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CMC cooling technique and then demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F burner outlet temperature. The CMC concept offers significant improvements in wall cooling effectiveness resulting in a large reduction in cooling air requirements. Therefore, more air is available for control of burner outlet temperature pattern in addition to the benefits of improved efficiency, reduced emissions, and lower smoke levels. The program was divided into four tasks. Task 1 defined component materials and localized design of the composite wall structure in conjunction with development of basic design models for the analysis of flow and heat transfer through the wall. Task 2 included implementation of the selected materials and validated design models during combustor preliminary design. Detail design of the selected combustor concept and its refinement with 3D aerothermal analysis were completed in Task 3. Task 4 covered detail drawings, process development and fabrication, and a series of burner rig tests. The purpose of this paper is to provide details of the investigation into the fundamental flow and heat transfer characteristics of the CMC wall structure as well as implementation of the fundamental analysis method for full-scale combustor design.

Experimental clean combustor program; noise measurement addendum, Phase 2

NASA Technical Reports Server (NTRS)

Emmerling, J. J.; Bekofske, K. L.

1976-01-01

Combustor noise measurements were performed using wave guide probes. Test results from two full scale annular combustor configurations in a combustor test rig are presented. A CF6-50 combustor represented a current design, and a double annular combustor represented the advanced clean combustor configuration. The overall acoustic power levels were found to correlate with the steady state heat release rate and inlet temperature. A theoretical analysis for the attenuation of combustor noise propagating through a turbine was extended from a subsonic relative flow condition to include the case of supersonic flow at the discharge side. The predicted attenuation from this analysis was compared to both engine data and extrapolated component combustor data. The attenuation of combustor noise through the CF6-50 turbine was found to be greater than 14 dB by both the analysis and the data.

Hypersonic Combustor Model Inlet CFD Simulations and Experimental Comparisons

NASA Technical Reports Server (NTRS)

Venkatapathy, E.; TokarcikPolsky, S.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

Numerous two-and three-dimensional computational simulations were performed for the inlet associated with the combustor model for the hypersonic propulsion experiment in the NASA Ames 16-Inch Shock Tunnel. The inlet was designed to produce a combustor-inlet flow that is nearly two-dimensional and of sufficient mass flow rate for large scale combustor testing. The three-dimensional simulations demonstrated that the inlet design met all the design objectives and that the inlet produced a very nearly two-dimensional combustor inflow profile. Numerous two-dimensional simulations were performed with various levels of approximations such as in the choice of chemical and physical models, as well as numerical approximations. Parametric studies were conducted to better understand and to characterize the inlet flow. Results from the two-and three-dimensional simulations were used to predict the mass flux entering the combustor and a mass flux correlation as a function of facility stagnation pressure was developed. Surface heat flux and pressure measurements were compared with the computed results and good agreement was found. The computational simulations helped determine the inlet low characteristics in the high enthalpy environment, the important parameters that affect the combustor-inlet flow, and the sensitivity of the inlet flow to various modeling assumptions.

NOx Emissions Characteristics and Correlation Equations of Two P and W's Axially Staged Sector Combustors Developed Under NASA Environmentally Responsible Aviation (ERA) Project

NASA Technical Reports Server (NTRS)

He, Zhuohui J.

2017-01-01

Two P&W (Pratt & Whitney)'s axially staged sector combustors have been developed under NASA's Environmentally Responsible Aviation (ERA) project. One combustor was developed under ERA Phase I, and the other was developed under ERA Phase II. Nitrogen oxides (NOx) emissions characteristics and correlation equations for these two sector combustors are reported in this article. The Phase I design was to optimize the NOx emissions reduction potential, while the Phase II design was more practical and robust. Multiple injection points and fuel staging strategies are used in the combustor design. Pilot-stage injectors are located on the front dome plate of the combustor, and main-stage injectors are positioned on the top and bottom (Phase I) or on the top only (Phase II) of the combustor liners downstream. Low power configuration uses only pilot-stage injectors. Main-stage injectors are added to high power configuration to help distribute fuel more evenly and achieve lean burn throughout the combustor yielding very low NOx emissions. The ICAO (International Civil Aviation Organization) landing-takeoff NOx emissions are verified to be 88 percent (Phase I) and 76 percent (Phase II) under the ICAO CAEP/6 (Committee on Aviation Environmental Protection 6th Meeting) standard, exceeding the ERA project goal of 75 percent reduction, and the combustors proved to have stable combustion with room to maneuver on fuel flow splits for operability.

Review of jet engine emissions

NASA Technical Reports Server (NTRS)

Grobman, J. S.

1972-01-01

A review of the emission characteristics of jet engines is presented. The sources and concentrations of the various constituents in the engine exhaust and the influence of engine operating conditions on emissions are discussed. Cruise emissions to be expected from supersonic engines are compared with emissions from subsonic engines. The basic operating principles of the gas turbine combustor are reviewed together with the effects of combustor operating conditions on emissions. The performance criteria that determine the design of gas turbine combustors are discussed. Combustor design techniques are considered that may be used to reduce emissions.

Quiet Clean Short-haul Experimental Engine (QCSEE). Double-annular clean combustor technology development report

NASA Technical Reports Server (NTRS)

Bahr, D. W.; Burrus, D. L.; Sabla, P. E.

1979-01-01

A sector combustor technology development program was conducted to define an advanced double annular dome combustor sized for use in the quiet clean short haul experimental engine (QCSEE). A design which meets the emission goals, and combustor performance goals of the QCSEE engine program was developed. Key design features were identified which resulted in substantial reduction in carbon monoxide and unburned hydrocarbon emission levels at ground idle operating conditions, in addition to very low nitric oxide emission levels at high power operating conditions. Their significant results are reported.

Energy efficient engine

NASA Technical Reports Server (NTRS)

Burrus, D.; Sabla, P. E.; Bahr, D. W.

1980-01-01

The feasibility of meeting or closely approaching the emissions goals established for the Energy Efficient Engine (E3) Project with an advanced design, single annular combustor was determined. A total of nine sector combustor configurations and one full-annular-combustor configuration were evaluated. Acceptable levels of carbon monoxide and hydrocarbon emissions were obtained with several of the sector combustor configurations tested, and several of the configurations tested demonstrated reduced levels of nitrogen oxides compared to conventional, single annular designs. None of the configurations tested demonstrated nitrogen oxide emission levels that meet the goal of the E3 Project.

CFD Analysis of Emissions for a Candidate N+3 Combustor

NASA Technical Reports Server (NTRS)

Ajmani, Kumud

2015-01-01

An effort was undertaken to analyze the performance of a model Lean-Direct Injection (LDI) combustor designed to meet emissions and performance goals for NASA's N+3 program. Computational predictions of Emissions Index (EINOx) and combustor exit temperature were obtained for operation at typical power conditions expected of a small-core, high pressure-ratio (greater than 50), high T3 inlet temperature (greater than 950K) N+3 combustor. Reacting-flow computations were performed with the National Combustion Code (NCC) for a model N+3 LDI combustor, which consisted of a nine-element LDI flame-tube derived from a previous generation (N+2) thirteen-element LDI design. A consistent approach to mesh-optimization, spraymodeling and kinetics-modeling was used, in order to leverage the lessons learned from previous N+2 flame-tube analysis with the NCC. The NCC predictions for the current, non-optimized N+3 combustor operating indicated a 74% increase in NOx emissions as compared to that of the emissions-optimized, parent N+2 LDI combustor.

CFD Analysis of Emissions for a Candidate N+3 Combustor

NASA Technical Reports Server (NTRS)

Ajmani, Kumud

2015-01-01

An effort was undertaken to analyze the performance of a model Lean-Direct Injection (LDI) combustor designed to meet emissions and performance goals for NASA's N+3 program. Computational predictions of Emissions Index (EINOx) and combustor exit temperature were obtained for operation at typical power conditions expected of a small-core, high pressure-ratio (greater than 50), high T3 inlet temperature (greater than 950K) N+3 combustor. Reacting-flow computations were performed with the National Combustion Code (NCC) for a model N+3 LDI combustor, which consisted of a nine-element LDI flame-tube derived from a previous generation (N+2) thirteen-element LDI design. A consistent approach to mesh-optimization, spray-modeling and kinetics-modeling was used, in order to leverage the lessons learned from previous N+2 flame-tube analysis with the NCC. The NCC predictions for the current, non-optimized N+3 combustor operating indicated a 74% increase in NOx emissions as compared to that of the emissions-optimized, parent N+2 LDI combustor.

Energy efficient engine pin fin and ceramic composite segmented liner combustor sector rig test report

NASA Technical Reports Server (NTRS)

Dubiel, D. J.; Lohmann, R. P.; Tanrikut, S.; Morris, P. M.

1986-01-01

Under the NASA-sponsored Energy Efficient Engine program, Pratt and Whitney has successfully completed a comprehensive test program using a 90-degree sector combustor rig that featured an advanced two-stage combustor with a succession of advanced segmented liners. Building on the successful characteristics of the first generation counter-parallel Finwall cooled segmented liner, design features of an improved performance metallic segmented liner were substantiated through representative high pressure and temperature testing in a combustor atmosphere. This second generation liner was substantially lighter and lower in cost than the predecessor configuration. The final test in this series provided an evaluation of ceramic composite liner segments in a representative combustor environment. It was demonstrated that the unique properties of ceramic composites, low density, high fracture toughness, and thermal fatigue resistance can be advantageously exploited in high temperature components. Overall, this Combustor Section Rig Test program has provided a firm basis for the design of advanced combustor liners.

Advanced catalytic combustors for low pollutant emissions, phase 1

NASA Technical Reports Server (NTRS)

Dodds, W. J.

1979-01-01

The feasibility of employing the known attractive and distinguishing features of catalytic combustion technology to reduce nitric oxide emissions from gas turbine engines during subsonic, stratospheric cruise operation was investigated. Six conceptual combustor designs employing catalytic combustion were defined and evaluated for their potential to meet specific emissions and performance goals. Based on these evaluations, two parallel-staged, fixed-geometry designs were identified as the most promising concepts. Additional design studies were conducted to produce detailed preliminary designs of these two combustors. Results indicate that cruise nitric oxide emissions can be reduced by an order of magnitude relative to current technology levels by the use of catalytic combustion. Also, these combustors have the potential for operating over the EPA landing-takeoff cycle and at cruise with a low pressure drop, high combustion efficiency and with a very low overall level of emission pollutants. The use of catalytic combustion, however, requires advanced technology generation in order to obtain the time-temperature catalytic reactor performance and durability required for practical aircraft engine combustors.

Effect of flame stabilizer design on performance and exhaust pollutants of a two-row 72-module swirl-can combustor

NASA Technical Reports Server (NTRS)

Biaglow, J. A.; Trout, A. M.

1976-01-01

A test program was conducted to evaluate the effects of four flame stabilizer designs on the performance and gaseous pollutant levels of an experimental full-annular swirl-can combustor. Combustor operating parameters, including inlet-air temperature, reference velocity, and fuel-air ratio, were set to simulate conditions in a 30:1 pressure ratio engine. Combustor inlet total pressure was held constant at 6 atm due to the facility limit. Combustor performance and gaseous pollutant levels were strongly affected by the geometry and resulting total pressure loss of the four flame stabilizer designs investigated. The addition of shrouds to two designs produced an 18 to 22% decrease in the combustion chamber pressure loss and thus resulted in doubling the exit temperature pattern factor and up to 42% higher levels of oxides of nitrogen. A previously developed oxides of nitrogen correlating parameter agreed with each model within an emission index of plus or minus 1 but was not capable of correlating all models together.

Ignition improvement by injector arrangement in a multi-fuel combustor for micro gas turbine

NASA Astrophysics Data System (ADS)

Antoshkiv, O.; Poojitganont, T.; Jeansirisomboon, S.; Berg, H. P.

2018-01-01

The novel combustor design also has an impact on the ignitor arrangement. The conventional ignitor system cannot guarantee optimal ignition performance in the usual radial position. The difficult ignitability of gaseous fuels was the main challenge for the ignitor system improvement. One way to improve the ignition performance significantly is a torch ignitor system in which the gaseous fuel is directly mixed with a large amount of the combustor air. To reach this goal, the ignition process was investigated in detail. The micro gas turbine (MGT) ignition was optimised considering three main procedures: torch ignitor operation, burner ignition and flame propagation between the neighbour injectors. A successful final result of the chain of ignition procedures depends on multiple aspects of the combustor design. Performed development work shows an important step towards designing modern high-efficiency low-emission combustors.

Study of research and development requirements of small gas-turbine combustors

NASA Technical Reports Server (NTRS)

Demetri, E. P.; Topping, R. F.; Wilson, R. P., Jr.

1980-01-01

A survey is presented of the major small-engine manufacturers and governmental users. A consensus was undertaken regarding small-combustor requirements. The results presented are based on an evaluation of the information obtained in the course of the study. The current status of small-combustor technology is reviewed. The principal problems lie in liner cooling, fuel injection, part-power performance, and ignition. Projections of future engine requirements and their effect on the combustor are discussed. The major changes anticipated are significant increases in operating pressure and temperature levels and greater capability of using heavier alternative fuels. All aspects of combustor design are affected, but the principal impact is on liner durability. An R&D plan which addresses the critical combustor needs is described. The plan consists of 15 recommended programs for achieving necessary advances in the areas of liner thermal design, primary-zone performance, fuel injection, dilution, analytical modeling, and alternative-fuel utilization.

Combustor materials requirements and status of ceramic matrix composites

NASA Technical Reports Server (NTRS)

Hecht, Ralph J.; Johnson, Andrew M.

1992-01-01

The HSCT combustor will be required to operate with either extremely rich or lean fuel/air ratios to reduce NO(x) emission. NASA High Speed Research (HSR) sponsored programs at Pratt & Whitney (P&W) and GE Aircraft Engines (GEAE) have been studying rich and lean burn combustor design approaches which are capable of achieving the aggressive HSCT NO(x) emission goals. In both of the combustor design approaches under study, high temperature (2400-3000 F) materials are necessary to meet the HSCT emission goals of 3-8 gm/kg. Currently available materials will not meet the projected requirements for the HSCT combustor. The development of new materials is an enabling technology for the successful introduction to service of the HSCT.

Composite Matrix Cooling Scheme for Small Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Paskin, Marc D.; Ross, Phillip T.; Mongia, Hukam C.; Acosta, Waldo A.

1990-01-01

The design, manufacture, and testing of a compliant metal/ceramic (CMC) wall cooling concept-implementing combustor for small gas turbine engines has been undertaken by a joint U.S. Army/NASA technology development program. CMC in principle promises greater wall cooling effectiveness than conventional designs and materials, thereby facilitating a substantial reduction in combustor cooling air requirements and furnishing greater airflow for the control of burner outlet temperature patterns as well as improving thermodynamic efficiency and reducing pollutant emissions and smoke levels. Rig test results have confirmed the projected benefits of the CMC concept at combustor outlet temperatures of the order of 2460 F, at which approximately 80 percent less cooling air than conventionally required was being employed by the CMC combustor.

Experimental clean combustor program, phase 2

NASA Technical Reports Server (NTRS)

Roberts, R.; Peduzzi, A.; Vitti, G. E.

1976-01-01

Combustor pollution reduction technology for commercial CTOL engines was generated and this technology was demonstrated in a full-scale JT9D engine in 1976. Component rig refinement of the two best combustor concepts were tested. These concepts are the vorbix combustor, and a hybrid combustor which combines the pilot zone of the staged premix combustor and the main zone of the swirl-can combustor. Both concepts significantly reduced all pollutant emissions relative to the JT9D-7 engine combustor. However, neither concept met all program goals. The hybrid combustor met pollution goals for unburned hydrocarbons and carbon monoxide but did not achieve the oxides of nitrogen goal. This combustor had significant performance deficiencies. The Vorbix combustor met goals for unburned hydrocarbons and oxides of nitrogen but did not achieve the carbon monoxide goal. Performance of the vorbix combustor approached the engine requirements. On the basis of these results, the vorbix combustor was selected for the engine demonstration program. A control study was conducted to establish fuel control requirements imposed by the low-emission combustor concepts and to identify conceptual control system designs. Concurrent efforts were also completed on two addendums: an alternate fuels addendum and a combustion noise addendum.

Advanced Low-Emissions Catalytic-Combustor Program, phase 1. [aircraft gas turbine engines

NASA Technical Reports Server (NTRS)

Sturgess, G. J.

1981-01-01

Six catalytic combustor concepts were defined, analyzed, and evaluated. Major design considerations included low emissions, performance, safety, durability, installations, operations and development. On the basis of these considerations the two most promising concepts were selected. Refined analysis and preliminary design work was conducted on these two concepts. The selected concepts were required to fit within the combustor chamber dimensions of the reference engine. This is achieved by using a dump diffuser discharging into a plenum chamber between the compressor discharge and the turbine inlet, with the combustors overlaying the prediffuser and the rear of the compressor. To enhance maintainability, the outer combustor case for each concept is designed to translate forward for accessibility to the catalytic reactor, liners and high pressure turbine area. The catalytic reactor is self-contained with air-cooled canning on a resilient mounting. Both selected concepts employed integrated engine-starting approaches to raise the catalytic reactor up to operating conditions. Advanced liner schemes are used to minimize required cooling air. The two selected concepts respectively employ fuel-rich initial thermal reaction followed by rapid quench and subsequent fuel-lean catalytic reaction of carbon monoxide, and, fuel-lean thermal reaction of some fuel in a continuously operating pilot combustor with fuel-lean catalytic reaction of remaining fuel in a radially-staged main combustor.

A Design Methodology for Rapid Implementation of Active Control Systems Across Lean Direct Injection Combustor Platforms

NASA Technical Reports Server (NTRS)

Baumann, William T.; Saunders, William R.; Vandsburger, Uri; Saus, Joseph (Technical Monitor)

2003-01-01

The VACCG team is comprised of engineers at Virginia Tech who specialize in the subject areas of combustion physics, chemical kinetics, dynamics and controls, and signal processing. Currently, the team's work on this NRA research grant is designed to determine key factors that influence combustion control performance through a blend of theoretical and experimental investigations targeting design and demonstration of active control for three different combustors. To validiate the accuracy of conclusions about control effectiveness, a sequence of experimental verifications on increasingly complex lean, direct injection combustors is underway. During the work period January 1, 2002 through October 15, 2002, work has focused on two different laboratory-scale combustors that allow access for a wide variety of measurements. As the grant work proceeds, one key goal will be to obtain certain knowledge about a particular combustor process using a minimum of sophisticated measurements, due to the practical limitations of measurements on full-scale combustors. In the second year, results obtained in the first year will be validated on test combustors to be identified in the first quarter of that year. In the third year, it is proposed to validate the results at more realistic pressure and power levels by utilizing the facilities at the Glenn Research Center.

Flow Coupling Effects in Jet-in-Crossflow Flowfields

NASA Technical Reports Server (NTRS)

Bain, D. B.; Smith, C. E.; Liscinsky, D. S.; Holdeman, J. D.

1996-01-01

The combustor designer is typically required to design liner orifices that effectively mix air jets with crossflow effluent. CFD combustor analysis is typically used in the design process; however the jets are usually assumed to enter the combustor with a uniform velocity and turbulence profile. The jet-mainstream flow coupling is usually neglected because of the computational expense. This CFD study was performed to understand the effect of jet-mainstream flow coupling, and to assess the accuracy of jet boundary conditions that are commonly used in combustor internal calculations. A case representative of a plenum-fed quick-mix section of a Rich Burn/Quick Mix/Lean Burn combustor (i.e. a jet-mainstream mass-flow ratio of about 3 and a jet-mainstream momentum-flux ratio of about 30) was investigated. This case showed that the jet velocity entering the combustor was very non-uniform, with a low normal velocity at the leading edge of the orifice and a high normal velocity at the trailing edge of the orifice. Three different combustor-only cases were analyzed with uniform inlet jet profile. None of the cases matched the plenum-fed calculations. To assess liner thickness effects, a thin-walled case was also analyzed. The CFD analysis showed the thin-walled jets had more penetration than the thick-walled jets.

Analysis of the impact of the use of broad specification fuels on combustors for commercial aircraft gas turbine engines

NASA Technical Reports Server (NTRS)

Szetela, E. J.; Lehmann, R. P.; Smith, A. L.

1979-01-01

An analytical study was conducted to assess the impact of the use of broad specification fuels with reduced hydrogen content on the design, performance, durability, emissions and operational characteristics of combustors for commercial aircraft gas turbine engines. The study was directed at defining necessary design revisions to combustors designed for use of Jet A when such are operated on ERBS (Experimental Referee Broad Specification Fuel) which has a nominal hydrogen content of 12.8 percent as opposed to 13.7 percent in current Jet A. The results indicate that improvements in combustor liner cooling, and/or materials, and methods of fuel atomization will be required if the hydrogen content of aircraft gas turbine fuel is decreased.

Materials for Advanced Turbine Engines (MATE): Project 3: Design, fabrication and evaluation of an oxide dispersion strengthened sheet alloy combustor liner, volume 1

NASA Technical Reports Server (NTRS)

Henricks, R. J.; Sheffler, K. D.

1984-01-01

The suitability of wrought oxide dispersion strengthened (ODS) superalloy sheet for gas turbine engine combustor applications was evaluated. Incoloy MA 956 (FeCrAl base) and Haynes Developmental Alloy (HDA) 8077 (NiCrAl base) were evaluated. Preliminary tests showed both alloys to be potentially viable combustor materials, with neither alloy exhibiting a significant advantage over the other. Both alloys demonstrated a +167C (300 F) advantage of creep and oxidation resistance with no improvement in thermal fatigue capability compared to a current generation combustor alloy (Hastelloy X). MA956 alloy was selected for further demonstration because it exhibited better manufacturing reproducibility than HDA8077. Additional property tests were conducted on MA956. To accommodate the limited thermal fatigue capability of ODS alloys, two segmented, mechanically attached, low strain ODS combustor design concepts having predicted fatigue lives or = 10,000 engine cycles were identified. One of these was a relatively conventional louvered geometry, while the other involved a transpiration cooled configuration. A series of 10,000 cycle combustor rig tests on subscale MA956 and Hastelloy X combustor components showed no cracking, thereby confirming the beneficial effect of the segmented design on thermal fatigue capability. These tests also confirmed the superior oxidation and thermal distortion resistance of the ODS alloy. A hybrid PW2037 inner burner liner containing MA956 and Hastelloy X components was designed and constructed.

Comparisons of a Three-Dimensional, Full Navier Stokes Computer Model with High Mach Number Combuster Test Data

NASA Technical Reports Server (NTRS)

Watkins, William B.

1990-01-01

Comparisons between scramjet combustor data and a three-dimensional full Navier-Stokes calculation have been made to verify and substantiate computational fluid dynamics (CFD) codes and application procedures. High Mach number scramjet combustor development will rely heavily on CFD applications to provide wind tunnel-equivalent data of quality sufficient to design, build and fly hypersonic aircraft. Therefore. detailed comparisons between CFD results and test data are imperative. An experimental case is presented, for which combustor wall static pressures were measured and flow-fieid interferograms were obtained. A computer model was done of the experiment, and counterpart parameters are compared with experiment. The experiment involved a subscale combustor designed and fabricated for the National Aero-Space Plane Program, and tested in the Calspan Corporation 96" hypersonic shock tunnel. The combustor inlet ramp was inclined at a 20 angle to the shock tunnel nozzle axis, and resulting combustor entrance flow conditions simulated freestream M=10. The combustor body and cowl walls were instrumented with static pressure transducers, and the combustor lateral walls contained windows through which flowfield holographic interferograms were obtained. The CFD calculation involved a three-dimensional time-averaged full Navier-Stokes code applied to the axial flow segment containing fuel injection and combustion. The full Navier-Stokes approach allowed for mixed supersonic and subsonic flow, downstream-upstream communication in subsonic flow regions, and effects of adverse pressure gradients. The code included hydrogen-air chemistry in the combustor segment which begins near fuel injection and continues through combustor exhaust. Combustor ramp and inlet segments on the combustor lateral centerline were modelled as two dimensional. Comparisons to be shown include calculated versus measured wall static pressures as functions of axial flow coordinate, and calculated path-averaged density contours versus an holographic Interferogram.

Design and Evaluation of a Single-Inlet Pulse Detonation Combustor

DTIC Science & Technology

2011-06-01

Kilogram/second m/s Meters/ second N Nitrogen NPS Naval Postgraduate School O Oxygen PDC Pulse Detonation Combustion PDE Pulse Detonation Engine...EVALUATION OF A SINGLE-INLET PULSE DETONATION COMBUSTOR by Danny Soria June 2011 Thesis Advisor: Christopher M. Brophy Second Reader: Garth V...COVERED Master’s Thesis 4. TITLE AND SUBTITLE Design and Evaluation of a Single-Inlet Pulse Detonation Combustor 6. AUTHOR(S) Danny Soria 5

National Combustion Code: A Multidisciplinary Combustor Design System

NASA Technical Reports Server (NTRS)

Stubbs, Robert M.; Liu, Nan-Suey

1997-01-01

The Internal Fluid Mechanics Division conducts both basic research and technology, and system technology research for aerospace propulsion systems components. The research within the division, which is both computational and experimental, is aimed at improving fundamental understanding of flow physics in inlets, ducts, nozzles, turbomachinery, and combustors. This article and the following three articles highlight some of the work accomplished in 1996. A multidisciplinary combustor design system is critical for optimizing the combustor design process. Such a system should include sophisticated computer-aided design (CAD) tools for geometry creation, advanced mesh generators for creating solid model representations, a common framework for fluid flow and structural analyses, modern postprocessing tools, and parallel processing. The goal of the present effort is to develop some of the enabling technologies and to demonstrate their overall performance in an integrated system called the National Combustion Code.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zhong, Bei-Jing; Wang, Jian-Hua

Excess enthalpy combustion is a promising approach to stabilize flame in micro-combustors. Using a Swiss-roll combustor configuration, excess enthalpy combustion can be conveniently achieved. In this work, three types of Swiss-roll combustors with double spiral-shaped channels were designed and fabricated. The combustors were tested using methane/air mixtures of various equivalence ratios. Both temperature distributions and extinction limits were determined for each combustor configuration at different methane mass flow rates. Results indicate that the Swiss-roll combustors developed in the current study greatly enhance combustion stability in center regions of the combustors. At the same time, excess enthalpy combustors of the Swiss-rollmore » configuration significantly extend the extinction limits of methane/air mixtures. In addition, the effects of combustor configurations and thermal insulation arrangements on temperature distributions and extinction limits were evaluated. With heat losses to the environment being significant, the use of thermal insulations further enhances the flame stability in center regions of the Swiss-roll combustors and extends flammable ranges. (author)« less

Multi-fuel combustor for gas turbine engines: Phase 1, Final report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Melconian, J.O.; Marden, W.W., III

An innovative can combustor configuration has been developed for gas turbine engines which has the potential of burning fuels ranging from gasoline to coal/water slurries at high efficiencies. The design is based on a Variable Residence Time (VRT) concept which allows large and agglomerated fuel particles adequate time to completely burn. High durability of the combustor is achieved by dual function use of the incoming air. For applications which require the burning of coal/water slurries, the design has the capability of removing the ash particles directly from the primary zone of the combustor. It is anticipated that because of themore » small size requirement of this combustor design, existing gas turbine engines could be retrofitted within the confines of the current engine envelope. In Phase 1, the feasibility of the concept was successfully demonstrated by three-dimensional mathematical modeling and water analogue tests. The Plexiglas model used in the water analogue tests was designed to fit the current production engine of a major manufacturer. 19 figs., 2 tabs.« less

An experimental and theoretical investigation of a fuel system tuner for the suppression of combustion driven oscillations

NASA Astrophysics Data System (ADS)

Scarborough, David E.

Manufacturers of commercial, power-generating, gas turbine engines continue to develop combustors that produce lower emissions of nitrogen oxides (NO x) in order to meet the environmental standards of governments around the world. Lean, premixed combustion technology is one technique used to reduce NOx emissions in many current power and energy generating systems. However, lean, premixed combustors are susceptible to thermo-acoustic oscillations, which are pressure and heat-release fluctuations that occur because of a coupling between the combustion process and the natural acoustic modes of the system. These pressure oscillations lead to premature failure of system components, resulting in very costly maintenance and downtime. Therefore, a great deal of work has gone into developing methods to prevent or eliminate these combustion instabilities. This dissertation presents the results of a theoretical and experimental investigation of a novel Fuel System Tuner (FST) used to damp detrimental combustion oscillations in a gas turbine combustor by changing the fuel supply system impedance, which controls the amplitude and phase of the fuel flowrate. When the FST is properly tuned, the heat release oscillations resulting from the fuel-air ratio oscillations damp, rather than drive, the combustor acoustic pressure oscillations. A feasibility study was conducted to prove the validity of the basic idea and to develop some basic guidelines for designing the FST. Acoustic models for the subcomponents of the FST were developed, and these models were experimentally verified using a two-microphone impedance tube. Models useful for designing, analyzing, and predicting the performance of the FST were developed and used to demonstrate the effectiveness of the FST. Experimental tests showed that the FST reduced the acoustic pressure amplitude of an unstable, model, gas-turbine combustor over a wide range of operating conditions and combustor configurations. Finally, combustor acoustic pressure amplitude measurements made in using the model combustor were used in conjunction with model predicted fuel system impedances to verify the developed design rules. The FST concept and design methodology presented in this dissertation can be used to design fuel system tuners for new and existing gas turbine combustors to reduce, or eliminate altogether, thermo-acoustic oscillations.

The E3 combustors: Status and challenges. [energy efficient turbofan engines

NASA Technical Reports Server (NTRS)

Sokolowski, D. E.; Rohde, J. E.

1981-01-01

The design, fabrication, and initial testing of energy efficient engine combustors, developed for the next generation of turbofan engines for commercial aircraft, are described. The combustor designs utilize an annular configuration with two zone combustion for low emissions, advanced liners for improved durability, and short, curved-wall, dump prediffusers for compactness. Advanced cooling techniques and segmented construction characterize the advanced liners. Linear segments are made from castable, turbine-type materials.

Low NOx Heavy Fuel Combustor Concept Program

NASA Technical Reports Server (NTRS)

Novick, A. S.; Troth, D. L.

1981-01-01

The development of the technology required to operate an industrial gas turbine combustion system on minimally processed, heavy petroleum or residual fuels having high levels of fuel-bound nitrogen (FBN) while producing acceptable levels of exhaust emissions is discussed. Three combustor concepts were designed and fabricated. Three fuels were supplied for the combustor test demonstrations: a typical middle distillate fuel, a heavy residual fuel, and a synthetic coal-derived fuel. The primary concept was an air staged, variable-geometry combustor designed to produce low emissions from fuels having high levels of FBN. This combustor used a long residence time, fuel-rich primary combustion zone followed by a quick-quench air mixer to rapidly dilute the fuel rich products for the fuel-lean final burnout of the fuel. This combustor, called the rich quench lean (RQL) combustor, was extensively tested using each fuel over the entire power range of the model 570 K engine. Also, a series of parameteric tests was conducted to determine the combustor's sensitivity to rich-zone equivalence ratio, lean-zone equivalence ratio, rich-zone residence time, and overall system pressure drop. Minimum nitrogen oxide emissions were measured at 50 to 55 ppmv at maximum continuous power for all three fuels. Smoke was less than a 10 SAE smoke number.

Pollution emissions from single swirl-can combustor modules at parametric test conditions

NASA Technical Reports Server (NTRS)

Mularz, E. J.; Wear, J. D.; Verbulecz, P. W.

1975-01-01

Exhaust pollutant emissions were measured from single swirl-can combustor modules operating over a pressure range of 69 to 276 N/sq cm (100 to 400 psia), over a fuel-air ratio range of 0.01 to 0.04, at an inlet air temperature of 733 K (860 F), and at a constant reference velocity of 23.2 m/sec). Many swirl-can module designs were evaluated; the 11 most promising designs exhibited oxides of nitrogen emission levels lower than that from conventional gas-turbine combustors. Although these single module test results are not necessarily indicative of the performance characteristics of a large array of modules, the results are very promixing and offer a number of module designs that should be tested in a full combustor.

Investigation of a low NOx full-scale annular combustor

NASA Technical Reports Server (NTRS)

1982-01-01

An atmospheric test program was conducted to evaluate a low NOx annular combustor concept suitable for a supersonic, high-altitude aircraft application. The lean premixed combustor, known as the vortex air blast (VAB) concept, was tested as a 22.0-cm diameter model in the early development phases to arrive at basic design and performance criteria. Final demonstration testing was carried out on a full scale combustor of 0.66-m diameter. Variable geometry dilution ports were incorporated to allow operation of the combustor across the range of conditions between idle (T(in) = 422 K, T(out) = 917 K) and cruise (T(in) = 833 K, T(out) - 1778 K). Test results show that the design could meet the program NOx goal of 1.0 g NO2/kg fuel at a one-atmospheric simulated cruise condition.

Analytical fuel property effects: Small combustors, phase 2

NASA Technical Reports Server (NTRS)

Hill, T. G.; Monty, J. D.; Morton, H. L.

1985-01-01

The effects of non-standard aviation fuels on a typical small gas turbine combustor were studied and the effectiveness of design changes intended to counter the effects of these fuels was evaluated. The T700/CT7 turboprop engine family was chosen as being representative of the class of aircraft power plants desired for this study. Fuel properties, as specified by NASA, are characterized by low hydrogen content and high aromatics levels. No. 2 diesel fuel was also evaluated in this program. Results demonstrated the anticipated higher than normal smoke output and flame radiation intensity with resulting increased metal temperatures on the baseline T700 combustor. Three new designs were evaluated using the non standard fuels. The three designs incorporated enhanced cooling features and smoke reduction features. All three designs, when burning the broad specification fuels, exhibited metal temperatures at or below the baseline combustor temperatures on JP-5. Smoke levels were acceptable but higher than predicted.

Pollutant emissions from and within a model gas turbine combustor at elevated pressures and temperatures

NASA Technical Reports Server (NTRS)

Drennan, S. A.; Peterson, C. O.; Khatib, F. M.; Sowa, W. A.; Samuelsen, G. S.

1993-01-01

Conventional and advanced gas turbine engines are coming under increased scrutiny regarding pollutant emissions. This, in turn, has created a need to obtain in-situ experimental data at practical conditions, as well as exhaust data, and to obtain the data in combustors that reflect modern designs. The in-situ data are needed to (1) assess the effects of design modifications on pollutant formation, and (2) develop a detailed data base on combustor performance for the development and verification of computer modeling. This paper reports on a novel high pressure, high temperature facility designed to acquire such data under controlled conditions and with access (optical and extractive) for in-situ measurements. To evaluate the utility of the facility, a model gas turbine combustor was selected which features practical hardware design, two rows of jets (primary and dilution) with four jets in each row, and advanced wall cooling techniques with laser drilled effusive holes. The dome is equipped with a flat-vaned swirler with vane angles of 60 degrees. Data are obtained at combustor pressures ranging from 2 to 10 atmospheres of pressure, levels of air preheat to 427 C, combustor reference velocities from 10.0 to 20.0 m/s, and an overall equivalence ratio of 0.3. Exit plane and in-situ measurements are presented for HC, O2, CO2, CO, and NO(x). The exit plane emissions of NO(x) correspond to levels reported from practical combustors and the in-situ data demonstrate the utility and potential for detailed flow field measurements.

Methods for reducing pollutant emissions from jet aircraft

NASA Technical Reports Server (NTRS)

Butze, H. F.

1971-01-01

Pollutant emissions from jet aircraft and combustion research aimed at reducing these emissions are defined. The problem of smoke formation and results achieved in smoke reduction from commercial combustors are discussed. Expermental results of parametric tests performed on both conventional and experimental combustors over a range of combustor-inlet conditions are presented. Combustor design techniques for reducing pollutant emissions are discussed. Improved fuel atomization resulting from the use of air-assist fuel nozzles has brought about significant reductions in hydrocarbon and carbon monoxide emissions at idle. Diffuser tests have shown that the combustor-inlet airflow profile can be controlled through the use of diffuser-wall bleed and that it may thus be possible to reduce emissions by controlling combustor airflow distribution. Emissions of nitric oxide from a shortlength annular swirl-can combustor were significantly lower than those from a conventional combustor operating at similar conditions.

Emission Modeling of an Interturbine Burner Based on Flameless Combustion

PubMed Central

2017-01-01

Since its discovery, the flameless combustion (FC) regime has been a promising alternative to reduce pollutant emissions of gas turbine engines. This combustion mode is characterized by well-distributed reaction zones, which potentially decreases temperature gradients, acoustic oscillations, and NOx emissions. Its attainment within gas turbine engines has proved to be challenging because previous design attempts faced limitations related to operational range and combustion efficiency. Along with an aircraft conceptual design, the AHEAD project proposed a novel hybrid engine. One of the key features of the proposed hybrid engine is the use of two combustion chambers, with the second combustor operating in the FC mode. This novel configuration would allow the facilitation of the attainment of the FC regime. The conceptual design was adapted to a laboratory scale combustor that was tested at elevated temperature and atmospheric pressure. In the current work, the emission behavior of this scaled combustor is analyzed using computational fluid dynamics (CFD) and chemical reactor network (CRN). The CFD was able to provide information with the flow field in the combustor, while the CRN was used to model and predict emissions. The CRN approach allowed the analysis of the NOx formation pathways, indicating that the prompt NOx was the dominant pathway in the combustor. The combustor design can be improved by modifying the mixing between fuel and oxidizer as well as the split between combustion and dilution air. PMID:29910533

A conceptual design of shock-eliminating clover combustor for large scale scramjet engine

NASA Astrophysics Data System (ADS)

Sun, Ming-bo; Zhao, Yu-xin; Zhao, Guo-yan; Liu, Yuan

2017-01-01

A new concept of shock-eliminating clover combustor is proposed for large scale scramjet engine to fulfill the requirements of fuel penetration, total pressure recovery and cooling. To generate the circular-to-clover transition shape of the combustor, the streamline tracing technique is used based on an axisymmetric expansion parent flowfield calculated using the method of characteristics. The combustor is examined using inviscid and viscous numerical simulations and a pure circular shape is calculated for comparison. The results showed that the combustor avoids the shock wave generation and produces low total pressure losses in a wide range of flight condition with various Mach number. The flameholding device for this combustor is briefly discussed.

Results of the pollution reduction technology program for turboprop engines

NASA Technical Reports Server (NTRS)

Mularz, E. J.

1976-01-01

A program was performed to evolve and demonstrate advanced combustor technology aimed at achieving the 1979 EPA standards for turboprop engines (Class P2). The engine selected for this program was the 501-D22A turboprop. Three combustor concepts were designed and tested in a combustor rig at the exact combustor operating conditions of the 50-D22A engine over the EPA landing-takeoff cycle. Each combustor concept exhibited pollutant emissions well below the EPA standards, achieving substantial reductions in unburned hydrocarbons, carbon monoxide, and smoke emissions compared with emissions from the production combustor of this engine. Oxides of nitrogen emissions remained well below the EPA standards, also.

Development of a short length combustor for a supersonic cruise turbofan engine using a 90 deg sector of a full annulus

NASA Technical Reports Server (NTRS)

Clements, T. R.

1972-01-01

A performance development program has been conducted on a short length, double-annular, ram-induction combustor. The combustor was designed for a large augmented turbofan engine capable of sustained flight speeds up to Mach 3.0. Performance tests were conducted at an inlet temperature and Mach number simulating engine sea level takeoff conditions. At the design temperature rise of 1600 F, combustion efficiency was 100%, pattern factor was 0.20, and combined diffuser-combustor pressure loss was 4.4% or 1.12 times the diffuser inlet velocity head. A temperature rise in excess of 2400 F with a combustion efficiency of 94% was demonstrated.

CFD-Based Design of a Filming Injector for N+3 Combustors

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Mongia, Hukam; Lee, Phil

2016-01-01

An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements coupled with a new fuel-filming injector design for next-generation N+3 combustors. The National Combustion Code (NCC) was used to perform non-reacting and two-phase reacting flow computations on a N+3 injector configuration, in a single-element and a five-element injector array. All computations were performed with a consistent approach towards mesh-generation, spray-, ignition- and kinetics-modeling with the NCC. Computational predictions of the aerodynamics of the injector were used to arrive at an optimal injector design that met effective area, aerodynamics, and fuel-air mixing criteria. LDI-3 emissions (EINOx, EICO and UHC) were compared with the previous generation LDI-2 combustor experimental data at representative engine cycle conditions.

Effect of Axially Staged Fuel Introduction on Performance of One-quarter Sector of Annular Turbojet Combustor

NASA Technical Reports Server (NTRS)

Zettle, Eugene V; Mark, Herman

1953-01-01

The design principle of injecting liquid fuel at more than one axial station in an annual turbojet combustor was investigated. Fuel was injected into the combustor as much as 5 inches downstream of the primary fuel injectors. Many fuel-injection configurations were examined and the performance results are presented for 11 configurations that best demonstrate the trends in performance obtained. The performance investigations were made at a constant combustor-inlet pressure of 15 inches of mercury absolute and at air flows up to 70 percent higher than values typical of current design practice. At these higher air flows, staging the fuel introduction improved the combustion efficiency considerably over that obtained in the combustor when no fuel staging was employed. At air flows currently encountered in turbojet engines, fuel staging was of minor value. Radial temperature distribution seemed relatively unaffected by the location of fuel-injection stations.

Coherent Anti-Stokes Raman Spectroscopic Thermometry in a Supersonic Combustor

NASA Technical Reports Server (NTRS)

Cutler, A. D.; Danehy, P. M.; Springer, R. R.; OByrne, S.; Capriotti, D. P.; DeLoach, R.

2003-01-01

An experiment has been conducted to acquire data for the validation of computational fluid dynamics codes used in the design of supersonic combustors. The flow in a supersonic combustor, consisting of a diverging duct with a single downstream-angled wail injector, is studied. Combustor entrance Mach number is 2 and enthalpy nominally corresponds to Mach 7 flight. The primary measurement technique is coherent anti-Stokes Raman spectroscopy, but surface pressures and temperatures have also been acquired. Modern design of experiment techniques have been used to maximize the quality of the data set (for the given level of effort) and to minimize systematic errors. Temperature maps are obtained at several planes in the flow for a case in which the combustor is piloted by injecting fuel upstream of the main injector and one case in which it is not piloted. Boundary conditions and uncertainties are characterized.

Introducing the VRT gas turbine combustor

NASA Technical Reports Server (NTRS)

Melconian, Jerry O.; Mostafa, Abdu A.; Nguyen, Hung Lee

1990-01-01

An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer model predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.

Flashback Arrestor for LPP, Low NOx Combustors

NASA Technical Reports Server (NTRS)

Kraemer, Gil; Lee, Chi-Ming

1998-01-01

Lean premixed, prevaporized (LPP) high temperature combustor designs as explored for the Advanced Subsonic Transport (AST) and High Speed Civil Transport (HSCT) combustors can achieve low NO(x), emission levels. An enabling device is needed to arrest flashback and inhibit preignition at high power conditions and during transients (surge and rapid spool down). A novel flashback arrestor design has demonstrated the ability to arrest flashback and inhibit preignition in a 4.6 cm diameter tubular reactor at full power inlet temperatures (725 C) using Jet-A fuel at 0.4 less than or equal To phi less than or equal to 3.5. Several low pressure loss (0.2 to 0.4% at 30 m/s) flashback arrestor designs were developed which arrested flashback at all of the test conditions. Flame holding was also inhibited off the flash arrestor face or within the downstream tube even velocities (less than or equal to 3 to 6 m/s), thus protecting the flashback arrestor and combustor components. Upstream flow conditions influence the specific configuration based on using either a 45% or 76% upstream geometric blockage. Stationary, lean premixed dry low NO(x) gas turbine combustors would also benefit from this low pressure drop flashback arrestor design which can be easily integrated into new and existing designs.

The VRT gas turbine combustor - Phase II

NASA Technical Reports Server (NTRS)

Melconian, Jerry O.; Mongia, Hukam C.; Nguyen, Hung L.

1992-01-01

An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual-function use of the incoming air. In Phase I, the feasibility of the concept was demonstrated by water analogue tests and 3D computer modeling. The flow pattern within the combustor was as predicted. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. In Phase II, hardware was designed, procured, and tested under conditions simulating typical supersonic civil aircraft cruise conditions to the limits of the rig. The test results confirmed many of the superior performance predictions of the VRT concept. The Hastelloy X liner showed no signs of distress after nearly six hours of tests using JP5 fuel.

75 FR 82370 - Approval and Promulgation of State Plans for Designated Facilities and Pollutants; State of...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-12-30

... Guidelines (EGs) applicable to existing Large Municipal Waste Combustors (LMWCs). These EGs apply to municipal waste combustors with a capacity to combust more than 250 tons per day of municipal solid waste... Municipal Waste Combustor (LMWC) Emissions From Existing Facilities AGENCY: Environmental Protection Agency...

40 CFR Table 1 to Subpart Cb of... - Nitrogen Oxides Guidelines for Designated Facilities

Code of Federal Regulations, 2010 CFR

2010-07-01

... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) STANDARDS OF PERFORMANCE FOR NEW STATIONARY SOURCES Emissions... (parts per million by volume) a Mass burn waterwall 205 205. Mass burn rotary waterwall 250 210. Refuse-derived fuel combustor 250 250. Fluidized bed combustor 180 180. Mass burn refractory combustors No limit...

40 CFR Table 1 to Subpart Cb of... - Nitrogen Oxides Guidelines for Designated Facilities

Code of Federal Regulations, 2012 CFR

2012-07-01

... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) STANDARDS OF PERFORMANCE FOR NEW STATIONARY SOURCES Emissions... (parts per million by volume) a Mass burn waterwall 205 205. Mass burn rotary waterwall 250 210. Refuse-derived fuel combustor 250 250. Fluidized bed combustor 180 180. Mass burn refractory combustors No limit...

40 CFR Table 1 to Subpart Cb of... - Nitrogen Oxides Guidelines for Designated Facilities

Code of Federal Regulations, 2014 CFR

2014-07-01

... AGENCY (CONTINUED) AIR PROGRAMS (CONTINUED) STANDARDS OF PERFORMANCE FOR NEW STATIONARY SOURCES Emissions... (parts per million by volume) a Mass burn waterwall 205 205. Mass burn rotary waterwall 250 210. Refuse-derived fuel combustor 250 250. Fluidized bed combustor 180 180. Mass burn refractory combustors No limit...

Dish stirling solar receiver combustor test program

NASA Technical Reports Server (NTRS)

Bankston, C. P.; Back, L. H.

1981-01-01

The operational and energy transfer characteristics of the Dish Stirling Solar Receiver (DSSR) combustor/heat exchanger system was evaluated. The DSSR is designed to operate with fossil fuel augmentation utilizing a swirl combustor and cross flow heat exchanger consisting of a single row of 4 closely spaced tubes that are curved into a conical shape. The performance of the combustor/heat exchanger system without a Stirling engine was studied over a range of operating conditions and output levels using water as the working fluid. Results show that the combustor may be started under cold conditions, controlled safety, and operated at a constant air/fuel ratio (10 percent excess air) over the required range of firing rates. Furthermore, nondimensional heat transfer coefficients based on total heat transfer are plotted versus Reynolds number and compared with literature data taken for single rows of closely spaced tubes perpendicular to cross flow. The data show enhanced heat transfer for the present geometry and test conditions. Analysis of the results shows that the present system meets specified thermal requirements, thus verifying the feasibility of the DSSR combustor design for final prototype fabrication.

Quiet Clean Short-haul Experimental Engine (QCSEE) clean combustor test report

NASA Technical Reports Server (NTRS)

1975-01-01

A component pressure test was conducted on a F101 PFRT combustor to evaluate the emissions levels of this combustor design at selected under the wing and over the wing operating conditions for the quiet clean short haul experimental engine (QCSEE). Emissions reduction techniques were evaluated which included compressor discharge bleed and sector burning in the combustor. The results of this test were utilized to compare the expected QCSEE emissions levels with the emission goals of the QCSEE engine program.

Combustion Dynamics Behavior in a Single-Element Lean Direct Injection (LDI) Gas Turbine Combustor

DTIC Science & Technology

2014-06-01

Constant mass inflow from a choked orifice Exit Boundary Condition Choked nozzle Diameter of combustor 50.8 mm Diameter of air plenum 25.4 mm A...schematic of the LDI combustor is shown in Fig. 1. It comprises an air inlet section, air plenum, swirler- venturi- injector assembly, combustion chamber...and exit nozzle . Air, heated with an 80 kW electrical heater, enters the combustor through a slotted choked orifice plate, designed to minimize

Wide range operation of advanced low NOx aircraft gas turbine combustors

NASA Technical Reports Server (NTRS)

Roberts, P. B.; Fiorito, R. J.; Butze, H. F.

1978-01-01

The paper summarizes the results of an experimental test rig program designed to define and demonstrates techniques which would allow the jet-induced circulation and vortex air blast combustors to operate stably with acceptable emissions at simulated engine idle without compromise to the low NOx emissions under the high-altitude supersonic cruise condition. The discussion focuses on the test results of the key combustor modifications for both the simulated engine idle and cruise conditions. Several range-augmentation techniques are demonstrated that allow the lean-reaction premixed aircraft gas turbine combustor to operate with low NOx emissons at engine cruise and acceptable CO and UHC levels at engine idle. These techniques involve several combinations, including variable geometry and fuel switching designs.

Combustion Dynamics and Control for Ultra Low Emissions in Aircraft Gas-Turbine Engines

NASA Technical Reports Server (NTRS)

DeLaat, John C.

2011-01-01

Future aircraft engines must provide ultra-low emissions and high efficiency at low cost while maintaining the reliability and operability of present day engines. The demands for increased performance and decreased emissions have resulted in advanced combustor designs that are critically dependent on efficient fuel/air mixing and lean operation. However, all combustors, but most notably lean-burning low-emissions combustors, are susceptible to combustion instabilities. These instabilities are typically caused by the interaction of the fluctuating heat release of the combustion process with naturally occurring acoustic resonances. These interactions can produce large pressure oscillations within the combustor and can reduce component life and potentially lead to premature mechanical failures. Active Combustion Control which consists of feedback-based control of the fuel-air mixing process can provide an approach to achieving acceptable combustor dynamic behavior while minimizing emissions, and thus can provide flexibility during the combustor design process. The NASA Glenn Active Combustion Control Technology activity aims to demonstrate active control in a realistic environment relevant to aircraft engines by providing experiments tied to aircraft gas turbine combustors. The intent is to allow the technology maturity of active combustion control to advance to eventual demonstration in an engine environment. Work at NASA Glenn has shown that active combustion control, utilizing advanced algorithms working through high frequency fuel actuation, can effectively suppress instabilities in a combustor which emulates the instabilities found in an aircraft gas turbine engine. Current efforts are aimed at extending these active control technologies to advanced ultra-low-emissions combustors such as those employing multi-point lean direct injection.

Results of the NASP Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test Program

NASA Technical Reports Server (NTRS)

Cavolowsky, John A.; Loomis, Mark P.; Deiwert, George S.

1995-01-01

This paper describes the test techniques and results from the National Aerospace Plane Government Work Package 53, the Ames Integrated Mixing Hypersonic Engine (AIMHYE) Scramjet Test program conducted in the NASA Ames 16-Inch Combustion Driven Shock Tunnel. This was a series of near full-scale scramjet combustor tests with the objective to obtain high speed combustor and nozzle data from an engine with injector configurations similar to the NASP E21 and E22a designs. The experimental test approach was to use a large combustor model (80-100% throat height) designed and fabricated for testing in the semi-free jet mode. The conditions tested were similar to the "blue book" conditions at Mach 12, 14, and 16. GWP 53 validated use of large, long test time impulse facilities, specifically the Ames 16-Inch Shock Tunnel, for high Mach number scramjet propulsion testing an integrated test rig (inlet, combustor, and nozzle). Discussion of key features of the test program will include: effects of the 2-D combustor inlet pressure profile; performance of large injectors' fueling system that included nozzlettes, base injection, and film cooling; and heat transfer measurements to the combustor. Significant instrumentation development and application efforts include the following: combustor force balance application for measurement of combustor drag for comparison with integrated point measurements of skin friction; nozzle metric strip for measuring thrust with comparison to integrated pressure measurements; and nonintrusive optical fiber-based diode laser absorption measurements of combustion products for determination of combustor performance. Direct measurements will be reported for specific test article configurations and compared with CFD solutions.

Challenges to Laser-Based Imaging Techniques in Gas Turbine Combustor Systems for Aerospace Applications

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Anderson, Robert C.; Zaller, Michelle M.; Hicks, Yolanda R.

1998-01-01

Increasingly severe constraints on emissions, noise and fuel efficiency must be met by the next generation of commercial aircraft powerplants. At NASA Lewis Research Center (LeRC) a cooperative research effort with industry is underway to design and test combustors that will meet these requirements. To accomplish these tasks, it is necessary to gain both a detailed understanding of the combustion processes and a precise knowledge of combustor and combustor sub-component performance at close to actual conditions. To that end, researchers at LeRC are engaged in a comprehensive diagnostic investigation of high pressure reacting flowfields that duplicate conditions expected within the actual engine combustors. Unique, optically accessible flame-tubes and sector rig combustors, designed especially for these tests. afford the opportunity to probe these flowfields with the most advanced, laser-based optical diagnostic techniques. However, these same techniques, tested and proven on comparatively simple bench-top gaseous flame burners, encounter numerous restrictions and challenges when applied in these facilities. These include high pressures and temperatures, large flow rates, liquid fuels, remote testing, and carbon or other material deposits on combustor windows. Results are shown that document the success and versatility of these nonintrusive optical diagnostics despite the challenges to their implementation in realistic systems.

Development of a Catalytic Combustor for Aircraft Gas Turbine Engines.

DTIC Science & Technology

1976-09-22

80 VI. DESIGN OF 7.6 CI CIANETE COMBUSTORS . . . . . . . . . . . 86 1. Design and Fabrication of CombusLors for Large Scale T est in...obtained for this program included round holes of different diameters, squares, rectangles, triangles, and other more complex hollow configurations

Study on the high speed scramjet characteristics at Mach 10 to 15 flight condition

NASA Astrophysics Data System (ADS)

Takahashi, M.; Itoh, K.; Tanno, H.; Komuro, T.; Sunami, T.; Sato, K.; Ueda, S.

A scramjet engine model, designed to establish steady and strong combustion at free-stream conditions corresponding to Mach 12 flight, was tested in a large free-piston driven shock tunnel. Combustion tests of a previous engine model showed that combustion heat release obtained in the combustor was not sufficient to maintain strong combustion. For a new scramjet engine model, the inlet compression ratio was increased to raise the static temperature and density of the flow at the combustor entrance. As a result of the aerodynamic design change, the pressure rise due to combustion increased and the duration of strong combustion conditions in the combustor was extended. A hyper-mixer injector designed to enhance mixing and combustion by introducing streamwise vortices was applied to the new engine model. The results showed that the hyper mixer injector was very effective in promoting combustion heat release and establishing steady and strong combustion in the combustor.

Active Combustion Control for Aircraft Gas-Turbine Engines-Experimental Results for an Advanced, Low-Emissions Combustor Prototype

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Kopasakis, George; Saus, Joseph R.; Chang, Clarence T.; Wey, Changlie

2012-01-01

Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. Mitigation of instabilities is an enabling technology for these low-emissions combustors. NASA Glenn Research Center s prior activity has demonstrated active control to suppress a high-frequency combustion instability in a combustor rig designed to emulate an actual aircraft engine instability experience with a conventional, rich-front-end combustor. The current effort is developing further understanding of the problem specifically as applied to future lean-burning, very low-emissions combustors. A prototype advanced, low-emissions aircraft engine combustor with a combustion instability has been identified and previous work has characterized the dynamic behavior of that combustor prototype. The combustor exhibits thermoacoustic instabilities that are related to increasing fuel flow and that potentially prevent full-power operation. A simplified, non-linear oscillator model and a more physics-based sectored 1-D dynamic model have been developed to capture the combustor prototype s instability behavior. Utilizing these models, the NASA Adaptive Sliding Phasor Average Control (ASPAC) instability control method has been updated for the low-emissions combustor prototype. Active combustion instability suppression using the ASPAC control method has been demonstrated experimentally with this combustor prototype in a NASA combustion test cell operating at engine pressures, temperatures, and flows. A high-frequency fuel valve was utilized to perturb the combustor fuel flow. Successful instability suppression was shown using a dynamic pressure sensor in the combustor for controller feedback. Instability control was also shown with a pressure feedback sensor in the lower temperature region upstream of the combustor. It was also demonstrated that the controller can prevent the instability from occurring while combustor operation was transitioning from a stable, low-power condition to a normally unstable high-power condition, thus enabling the high-power condition.

An Adaptive Instability Suppression Controls Method for Aircraft Gas Turbine Engine Combustors

NASA Technical Reports Server (NTRS)

Kopasakis, George; DeLaat, John C.; Chang, Clarence T.

2008-01-01

An adaptive controls method for instability suppression in gas turbine engine combustors has been developed and successfully tested with a realistic aircraft engine combustor rig. This testing was part of a program that demonstrated, for the first time, successful active combustor instability control in an aircraft gas turbine engine-like environment. The controls method is called Adaptive Sliding Phasor Averaged Control. Testing of the control method has been conducted in an experimental rig with different configurations designed to simulate combustors with instabilities of about 530 and 315 Hz. Results demonstrate the effectiveness of this method in suppressing combustor instabilities. In addition, a dramatic improvement in suppression of the instability was achieved by focusing control on the second harmonic of the instability. This is believed to be due to a phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling. These results may have implications for future research in combustor instability control.

40 CFR Table 1 to Subpart Fff of... - Municipal Waste Combustor Units (MWC Units) Excluded From Subpart FFF 1

Code of Federal Regulations, 2013 CFR

2013-07-01

... 40 Protection of Environment 9 2013-07-01 2013-07-01 false Municipal Waste Combustor Units (MWC... FOR DESIGNATED FACILITIES AND POLLUTANTS Federal Plan Requirements for Large Municipal Waste... Part 62—Municipal Waste Combustor Units (MWC Units) Excluded From Subpart FFF 1 State MWC units Alabama...

Broad specification fuels technology program, phase 1

NASA Technical Reports Server (NTRS)

Lohmann, R. P.; Jeroszko, R. A.

1982-01-01

An experimental evaluation was conducted to assess the impact of the use of broadened properties fuels on combustor design concepts. Emphasis was placed on establishing the viability of design modifications to current combustor concepts and the use of advanced technology concepts to facilitate operation on Experimental Referee Broad Specification (ERBS) fuel while meeting exhaust emissions and performance specifications and maintaining acceptable durability. Three different combustor concepts, representative of progressively more aggressive technology levels, were evaluated. When operated on ERBS rather than Jet A fuel, a single stage combustor typical of that in the most recent versions of the JT9D-7 engine was found to produce excess carbon monoxide emissions at idle and elevated liner temperatures at high power levels that were projected to reduced liner life by 13 percent. The introduction of improved component technology, such as refined fuel injectors and advanced liner cooling concepts were shown to have the potential of enhancing the fuel flexibility of the single stage combustor.

Introducing the VRT gas turbine combustor

DOE Office of Scientific and Technical Information (OSTI.GOV)

Melconian, J.O.; Mostafa, A.A.; Nguyen, H.L.

An innovative annular combustor configuration is being developed for aircraft and other gas turbine engines. This design has the potential of permitting higher turbine inlet temperatures by reducing the pattern factor and providing a major reduction in NO(x) emission. The design concept is based on a Variable Residence Time (VRT) technique which allows large fuel particles adequate time to completely burn in the circumferentially mixed primary zone. High durability of the combustor is achieved by dual function use of the incoming air. The feasibility of the concept was demonstrated by water analogue tests and 3-D computer modeling. The computer modelmore » predicted a 50 percent reduction in pattern factor when compared to a state of the art conventional combustor. The VRT combustor uses only half the number of fuel nozzles of the conventional configuration. The results of the chemical kinetics model require further investigation, as the NO(x) predictions did not correlate with the available experimental and analytical data base.« less

CFD Evaluation of a 3rd Generation LDI Combustor

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Mongia, Hukam; Lee, Phil

2017-01-01

An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements for next-generation LDI-3 combustor design. The National Combustion Code (NCC) was used to perform non-reacting and two-phase reacting flow computations for a nineteen-element injector array arranged in a three-module, 7-5-7 element configuration. All computations were performed with a consistent approach of mesh-optimization, spray-modeling, ignition and kinetics-modeling with the NCC. Computational predictions of the aerodynamics of the injector were used to arrive at an optimal injector design that meets effective area and fuel-air mixing criteria. LDI-3 emissions (EINOx, EICO and UHC) were compared with the previous generation LDI-2 combustor experimental data at representative engine cycle conditions.

Novel designs of fluidized bed combustors for low pollutant emissions

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lin, W.; Bleek, C.M. van den; Dam-Johansen, K.

1995-12-31

It is known that NH{sub 3}, released during the devolatilization of fuel, is an important precursor for NO formation in fluidized bed combustors. On the other hand, NH{sub 3} may be used as a reducing agent in the thermal DeNO{sub x} process to reduce NO{sub x} emission levels. In this paper, a new concept of fluidized bed combustors is proposed based on the idea of in situ reduction of NO{sub x} by self-produced NH{sub 3} from fuel without lowering the sulfur capture level. This design is intended to separate the NH{sub 3} release process under reducing conditions from the charmore » combustion process under oxidizing conditions; this self-released NH{sub 3}, together with some combustibles, is mixed with gaseous combustion products in the upper part of the combustor for a further reduction of the NO{sub x} formed during combustion. Furthermore, the combustion of the combustibles may cause the temperature to rise in this upper zone and thereby reduce the emission of N{sub 2}O. The applications of this design to bubbling and circulating fluidized bed combustors are described and the mechanisms of the main reactions involved discussed.« less

NASA Broad Specification Fuels Combustion Technology program - Pratt and Whitney Aircraft Phase I results and status

NASA Technical Reports Server (NTRS)

Lohmann, R. P.; Fear, J. S.

1982-01-01

In connection with increases in the cost of fuels and the reduced availability of high quality petroleum crude, a modification of fuel specifications has been considered to allow acceptance of poorer quality fuels. To obtain the information upon which a selection of appropriate fuels for aircraft can be based, the Broad Specification Fuels Combustion Technology program was formulated by NASA. A description is presented of program-related investigations conducted by an American aerospace company. The specific objective of Phase I of this program has been to evaluate the impact of the use of broadened properties fuels on combustor design through comprehensive combustor rig testing. Attention is given to combustor concepts, experimental evaluation, results obtained with single stage combustors, the stage combustor concept, and the capability of a variable geometry combustor.

NASA/General Electric broad-specification fuels combustion technology program - Phase I results and status

NASA Technical Reports Server (NTRS)

Dodds, W. J.; Ekstedt, E. E.; Bahr, D. W.; Fear, J. S.

1982-01-01

A program is being conducted to develop the technology required to utilize fuels with broadened properties in aircraft gas turbine engines. The first phase of this program consisted of the experimental evaluation of three different combustor concepts to determine their potential for meeting several specific emissions and performance goals, when operated on broadened property fuels. The three concepts were a single annular combustor; a double annular combustor; and a short single annular combustor with variable geometry. All of these concepts were sized for the General Electric CF6-80 engine. A total of 24 different configurations of these concepts were evaluated in a high pressure test facility, using four test fuels having hydrogen contents between 11.8 and 14%. Fuel effects on combustor performance, durability and emissions, and combustor design features to offset these effects were demonstrated.

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

NASA Technical Reports Server (NTRS)

Diehl, L. A.; Trout, A. M.

1976-01-01

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.

Wide range operation of advanced low NOx combustors for supersonic high-altitude aircraft gas turbines

NASA Technical Reports Server (NTRS)

Roberts, P. B.; Fiorito, R. J.

1977-01-01

An initial rig program tested the Jet Induced Circulation (JIC) and Vortex Air Blast (VAB) systems in small can combustor configurations for NOx emissions at a simulated high altitude, supersonic cruise condition. The VAB combustor demonstrated the capability of meeting the NOx goal of 1.0 g NO2/kg fuel at the cruise condition. In addition, the program served to demonstrate the limited low-emissions range available from the lean, premixed combustor. A follow-on effort was concerned with the problem of operating these lean, premixed combustors with acceptable emissions at simulated engine idle conditions. Various techniques have been demonstrated that allow satisfactory operation on both the JIC and VAB combustors at idle with CO emissions below 20 g/kg fuel. The VAB combustor was limited by flashback/autoignition phenomena at the cruise conditions to a pressure of 8 atmospheres. The JIC combustor was operated up to the full design cruise pressure of 14 atmospheres without encountering an autoignition limitation although the NOx levels, in the 2-3 g NO2/kg fuel range, exceeded the program goal.

Supplementary steam - A viable hydrogen power generation concept

NASA Technical Reports Server (NTRS)

Wright, D. E.; Lee, J. C.

1979-01-01

Technical and economic aspects of a supplementary steam generation for peaking power applications are discussed. Preliminary designs of the hydrogen/oxygen combustors to be used for such applications are described. The integration of the hydrogen/oxygen steam-generating equipment into a typical coal-fired steam station is studied. The basic steam generation system was designed as a 20 MW supplementary system to be added to the existing 160 MW system. An analysis of the operating and design requirements of the supplementary system is conducted. Estimates were made for additional steam and fuel supply lines and for additional control required to operate the combustors and to integrate the combustor system into the facility.

Systems Design and Experimental Evaluation of a High-Altitude Relight Test Facility

NASA Astrophysics Data System (ADS)

Paxton, Brendan

Novel advances in gas turbine engine combustor technology, led by endeavors into fuel efficiency and demanding environmental regulations, have been fraught with performance and safety concerns. While the majority of low emissions gas turbine engine combustor technology has been necessary for power generation applications, the push for ultra-low NOx combustion in aircraft jet engines has been ever present. Recent state-of-the-art combustor designs notably tackle historic emissions challenges by operating at fuel-lean conditions, which are characterized by an increase in the amount of air flow sent to the primary combustion zone. While beneficial in reducing NOx emissions, the fuel-lean mechanisms that characterize these combustor designs rely heavily upon high-energy and high-velocity air flows to sufficiently mix and atomize fuel droplets, ultimately leading to flame stability concerns during low-power operation. When operating at high-altitude conditions, these issues are further exacerbated by the presence of low ambient air pressures and temperatures, which can lead to engine flame-out situations and hamper engine relight attempts. To aid academic and industrial research ventures into improving the high-altitude lean blow-out and relight performance of modern gas turbine engine combustor technologies, the High-Altitude Relight Test Facility (HARTF) was designed and constructed at the University of Cincinnati (UC) Combustion and Fire Research Laboratory (CFRL). Following its construction, an experimental evaluation of its abilities to facilitate optically-accessible ignition, combustion, and spray testing for gas turbine engine combustor hardware at simulated high-altitude conditions was performed. In its evaluation, performance limit references were established through testing of the HARTF vacuum and cryogenic air-chilling capabilities. These tests were conducted with regard to end-user control---the creation and the maintenance of a realistic high-altitude environment simulation. To evaluate future testing applications, as well as to understand the abilities of the HARTF to accommodate different sizes and configurations of industrial gas turbine engine combustor hardware, ignition testing was conducted at challenging high-altitude windmilling conditions with a linearly-arranged five-swirler array, replicating the implementation of a multi-cup combustor sector.

Effects of broadened property fuels on radiant heat flux to gas turbine combustor liners

NASA Technical Reports Server (NTRS)

Haggard, J. B., Jr.

1983-01-01

The effects of fuel type, inlet air pressure, inlet air temperature, and fuel/air ratio on the combustor radiation were investigated. Combustor liner radiant heat flux measurements were made in the spectral region between 0.14 and 6.5 microns at three locations in a modified commercial aviation can combustor. Two fuels, Jet A and a heavier distillate research fuel called ERBS were used. The use of ERBS fuel as opposed to Jet A under similar operating conditions resulted in increased radiation to the combustor liner and hence increased backside liner temperature. This increased radiation resulted in liner temperature increases always less than 73 C. The increased radiation is shown by way of calculations to be the result of increased soot concentrations in the combustor. The increased liner temperatures indicated can substantially affect engine maintenance costs by reducing combustor liner life up to 1/3 because of the rapid decay in liner material properties when operated beyond their design conditions.

A Comparison of Combustor-Noise Models

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.

2012-01-01

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.

Rapid mix concepts for low emission combustors in gas turbine engines

NASA Technical Reports Server (NTRS)

Talpallikar, Milind V.; Smith, Clifford E.; Lai, Ming-Chia

1990-01-01

NASA LeRC has identified the Rich burn/Quick mix/Lean burn (RQL) combustor as a potential gas turbine combustor concept to reduce NOx emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NOx levels, NASA LeRC soon will test a flametube version of an RQL combustor. The critical technology needed for the RQL combustor is a method of quickly mixing combustion air with rich burn gases. Two concepts were proposed to enhance jet mixing in a circular cross-section: the Asymmetric Jet Penetration (AJP) concept; and the Lobed Mixer (LM) concept. In Phase 1, two preliminary configurations of the AJP concept were compared with a conventional 12-jet radial-inflow slot design. The configurations were screened using an advanced 3-D Computational Fluid Dynamics (CFD) code named REFLEQS. Both non-reacting and reacting analyses were performed. For an objective comparison, the conventional design was optimized by parametric variation of the jet-to-mainstream momentum flux (J) ratio. The optimum J was then employed in the AJP simulations. Results showed that the three-jet AJP configuration was superior in overall mixedness compared to the conventional design. However, in regards to NOx emissions, the AJP configuration was inferior. The higher emission level for AJP was caused by a single hot spot located in the wake of the central jet as it entered the combustor. Ways of maintaining good mixedness while eliminating the hot spot were identified for Phase 2 study. Overall, Phase 1 showed the viability of using CFD analyses to evaluate quick-mix concepts. A high probability exists that advancing mixing concepts will reduce NOx emissions in RQL combustors, and should be explored in Phase 2, by parallel numerical and experimental work.

Compliant Metal Enhanced Convection Cooled Reverse-Flow Annular Combustor

NASA Technical Reports Server (NTRS)

Paskin, Marc D.; Acosta, Waldo A.

1994-01-01

A joint Army/NASA program was conducted to design, fabricate, and test an advanced, reverse-flow, small gas turbine combustor using a compliant metal enhanced (CME) convection wall cooling concept. The objectives of this effort were to develop a design method (basic design data base and analysis) for the CME cooling technique and tben demonstrate its application to an advanced cycle, small, reverse-flow combustor with 3000 F (1922 K) burner outlet temperature (BOT). The CME concept offers significant improvements in wall cooling effectiveness resulting in a large reduction in cooling air requirements. Therefore, more air is available for control of burner outlet temperature pattern in addition to the benefit of improved efficiency, reduced emissions, and smoke levels. Rig test results demonstrated the benefits and viability of the CME concept meeting or exceeding the aerothermal performance and liner wall temperature characteristics of similar lower temperature-rise combustors, achieving 0.15 pattern factor at 3000 F (1922 K) BOT, while utilizing approximately 80 percent less cooling air than conventional, film-cooled combustion systems.

LDV measurements in an annular combustor model. M.S. Thesis

NASA Technical Reports Server (NTRS)

Barron, Dean A.

1986-01-01

The design and setup of a Laser Doppler Velocimeter (LDV) system used to take velocity measurements in an annular combustor model are covered. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. The LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

LDV Measurements in an Annular Combustor Model

NASA Technical Reports Server (NTRS)

Barron, Dean A.

1996-01-01

This thesis covers the design and setup of a laser doppler velocimeter (LDV) system used to take velocity measurements in an annular combustor model. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.

Lean, Premixed-Prevaporized (LPP) combustor conceptual design study

NASA Technical Reports Server (NTRS)

Dickman, R. A.; Dodds, W. J.; Ekstedt, E. E.

1979-01-01

Four combustion systems were designed and sized for the energy efficient engine. A fifth combustor was designed for the cycle and envelope of the twin-spool, high bypass ratio, high pressure ratio turbofan engine. Emission levels, combustion performance, life, and reliability assessments were made for these five combustion systems. Results of these design studies indicate that cruise NOx emission can be reduced by the use of lean, premixed-prevaporaized combustion and airflow modulation.

Linear test bed. Volume 1: Test bed no. 1. [aerospike test bed with segmented combustor

NASA Technical Reports Server (NTRS)

1972-01-01

The Linear Test Bed program was to design, fabricate, and evaluation test an advanced aerospike test bed which employed the segmented combustor concept. The system is designated as a linear aerospike system and consists of a thrust chamber assembly, a power package, and a thrust frame. It was designed as an experimental system to demonstrate the feasibility of the linear aerospike-segmented combustor concept. The overall dimensions are 120 inches long by 120 inches wide by 96 inches in height. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure, at a mixture ratio of 5.5. At the design conditions, the sea level thrust is 200,000 pounds. The complete program including concept selection, design, fabrication, component test, system test, supporting analysis and posttest hardware inspection is described.

NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts

NASA Technical Reports Server (NTRS)

Manthey, Lri

2001-01-01

Topics discussed include: UEET Overview; Technology Benefits; Emissions Overview; P&W Low Emissions Combustor Development; GE Low Emissions Combustor Development; Rolls-Royce Low Emissions Combustor Development; Honeywell Low Emissions Combustor Development; NASA Multipoint LDI Development; Stanford Activities In Concepts for Advanced Gas Turbine Combustors; Large Eddy Simulation (LES) of Gas Turbine Combustion; NASA National Combustion Code Simulations; Materials Overview; Thermal Barrier Coatings for Airfoil Applications; Disk Alloy Development; Turbine Blade Alloy; Ceramic Matrix Composite (CMC) Materials Development; Ceramic Matrix Composite (CMC) Materials Characterization; Environmental Barrier Coatings (EBC) for Ceramic Matrix Composite (CMC) Materials; Ceramic Matrix Composite Vane Rig Testing and Design; Ultra-High Temperature Ceramic (UHTC) Development; Lightweight Structures; NPARC Alliance; Technology Transfer and Commercialization; and Turbomachinery Overview; etc.

Multi-Dimensional Measurements of Combustion Species in Flame Tube and Sector Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Hicks, Yolanda Royce

1996-01-01

The higher temperature and pressure cycles of future aviation gas turbine combustors challenge designers to produce combustors that minimize their environmental impact while maintaining high operation efficiency. The development of low emissions combustors includes the reduction of unburned hydrocarbons, smoke, and particulates, as well as the reduction of oxides of nitrogen (NO(x)). In order to better understand and control the mechanisms that produce emissions, tools are needed to aid the development of combustor hardware. Current methods of measuring species within gas turbine combustors use extractive sampling of combustion gases to determine major species concentrations and to infer the bulk flame temperature. These methods cannot be used to measure unstable combustion products and have poor spatial and temporal resolution. The intrusive nature of gas sampling may also disturb the flow structure within a combustor. Planar laser-induced fluorescence (PLIF) is an optical technique for the measurement of combustion species. In addition to its non-intrusive nature, PLIF offers these advantages over gas sampling: high spatial resolution, high temporal resolution, the ability to measure unstable species, and the potential to measure combustion temperature. This thesis considers PLIF for in-situ visualization of combustion species as a tool for the design and evaluation of gas turbine combustor subcomponents. This work constitutes the first application of PLIF to the severe environment found in liquid-fueled, aviation gas turbine combustors. Technical and applied challenges are discussed. PLIF of OH was used to observe the flame structure within the post flame zone of a flame tube combustor, and within the flame zone of a sector combustor, for a variety of fuel injector configurations. OH was selected for measurement because it is a major combustion intermediate, playing a key role in the chemistry of combustion, and because its presence within the flame zone can serve as a qualitative marker of flame temperature. All images were taken in the environment of actual engines during flight, using actual jet fuel. The results of the PLIF study led directly to the modification of a fuel injector.

Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2015-01-01

Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at highpressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NOx emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8%.

Parametric Study of Pulse-Combustor-Driven Ejectors at High-Pressure

NASA Technical Reports Server (NTRS)

Yungster, Shaye; Paxson, Daniel E.; Perkins, Hugh D.

2015-01-01

Pulse-combustor configurations developed in recent studies have demonstrated performance levels at high-pressure operating conditions comparable to those observed at atmospheric conditions. However, problems related to the way fuel was being distributed within the pulse combustor were still limiting performance. In the first part of this study, new configurations are investigated computationally aimed at improving the fuel distribution and performance of the pulse-combustor. Subsequent sections investigate the performance of various pulse-combustor driven ejector configurations operating at high pressure conditions, focusing on the effects of fuel equivalence ratio and ejector throat area. The goal is to design pulse-combustor-ejector configurations that maximize pressure gain while achieving a thermal environment acceptable to a turbine, and at the same time maintain acceptable levels of NO(x) emissions and flow non-uniformities. The computations presented here have demonstrated pressure gains of up to 2.8.

Parametric test results of a swirl-can combustor

NASA Technical Reports Server (NTRS)

Niedzwiecki, R. W.; Jones, R. E.

1973-01-01

Pollutant levels of oxides of nitrogen, unburned hydrocarbons, and carbon monoxide were measured for three models of an experimental, annular swirl can combustor. The combustor was 1.067 meters in outer diameter, incorporated 120 modules, and was specifically designed for elevated exit temperature performance. Test conditions included combustor inlet temperatures of 589, 756 and 839 K, inlet pressures of 3 to 6.4 atmospheres, reference velocities of 21 to 38 meters per second and combustor equivalence ratios, based on total combustor flows of 0.206 to 1.028. Maximum oxides of nitrogen emission index values occurred at an equivalence ratio of 0.7 with lower values measured for both higher and lower equivalence ratios. Oxides of nitrogen concentrations, to the 0.7 level with 756 K inlet air, were correlated for the three models by a combined parameter consisting of measured flow and geometric parameters. Effects of the individual parameters comprising the correlation are also presented.

Advanced technology for reducing aircraft engine pollution

NASA Technical Reports Server (NTRS)

Jones, R. E.

1973-01-01

The proposed EPA regulations covering emissions of gas turbine engines will require extensive combustor development. The NASA is working to develop technology to meet these goals through a wide variety of combustor research programs conducted in-house, by contract, and by university grant. In-house efforts using the swirl-can modular combustor have demonstrated sizable reduction in NO emission levels. Testing to reduce idle pollutants has included the modification of duplex fuel nozzles to air-assisted nozzles and an exploration of the potential improvements possible with combustors using fuel staging and variable geometry. The Experimental Clean Combustor Program, a large contracted effort, is devoted to the testing and development of combustor concepts designed to achieve a large reduction in the levels of all emissions. This effort is planned to be conducted in three phases with the final phase to be an engine demonstration of the best reduced emission concepts.

Energy efficient engine sector combustor rig test program

NASA Technical Reports Server (NTRS)

Dubiel, D. J.; Greene, W.; Sundt, C. V.; Tanrikut, S.; Zeisser, M. H.

1981-01-01

Under the NASA-sponsored Energy Efficient Engine program, Pratt & Whitney Aircraft has successfully completed a comprehensive combustor rig test using a 90-degree sector of an advanced two-stage combustor with a segmented liner. Initial testing utilized a combustor with a conventional louvered liner and demonstrated that the Energy Efficient Engine two-stage combustor configuration is a viable system for controlling exhaust emissions, with the capability to meet all aerothermal performance goals. Goals for both carbon monoxide and unburned hydrocarbons were surpassed and the goal for oxides of nitrogen was closely approached. In another series of tests, an advanced segmented liner configuration with a unique counter-parallel FINWALL cooling system was evaluated at engine sea level takeoff pressure and temperature levels. These tests verified the structural integrity of this liner design. Overall, the results from the program have provided a high level of confidence to proceed with the scheduled Combustor Component Rig Test Program.

Linear aerospike engine study. [for reusable launch vehicles

NASA Technical Reports Server (NTRS)

Diem, H. G.; Kirby, F. M.

1977-01-01

Parametric data on split-combustor linear engine propulsion systems are presented for use in mixed-mode single-stage-to-orbit (SSTO) vehicle studies. Preliminary design data for two selected engine systems are included. The split combustor was investigated for mixed-mode operations with oxygen/hydrogen propellants used in the inner combustor in Mode 2, and in conjunction with either oxygen/RP-1, oxygen/RJ-5, O2/CH4, or O2/H2 propellants in the outer combustor for Mode 1. Both gas generator and staged combustion power cycles were analyzed for providing power to the turbopumps of the inner and outer combustors. Numerous cooling circuits and cooling fluids (propellants) were analyzed and hydrogen was selected as the preferred coolant for both combustors and the linear aerospike nozzle. The maximum operating chamber pressure was determined to be limited by the availability of hydrogen coolant pressure drop in the coolant circuit.

The large-amplitude combustion oscillation in a single-side expansion scramjet combustor

NASA Astrophysics Data System (ADS)

Ouyang, Hao; Liu, Weidong; Sun, Mingbo

2015-12-01

The combustion oscillation in scramjet combustor is believed not existing and ignored for a long time. Compared with the flame pulsation, the large-amplitude combustion oscillation in scramjet combustor is indeed unfamiliar and difficult to be observed. In this study, the specifically designed experiments are carried out to investigate this unusual phenomenon in a single-side expansion scramjet combustor. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The obtained results show that the large-amplitude combustion oscillation can exist in scramjet combustor, which is not occasional and can be reproduced. Under the given conditions of this study, moreover, the large-amplitude combustion oscillation is regular and periodic, whose principal frequency is about 126 Hz. The proceeding of the combustion oscillation is accompanied by the transformation of the flame-holding pattern and combustion mode transition between scramjet mode combustion and ramjet mode combustion.

Analytical and experimental evaluations of the effect of broad property fuels on combustors for commercial aircraft gas turbine engines

NASA Technical Reports Server (NTRS)

Smith, A. L.

1980-01-01

The impacts of broad property fuels on the design, performance, durability, emissions, and operational characteristics of current and advanced combustors for commercial aircraft gas turbine engines were studied. The effect of fuel thermal stability on engine and airframe fuel system was evaluated. Tradeoffs between fuel properties, exhaust emissions, and combustor life were also investigated. Results indicate major impacts of broad property fuels on allowable metal temperatures in fuel manifolds and injector support, combustor cyclic durability, and somewhat lesser impacts on starting characteristics, lightoff, emissions, and smoke.

Fuel and Combustor Concerns for Future Commercial Combustors

NASA Technical Reports Server (NTRS)

Chang, Clarence T.

2017-01-01

Civil aircraft combustor designs will move from rich-burn to lean-burn due to the latter's advantage in low NOx and nvPM emissions. However, the operating range of lean-burn is narrower, requiring premium mixing performance from the fuel injectors. As the OPR increases, the corresponding combustor inlet temperature increase can benefit greatly with fuel composition improvements. Hydro-treatment can improve coking resistance, allowing finer fuel injection orifices to speed up mixing. Selective cetane number control across the fuel carbon-number distribution may allow delayed ignition at high power while maintaining low-power ignition characteristics.

Initiation Mechanisms of Low-loss Swept-ramp Obstacles for Deflagration to Detonation Transition in Pulse Detonation Combustors

DTIC Science & Technology

2009-12-01

minimal pressure losses. 15. NUMBER OF PAGES 113 14. SUBJECT TERMS Pulse Detonation Combustors, PDC, Pulse Detonation Engines, PDE , PDE ...Postgraduate School PDC Pulse Detonation Combustor PDE Pulse Detonation Engine RAM Random Access Memory RDT Research, Design and Test RPL...inhibiting the implementation of this advanced propulsion system. The primary advantage offered by pulse detonation engines ( PDEs ) is the high efficiency

Applied analytical combustion/emissions research at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Deur, J. M.; Kundu, K. P.; Nguyen, H. L.

1992-01-01

Emissions of pollutants from future commercial transports are a significant concern. As a result, the Lewis Research Center (LeRC) is investigating various low emission combustor technologies. As part of this effort, a combustor analysis code development program was pursued to guide the combustor design process, to identify concepts having the greatest promise, and to optimize them at the lowest cost in the minimum time.

Applied Analytical Combustion/emissions Research at the NASA Lewis Research Center - a Progress Report

NASA Technical Reports Server (NTRS)

Deur, J. M.; Kundu, K. P.; Nguyen, H. L.

1992-01-01

Emissions of pollutants from future commercial transports are a significant concern. As a result, the Lewis Research Center (LeRC) is investigating various low emission combustor technologies. As part of this effort, a combustor analysis code development program was pursued to guide the combustor design process, to identify concepts having the greatest promise, and to optimize them at the lowest cost in the minimum time.

Combustion oscillation: Chemical control showing mechanistic link to recirculation zone purge time

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gemmen, R.S.; Richards, G.A.; Yip, M.J.

1995-12-01

Active control mechanisms are being examined for lean premix combustion applications, such as gas turbine generators. Lean premix combustors are susceptible to large combustion oscillations, particularly when driven very lean to achieve low NOx. While past design work has been focussed on understanding the source of the oscillation and modifying the combustor to avoid such oscillations, commercial combustion designers have more recently considered applying new control elements. As part of the U.S. Department of Energy`s Advanced Gas Turbine Systems Program, the Morgantown Energy Technology Center is investigating various active control techniques. This paper presents results from experiments studying the effectmore » of pilot fuel modulation on combustor oscillation and pollutant emissions for a pilot stabilized dump swirl combustor, typical of gas turbine combustors. The results show that a significant level of attenuation can be achieved in the combustor pressure oscillation (50 to 90 percent) while only moderately affecting pollutant emissions. The control mechanism producing the attenuation is shown to be purely chemical in nature, rather than fluid mechanic. In addition, the frequency region over which control is obtained is shown to be related to the recirculation zone purge time. For this reason, control can be achieved at control frequencies much lower than the frequency of oscillation.« less

Program user's manual for optimizing the design of a liquid or gaseous propellant rocket engine with the automated combustor design code AUTOCOM

NASA Technical Reports Server (NTRS)

Reichel, R. H.; Hague, D. S.; Jones, R. T.; Glatt, C. R.

1973-01-01

This computer program manual describes in two parts the automated combustor design optimization code AUTOCOM. The program code is written in the FORTRAN 4 language. The input data setup and the program outputs are described, and a sample engine case is discussed. The program structure and programming techniques are also described, along with AUTOCOM program analysis.

National Combustion Code, a Multidisciplinary Combustor Design System, Will Be Transferred to the Commercial Sector

NASA Technical Reports Server (NTRS)

Steele, Gynelle C.

1999-01-01

The NASA Lewis Research Center and Flow Parametrics will enter into an agreement to commercialize the National Combustion Code (NCC). This multidisciplinary combustor design system utilizes computer-aided design (CAD) tools for geometry creation, advanced mesh generators for creating solid model representations, a common framework for fluid flow and structural analyses, modern postprocessing tools, and parallel processing. This integrated system can facilitate and enhance various phases of the design and analysis process.

HSCT Sector Combustor Hardware Modifications for Improved Combustor Design

NASA Technical Reports Server (NTRS)

Greenfield, Stuart C.; Heberling, Paul V.; Moertle, George E.

2005-01-01

An alternative to the stepped-dome design for the lean premixed prevaporized (LPP) combustor has been developed. The new design uses the same premixer types as the stepped-dome design: integrated mixer flameholder (IMFH) tubes and a cyclone swirler pilot. The IMFH fuel system has been taken to a new level of development. Although the IMFH fuel system design developed in this Task is not intended to be engine-like hardware, it does have certain characteristics of engine hardware, including separate fuel circuits for each of the fuel stages. The four main stage fuel circuits are integrated into a single system which can be withdrawn from the combustor as a unit. Additionally, two new types of liner cooling have been designed. The resulting lean blowout data was found to correlate well with the Lefebvre parameter. As expected, CO and unburned hydrocarbons emissions were shown to have an approximately linear relationship, even though some scatter was present in the data, and the CO versus flame temperature data showed the typical cupped shape. Finally, the NOx emissions data was shown to agree well with a previously developed correlation based on emissions data from Configuration 3 tests performed at GEAE. The design variations of the cyclone swirler pilot that were investigated in this study did not significantly change the NOx emissions from the baseline design (GEAE Configuration 3) at supersonic cruise conditions.

User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 1: User's manual

NASA Technical Reports Server (NTRS)

Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

1991-01-01

The user's manual for the rocket combustor interactive design (ROCCID) computer program is presented. The program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial, and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can easily be added. The analysis model in ROCCID can account for the influence of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

Core Noise: Overview of Upcoming LDI Combustor Test

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.

2012-01-01

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.

Engineering analyses and design calculations of NASA, Langley Research Center hydrogen-air-vitiated heater with oxygen replenishment

NASA Technical Reports Server (NTRS)

1973-01-01

The technical basis is presented for the design of the hydrogen-air-vitiated heater. The heater liner is subjected to a maximum thermal environment at a specified condition, where the combustion gas temperature, pressure and flow rate are 5000 F, 750 psia, and 11.0 lb/sec, respectively, and results in a heat flux of the order of 275 BTU/sec-sq ft. Cooling and stress analyses indicate that water is the logical choice for cooling of the combustor liner. A mixing analysis was undertaken to establish a good combination of combustor length and injector configuration. The analysis, using a conservative analytical approach, indicates a combustor length of the order of 5 ft combined with discrete fuel and oxidizer injection at an approximate 2-1/2 inch radial combustor position, and results in uniform combustion products at the heater exit for all specified envelope conditions.

Turbulent Recirculating Flows in Isothermal Combustor Geometries

NASA Technical Reports Server (NTRS)

Lilley, D.; Rhode, D.

1985-01-01

Computer program developed that provides mathematical solution to design and construction of combustion chambers for jet engines. Improved results in areas of combustor flow fields accomplished by this computerprogram solution, cheaper and quicker than experiments involving real systems for models.

Emissions of nitrogen oxides from an experimental hydrogen-fueled gas turbine combustor

NASA Technical Reports Server (NTRS)

Norgren, C. T.; Ingebo, R. D.

1974-01-01

The effect of operating variables of a hydrogen fueled combustor on exhaust concentrations of total oxides of nitrogen was determined at inlet-air temperature levels up to 810 K, pressure of 414,000N/sa m, and reference velocity of 21.3 m/sec. The combustor, which was originally designed for hydrocarbon fuel produced a NO(x) concentration of 380 ppm with hydrogen at 810 K inlet-air temperature. A reduction in NO(x) of about 30 % was obtained by modification to a lean or rich primary zone. The lowest NO(x) levels obtained with hydrogen were equivalent to those of the reference combustor burning hydrocarbon fuels.

Planar measurement of flow field parameters in a nonreacting supersonic combustor using laser-induced iodine fluorescence

NASA Technical Reports Server (NTRS)

Hartfield, Roy J., Jr.; Hollo, Steven D.; Mcdaniel, James C.

1990-01-01

A nonintrusive optical technique, laser-induced iodine fluorescence, has been used to obtain planar measurements of flow field parameters in the supersonic mixing flow field of a nonreacting supersonic combustor. The combustor design used in this work was configured with staged transverse sonic injection behind a rearward-facing step into a Mach 2.07 free stream. A set of spatially resolved measurements of temperature and injectant mole fraction has been generated. These measurements provide an extensive and accurate experimental data set required for the validation of computational fluid dynamic codes developed for the calculation of highly three-dimensional combustor flow fields.

Estimating the uncertainty in thermochemical calculations for oxygen-hydrogen combustors

NASA Astrophysics Data System (ADS)

Sims, Joseph David

The thermochemistry program CEA2 was combined with the statistical thermodynamics program PAC99 in a Monte Carlo simulation to determine the uncertainty in several CEA2 output variables due to uncertainty in thermodynamic reference values for the reactant and combustion species. In all, six typical performance parameters were examined, along with the required intermediate calculations (five gas properties and eight stoichiometric coefficients), for three hydrogen-oxygen combustors: a main combustor, an oxidizer preburner and a fuel preburner. The three combustors were analyzed in two different modes: design mode, where, for the first time, the uncertainty in thermodynamic reference values---taken from the literature---was considered (inputs to CEA2 were specified and so had no uncertainty); and data reduction mode, where inputs to CEA2 did have uncertainty. The inputs to CEA2 were contrived experimental measurements that were intended to represent the typical combustor testing facility. In design mode, uncertainties in the performance parameters were on the order of 0.1% for the main combustor, on the order of 0.05% for the oxidizer preburner and on the order of 0.01% for the fuel preburner. Thermodynamic reference values for H2O were the dominant sources of uncertainty, as was the assigned enthalpy for liquid oxygen. In data reduction mode, uncertainties in performance parameters increased significantly as a result of the uncertainties in experimental measurements compared to uncertainties in thermodynamic reference values. Main combustor and fuel preburner theoretical performance values had uncertainties of about 0.5%, while the oxidizer preburner had nearly 2%. Associated experimentally-determined performance values for all three combustors were 3% to 4%. The dominant sources of uncertainty in this mode were the propellant flowrates. These results only apply to hydrogen-oxygen combustors and should not be generalized to every propellant combination. Species for a hydrogen-oxygen system are relatively simple, thereby resulting in low thermodynamic reference value uncertainties. Hydrocarbon combustors, solid rocket motors and hybrid rocket motors have combustion gases containing complex molecules that will likely have thermodynamic reference values with large uncertainties. Thus, every chemical system should be analyzed in a similar manner as that shown in this work.

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

NASA Technical Reports Server (NTRS)

Clements, T. R.

1973-01-01

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.

A CFD Study of Jet Mixing in Reduced Flow Areas for Lower Combustor Emissions

NASA Technical Reports Server (NTRS)

Smith, C. E.; Talpallikar, M. V.; Holdeman, J. D.

1991-01-01

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor has the potential of significantly reducing NO(x) emissions in combustion chambers of High Speed Civil Transport aircraft. Previous work on RQL combustors for industrial applications suggested the benefit of necking down the mixing section. A 3-D numerical investigation was performed to study the effects of neckdown on NO(x) emissions and to develop a correlation for optimum mixing designs in terms of neckdown area ratio. The results of the study showed that jet mixing in reduced flow areas does not enhance mixing, but does decrease residence time at high flame temperatures, thus reducing NO(x) formation. By necking down the mixing flow area by 4, a potential NO(x) reduction of 16:1 is possible for annual combustors. However, there is a penalty that accompanies the mixing neckdown: reduced pressure drop across the combustor swirler. At conventional combustor loading parameters, the pressure drop penalty does not appear to be excessive.

Pollution reduction technology program for turboprop engines

NASA Technical Reports Server (NTRS)

Tomlinson, J. G.

1977-01-01

The reduction of CO, HC, and smoke emissions while maintaining acceptable NO(x) emissions without affecting fuel consumption, durability, maintainability, and safety was accomplished. Component combustor concept screening directed toward the demonstration of advanced combustor technology required to meet the EPA exhaust emissions standards for class P2 turboprop engines was covered. The combustion system for the Allison 501-D22A engine was used, and three combustor design concepts - reverse flow, prechamber, and staged fuel were evaluated.

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/Lean Combustor for High Speed Civil Transport Aircraft Engines

NASA Technical Reports Server (NTRS)

Rosfjord, T. J.; Padget, F. C.; Tacina, Robert R. (Technical Monitor)

2001-01-01

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of a HSCT engine cycle, Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NOx. The spatial profiles of NOx and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NOx. Based on this study, it was also concluded that (1) While NOx formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NOx exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) A RQL combustor can achieve the emissions goal of EINOX = 5 at the Supersonic Cruise operating condition for a HSCT engine.

Experimental Assessment of the Emissions Control Potential of a Rich/Quench/ Lean Combustor for High Speed Civil Transport Aircraft Engines

NASA Technical Reports Server (NTRS)

Tacina, Robert R. (Technical Monitor); Rosfjord, T. J.; Padget, F. C.

2001-01-01

In support of Pratt & Whitney efforts to define the Rich burn/Quick mix/Lean burn (RQL) combustor for the High Speed Civil Transport (HSCT) aircraft engine, UTRC conducted a flametube-scale study of the RQL concept. Extensive combustor testing was performed at the Supersonic Cruise (SSC) condition of an HSCT engine cycle. Data obtained from probe traverses near the exit of the mixing section confirmed that the mixing section was the critical component in controlling combustor emissions. Circular-hole configurations, which produced rapidly-, highly-penetrating jets, were most effective in limiting NO(x). The spatial profiles of NO(x) and CO at the mixer exit were not directly interpretable using a simple flow model based on jet penetration, and a greater understanding of the flow and chemical processes in this section are required to optimize it. Neither the rich-combustor equivalence ratio nor its residence time was a direct contributor to the exit NO(x). Based on this study, it was also concluded that: (1) While NO(x) formation in both the mixing section and the lean combustor contribute to the overall emission, the NOx formation in the mixing section dominates. The gas composition exiting the rich combustor can be reasonably represented by the equilibrium composition corresponding to the rich combustor operating condition. Negligible NO(x) exits the rich combustor. (2) At the SSC condition, the oxidation processes occurring in the mixing section consume 99 percent of the CO exiting the rich combustor. Soot formed in the rich combustor is also highly oxidized, with combustor exit SAE Smoke Number <3. (3) Mixing section configurations which demonstrated enhanced emissions control at SSC also performed better at part-power conditions. Data from mixer exit traverses reflected the expected mixing behavior for off-design jet to crossflow momentum-flux ratios. (4) Low power operating conditions require that the RQL combustor operate as a lean-lean combustor to achieve low CO and high efficiency. (5) An RQL combustor can achieve the emissions goal of EINO(x) = 5 at the Supersonic Cruise operating condition for an HSCT engine.

Aircraft gas turbine low-power emissions reduction technology program

NASA Technical Reports Server (NTRS)

Dodds, W. J.; Gleason, C. C.; Bahr, D. W.

1978-01-01

Advanced aircraft turbine engine combustor technology was used to reduce low-power emissions of carbon monoxide and unburned hydrocarbons to levels significantly lower than those which were achieved with current technology. Three combustor design concepts, which were designated as the hot-wall liner concept, the recuperative-cooled liner concept, and the catalyst converter concept, were evaluated in a series of CF6-50 engine size 40 degree-sector combustor rig tests. Twenty-one configurations were tested at operating conditions spanning the design condition which was an inlet temperature and pressure of 422 K and 304 kPa, a reference velocity of 23 m/s and a fuel-air-ration of 10.5 g/kg. At the design condition typical of aircraft turbine engine ground idle operation, the best configurations of all three concepts met the stringent emission goals which were 10, 1, and 4 g/kg for CO, HC, and Nox, respectively.

Forebody and Inlet Design for the HIFiRE 2 Flight Test

NASA Technical Reports Server (NTRS)

Ferlemann, Paul G.

2008-01-01

A forebody and inlet have been designed for the HIFiRE 2 scramjet flight test. The test will explore the operating, performance, and stability characteristics of a simple hydrocarbon-fueled scramjet combustor as it transitions from dual-mode to scramjet-mode operation and during supersonic combustion at Mach 8+ flight conditions. Requirements for the compression system were derived from inlet starting and combustor inflow requirements as well as physical size constraints. The design process is described. A planar, fixed geometry, mixed compression concept was used to produce laterally uniform flow at the inlet entrance and a conservative amount of internal contraction with respect to inlet starting. A grid sensitivity study was performed so that important flow physics caused by three-dimensional shock boundary layer interactions could be captured with confidence. Results from low Mach number operability studies, nominal trajectory cases, and high dynamic pressure heat load cases are discussed. The forebody and inlet solutions provide information for on-going combustor calculations, mass capture across the trajectory for fuel system design, and surface heating rates for thermal/structural analysis. The design has a one freestream Mach number margin for inlet starting, exceeds the high Mach number combustor entrance pressure requirement, produces high quality flow at the inlet exit for all Mach numbers and vehicle attitudes in the design space, and fits inside the booster shroud.

User's manual for rocket combustor interactive design (ROCCID) and analysis computer program. Volume 2: Appendixes A-K

NASA Technical Reports Server (NTRS)

Muss, J. A.; Nguyen, T. V.; Johnson, C. W.

1991-01-01

The appendices A-K to the user's manual for the rocket combustor interactive design (ROCCID) computer program are presented. This includes installation instructions, flow charts, subroutine model documentation, and sample output files. The ROCCID program, written in Fortran 77, provides a standardized methodology using state of the art codes and procedures for the analysis of a liquid rocket engine combustor's steady state combustion performance and combustion stability. The ROCCID is currently capable of analyzing mixed element injector patterns containing impinging like doublet or unlike triplet, showerhead, shear coaxial and swirl coaxial elements as long as only one element type exists in each injector core, baffle, or barrier zone. Real propellant properties of oxygen, hydrogen, methane, propane, and RP-1 are included in ROCCID. The properties of other propellants can be easily added. The analysis models in ROCCID can account for the influences of acoustic cavities, helmholtz resonators, and radial thrust chamber baffles on combustion stability. ROCCID also contains the logic to interactively create a combustor design which meets input performance and stability goals. A preliminary design results from the application of historical correlations to the input design requirements. The steady state performance and combustion stability of this design is evaluated using the analysis models, and ROCCID guides the user as to the design changes required to satisfy the user's performance and stability goals, including the design of stability aids. Output from ROCCID includes a formatted input file for the standardized JANNAF engine performance prediction procedure.

MUNICIPAL SOLID WASTE COMBUSTOR ASH DEMONSTRATION PROGRAM - "THE BOATHOUSE"

EPA Science Inventory

The report presents the results of a research program designed to examine the engineering and environmental acceptability of using municipal solid waste (MSW) combustor ash as an aggregate substitute in the manufacture of construction quality cement blocks. 50 tons of MSW combust...

Low NOx heavy fuel combustor concept program

NASA Technical Reports Server (NTRS)

White, D. J.; Lecren, R. T.; Batakis, A. P.

1981-01-01

A total of twelve low NOx combustor configurations, embodying three different combustion concepts, were designed and fabricated as modular units. These configurations were evaluated experimentally for exhaust emission levels and for mechanical integrity. Emissions data were obtained in depth on two of the configurations.

Performance of a Model Rich Burn-quick Mix-lean Burn Combustor at Elevated Temperature and Pressure

NASA Technical Reports Server (NTRS)

Peterson, Christopher O.; Sowa, William A.; Samuelsen, G. S.

2002-01-01

As interest in pollutant emission from stationary and aero-engine gas turbines increases, combustor engineers must consider various configurations. One configuration of increasing interest is the staged, rich burn - quick mix - lean burn (RQL) combustor. This report summarizes an investigation conducted in a recently developed high pressure gas turbine combustor facility. The model RQL combustor was plenum fed and modular in design. The fuel used for this study is Jet-A which was injected from a simplex atomizer. Emission (CO2, CO, O2, UHC, NOx) measurements were obtained using a stationary exit plane water-cooled probe and a traversing water-cooled probe which sampled from the rich zone exit and the lean zone entrance. The RQL combustor was operated at inlet temperatures ranging from 367 to 700 K, pressures ranging from 200 to 1000 kPa, and combustor reference velocities ranging from 10 to 20 m/s. Variations were also made in the rich zone and lean zone equivalence ratios. Several significant trends were observed. NOx production increased with reaction temperature, lean zone equivalence ratio and residence time and decreased with increased rich zone equivalence ratio. NOx production in the model RQL combustor increased to the 0.4 power with increased pressure. This correlation, compared to those obtained for non-staged combustors (0.5 to 0.7), suggests a reduced dependence on NOx on pressure for staged combustors. Emissions profiles suggest that rich zone mixing is not uniform and that the rich zone contributes on the order of 16 percent to the total NOx produced.

Small, modular, low-cost coal-fired power plants for the international market

DOE Office of Scientific and Technical Information (OSTI.GOV)

Zauderer, B.; Frain, B.; Borck, B.

1997-12-31

This paper presents recent operating results of Coal Tech`s second generation, air cooled, slagging coal combustor, and its application to power plants in the 1 to 20 MW range. This 20 MMBtu/hour combustor was installed in a new demonstration plant in Philadelphia, PA in 1995. It contains the combustion components of a 1 MWe coal fired power plant, a 17,500 lb/hour steam boiler, coal storage and feed components, and stack gas cleanup components. The plant`s design incorporates improvements resulting from 2,000 hours of testing between 1987 and 1993 on a first generation, commercial scale, air cooled combustor of equal thermalmore » rating. Since operations began in early 1996, a total of 51 days of testing have been successfully completed. Major results include durability of the combustor`s refractory wall, excellent combustion with high ash concentration in the fuel, removal of 95% to 100% of the slag in the combustor, very little ash deposition in the boiler, major reduction of in-plant parasitic power, and simplified power system control through the use of modular designs of sub-systems and computer control. Rapid fuel switching between oil, gas, and coal and turndown of up to a factor of three was accomplished. All these features have been incorporated in advanced coal fired plant designs in the 1 to 20 MWe range. Incremental capital costs are only $100 to $200/kW higher than comparable rated gas or oil fired steam generating systems. Most of its components and subsystems can be factory assembled for very rapid field installation. The low capital, low operating costs, fuel flexibility, and compatibility with very high ash fuels, make this power system very attractive in regions of the world having domestic supplies of these fuels.« less

Pollution technology program, can-annular combustor engines

NASA Technical Reports Server (NTRS)

Roberts, R.; Fiorentino, A. J.; Greene, W.

1976-01-01

A Pollution Reduction Technology Program to develop and demonstrate the combustor technology necessary to reduce exhaust emissions for aircraft engines using can-annular combustors is described. The program consisted of design, fabrication, experimental rig testing and assessment of results and was conducted in three program elements. The combustor configurations of each program element represented increasing potential for meeting the 1979 Environmental Protection Agency (EPA) emission standards, while also representing increasing complexity and difficulty of development and adaptation to an operational engine. Experimental test rig results indicate that significant reductions were made to the emission levels of the baseline JT8D-17 combustor by concepts in all three program elements. One of the Element I single-stage combustors reduced carbon monoxide to a level near, and total unburned hydrocarbons (THC) and smoke to levels below the 1979 EPA standards with little or no improvement in oxides of nitrogen. The Element II two-stage advanced Vorbix (vortex burning and mixing) concept met the standard for THC and achieved significant reductions in CO and NOx relative to the baseline. Although the Element III prevaporized-premixed concept reduced high power NOx below the Element II results, there was no improvement to the integrated EPA parameter relative to the Vorbix combustor.

Design Considerations of ISTAR Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2003-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system that produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Design Considerations of Istar Hydrocarbon Fueled Combustor Operating in Air Augmented Rocket, Ramjet and Scramjet Modes

NASA Technical Reports Server (NTRS)

Andreadis, Dean; Drake, Alan; Garrett, Joseph L.; Gettinger, Christopher D.; Hoxie, Stephen S.

2002-01-01

The development and ground test of a rocket-based combined cycle (RBCC) propulsion system is being conducted as part of the NASA Marshall Space Flight Center (MSFC) Integrated System Test of an Airbreathing Rocket (ISTAR) program. The eventual flight vehicle (X-43B) is designed to support an air-launched self-powered Mach 0.7 to 7.0 demonstration of an RBCC engine through all of its airbreathing propulsion modes - air augmented rocket (AAR), ramjet (RJ), and scramjet (SJ). Through the use of analytical tools, numerical simulations, and experimental tests the ISTAR program is developing and validating a hydrocarbon-fueled RBCC combustor design methodology. This methodology will then be used to design an integrated RBCC propulsion system thai: produces robust ignition and combustion stability characteristics while maximizing combustion efficiency and minimizing drag losses. First order analytical and numerical methods used to design hydrocarbon-fueled combustors are discussed with emphasis on the methods and determination of requirements necessary to establish engine operability and performance characteristics.

Cars Thermometry in a Supersonic Combustor for CFD Code Validation

NASA Technical Reports Server (NTRS)

Cutler, A. D.; Danehy, P. M.; Springer, R. R.; DeLoach, R.; Capriotti, D. P.

2002-01-01

An experiment has been conducted to acquire data for the validation of computational fluid dynamics (CFD) codes used in the design of supersonic combustors. The primary measurement technique is coherent anti-Stokes Raman spectroscopy (CARS), although surface pressures and temperatures have also been acquired. Modern- design- of-experiment techniques have been used to maximize the quality of the data set (for the given level of effort) and minimize systematic errors. The combustor consists of a diverging duct with single downstream- angled wall injector. Nominal entrance Mach number is 2 and enthalpy nominally corresponds to Mach 7 flight. Temperature maps are obtained at several planes in the flow for two cases: in one case the combustor is piloted by injecting fuel upstream of the main injector, the second is not. Boundary conditions and uncertainties are adequately characterized. Accurate CFD calculation of the flow will ultimately require accurate modeling of the chemical kinetics and turbulence-chemistry interactions as well as accurate modeling of the turbulent mixing

Experimental clean combustor program: Noise study

NASA Technical Reports Server (NTRS)

Sofrin, T. G.; Riloff, N., Jr.

1976-01-01

Under a Noise Addendum to the NASA Experimental Clean Combustor Program (ECCP) internal pressure fluctuations were measured during tests of JT9D combustor designs conducted in a burner test rig. Measurements were correlated with burner operating parameters using an expression relating farfield noise to these parameters. For a given combustor, variation of internal noise with operating parameters was reasonably well predicted by this expression but the levels were higher than farfield predictions and differed significantly among several combustors. For two burners, discharge stream temperature fluctuations were obtained with fast-response thermocouples to allow calculation of indirect combustion noise which would be generated by passage of the temperature inhomogeneities through the high pressure turbine stages of a JT9D turbofan engine. Using a previously developed analysis, the computed indirect combustion noise was significantly lower than total low frequency core noise observed on this and several other engines.

DEVELOPMENT OF A VORTEX CONTAINMENT COMBUSTOR FOR COAL COMBUSTION SYSTEMS

EPA Science Inventory

The report describes the development of a vortex containment combustor (VCC) for coal combustion systems, designed to solve major problems facing the conversion of oil- and gas-fired boilers to coal (e.g., derating, inorganic impurities in coal, and excessive formation of NOx and...

DEVELOPMENT OF A VORTEX CONTAINMENT COMBUSTOR FOR COAL COMBUSTION SYTEMS

EPA Science Inventory

The report describes the development of a vortex containment combustor (VCC) for coal combustion systems, designed to solve major problems facing the conversion of oil- and gas-fired boilers to coal (e.g., derating, inorganic impurities in coal, and excessive formation of NOx and...

MATE (Materials for Advanced Turbine Engines) Program, Project 3. Volume 2: Design, fabrication and evaluation of an oxide dispersion strengthened sheet alloy combustor liner

NASA Technical Reports Server (NTRS)

Bose, S.; Sheffler, K. D.

1988-01-01

The suitability of wrought oxide dispersion strengthened (ODS) superalloy sheet for gas turbine engine combustor applications was evaluated. Two yttria (Y2O3) dispersion strengthened alloys were evaluated; Incoloy MA956 and Haynes Development Alloy (HDA) 8077 (NiCrAl base). Preliminary tests showed both alloys to be potentially viable combustor materials, with neither alloy exhibiting a significant advantage over the other. MA956 was selected as the final alloy based on manufacturing reproducibility for evaluation as a burner liner. A hybrid PW2037 inner burner liner containing MA956 and Hastelloy X components and using a louvered configuration was designed and constructed. The louvered configuration was chosen because of field experience and compatibility with the bill of material PW2037 design. The simulated flight cycle for the ground based engine tests consisted of 4.5 min idle, 1.5 min takeoff and intermediate conditions in a PW2037 engine with average uncorrected combustor exit temperature of 1527 C. Post test evaluation consisting of visual observations and fluorescent penetrant inspections was conducted after 500 cycles of testing. No loss of integrity in the burner liner was shown.

Hot section viewing system

NASA Technical Reports Server (NTRS)

Morey, W. W.

1984-01-01

This report covers the development and testing of a prototype combustor viewing system. The system allows one to see and record images from the inside of an operating gas turbine combustor. The program proceeded through planned phases of conceptual design, preliminary testing to resolve problem areas, prototype design and fabrication, and rig testing. Successful tests were completed with the viewing system in the laboratory, in a high pressure combustor rig, and on a Pratt and Whitney PW20307 jet engine. Both film and video recordings were made during the tests. Digital image analysis techniques were used to enhance images and bring out special effects. The use of pulsed laser illumination was also demonstrated as a means for observing liner surfaces in the presence of luminous flame.

A Numerical Investigation of the Startup Transient in a Wave Rotor

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.

1996-01-01

The startup process is investigated for a hypothetical four-port wave rotor, envisioned as a topping cycle for a small gas turbine engine. The investigation is conducted numerically using a multi-passage, one-dimensional CFD-based wave rotor simulation in combination with lumped volume models for the combustor, exhaust valve plenum, and rotor center cavity components. The simulation is described and several startup transients are presented which illustrate potential difficulties for the specific cycle design investigated. In particular it is observed that, prior to combustor light-off, or just after, the flow through the combustor loop is reversed from the design direction. The phenomenon is demonstrated and several possible modifications techniques are discussed which avoid or overcome the problem.

High Frequency Adaptive Instability Suppression Controls in a Liquid-Fueled Combustor

NASA Technical Reports Server (NTRS)

Kopasakis, George

2003-01-01

This effort extends into high frequency (>500 Hz), an earlier developed adaptive control algorithm for the suppression of thermo-acoustic instabilities in a liquidfueled combustor. The earlier work covered the development of a controls algorithm for the suppression of a low frequency (280 Hz) combustion instability based on simulations, with no hardware testing involved. The work described here includes changes to the simulation and controller design necessary to control the high frequency instability, augmentations to the control algorithm to improve its performance, and finally hardware testing and results with an experimental combustor rig developed for the high frequency case. The Adaptive Sliding Phasor Averaged Control (ASPAC) algorithm modulates the fuel flow in the combustor with a control phase that continuously slides back and forth within the phase region that reduces the amplitude of the instability. The results demonstrate the power of the method - that it can identify and suppress the instability even when the instability amplitude is buried in the noise of the combustor pressure. The successful testing of the ASPAC approach helped complete an important NASA milestone to demonstrate advanced technologies for low-emission combustors.

Testing of felt-ceramic materials for combustor applications

NASA Technical Reports Server (NTRS)

Venkat, R. S.; Roffe, G.

1983-01-01

The feasibility of using composite felt ceramic materials as combustor liners was experimentally studied. The material consists of a porous felt pad sandwiched between a layer of ceramic and one of solid metal. Flat, rectangular test panels, which encompassed several design variations of the basic composite material, were tested, two at a time, in a premixed gas turbine combustor as sections of the combustor wall. Tests were conducted at combustor inlet conditions of 0.5 MPa and 533 K with a reference velocity of 25 m/s. The panels were subjected to a hot gas temperature of 2170 K with 1% of the total airflow used to film cool the ceramic surface of the test panel. In general, thin ceramic layers yield low ceramic stress levels with high felt ceramic interface temperatures. On the other hand, thick ceramic layers result in low felt ceramic interface temperatures but high ceramic stress levels. Extensive thermal cycling appears to cause material degradation, but for a limited number of cycles, the survivability of felt ceramic materials, even under extremely severe combustor operating conditions, was conclusively demonstrated.

Combustor kinetic energy efficiency analysis of the hypersonic research engine data

NASA Astrophysics Data System (ADS)

Hoose, K. V.

1993-11-01

A one-dimensional method for measuring combustor performance is needed to facilitate design and development scramjet engines. A one-dimensional kinetic energy efficiency method is used for measuring inlet and nozzle performance. The objective of this investigation was to assess the use of kinetic energy efficiency as an indicator for scramjet combustor performance. A combustor kinetic energy efficiency analysis was performed on the Hypersonic Research Engine (HRE) data. The HRE data was chosen for this analysis due to its thorough documentation and availability. The combustor, inlet, and nozzle kinetic energy efficiency values were utilized to determine an overall engine kinetic energy efficiency. Finally, a kinetic energy effectiveness method was developed to eliminate thermochemical losses from the combustion of fuel and air. All calculated values exhibit consistency over the flight speed range. Effects from fuel injection, altitude, angle of attack, subsonic-supersonic combustion transition, and inlet spike position are shown and discussed. The results of analyzing the HRE data indicate that the kinetic energy efficiency method is effective as a measure of scramjet combustor performance.

Creativity in ergonomic design: a supplemental value-adding source for product and service development.

PubMed

Zeng, Liang; Proctor, Robert W; Salvendy, Gavriel

2010-08-01

This article investigates the role of creativity in ergonomic design and the generic process of developing creative products and services. Creativity is gaining increased emphasis in both academia and industry. More than 50 years of research in creativity indicates that creativity is key to product and service innovation. Nevertheless, there is scarcely any comprehensive review dedicated to appraising the complex construct of creativity, the underlying cognitive process, and the role of creativity in product and service development. We review relevant literature regarding creativity, creative cognition, and the engineering design process to appraise the role of creativity in ergonomic design and to construct a conceptual model of creative product and service development. A framework of ergodesign creativity is advanced that highlights the central role of creativity in synergistically addressing the four dimensions of ergonomic design: functionality, safety, usability, and affectivity. A conceptual model of creative design process is then constructed that is goal oriented and is initiated by active problem finding and problem formulating. This process is carried out in a recursive and dynamic way, facilitated by creative thinking strategies. It is proposed that ergodesign creativity can add supplemental value to products and services, which subsequently affects consumer behavior and helps organizations gain competitive advantage. The proposed conceptual framework of ergodesign creativity and creative design process can serve as the ground for future theory development. Propositions advanced in this study should facilitate designers generating products and services that are creative and commercially competitive.

Flame stabilization and mixing characteristics in a Stagnation Point Reverse Flow combustor

NASA Astrophysics Data System (ADS)

Bobba, Mohan K.

A novel combustor design, referred to as the Stagnation Point Reverse-Flow (SPRF) combustor, was recently developed that is able to operate stably at very lean fuel-air mixtures and with low NOx emissions even when the fuel and air are not premixed before entering the combustor. The primary objective of this work is to elucidate the underlying physics behind the excellent stability and emissions performance of the SPRF combustor. The approach is to experimentally characterize velocities, species mixing, heat release and flame structure in an atmospheric pressure SPRF combustor with the help of various optical diagnostic techniques: OH PLIF, chemiluminescence imaging, PIV and Spontaneous Raman Scattering. Results indicate that the combustor is primarily stabilized in a region downstream of the injector that is characterized by low average velocities and high turbulence levels; this is also the region where most of the heat release occurs. High turbulence levels in the shear layer lead to increased product entrainment levels, elevating the reaction rates and thereby enhancing the combustor stability. The effect of product entrainment on chemical timescales and the flame structure is illustrated with simple reactor models. Although reactants are found to burn in a highly preheated (1300 K) and turbulent environment due to mixing with hot product gases, the residence times are sufficiently long compared to the ignition timescales such that the reactants do not autoignite. Turbulent flame structure analysis indicates that the flame is primarily in the thin reaction zones regime throughout the combustor, and it tends to become more flamelet like with increasing distance from the injector. Fuel-air mixing measurements in case of non-premixed operation indicate that the fuel is shielded from hot products until it is fully mixed with air, providing nearly premixed performance without the safety issues associated with premixing. The reduction in NOx emissions in the SPRF combustor are primarily due to its ability to stably operate under ultra lean (and nearly premixed) condition within the combustor. Further, to extend the usefulness of this combustor configuration to various applications, combustor geometry scaling rules were developed with the help of simplified coaxial and opposed jet models.

Investigation of Combustion Control in a Dump Combustor Using the Feedback Free Fluidic Oscillator

NASA Technical Reports Server (NTRS)

Meier, Eric J.; Casiano, Matthew J.; Anderson, William E.; Heister, Stephen D.

2015-01-01

A feedback free fluidic oscillator was designed and integrated into a single element rocket combustor with the goal of suppressing longitudinal combustion instabilities. The fluidic oscillator uses internal fluid dynamics to create an unsteady outlet jet at a specific frequency. An array of nine fluidic oscillators was tested to mimic modulated secondary oxidizer injection into the combustor dump plane. The combustor has a coaxial injector that uses gaseous methane and decomposed hydrogen peroxide with an overall O/F ratio of 11.7. A sonic choke plate on an actuator arm allows for continuous adjustment of the oxidizer post acoustics enabling the study of a variety of instability magnitudes. The fluidic oscillator unsteady outlet jet performance is compared against equivalent steady jet injection and a baseline design with no secondary oxidizer injection. At the most unstable operating conditions, the unsteady outlet jet saw a 67% reduction in the instability pressure oscillation magnitude when compared to the steady jet and baseline data. Additionally, computational fluid dynamics analysis of the combustor gives insight into the flow field interaction of the fluidic oscillators. The results indicate that open loop high frequency propellant modulation for combustion control can be achieved through fluidic devices that require no moving parts or electrical power to operate.

An Overview of Spray Modeling With OpenNCC and its Application to Emissions Predictions of a LDI Combustor at High Pressure

NASA Technical Reports Server (NTRS)

Raju, M. S.

2016-01-01

The open national combustion code (Open- NCC) is developed with the aim of advancing the current multi-dimensional computational tools used in the design of advanced technology combustors. In this paper we provide an overview of the spray module, LSPRAY-V, developed as a part of this effort. The spray solver is mainly designed to predict the flow, thermal, and transport properties of a rapidly evaporating multi-component liquid spray. The modeling approach is applicable over a wide-range of evaporating conditions (normal, superheat, and supercritical). The modeling approach is based on several well-established atomization, vaporization, and wall/droplet impingement models. It facilitates large-scale combustor computations through the use of massively parallel computers with the ability to perform the computations on either structured & unstructured grids. The spray module has a multi-liquid and multi-injector capability, and can be used in the calculation of both steady and unsteady computations. We conclude the paper by providing the results for a reacting spray generated by a single injector element with 600 axially swept swirler vanes. It is a configuration based on the next-generation lean-direct injection (LDI) combustor concept. The results include comparisons for both combustor exit temperature and EINOX at three different fuel/air ratios.

76 FR 22822 - Approval and Promulgation of State Plans for Designated Facilities and Pollutants: Florida...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-04-25

... Solid Waste Incinerator (OSWI) units from the State of Florida; Large Municipal Waste Combustor (LMWC), Small Municipal Waste Combustor (SMWC), and OSWI units from Jefferson County, Kentucky; LMWC, SMWC, and..., North Carolina; LMWC, SMWC, Hospital/Medical/Infectious Waste Incinerator (HMIWI), and OSWI units from...

Adaptation of Combustion Principles to Aircraft Propulsion. Volume I; Basic Considerations in the Combustion of Hydrocarbon Fuels with Air

NASA Technical Reports Server (NTRS)

Barnett, Henry C (Editor); Hibbard, Robert R (Editor)

1955-01-01

The report summarizes source material on combustion for flight-propulsion engineers. First, several chapters review fundamental processes such as fuel-air mixture preparation, gas flow and mixing, flammability and ignition, flame propagation in both homogenous and heterogenous media, flame stabilization, combustion oscillations, and smoke and carbon formation. The practical significance and the relation of these processes to theory are presented. A second series of chapters describes the observed performance and design problems of engine combustors of the principal types. An attempt is made to interpret performance in terms of the fundamental processes and theories previously reviewed. Third, the design of high-speed combustion systems is discussed. Combustor design principles that can be established from basic considerations and from experience with actual combustors are described. Finally, future requirements for aircraft engine combustion systems are examined.

Evaluation of a bulk calorimeter and heat balance for determination of supersonic combustor efficiency

NASA Technical Reports Server (NTRS)

Mcclinton, C. R.; Anderson, G. Y.

1980-01-01

Results are presented from the shakedown and evaluation test of a bulk calorimeter. The calorimeter is designed to quench the combustion at the exit of a direct-connect, hydrogen fueled, scramjet combustor model, and to provide the measurements necessary to perform an analysis of combustion efficiency. Results indicate that the calorimeter quenches reaction, that reasonable response times are obtained, and that the calculated combustion efficiency is repeatable within + or -3 percent and varies in a regular way with combustor model parameters such as injected fuel equivalence ratio.

Method for operating a combustor in a fuel cell system

DOEpatents

Clingerman, Bruce J.; Mowery, Kenneth D.

2002-01-01

In one aspect, the invention provides a method of operating a combustor to heat a fuel processor to a desired temperature in a fuel cell system, wherein the fuel processor generates hydrogen (H.sub.2) from a hydrocarbon for reaction within a fuel cell to generate electricity. More particularly, the invention provides a method and select system design features which cooperate to provide a start up mode of operation and a smooth transition from start-up of the combustor and fuel processor to a running mode.

Design of thermal protection system for 8 foot HTST combustor

NASA Technical Reports Server (NTRS)

Moskowitz, S.

1973-01-01

The combustor in the 8-foot high temperature structures tunnel at the NASA-Langley Research Center has encountered cracking over a period of 50-250 tunnel tests within a limited range of the required operating envelope. A program was conducted which analyzed the failed combustor liner hardware and determined that the mechanism of failure was vibratory fatigue. A vibration damper system using wave springs located axially between the liner T-bar and the liner support was designed as an intermediate solution to extend the life of the current two-pass regenerative air-cooled liner. The effects of liner wall thickness, cooling air passage height, stiffener ring geometry, reflective coatings, and liner material selection were investigated for these designs. Preliminary layout design arrangements including the external water-cooling system requirements, weight estimates, installation requirements and preliminary estimates of manufacturing costs were prepared for the most promissing configurations. A state-of-the-art review of thermal barrier coatings and an evaluation of reflective coatings for the gasside surface of air-cooled liners are included.

Numerical Study of Low Emission Gas Turbine Combustor Concepts

NASA Technical Reports Server (NTRS)

Yang, Song-Lin

2002-01-01

To further reduce pollutant emissions, such as CO, NO(x), UHCs, etc., in the next few decades, innovative concepts of gas turbine combustors must be developed. Several concepts, such as the LIPP (Lean- Premixed- Prevaporized), RQL (Rich-Burn Quick-Quench Lean-Burn), and LDI (Lean-Direct-Injection), have been under study for many years. To fully realize the potential of these concepts, several improvements, such as inlet geometry, air swirler, aerothermochemistry control, fuel preparation, fuel injection and injector design, etc., must be made, which can be studied through the experimental method and/or the numerical technique. The purpose of this proposal is to use the CFD technique to study, and hence, to guide the design process for low emission gas turbine combustors. A total of 13 technical papers have been (or will be) published.

Scale and geometry effects on heat-recirculating combustors

NASA Astrophysics Data System (ADS)

Chen, Chien-Hua; Ronney, Paul D.

2013-10-01

A simple analysis of linear and spiral counterflow heat-recirculating combustors was conducted to identify the dimensionless parameters expected to quantify the performance of such devices. A three-dimensional (3D) numerical model of spiral counterflow 'Swiss roll' combustors was then used to confirm and extend the applicability of the identified parameters. It was found that without property adjustment to maintain constant values of these parameters, at low Reynolds number (Re) smaller-scale combustors actually showed better performance (in terms of having lower lean extinction limits at the same Re) due to lower heat loss and internal wall-to-wall radiation effects, whereas at high Re, larger-scale combustors showed better performance due to longer residence time relative to chemical reaction time. By adjustment of property values, it was confirmed that four dimensionless parameters were sufficient to characterise combustor performance at all scales: Re, a heat loss coefficient (α), a Damköhler number (Da) and a radiative transfer number (R). The effect of diffusive transport effect (i.e. Lewis number) was found to be significant only at low Re. Substantial differences were found between the performance of linear and spiral combustors; these were explained in terms of the effects of the area exposed to heat loss to ambient and the sometimes detrimental effect of increasing heat transfer to adjacent outlet turns of the spiral exchanger. These results provide insight into the optimal design of small-scale combustors and choice of operation conditions.

A Modified Through-Flow Wave Rotor Cycle with Combustor Bypass Ducts

NASA Technical Reports Server (NTRS)

Paxson Daniel E.; Nalim, M. Razi

1998-01-01

A wave rotor cycle is described which avoids the inherent problem of combustor exhaust gas recirculation (EGR) found in four-port, through-flow wave rotor cycles currently under consideration for topping gas turbine engines. The recirculated hot gas is eliminated by the judicious placement of a bypass duct which transfers gas from one end of the rotor to the other. The resulting cycle, when analyzed numerically, yields an absolute mean rotor temperature 18% below the already impressive value of the conventional four-port cycle (approximately the turbine inlet temperature). The absolute temperature of the gas leading to the combustor is also reduced from the conventional four-port design by 22%. The overall design point pressure ratio of this new bypass cycle is approximately the same as the conventional four-port cycle. This paper will describe the EGR problem and the bypass cycle solution including relevant wave diagrams. Performance estimates of design and off-design operation of a specific wave rotor will be presented. The results were obtained using a one-dimensional numerical simulation and design code.

For operation of the Computer Software Management and Information Center (COSMIC)

NASA Technical Reports Server (NTRS)

Carmon, J. L.

1983-01-01

Computer programs for large systems of normal equations, an interactive digital signal process, structural analysis of cylindrical thrust chambers, swirling turbulent axisymmetric recirculating flows in practical isothermal combustor geometrics, computation of three dimensional combustor performance, a thermal radiation analysis system, transient response analysis, and a software design analysis are summarized.

40 CFR Table 1 to Subpart Fff of... - Municipal Waste Combustor Units (MWC Units) Excluded From Subpart FFF 1

Code of Federal Regulations, 2012 CFR

2012-07-01

... municipal solid waste at the following MWC sites: (a) Foster Wheeler Charleston Resource Recovery Facility... 40 Protection of Environment 9 2012-07-01 2012-07-01 false Municipal Waste Combustor Units (MWC... FOR DESIGNATED FACILITIES AND POLLUTANTS Federal Plan Requirements for Large Municipal Waste...

40 CFR Table 1 to Subpart Fff of... - Municipal Waste Combustor Units (MWC Units) Excluded From Subpart FFF 1

Code of Federal Regulations, 2014 CFR

2014-07-01

... municipal solid waste at the following MWC sites: (a) Foster Wheeler Charleston Resource Recovery Facility... 40 Protection of Environment 9 2014-07-01 2014-07-01 false Municipal Waste Combustor Units (MWC... FOR DESIGNATED FACILITIES AND POLLUTANTS Federal Plan Requirements for Large Municipal Waste...

76 FR 22861 - Approval and Promulgation of State Plans for Designated Facilities and Pollutants: Florida...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-04-25

... Waste Incinerator (OSWI) units from the State of Florida; Large Municipal Waste Combustor (LMWC), Small Municipal Waste Combustor (SMWC), and OSWI units from Jefferson County, Kentucky; LMWC, SMWC, and OSWI units...; LMWC, SMWC, Hospital/Medical/Infectious Waste Incinerator (HMIWI), and OSWI units from Buncombe County...

Laser velocimetry measurements in a gas turbine research combustor

NASA Technical Reports Server (NTRS)

Driscoll, J. F.; Pelaccio, D. G.

1979-01-01

The effects of turbulence on the production of pollutant species in a gas-turbine research combustor are studied using laser diffraction velocimetry (LDV) techniques. Measurements that were made in the primary combustion zone include mean velocity, rms velocity fluctuations, velocity probability distributions, and autocorrelation functions. A unique combustor design provides relatively uniform flow conditions and independent control of drop size, equivalence ratio, inlet temperature, and combustor pressure. Parameters which characterize the nature of the spray combustion (i.e., whether single droplet or group combustion occurs), were determined from the LDV data. Turbulent diffusivity (eddy viscosity) reaches a value of 2930 sq cm/sec, corresponding to a convective integral length scale of 1.8 cm. The group combustion number, based on turbulent diffusivity, is measured to be 6.2

Small Engine Technology (SET) - Task 4, Regional Turboprop/Turbofan Engine Advanced Combustor Study

NASA Technical Reports Server (NTRS)

Reynolds, Robert; Srinivasan, Ram; Myers, Geoffrey; Cardenas, Manuel; Penko, Paul F. (Technical Monitor)

2003-01-01

Under the SET Program Task 4 - Regional Turboprop/Turbofan Engine Advanced Combustor Study, a total of ten low-emissions combustion system concepts were evaluated analytically for three different gas turbine engine geometries and three different levels of oxides of nitrogen (NOx) reduction technology, using an existing AlliedSignal three-dimensional (3-D) Computational Fluid Dynamics (CFD) code to predict Landing and Takeoff (LTO) engine cycle emission values. A list of potential Barrier Technologies to the successful implementation of these low-NOx combustor designs was created and assessed. A trade study was performed that ranked each of the ten study configurations on the basis of a number of manufacturing and durability factors, in addition to emissions levels. The results of the trade study identified three basic NOx-emissions reduction concepts that could be incorporated in proposed follow-on combustor technology development programs aimed at demonstrating low-NOx combustor hardware. These concepts are: high-flow swirlers and primary orifices, fuel-preparation cans, and double-dome swirlers.

Assessment of Reduced-Kinetics Mechanisms for Combustion of Jet Fuel in CFD Applications

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Kundu, Krihna P.; Yungster, Shaye J.

2014-01-01

A computational effort was undertaken to analyze the details of fluid flow in Lean-Direct Injection (LDI) combustors for next-generation LDI design. The National Combustor Code (NCC) was used to perform reacting flow computations on single-element LDI injector configurations. The feasibility of using a reduced chemical-kinetics approach, which optimizes the reaction rates and species to model the emissions characteristics typical of lean-burning gas-turbine combustors, was assessed. The assessments were performed with Reynolds- Averaged Navier-Stokes (RANS) and Time-Filtered Navier Stokes (TFNS) time-integration, with a Lagrangian spray model with the NCC code. The NCC predictions for EINOx and combustor exit temperature were compared with experimental data for two different single-element LDI injector configurations, with 60deg and 45deg axially swept swirler vanes. The effects of turbulence-chemistry interaction on the predicted flow in a typical LDI combustor were studied with detailed comparisons of NCC TFNS with experimental data.

User's manual for a TEACH computer program for the analysis of turbulent, swirling reacting flow in a research combustor

NASA Technical Reports Server (NTRS)

Chiappetta, L. M.

1983-01-01

Described is a computer program for the analysis of the subsonic, swirling, reacting turbulent flow in an axisymmetric, bluff-body research combustor. The program features an improved finite-difference procedure designed to reduce the effects of numerical diffusion and a new algorithm for predicting the pressure distribution within the combustor. A research version of the computer program described in the report was supplied to United Technologies Research Center by Professor A. D. Gosman and his students, R. Benodeker and R. I. Issa, of Imperial College, London. The Imperial College staff also supplied much of the program documentation. Presented are a description of the mathematical model for flow within an axisymmetric bluff-body combustor, the development of the finite-difference procedure used to represent the system of equations, an outline of the algorithm for determining the static pressure distribution within the combustor, a description of the computer program including its input format, and the results for representative test cases.

Emission Characteristics of A P and W Axially Staged Sector Combustor

NASA Technical Reports Server (NTRS)

He, Zhuohui J.; Wey, Changlie; Chang, Clarence T.; Lee, Chi Ming; Surgenor, Angela D.; Kopp-Vaughan, Kristin; Cheung, Albert

2016-01-01

Emission characteristics of a three-cup P and W Axially Controlled Stoichiometry (ACS) sector combustor are reported in this article. Multiple injection points and fuel staging strategies are used in this combustor design. Pilot-stage injectors are located on the front dome plate of the combustor, and main-stage injectors are positioned on the top and bottom of the combustor liners downstream. Low power configuration uses only pilot-stage injectors. Main-stage injectors are added to high power configuration to help distribute fuel more evenly and achieve overall lean burn yielding very low NOx emissions. Combustion efficiencies at four ICAO LTO conditions were all above 99%. Three EINOx emissions correlation equations were developed based on the experimental data to describe the NOx emission trends of this combustor concept. For the 7% and 30% engine power conditions, NOx emissions are obtained with the low power configuration, and the EINOx values are 6.16 and 6.81. The high power configuration was used to assess 85% and 100% engine power NOx emissions, with measured EINOx values of 4.58 and 7.45, respectively. The overall landing-takeoff cycle NOx emissions are about 12% relative to ICAO CAEP/6 level.

Analysis of a Preloaded Bolted Joint in a Ceramic Composite Combustor

NASA Technical Reports Server (NTRS)

Hissam, D. Andy; Bower, Mark V.

2003-01-01

This paper presents the detailed analysis of a preloaded bolted joint incorporating ceramic materials. The objective of this analysis is to determine the suitability of a joint design for a ceramic combustor. The analysis addresses critical factors in bolted joint design including preload, preload uncertainty, and load factor. The relationship between key joint variables is also investigated. The analysis is based on four key design criteria, each addressing an anticipated failure mode. The criteria are defined in terms of margin of safety, which must be greater than zero for the design criteria to be satisfied. Since the proposed joint has positive margins of safety, the design criteria are satisfied. Therefore, the joint design is acceptable.

Application of jet-shear-layer mixing and effervescent atomization to the development of a low-NO(x) combustor. Ph.D. Thesis - Purdue Univ.

NASA Technical Reports Server (NTRS)

Colantonio, Renato Olaf

1993-01-01

An investigation was conducted to develop appropriate technologies for a low-NO(x), liquid-fueled combustor. The combustor incorporates an effervescent atomizer used to inject fuel into a premixing duct. Only a fraction of the combustion air is used in the premixing process to avoid autoignition and flashback problems. This fuel-rich mixture is introduced into the remaining combustion air by a rapid jet-shear-layer-mixing process involving radial fuel-air jets impinging on axial air jets in the primary combustion zone. Computational analysis was used to provide a better understanding of the fluid dynamics that occur in jet-shear-layer mixing and to facilitate a parametric analysis appropriate to the design of an optimum low-NO(x) combustor. A number of combustor configurations were studied to assess the key combustor technologies and to validate the modeling code. The results from the experimental testing and computational analysis indicate a low-NO(x) potential for the jet-shear-layer combustor. Key parameters found to affect NO(x) emissions are the primary combustion zone fuel-air ratio, the number of axial and radial jets, the aspect ratio and radial location of the axial air jets, and the radial jet inlet hole diameter. Each of these key parameters exhibits a low-NO(x) point from which an optimized combustor was developed. Using the parametric analysis, NO(x) emissions were reduced by a factor of 3 as compared with the emissions from conventional, liquid-fueled combustors operating at cruise conditions. Further development promises even lower NO(x) with high combustion efficiency.

Advanced Catalytic Combustors for Low Pollutant Emissions

DTIC Science & Technology

1979-11-01

liner materials in both designs are sheet or forged HS188 . This alloy was designed for stability of the microstructure and properties during heat...625 Inco 718 I: Inca 625 HS188 L605 Hast X Figure 37. Catalytic Combustor Materials. 111 &I ~ 4 L. ,.Add. R service. The hot corrosion resistance of... HS188 is similar to L605 and somewhat F better than Hastelloy X, another common liner material. HS188 exhibits good low cyclo fatigue resistance up to

Active Pattern Factor Control for Gas Turbine Engines

NASA Technical Reports Server (NTRS)

May, James E.

1998-01-01

Small variations in fuel/air mixture ratios within gas turbine combustors can result in measurable, and potentially detrimental, exit thermal gradients. Thermal gradients can increase emissions, as well as shorten the design life of downstream turbomachinery, particularly stator vanes. Uniform temperature profiles are usually sought through careful design and manufacturing of related combustor components. However, small componentto-component variations as well as numerous aging effects degrade system performance. To compensate for degraded thermal performance, researchers are investigating active, closed-loop control schemes.

Experimental investigations on effect of different materials and varying depths of one turn exhaust channel swiss roll combustor on its thermal performance

NASA Astrophysics Data System (ADS)

Mane Deshmukh, Sagar B.; Krishnamoorthy, A.; Bhojwani, V. K.; Pawane, Ashwini

2017-05-01

More energy density of hydrocarbon fuels compared to advanced batteries available in the market demands for development of systems which will use hydrocarbon fuels at small scale to generate power in small quantity (i.e. in few watts) and device efficiency should be reasonably good, but the basic requirement is to generate heat from the fuels like methane, propane, hydrogen, LPG and converting into power. Swiss roll combustor has proved to be best combustor at small scale. Present work is carried out on one turn exhaust channel and half turn of inlet mixture channel Swiss roll combustor. Purpose of keeping exhaust channel length more than the inlet mixture channel to ensure sufficient time for heat exchange between burned and unburned gases, which is not reported in earlier studies. Experimental study mentions effects of different design parameters like materials of combustor, various depths, equivalence ratio, mass flow rates of liquefied petroleum gas (LPG), volume of combustion space and environmental conditions (with insulation and without insulation to combustors) on fuel lean limit and fuel rich limit, temperature profile obtained on all external surfaces, in the main combustion chamber, in the channel carrying unburned gas mixture and burned gas mixture, heat loss to atmosphere from all the walls of combustor, flame location. Different combustor materials tested were stainless steel, Aluminum, copper, brass, bronze, Granite. Depths considered were 22mm, 15mm, 10mm and 5mm. It was observed that flame stability inside the combustion chamber is affected by materials, depths and flow rates. Unburned mixture carrying channel was kept below quenching distance of flame to avoid flash back. Burned gas carrying channel dimension was more than the quenching distance. Considerable temperature rise was observed with insulation to combustors. But combustors with more thermal conductivity showed more heat loss to atmosphere which led to instability of flame.

77 FR 32022 - Direct Final Negative Declaration and Withdrawal of Large Municipal Waste Combustors State Plan...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-31

... Negative Declaration and Withdrawal of Large Municipal Waste Combustors State Plan for Designated.... SUMMARY: EPA is taking direct final action to approve Illinois' negative declaration and request for EPA..., the state may submit a letter of certification to that effect, or a negative declaration, in lieu of a...

The jet engine design that can drastically reduce oxides of nitrogen

NASA Technical Reports Server (NTRS)

Ferri, A.; Agnone, A.

1977-01-01

The NOx pollution problem of hydrogen fueled turbojets and supersonic combustion ramjets (scramjets) was investigated to determine means of substantially alleviating the problem. Since the NOx reaction rates are much slower than the energy producing reactions, the NOx production depends mainly on the maximum local temperatures in the combustor and the NOx concentration is far from equilibrium at the end of a typical combustor (L approximately 1 ft). In diffusion flames, as used in present turbojets and scramjets combustor designs, the maximum local temperature occurs at the flame and is equal to the stoichiometric value. Whereas, in the heat conduction flames, wherein the flame propagates due to a heat conduction process away from the flame to the cooler oncoming premixed unburnt gases, the maximum temperature is lower than in the diffusion flame. Hence the corresponding pollution index is also lower.

Low-nitrogen oxides combustion of dried sludge using a pilot-scale cyclone combustor with recirculation.

PubMed

Shim, Sung Hoon; Jeong, Sang Hyun; Lee, Sang-Sup

2015-04-01

Recently, numerical and experimental studies have been conducted to develop a moderate or intense low-oxygen dilution (MILD) combustion technology for solid fuels. The study results demonstrated that intense recirculation inside the furnace by high-momentum air is a key parameter to achieve the MILD combustion of solid fuels. However, the high-velocity air requires a significant amount of electricity consumption. A cyclone-type MILD combustor was therefore designed and constructed in the authors' laboratory to improve the recirculation inside the combustor. The laboratory-scale tests yielded promising results for the MILD combustion of dried sewage sludge. To achieve pilot-scale MILD combustion of dried sludge in this study, the effects of geometric parameters such as the venturi tube configuration, the air injection location, and the air nozzle diameter were investigated. With the optimized geometric and operational conditions, the pilot-scale cyclone combustor demonstrated successful MILD combustion of dried sludge at a rate of 75 kg/hr with an excess air ratio of 1.05. A horizontal cyclone combustor with recirculation demonstrated moderate or intense low-oxygen dilution (MILD) combustion of dried sewage sludge at a rate of 75 kg/hr. Optimizing only geometric and operational conditions of the combustor reduced nitrogen oxide (NOx) emissions to less than 75 ppm. Because the operating cost of the MILD combustor is much lower than that of the selective catalytic reduction (SCR) applied to the conventional combustor, MILD combustion technology with the cyclone type furnace is an eligible option for reducing NOx emissions from the combustion of dried sewage sludge.

Flow interaction in the combustor-diffusor system of industrial gas turbines

DOE Office of Scientific and Technical Information (OSTI.GOV)

Agrawal, A.K.; Kapat, J.S.; Yang, T.

1996-05-01

This paper presents an experimental/computational study of cold flow in the combustor-diffuser system of industrial gas turbines to address issues relating to flow interactions and pressure losses in the pre- and dump diffusers. The present configuration with can annular combustors differs substantially from the aircraft engines which typically use a 360 degree annular combustor. Experiments were conducted in a one-third scale, annular 360-degree model using several can combustors equispaced around the turbine axis. A 3-D computational fluid dynamics analysis employing the multidomain procedure was performed to supplement the flow measurements. The measured data correlated well with the computations. The airflowmore » in the dump diffuser adversely affected the prediffuser flow by causing it to accelerate in the outer region at the prediffuser exit. This phenomenon referred to as the sink-effect also caused a large fraction of the flow to bypass much of the dump diffuser and go directly from the prediffuser exit to the bypass air holes on the combustor casing, thereby, rendering the dump diffuser ineffective in diffusing the flow. The dump diffuser was occupied by a large recirculation region which dissipated the flow kinetic energy. Approximately 1.2 dynamic head at the prediffuser inlet was lost in the combustor-diffuser system; much of it in the dump diffuser where the fluid passed through the narrow gaps and pathways. Strong flow interactions in the combustor-diffuser system indicate the need for design modifications which could not be addressed by empirical correlations based on simple flow configurations.« less

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James R.

2007-01-01

This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

Numerical study of effect of compressor swirling flow on combustor design in a MTE

NASA Astrophysics Data System (ADS)

Mu, Yong; Wang, Chengdong; Liu, Cunxi; Liu, Fuqiang; Hu, Chunyan; Xu, Gang; Zhu, Junqiang

2017-08-01

An effect of the swirling flow on the combustion performance is studied by the computational fluid dynamics (CFD) in a micro-gas turbine with a centrifugal compressor, dump diffuser and forward-flow combustor. The distributions of air mass and the Temperature Pattern Factor (as: Overall Temperature Distribution Factor -OTDF) in outlet are investigated with two different swirling angles of compressed air as 0° and 15° in three combustors. The results show that the influences of swirling flow on the air distribution and OTDF cannot be neglected. Compared with no-swirling flow, the air through outer liner is more, and the air through the inner liner is less, and the pressure loss is bigger under the swirling condition in the same combustor. The Temperature Pattern Factor changes under the different swirling conditions.

Design and evaluation of combustors for reducing aircraft engine pollution

NASA Technical Reports Server (NTRS)

Jones, R. E.; Grobman, J.

1973-01-01

Various techniques and test results are briefly described and referenced for detail. The effort arises from the increasing concern for the measurement and control of emissions from gas turbine engines. The greater part of this research is focused on reducing the oxides of nitrogen formed during takeoff and cruise in both advanced CTOL, high pressure ratio engines, and advanced supersonic aircraft engines. The experimental approaches taken to reduce oxides of nitrogen emissions include the use of: multizone combustors incorporating reduced dwell time, fuel-air premixing, air atomization, fuel prevaporization, water injection, and gaseous fuels. In the experiments conducted to date, some of these techniques were more successful than others in reducing oxides of nitrogen emissions. Tests are being conducted on full-annular combustors at pressures up to 6 atmospheres and on combustor segments at pressures up to 30 atmospheres.

Advanced technology for reducing aircraft engine pollution

NASA Technical Reports Server (NTRS)

Jones, R. E.

1973-01-01

Combustor research programs are described whose purpose is to demonstrate significantly lower exhaust emission levels. The proposed EPA regulations covering the allowable levels of emissions will require a major technological effort if these levels are to be met by 1979. Pollution reduction technology is being pursued by NASA through a combination of in-house research, contracted progams, and university grants. In-house research with the swirl-can modular combustor and the double-annular combustor has demonstrated significant reduction in the level of NO(x) emissions. The work is continuing in an attempt to further reduce these levels by improvements in module design and in air-fuel scheduling. Research on the reduction of idle emissions has included the conversion of conventional duplex fuel nozzles to air-assisted nozzles and exploration of the potential improvements possible with fuel staging and variable combustor geometry.

CFD Based Design of a Filming Injector for N+3 Combustors

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Mongia, Hukam; Lee, Phil

2016-01-01

An effort was undertaken to perform CFD analysis of fluid flow in Lean-Direct Injection (LDI) combustors with axial swirl-venturi elements for next-generation LDI-3 combustor design. The National Combustion Code (NCC) was used to perform non-reacting and two-phase reacting flow computations for a newly-designed pre-filming type fuel injector LDI-3 injector, in a single-injector and a five-injector array configuration. All computations were performed with a consistent approach of mesh-optimization, spray-modeling, ignition and kinetics-modeling. Computational predictions of the aerodynamics of the single-injector were used to arrive at an optimized main-injector design that meets effective area and fuel-air mixing criteria. Emissions (EINOx) characteristics were predicted for a medium-power engine cycle condition, and will be compared with data when it is made available from experimental measurements. The use of a PDF-like turbulence-chemistry interaction model with NCC's Time-Filtered Navier-Stokes (TFNS) solver is shown to produce a significant impact on the CFD results, when compared with a laminar-chemistry TFNS approach for the five-injector computations.

Advanced Optical Diagnostic Methods for Describing Fuel Injection and Combustion Flowfield Phenomena

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Hicks, Yolanda R.; Anderson, Robert C.

2004-01-01

Over the past decade advanced optical diagnostic techniques have evolved and matured to a point where they are now widely applied in the interrogation of high pressure combusting flows. At NASA Glenn Research Center (GRC), imaging techniques have been used successfully in on-going work to develop the next generation of commercial aircraft gas turbine combustors. This work has centered on providing a means by which researchers and designers can obtain direct visual observation and measurements of the fuel injection/mixing/combustion processes and combustor flowfield in two- and three-dimensional views at actual operational conditions. Obtaining a thorough understanding of the chemical and physical processes at the extreme operating conditions of the next generation of combustors is critical to reducing emissions and increasing fuel efficiency. To accomplish this and other tasks, the diagnostic team at GRC has designed and constructed optically accessible, high pressurer high temperature flame tubes and sectar rigs capable of optically probing the 20-60 atm flowfields of these aero-combustors. Among the techniques employed at GRC are planar laser-induced fluorescence (PLIF) for imaging molecular species as well as liquid and gaseous fuel; planar light scattering (PLS) for imaging fuel sprays and droplets; and spontaneous Raman scattering for species and temperature measurement. Using these techniques, optical measurements never before possible have been made in the actual environments of liquid fueled gas turbines. 2-D mapping of such parameters as species (e.g. OH-, NO and kerosene-based jet fuel) distribution, injector spray angle, and fuel/air distribution are just some of the measurements that are now routinely made. Optical imaging has also provided prompt feedback to researchers regarding the effects of changes in the fuel injector configuration on both combustor performance and flowfield character. Several injector design modifications and improvements have resulted from this feedback. Alternate diagnostic methods are constantly being evaluated as to their suitability as a diagnostic tool in these environments. A new method currently under examination is background oriented Schlieren (BOS) for examining the fuel/air mixing processes. While ratioing the Stokes and anti-Stokes nitrogen lines obtained from spontaneous Raman is being refined for temperature measurement. While the primary focus of the GRC diagnostic work remains optical species measurement and flow stream characterization, an increased emphasis has been placed on our involvement in flame code validation efforts. A functional combustor code should shorten and streamline future combustor design. Quantitative measurements of flow parameters such as temperature, species concentration, drop size and velocity using such methods as Raman and phase Doppler anemometry will provide data necessary in this effort.

Creativity and Creative Teams

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Bauer, Steven X. S.; Hunter, Craig A.

2001-01-01

A review of the linkage between knowledge, creativity, and design is presented and related to the best practices of multidisciplinary design teams. The discussion related to design and design teams is presented in the context of both the complete aerodynamic design community and specifically the work environment at the NASA Langley Research Center. To explore ways to introduce knowledge and creativity into the research and design environment at NASA Langley Research Center a creative design activity was executed within the context of a national product development activity. The success of the creative design team activity gave rise to a need to communicate the experience in a straightforward and managed approach. As a result the concept of creative potential its formulated and assessed with a survey of a small portion of the aeronautics research staff at NASA Langley Research Center. The final section of the paper provides recommendations for future creative organizations and work environments.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Technical Reports Server (NTRS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-01-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Astrophysics Data System (ADS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-07-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

Effect of vane twist on the performance of dome swirlers for gas turbine airblast atomizers

NASA Astrophysics Data System (ADS)

Micklow, Gerald J.; Dogra, Anju S.; Nguyen, H. Lee

1990-06-01

For advanced gas turbine engines, two combustor systems, the lean premixed/prevaporized (LPP) and the rich burn/quick quench/lean burn (RQL) offer great potential for reducing NO(x) emissions. An important consideration for either concept is the development of an advanced fuel injection system that will provide a stable, efficient, and very uniform combustion system over a wide operating range. High-shear airblast fuel injectors for gas turbine combustors have exhibited superior atomization and mixing compared with pressure-atomizing fuel injectors. This improved mixing has lowered NO(x) emissions and the pattern factor, and has enabled combustors to alternate fuels while maintaining a stable, efficient combustion system. The performance of high-shear airblast fuel injectors is highly dependent on the design of the dome swirl vanes. The type of swirl vanes most widely used in gas turbine combustors are usually flat for ease of manufacture, but vanes with curvature will, in general, give superior aerodynamic performance. The design and performance of high-turning, low-loss curved dome swirl vanes with twist along the span are investigated. The twist induces a secondary vortex flow pattern which will improve the atomization of the fuel, thereby producing a more uniform fuel-air distribution. This uniform distribution will increase combustion efficiency while lowering NO(x) emissions. A systematic swirl vane design system is presented based on one-, two-, and three-dimensional flowfield calculations, with variations in vane-turning angle, rate of turning, vane solidity, and vane twist as design parameters.

Creative user-centered visualization design for energy analysts and modelers.

PubMed

Goodwin, Sarah; Dykes, Jason; Jones, Sara; Dillingham, Iain; Dove, Graham; Duffy, Alison; Kachkaev, Alexander; Slingsby, Aidan; Wood, Jo

2013-12-01

We enhance a user-centered design process with techniques that deliberately promote creativity to identify opportunities for the visualization of data generated by a major energy supplier. Visualization prototypes developed in this way prove effective in a situation whereby data sets are largely unknown and requirements open - enabling successful exploration of possibilities for visualization in Smart Home data analysis. The process gives rise to novel designs and design metaphors including data sculpting. It suggests: that the deliberate use of creativity techniques with data stakeholders is likely to contribute to successful, novel and effective solutions; that being explicit about creativity may contribute to designers developing creative solutions; that using creativity techniques early in the design process may result in a creative approach persisting throughout the process. The work constitutes the first systematic visualization design for a data rich source that will be increasingly important to energy suppliers and consumers as Smart Meter technology is widely deployed. It is novel in explicitly employing creativity techniques at the requirements stage of visualization design and development, paving the way for further use and study of creativity methods in visualization design.

Effect of structural heat conduction on the performance of micro-combustors and micro-thrusters

NASA Astrophysics Data System (ADS)

Leach, Timothy Thierry

This thesis investigates the effect of gas-structure interaction on the design and performance of miniaturized combustors with characteristic dimensions less than a few millimeters. These are termed 'micro-combustors' and are intended for use in devices ranging from micro-scale rocket motors for micro, nano, and pico-satellite propulsion, to micro-scale engines for micro-Unmanned Air Vehicle (UAV) propulsion and compact power generation. Analytical models for the propagation of a premixed laminar flame in a micro-channel are developed. The models' predictions are compared to the results of more detailed numerical simulations that incorporate multi-step chemistry, distributed heat transfer between the reacting gas and the combustor structure, heat transfer between the combustor and the environment, and heat transfer within the combustor structure. The results of the modeling and simulation efforts are found to be in good qualitative agreement and demonstrate that the behavior of premixed laminar flames in micro-channels is governed by heat transfer within the combustor structure and heat loss to the environment. The key findings of this work are as follows: First, heat transfer through the micro-combustor's structure tends to increase the flame speed and flame thickness. The increase in flame thickness with decreasing passage height suggests that micro-scale combustors will need to be longer than their conventional-scale counterparts. However, the increase in flame speed more than compensates for this effect and the net effect is that miniaturizing a combustor can increase its power density substantially. Second, miniaturizing chemical rocket thrusters can substantially increase thrust/weight ratio but comes at the price of reduced specific impulse (i.e. overall efficiency). Third, heat transfer through the combustor's structure increases steady-state and transient flame stability. This means that micro-scale combustors will be more stable than their conventional-scale counterparts. Fourth and finally, the extended temperature profile associated with the broadened flame causes a different set of elementary reactions to dominate the operation of the overall reaction mechanism at the micro-scale. This suggests that new chemical mechanisms may need to be developed in order to accurately simulate combustion at small-scales. It also calls into question the efficacy of single-step mechanisms presently used by other researchers.

Fuel property effects on USAF gas turbine engine combustors and afterburners

NASA Technical Reports Server (NTRS)

Reeves, C. M.

1984-01-01

Since the early 1970s, the cost and availability of aircraft fuel have changed drastically. These problems prompted a program to evaluate the effects of broadened specification fuels on current and future aircraft engine combustors employed by the USAF. Phase 1 of this program was to test a set of fuels having a broad range of chemical and physical properties in a select group of gas turbine engine combustors currently in use by the USAF. The fuels ranged from JP4 to Diesel Fuel number two (DF2) with hydrogen content ranging from 14.5 percent down to 12 percent by weight, density ranging from 752 kg/sq m to 837 kg/sq m, and viscosity ranging from 0.830 sq mm/s to 3.245 sq mm/s. In addition, there was a broad range of aromatic content and physical properties attained by using Gulf Mineral Seal Oil, Xylene Bottoms, and 2040 Solvent as blending agents in JP4, JP5, JP8, and DF2. The objective of Phase 2 was to develop simple correlations and models of fuel effects on combustor performance and durability. The major variables of concern were fuel chemical and physical properties, combustor design factors, and combustor operating conditions.

Emission response from extended length, variable geometry gas turbine combustor

DOE Office of Scientific and Technical Information (OSTI.GOV)

Troth, D.L.; Verdouw, A.J.; Tomlinson, J.G.

1974-01-01

A program to analyze, select, and experimentally evaluate low emission combustors for aircraft gas turbine engines is conducted to demonstrate a final combustor concept having a 50 percent reduction in total mass emissions (carbon monoxide, unburnt hydrocarbons, oxides of nitrogen, and exhaust smoke) without an increase in any specific pollutant. Research conducted under an Army Contract established design concepts demonstrating significant reductions in CO and UHC emissions. Two of these concepts were an extended length intermediate zone to consume CO and UHC and variable geometry to control the primary zone fuel air ratio over varying power conditions. Emission reduction featuresmore » were identified by analytical methods employing both reaction kinetics and empirical correlations. Experimental results were obtained on a T63 component combustor rig operating at conditions simulating the engine over the complete power operating range with JP-4 fuel. A combustor incorporating both extended length and variable geometry was evaluated and the performance and emission results are reported. These results are compared on the basis of a helicopter duty cycle and the EPA 1979 turboprop regulation landing take off cycle. The 1979 EPA emission regulations for P2 class engines can be met with the extended length variable geometry combustor on the T63 turboprop engine.« less

CFD analysis of a scramjet combustor with cavity based flame holders

NASA Astrophysics Data System (ADS)

Kummitha, Obula Reddy; Pandey, Krishna Murari; Gupta, Rajat

2018-03-01

Numerical analysis of a scramjet combustor with different cavity flame holders has been carried out using ANSYS 16 - FLUENT tool. In this research article the internal fluid flow behaviour of the scramjet combustor with different cavity based flame holders have been discussed in detail. Two dimensional Reynolds-Averaged Navier-Stokes governing(RANS) equations and shear stress turbulence (SST) k - ω model along with finite rate/eddy dissipation chemistry turbulence have been considered for modelling chemical reacting flows. Due to the advantage of less computational time, global one step reaction mechanism has been used for combustion modelling of hydrogen and air. The performance of the scramjet combustor with two different cavities namely spherical and step cavity has been compared with the standard DLR scramjet. From the comparison of numerical results, it is found that the development of recirculation regions and additional shock waves from the edge of cavity flame holder is increased. And also it is observed that with the cavity flame holder the residence time of air in the scramjet combustor is also increased and achieved stabilized combustion. From this research analysis, it has been found that the mixing and combustion efficiency of scramjet combustor with step cavity design is optimum as compared to other models.

Time-Resolved Optical Measurements of Fuel-Air Mixedness in Windowless High Speed Research Combustors

NASA Technical Reports Server (NTRS)

Nguyen, Quang-Viet

1998-01-01

Fuel distribution measurements in gas turbine combustors are needed from both pollution and fuel-efficiency standpoints. In addition to providing valuable data for performance testing and engine development, measurements of fuel distributions uniquely complement predictive numerical simulations. Although equally important as spatial distribution, the temporal distribution of the fuel is an often overlooked aspect of combustor design and development. This is due partly to the difficulties in applying time-resolved diagnostic techniques to the high-pressure, high-temperature environments inside gas turbine engines. Time-resolved measurements of the fuel-to-air ratio (F/A) can give researchers critical insights into combustor dynamics and acoustics. Beginning in early 1998, a windowless technique that uses fiber-optic, line-of-sight, infrared laser light absorption to measure the time-resolved fluctuations of the F/A (refs. 1 and 2) will be used within the premixer section of a lean-premixed, prevaporized (LPP) combustor in NASA Lewis Research Center's CE-5 facility. The fiber-optic F/A sensor will permit optical access while eliminating the need for film-cooled windows, which perturb the flow. More importantly, the real-time data from the fiber-optic F/A sensor will provide unique information for the active feedback control of combustor dynamics. This will be a prototype for an airborne sensor control system.

In-situ measurement of residence time distributions in a turbulent oxy-fuel gas-flame combustor

NASA Astrophysics Data System (ADS)

Bürkle, Sebastian; Becker, Lukas G.; Agizza, Maria Angela; Dreizler, Andreas; Ebert, Volker; Wagner, Steven

2017-07-01

For improving the design of combustors, the knowledge of residence-time distributions (RTD) is important as they influence exhaust gas compositions. Measuring RTDs in combustors is challenging, due to high temperatures, chemical reactions, the presence of particles or corrosive species as well as high turbulence levels. This paper presents a technique for the in situ measurement of RTDs in combustors. Based on tunable diode laser absorption spectroscopy (TDLAS), the temporal evolution of the concentration of tracers is tracked simultaneously at the combustion chamber inlet and outlet. Using either air or mixtures of oxygen and carbon dioxide (oxy-fuel atmosphere) as oxidants, the method is applied to reacting and non-reacting operating conditions in a 20-kWth methane combustor. For reacting conditions, hydrogen chloride is used as a tracer, whereas for non-reacting conditions methane was chosen. Depending on the tracer, for a repetition rate of approximately 2 kHz detection limits of 16-40 ppmV are achieved. For deriving RTDs, low-pass filtering is compared to reactor network modeling. Different RTDs observed for varying operating conditions are discussed.

The NASA Ames Hypersonic Combustor-Model Inlet CFD Simulations and Experimental Comparisons

NASA Technical Reports Server (NTRS)

Venkatapathy, E.; Tokarcik-Polsky, S.; Deiwert, G. S.; Edwards, Thomas A. (Technical Monitor)

1995-01-01

Computations have been performed on a three-dimensional inlet associated with the NASA Ames combustor model for the hypersonic propulsion experiment in the 16-inch shock tunnel. The 3-dimensional inlet was designed to have the combustor inlet flow nearly two-dimensional and of sufficient mass flow necessary for combustion. The 16-inch shock tunnel experiment is a short duration test with test time of the order of milliseconds. The flow through the inlet is in chemical non-equilibrium. Two test entries have been completed and limited experimental results for the inlet region of the combustor-model are available. A number of CFD simulations, with various levels of simplifications such as 2-D simulations, 3-D simulations with and without chemical reactions, simulations with and without turbulent conditions, etc., have been performed. These simulations have helped determine the model inlet flow characteristics and the important factors that affect the combustor inlet flow and the sensitivity of the flow field to these simplifications. In the proposed paper, CFD modeling of the hypersonic inlet, results from the simulations and comparison with available experimental results will be presented.

Evaluation of fuel injection configurations to control carbon and soot formation in small GT combustors

NASA Technical Reports Server (NTRS)

Rosfjord, T. J.; Briehl, D.

1982-01-01

An experimental program to investigate hardware configurations which attempt to minimize carbon formation and soot production without sacrificing performance in small gas turbine combustors has been conducted at the United Technologies Research Center. Four fuel injectors, embodying either airblast atomization, pressure atomization, or fuel vaporization techniques, were combined with nozzle air swirlers and injector sheaths, and evaluated at test conditions which included and extended beyond standard small gas turbine combustor operation. Extensive testing was accomplished with configurations embodying either a spill return or a T-vaporizer injector. Minimal carbon deposits were observed on the spill return nozzle for tests using either Jet A or ERBS test fuel. A more extensive film of soft carbon was observed on the vaporizer after operation at standard engine conditions, with large carbonaceous growths forming on the device during off-design operation at low combustor inlet temperature. Test results indicated that smoke emission levels depended on the combustor fluid mechanics (especially the mixing rates near the injector), the atomization quality of the injector and the fuel hydrogen content.

Improved Controllers For Heaters In Toxic-Gas Combustors

NASA Technical Reports Server (NTRS)

Wishard, James; Lamb, James; Fortier, Edward; Velasquez, Hugo; Waltman, Doug

1995-01-01

Commercial electronic proportional controllers installed in place of mechanical power controllers for electric heaters in toxic-gas combustors at NASA's Jet Propulsion Laboratory. Designed to maintain temperature of heater at preset value by turning power fully on or fully off when temperature falls below or rises above that value, respectively. Solid-state power controllers overcome deficiencies of mechanical power controllers.

Fuel Property Effects on Combustor Performance.

DTIC Science & Technology

1980-03-01

emissions measurements are given in a previous report by Moses and Nae- geli . (4) 8 B. Combustor Rigs A Phillips-designed, 2-inch-diameter, high...333 311 80 256 262 270 294 331 176 342 321 85 261 268 277 308 340 188 352 333 90 267 275 295 324 351 200 364 346 95 274 297 322 347 367 215 382 364

Low NO sub x heavy fuel combustor concept program phase 1A gas tests

NASA Technical Reports Server (NTRS)

Cutrone, M. B.; Beebe, K. W.; Cutrone, M. B.

1982-01-01

The emissions performance of a rich lean combustor (developed for liquid fuels) for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf were assessed. The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. Although meeting NOx goals for the 167 Btu/scf gas, NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx generation from NH3 was significant at ammonia concentrations significantly less tha 0.5%. These levels occur depending on fuel gas cleanup system design, However, NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher.

The Effect of Air Preheat at Atmospheric Pressure on the Formation of NO(x) in the Quick-Mix Sections of an Axially Staged Combustor

NASA Technical Reports Server (NTRS)

Vardakas, M. A.; Leong, M. Y.; Brouwer, J.; Samuelsen, G. S.; Holdeman, J. D.

1999-01-01

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor concept has been proposed to minimize the formation of nitrogen oxides (NO(x)) in gas turbine systems. The success of this combustor strategy is dependent upon the efficiency of the mixing section bridging the fuel-rich and fuel-lean stages. Note that although these results were obtained from an experiment designed to study an RQL mixer, the link between mixing and NOx signatures is considerably broader than this application, in that the need to understand this link exists in most advanced combustors. The experiment reported herein was designed to study the effects of inlet air temperature on NO(x) formation in a mixing section. The results indicate that NO(x) emission is increased for all preheated cases compared to non-preheated cases. When comparing the various mixing modules, the affect of jet penetration is important, as this determines where NO(x) concentrations peak, and affects overall NO(x) production. Although jet air comprises 70 percent of the total airflow, the impact that jet air preheat has on overall NO(x) emissions is small compared to preheating both main and jet air flow.

Evaluation of Water Injection Effect on NO(x) Formation for a Staged Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

Fan, L.; Yang, S. L.; Kundu, K. P.

1996-01-01

NO(x) emission control by water injection on a staged turbine combustor (STC) was modeled using the KIVA-2 code with modification. Water is injected into the rich-burn combustion zone of the combustor by a single nozzle. Parametric study for different water injection patterns was performed. Results show NO(x) emission will decrease after water being injected. Water nozzle location also has significant effect for NO formation and fuel ignition. The chemical kinetic model is also sensitive to the excess water. Through this study, a better understanding of the physics and chemical kinetics is obtained, this will enhance the STC design process.

Final analysis and design of a thermal protection system for 8-foot HTST combustor

NASA Technical Reports Server (NTRS)

Moskowitz, S.

1973-01-01

The cylindrical shell combustor with T-bar supports in the 8-foot HTST at the NASA-Langley Research Center encountered vibratory fatigue cracking over a period of 50-250 tunnel tests within a limited range of the required operating envelope. A preliminary design study provided several suitable thermal protection system designs for the combustor, one of which was a two-pass regenerative type air-cooled omega-shaped segment liner. A final design layout of the omega segment liner was prepared and analyzed for steady-state and transient conditions. The design of a support system for the fuel spray bar assembly was also included. Detail drawings suitable for fabrication purposes were also prepared. Liner design problems defined during the preliminary study included (1) the ingress of gas into the attachment bulb section of the omega segment, (2) the large thermal gradient along the leg of the omega bulb attachment section and, (3) the local peak metal temperature at the radius between the liner ID and the leg of the bulb attachment. These were resolved during the final design task. Analyses of the final design of the omega segment liner indicated that all design goals were met and the design provided the capability of operating over the required test envelope with a life expectancy substantially above the goal of 1500 cycles.

Conceptual model of turbulent flameholding for scramjet combustors

NASA Technical Reports Server (NTRS)

Huber, P. W.

1980-01-01

New concepts and approaches to scramjet combustor design are presented. Blowoff was from failure of the recirculation-zone (RZ) flame to reach the dividing streamline (DS) at the rear stagnation zone. Increased turbulent exchange across the DS helped flameholding due to forward movement of the flame anchor point inside the RZ. Modeling of the blowoff phenomenon was based on a mass conservation concept involving the traverse of a flame element across the RZ and a flow element along the DS. The scale required to achieve flameholding, predicted by the model, showed a strong adverse effect of low pressure and low fuel equivalence ratio, moderate effect of flight Mach number, and little effect of temperature recovery factor. Possible effects of finite rate chemistry on flameholding and flamespreading in scramjets are discussed and recommendations for approaches to engine combustor design as well as for needed research to reduce uncertainties in the concepts are made.

Fostering creativity in product and service development: validation in the domain of information technology.

PubMed

Zeng, Liang; Proctor, Robert W; Salvendy, Gavriel

2011-06-01

This research is intended to empirically validate a general model of creative product and service development proposed in the literature. A current research gap inspired construction of a conceptual model to capture fundamental phases and pertinent facilitating metacognitive strategies in the creative design process. The model also depicts the mechanism by which design creativity affects consumer behavior. The validity and assets of this model have not yet been investigated. Four laboratory studies were conducted to demonstrate the value of the proposed cognitive phases and associated metacognitive strategies in the conceptual model. Realistic product and service design problems were used in creativity assessment to ensure ecological validity. Design creativity was enhanced by explicit problem analysis, whereby one formulates problems from different perspectives and at different levels of abstraction. Remote association in conceptual combination spawned more design creativity than did near association. Abstraction led to greater creativity in conducting conceptual expansion than did specificity, which induced mental fixation. Domain-specific knowledge and experience enhanced design creativity, indicating that design can be of a domain-specific nature. Design creativity added integrated value to products and services and positively influenced customer behavior. The validity and value of the proposed conceptual model is supported by empirical findings. The conceptual model of creative design could underpin future theory development. Propositions advanced in this article should provide insights and approaches to facilitate organizations pursuing product and service creativity to gain competitive advantage.

Liner cooling research at NASA Lewis Research Center. [for gas turbine combustion chambers

NASA Technical Reports Server (NTRS)

Acosta, Waldo A.

1987-01-01

Described are recently completed and current advanced liner research applicable to advanced small gas turbine engines. Research relating to the evolution of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently under way at NASA Lewis Research Center. As part of this research, a reverse-flow combustor geometry was maintained while different advanced liner wall cooling techniques were investigated and compared to a baseline combustor. The performance of the combustors featuring counterflow film-cooled (CFFC) panels, transpiration cooled liner walls (TRANS), and compliant metal/ceramic (CMC) walls was obtained over a range of simulated flight conditions of a 16:1 pressure ratio gas turbine engine and fuel/air ratios up to 0.034. All the combustors featured an identical fuel injection system, identical geometric configuration outline, and similar designed internal aerothermodynamics.

Two-stage combustion for reducing pollutant emissions from gas turbine combustors

NASA Technical Reports Server (NTRS)

Clayton, R. M.; Lewis, D. H.

1981-01-01

Combustion and emission results are presented for a premix combustor fueled with admixtures of JP5 with neat H2 and of JP5 with simulated partial-oxidation product gas. The combustor was operated with inlet-air state conditions typical of cruise power for high performance aviation engines. Ultralow NOx, CO and HC emissions and extended lean burning limits were achieved simultaneously. Laboratory scale studies of the non-catalyzed rich-burning characteristics of several paraffin-series hydrocarbon fuels and of JP5 showed sooting limits at equivalence ratios of about 2.0 and that in order to achieve very rich sootless burning it is necessary to premix the reactants thoroughly and to use high levels of air preheat. The application of two-stage combustion for the reduction of fuel NOx was reviewed. An experimental combustor designed and constructed for two-stage combustion experiments is described.

Opportunities for research in aerothermodynamics

NASA Technical Reports Server (NTRS)

Graham, R. W.

1983-01-01

"Aerothermodynamics' involves the disciplines of chemistry, thermodynamics, fluid mechanics and heat transfer which have collaborative importance in propulsion systems. There are growing opportunities for the further application of these disciplines to improve the methodology for the design of advanced gas turbines; particularly, the combustor and turbine. Design procedures follow empirical or cut and try guidelines. The tremendous advances in computational analysis and in instrumentation techniques hold promise for research answers to complex physical processes that are currently not well understood. The transfer of basic research understanding to engineering design should result in shorter, less expensive development commitments for engines. The status and anticipated opportunities in research topics relevant to combustors and turbines is reviewed.

An exploration of students' perceptions and attitudes towards creativity in engineering education

NASA Astrophysics Data System (ADS)

Waller, David R.

This study used a mixed methods approach to develop a broad and deep understanding of students’ perceptions towards creativity in engineering education. Studies have shown that students’ attitudes can have an impact on their motivation to engage in creative behavior. Using an ex-post facto independent factorial design, attitudes of value towards creativity, time for creativity, and creativity stereotypes were measured and compared across gender, year of study, engineering discipline, preference for open-ended problem solving, and confidence in creative abilities. Participants were undergraduate engineering students at Queen’s University from all years of study. A qualitative phenomenological methodology was adopted to study students’ understandings and experiences with engineering creativity. Eleven students participated in oneon- one interviews that provided depth and insight into how students experience and define engineering creativity, and the survey included open-ended items developed using the 10 Maxims of Creativity in Education as a guiding framework. The findings from the survey suggested that students had high value for creativity, however students in fourth year or higher had less value than those in other years. Those with preference for open-ended problem solving and high confidence valued creative more than their counterparts. Students who preferred open-ended problem solving and students with high confidence reported that time was less of a hindrance to their creativity. Males identified more with creativity stereotypes than females, however overall they were both low. Open-ended survey and interview results indicated that students felt they experienced creativity in engineering design activities. Engineering creativity definitions had two elements: creative action and creative characteristic. Creative actions were associated with designing, and creative characteristics were predominantly associated with novelty. Other barriers that emerged from the qualitative analysis were lack of opportunity, lack of assessment, and discomfort with creativity. It was concluded that a universal definition is required to establish clear and aligned understandings of engineering creativity. Instructors may want to consider demonstrating value by assessing creativity and establishing clear criteria in design projects. It is recommended that students be given more opportunities for practice through design activities and that they be introduced to design and creative thinking concepts early in their engineering education.

Stochastic modelling of turbulent combustion for design optimization of gas turbine combustors

NASA Astrophysics Data System (ADS)

Mehanna Ismail, Mohammed Ali

The present work covers the development and the implementation of an efficient algorithm for the design optimization of gas turbine combustors. The purpose is to explore the possibilities and indicate constructive suggestions for optimization techniques as alternative methods for designing gas turbine combustors. The algorithm is general to the extent that no constraints are imposed on the combustion phenomena or on the combustor configuration. The optimization problem is broken down into two elementary problems: the first is the optimum search algorithm, and the second is the turbulent combustion model used to determine the combustor performance parameters. These performance parameters constitute the objective and physical constraints in the optimization problem formulation. The examination of both turbulent combustion phenomena and the gas turbine design process suggests that the turbulent combustion model represents a crucial part of the optimization algorithm. The basic requirements needed for a turbulent combustion model to be successfully used in a practical optimization algorithm are discussed. In principle, the combustion model should comply with the conflicting requirements of high fidelity, robustness and computational efficiency. To that end, the problem of turbulent combustion is discussed and the current state of the art of turbulent combustion modelling is reviewed. According to this review, turbulent combustion models based on the composition PDF transport equation are found to be good candidates for application in the present context. However, these models are computationally expensive. To overcome this difficulty, two different models based on the composition PDF transport equation were developed: an improved Lagrangian Monte Carlo composition PDF algorithm and the generalized stochastic reactor model. Improvements in the Lagrangian Monte Carlo composition PDF model performance and its computational efficiency were achieved through the implementation of time splitting, variable stochastic fluid particle mass control, and a second order time accurate (predictor-corrector) scheme used for solving the stochastic differential equations governing the particles evolution. The model compared well against experimental data found in the literature for two different configurations: bluff body and swirl stabilized combustors. The generalized stochastic reactor is a newly developed model. This model relies on the generalization of the concept of the classical stochastic reactor theory in the sense that it accounts for both finite micro- and macro-mixing processes. (Abstract shortened by UMI.)

Combustion system CFD modeling at GE Aircraft Engines

NASA Technical Reports Server (NTRS)

Burrus, D.; Mongia, H.; Tolpadi, Anil K.; Correa, S.; Braaten, M.

1995-01-01

This viewgraph presentation discusses key features of current combustion system CFD modeling capabilities at GE Aircraft Engines provided by the CONCERT code; CONCERT development history; modeling applied for designing engine combustion systems; modeling applied to improve fundamental understanding; CONCERT3D results for current production combustors; CONCERT3D model of NASA/GE E3 combustor; HYBRID CONCERT CFD/Monte-Carlo modeling approach; and future modeling directions.

Combustion system CFD modeling at GE Aircraft Engines

NASA Astrophysics Data System (ADS)

Burrus, D.; Mongia, H.; Tolpadi, Anil K.; Correa, S.; Braaten, M.

1995-03-01

This viewgraph presentation discusses key features of current combustion system CFD modeling capabilities at GE Aircraft Engines provided by the CONCERT code; CONCERT development history; modeling applied for designing engine combustion systems; modeling applied to improve fundamental understanding; CONCERT3D results for current production combustors; CONCERT3D model of NASA/GE E3 combustor; HYBRID CONCERT CFD/Monte-Carlo modeling approach; and future modeling directions.

Fatigue life prediction of liquid rocket engine combustor with subscale test verification

NASA Astrophysics Data System (ADS)

Sung, In-Kyung

Reusable rocket systems such as the Space Shuttle introduced a new era in propulsion system design for economic feasibility. Practical reusable systems require an order of magnitude increase in life. To achieve this improved methods are needed to assess failure mechanisms and to predict life cycles of rocket combustor. A general goal of the research was to demonstrate the use of subscale rocket combustor prototype in a cost-effective test program. Life limiting factors and metal behaviors under repeated loads were surveyed and reviewed. The life prediction theories are presented, with an emphasis on studies that used subscale test hardware for model validation. From this review, low cycle fatigue (LCF) and creep-fatigue interaction (ratcheting) were identified as the main life limiting factors of the combustor. Several life prediction methods such as conventional and advanced viscoplastic models were used to predict life cycle due to low cycle thermal stress, transient effects, and creep rupture damage. Creep-fatigue interaction and cyclic hardening were also investigated. A prediction method based on 2D beam theory was modified using 3D plate deformation theory to provide an extended prediction method. For experimental validation two small scale annular plug nozzle thrusters were designed, built and tested. The test article was composed of a water-cooled liner, plug annular nozzle and 200 psia precombustor that used decomposed hydrogen peroxide as the oxidizer and JP-8 as the fuel. The first combustor was tested cyclically at the Advanced Propellants and Combustion Laboratory at Purdue University. Testing was stopped after 140 cycles due to an unpredicted failure mechanism due to an increasing hot spot in the location where failure was predicted. A second combustor was designed to avoid the previous failure, however, it was over pressurized and deformed beyond repair during cold-flow test. The test results are discussed and compared to the analytical and numerical predictions. A detailed comparison was not performed, however, due to the lack of test data resulting from a failure of the test article. Some theoretical and experimental aspects such as fin effect and round corner were found to reduce the discrepancy between prediction and test results.

Linkographic Evidence for Concurrent Divergent and Convergent Thinking in Creative Design

ERIC Educational Resources Information Center

Goldschmidt, Gabriela

2016-01-01

For a long time, the creativity literature has stressed the role of divergent thinking in creative endeavor. More recently, it has been recognized that convergent thinking also has a role in creativity, and the design literature, which sees design as a creative activity a priori, has largely adopted this view: Divergent and convergent thinking are…

Turbulent heat transfer prediction method for application to scramjet engines

NASA Technical Reports Server (NTRS)

Pinckney, S. Z.

1974-01-01

An integral method for predicting boundary layer development in turbulent flow regions on two-dimensional or axisymmetric bodies was developed. The method has the capability of approximating nonequilibrium velocity profiles as well as the local surface friction in the presence of a pressure gradient. An approach was developed for the problem of predicting the heat transfer in a turbulent boundary layer in the presence of a high pressure gradient. The solution was derived with particular emphasis on its applicability to supersonic combustion; thus, the effects of real gas flows were included. The resulting integrodifferential boundary layer method permits the estimation of cooling reguirements for scramjet engines. Theoretical heat transfer results are compared with experimental combustor and noncombustor heat transfer data. The heat transfer method was used in the development of engine design concepts which will produce an engine with reduced cooling requirements. The Langley scramjet engine module was designed by utilizing these design concepts and this engine design is discussed along with its corresponding cooling requirements. The heat transfer method was also used to develop a combustor cooling correlation for a combustor whose local properties are computed one dimensionally by assuming a linear area variation and a given heat release schedule.

An Idealized, Single Radial Swirler, Lean-Direct-Injection (LDI) Concept Meshing Script

NASA Technical Reports Server (NTRS)

Iannetti, Anthony C.; Thompson, Daniel

2008-01-01

To easily study combustor design parameters using computational fluid dynamics codes (CFD), a Gridgen Glyph-based macro (based on the Tcl scripting language) dubbed BladeMaker has been developed for the meshing of an idealized, single radial swirler, lean-direct-injection (LDI) combustor. BladeMaker is capable of taking in a number of parameters, such as blade width, blade tilt with respect to the perpendicular, swirler cup radius, and grid densities, and producing a three-dimensional meshed radial swirler with a can-annular (canned) combustor. This complex script produces a data format suitable for but not specific to the National Combustion Code (NCC), a state-of-the-art CFD code developed for reacting flow processes.

Emissions Prediction and Measurement for Liquid-Fueled TVC Combustor with and without Water Injection

NASA Technical Reports Server (NTRS)

Brankovic, A.; Ryder, R. C., Jr.; Hendricks, R. C.; Liu, N.-S.; Shouse, D. T.; Roquemore, W. M.

2005-01-01

An investigation is performed to evaluate the performance of a computational fluid dynamics (CFD) tool for the prediction of the reacting flow in a liquid-fueled combustor that uses water injection for control of pollutant emissions. The experiment consists of a multisector, liquid-fueled combustor rig operated at different inlet pressures and temperatures, and over a range of fuel/air and water/fuel ratios. Fuel can be injected directly into the main combustion airstream and into the cavities. Test rig performance is characterized by combustor exit quantities such as temperature and emissions measurements using rakes and overall pressure drop from upstream plenum to combustor exit. Visualization of the flame is performed using gray scale and color still photographs and high-frame-rate videos. CFD simulations are performed utilizing a methodology that includes computer-aided design (CAD) solid modeling of the geometry, parallel processing over networked computers, and graphical and quantitative post-processing. Physical models include liquid fuel droplet dynamics and evaporation, with combustion modeled using a hybrid finite-rate chemistry model developed for Jet-A fuel. CFD and experimental results are compared for cases with cavity-only fueling, while numerical studies of cavity and main fueling was also performed. Predicted and measured trends in combustor exit temperature, CO and NOx are in general agreement at the different water/fuel loading rates, although quantitative differences exist between the predictions and measurements.

Determination of decay coefficients for combustors with acoustic absorbers

NASA Technical Reports Server (NTRS)

Mitchell, C. E.; Espander, W. R.; Baer, M. R.

1972-01-01

An analytical technique for the calculation of linear decay coefficients in combustors with acoustic absorbers is presented. Tuned circumferential slot acoustic absorbers were designed for the first three transverse modes of oscillation, and decay coefficients for these absorbers were found as a function of backing distance for seven different chamber configurations. The effectiveness of the absorbers for off-design values of the combustion response and acoustic mode is also investigated. Results indicate that for tuned absorbers the decay coefficient increases approximately as the cube of the backing distance. For most off-design situations the absorber still provides a damping effect. However, if an absorber designed for some higher mode of oscillation is used to damp lower mode oscillations, a driving effect is frequently found.

Can Water-Injected Turbomachines Provide Cost-Effective Emissions and Maintenance Reductions?

NASA Technical Reports Server (NTRS)

Hendricks, Robert C.; Daggett, David L.; Shouse, Dale T.; Roquemore, William M.; Brankovic, Andreja; Ryder, Robert C., Jr.

2011-01-01

An investigation has been performed to evaluate the effect of water injection on the performance of the Air Force Research Laboratory (AFRL, Wright-Patterson Air Force Base (WPAFB)) experimental trapped vortex combustor (TVC) over a range of fuel-to-air and water-to-fuel ratios. Performance is characterized by combustor exit quantities: temperature and emissions measurements using rakes, and overall pressure drop, from upstream plenum to combustor exit. Combustor visualization is performed using gray-scale and color still photographs and high-frame-rate videos. A parallel investigation evaluated the performance of a computational fluid dynamics (CFD) tool for the prediction of the reacting flow in a liquid fueled combustor (e.g., TVC) that uses water injection for control of pollutant emissions and turbine inlet temperature. Generally, reasonable agreement is found between data and NO(x) computations. Based on a study assessing the feasibility and performance impact of using water injection on a Boeing 747-400 aircraft to reduce NO(x) emissions during takeoff, retrofitting does not appear to be cost effective; however, an operator of a newly designed engine and airframe might be able to save up to 1.0 percent in operating costs. Other challenges of water injection will be discussed.

Experimental clean combustor program, phase 1. [aircraft exhaust/gas analysis - gas turbine engines

NASA Technical Reports Server (NTRS)

Roberts, R.; Peduzzi, A.; Vitti, G. E.

1975-01-01

A program of screening three low emission combustors for conventional takeoff and landing, by testing and analyzing thirty-two configurations is presented. Configurations were tested that met the emission goals at idle operating conditions for carbon monoxide and for unburned hydrocarbons (emission index values of 20 and 4, respectively). Configurations were also tested that met a smoke number goal of 15 at sea-level take-off conditions. None of the configurations met the goal for oxides of nitrogen emissions at sea-level take-off conditions. The best configurations demonstrated oxide of nitrogen emission levels that were approximately 61 percent lower than those produced by the JT9D-7 engine, but these levels were still approximately 24 percent above the goal of an emission index level of 10. Additional combustor performance characteristics, including lean blowout, exit temperature pattern factor and radial profile, pressure loss, altitude stability, and altitude relight characteristics were documented. The results indicate the need for significant improvement in the altitude stability and relight characteristics. In addition to the basic program for current aircraft engine combustors, seventeen combustor configurations were evaluated for advanced supersonic technology applications. The configurations were tested at cruise conditions, and a conceptual design was evolved.

Performance characteristics of a slagging gasifier for MHD combustor systems

NASA Technical Reports Server (NTRS)

Smith, K. O.

1979-01-01

The performance of a two stage, coal combustor concept for magnetohydrodynamic (MHD) systems was investigated analytically. The two stage MHD combustor is comprised of an entrained flow, slagging gasifier as the first stage, and a gas phase reactor as the second stage. The first stage was modeled by assuming instantaneous coal devolatilization, and volatiles combustion and char gasification by CO2 and H2O in plug flow. The second stage combustor was modeled assuming adiabatic instantaneous gas phase reactions. Of primary interest was the dependence of char gasification efficiency on first stage particle residence time. The influence of first stage stoichiometry, heat loss, coal moisture, coal size distribution, and degree of coal devolatilization on gasifier performance and second stage exhaust temperature was determined. Performance predictions indicate that particle residence times on the order of 500 msec would be required to achieve gasification efficiencies in the range of 90 to 95 percent. The use of a finer coal size distribution significantly reduces the required gasifier residence time for acceptable levels of fuel use efficiency. Residence time requirements are also decreased by increased levels of coal devolatilization. Combustor design efforts should maximize devolatilization by minimizing mixing times associated with coal injection.

Optimization of Orifice Geometry for Cross-Flow Mixing in a Cylindrical Duct

NASA Technical Reports Server (NTRS)

Kroll, J. T.; Sowa, W. A.; Samuelsen, G. S.

1996-01-01

Mixing of gaseous jets in a cross-flow has significant applications in engineering, one example of which is the dilution zone of a gas turbine combustor. Despite years of study, the design of the jet injection in combustors is largely based on practical experience. The emergence of NO(x) regulations for stationary gas turbines and the anticipation of aero-engine regulations requires an improved understanding of jet mixing as new combustor concepts are introduced. For example, the success of the staged combustor to reduce the emission of NO(x) is almost entirely dependent upon the rapid and complete dilution of the rich zone products within the mixing section. It is these mixing challenges to which the present study is directed. A series of experiments was undertaken to delineate the optimal mixer orifice geometry. A cross-flow to core-flow momentum-flux ratio of 40 and a mass flow ratio of 2.5 were selected as representative of a conventional design. An experimental test matrix was designed around three variables: the number of orifices, the orifice length-to- width ratio, and the orifice angle. A regression analysis was performed on the data to arrive at an interpolating equation that predicted the mixing performance of orifice geometry combinations within the range of the test matrix parameters. Results indicate that the best mixing orifice geometry tested involves eight orifices with a long-to-short side aspect ratio of 3.5 at a twenty-three degree inclination from the center-line of the mixing section.

One-Dimensional Spontaneous Raman Measurements of Temperature Made in a Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

Hicks, Yolanda R.; Locke, Randy J.; DeGroot, Wilhelmus A.; Anderson, Robert C.

2002-01-01

The NASA Glenn Research Center is working with the aeronautics industry to develop highly fuel-efficient and environmentally friendly gas turbine combustor technology. This effort includes testing new hardware designs at conditions that simulate the high-temperature, high-pressure environment expected in the next-generation of high-performance engines. Glenn has the only facilities in which such tests can be performed. One aspect of these tests is the use of nonintrusive optical and laser diagnostics to measure combustion species concentration, fuel/air ratio, fuel drop size, and velocity, and to visualize the fuel injector spray pattern and some combustion species distributions. These data not only help designers to determine the efficacy of specific designs, but provide a database for computer modelers and enhance our understanding of the many processes that take place within a combustor. Until recently, we lacked one critical capability, the ability to measure temperature. This article summarizes our latest developments in that area. Recently, we demonstrated the first-ever use of spontaneous Raman scattering to measure combustion temperatures within the Advanced Subsonics Combustion Rig (ASCR) sector rig. We also established the highest rig pressure ever achieved for a continuous-flow combustor facility, 54.4 bar. The ASCR facility can provide operating pressures from 1 to 60 bar (60 atm). This photograph shows the Raman system setup next to the ASCR rig. The test was performed using a NASA-concept fuel injector and Jet-A fuel over a range of air inlet temperatures, pressures, and fuel/air ratios.

Pulse Detonation Engine Test Bed Developed

NASA Technical Reports Server (NTRS)

Breisacher, Kevin J.

2002-01-01

A detonation is a supersonic combustion wave. A Pulse Detonation Engine (PDE) repetitively creates a series of detonation waves to take advantage of rapid burning and high peak pressures to efficiently produce thrust. NASA Glenn Research Center's Combustion Branch has developed a PDE test bed that can reproduce the operating conditions that might be encountered in an actual engine. It allows the rapid and cost-efficient evaluation of the technical issues and technologies associated with these engines. The test bed is modular in design. It consists of various length sections of both 2- and 2.6- in. internal-diameter combustor tubes. These tubes can be bolted together to create a variety of combustor configurations. A series of bosses allow instrumentation to be inserted on the tubes. Dynamic pressure sensors and heat flux gauges have been used to characterize the performance of the test bed. The PDE test bed is designed to utilize an existing calorimeter (for heat load measurement) and windowed (for optical access) combustor sections. It uses hydrogen as the fuel, and oxygen and nitrogen are mixed to simulate air. An electronic controller is used to open the hydrogen and air valves (or a continuous flow of air is used) and to fire the spark at the appropriate times. Scheduled tests on the test bed include an evaluation of the pumping ability of the train of detonation waves for use in an ejector and an evaluation of the pollutants formed in a PDE combustor. Glenn's Combustion Branch uses the National Combustor Code (NCC) to perform numerical analyses of PDE's as well as to evaluate alternative detonative combustion devices. Pulse Detonation Engine testbed.

A numerical study of mixing in supersonic combustors with hypermixing injectors

NASA Technical Reports Server (NTRS)

Lee, J.

1993-01-01

A numerical study was conducted to evaluate the performance of wall mounted fuel-injectors designed for potential Supersonic Combustion Ramjet (SCRAM-jet) engine applications. The focus of this investigation was to numerically simulate existing combustor designs for the purpose of validating the numerical technique and the physical models developed. Three different injector designs of varying complexity were studied to fully understand the computational implications involved in accurate predictions. A dual transverse injection system and two streamwise injector designs were studied. The streamwise injectors were designed with swept ramps to enhance fuel-air mixing and combustion characteristics at supersonic speeds without the large flow blockage and drag contribution of the transverse injection system. For this study, the Mass-Average Navier-Stokes equations and the chemical species continuity equations were solved. The computations were performed using a finite-volume implicit numerical technique and multiple block structured grid system. The interfaces of the multiple block structured grid systems were numerically resolved using the flux-conservative technique. Detailed comparisons between the computations and existing experimental data are presented. These comparisons show that numerical predictions are in agreement with the experimental data. These comparisons also show that a number of turbulence model improvements are needed for accurate combustor flowfield predictions.

A numerical study of mixing in supersonic combustors with hypermixing injectors

NASA Technical Reports Server (NTRS)

Lee, J.

1992-01-01

A numerical study was conducted to evaluate the performance of wall mounted fuel-injectors designed for potential Supersonic Combustion Ramjet (SCRAM-jet) engine applications. The focus of this investigation was to numerically simulate existing combustor designs for the purpose of validating the numerical technique and the physical models developed. Three different injector designs of varying complexity were studied to fully understand the computational implications involved in accurate predictions. A dual transverse injection system and two streamwise injector designs were studied. The streamwise injectors were designed with swept ramps to enhance fuel-air mixing and combustion characteristics at supersonic speeds without the large flow blockage and drag contribution of the transverse injection system. For this study, the Mass-Averaged Navier-Stokes equations and the chemical species continuity equations were solved. The computations were performed using a finite-volume implicit numerical technique and multiple block structured grid system. The interfaces of the multiple block structured grid systems were numerically resolved using the flux-conservative technique. Detailed comparisons between the computations and existing experimental data are presented. These comparisons show that numerical predictions are in agreement with the experimental data. These comparisons also show that a number of turbulence model improvements are needed for accurate combustor flowfield predictions.

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

NASA Technical Reports Server (NTRS)

Lohmann, R. P.; Szetela, E. J.; Vranos, A.

1978-01-01

The impact of the use of broad specification fuels on the design, performance durability, emissions and operational characteristics of combustors for commercial aircraft gas turbine engines was assessed. Single stage, vorbix and lean premixed prevaporized combustors, in the JT9D and an advanced energy efficient engine cycle were evaluated when operating on Jet A and ERBS (Experimental Referee Broad Specification) fuels. Design modifications, based on criteria evolved from a literature survey, were introduced and their effectiveness at offsetting projected deficiencies resulting from the use of ERBS was estimated. The results indicate that the use of a broad specification fuel such as ERBS, will necessitate significant technology improvements and redesign if deteriorated performance, durability and emissions are to be avoided. Higher radiant heat loads are projected to seriously compromise liner life while the reduced thermal stability of ERBS will require revisions to the engine-airframe fuel system to reduce the thermal stress on the fuel. Smoke and emissions output are projected to increase with the use of broad specification fuels. While the basic geometry of the single stage and vorbix combustors are compatible with the use of ERBS, extensive redesign of the front end of the lean premixed prevaporized burner will be required to achieve satisfactory operation and optimum emissions.

Non-Intrusive Laser-Induced Imaging for Speciation and Patternation in High Pressure Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Zaller, Michelle M.; Hicks, Yolanda R.; Anderson, Robert C.

1999-01-01

The next generation of was turbine combustors for aerospace applications will be required to meet increasingly stringent constraints on fuel efficiency, noise abatement, and emissions. The power plants being designed to meet these constraints will operate at extreme conditions of temperature and pressure, thereby generating unique challenges to the previously employed diagnostic methodologies. Current efforts at NASA Glenn Research Center (GRC) utilize optically accessible, high pressure flametubes and sector combustor rigs to probe, via advanced nonintrusive laser techniques, the complex flowfields encountered in advanced combustor designs. The fuel-air mixing process is of particular concern for lowering NO(x) emissions generated in lean, premixed engine concepts. Using planar laser-induced fluorescence (PLIF) we have obtained real-time, detailed imaging of the fuel spray distribution for a number of fuel injector over a wide range of operational conditions that closely match those expected in the proposed propulsion systems. Using a novel combination of planar imaging, of fuel fluorescence and computational analysis that allows an examination of the flowfield from any perspective, we have produced spatially and temporally resolved fuel-air distribution maps. These maps provide detailed insight into the fuel injection at actual conditions never before possible, thereby greatly enhancing the evaluation of fuel injector performance and combustion phenomena.

Nonintrusive laser-induced imaging for speciation and patternation in high-pressure gas turbine combustors

NASA Astrophysics Data System (ADS)

Locke, Randy J.; Zaller, Michelle M.; Hicks, Yolanda R.; Anderson, Robert C.

1999-10-01

The next generation of ga turbine combustors for aerospace applications will be required to meet increasingly stringent constraints on fuel efficiency, noise abatement, and emissions. The power plants being designed to meet these constraints will operate at extreme conditions of temperature and pressure, thereby generating unique challenges to the previously employed diagnostic methodologies. Current efforts at NASA Glenn Research Center GRC utilize optically accessible, high-pressure flametubes and sector combustor rigs to probe, via advanced nonintrusive laser techniques, the complex flowfields encountered in advanced combustor designs. The fuel-air mixing process is of particular concern for lowering NOx emissions generated in lean, premixed engine concepts. Using planar laser-induced fluorescence we have obtained real- time, detailed imaging of the fuel spray distribution for a number of fuel injectors over a wide range of operational conditions that closely match those expected in the proposed propulsion systems. Using a novel combination of planar imaging of fuel fluorescence and computational analysis that allows an examination of the flowfield from any perspective, we have produced spatially and temporally resolved fuel-air distribution maps. These maps provide detailed insight into the fuel injection process at actual conditions never before possible, thereby greatly enhancing the evaluation of fuel injector performance and combustion phenomena.

Design verification and cold-flow modeling test report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Not Available

1993-07-01

This report presents a compilation of the following three test reports prepared by TRW for Alaska Industrial Development and Export Authority (AIDEA) as part of the Healy Clean Coal Project, Phase 1 Design of the TRW Combustor and Auxiliary Systems, which is co-sponsored by the Department of Energy under the Clean Coal Technology 3 Program: (1) Design Verification Test Report, dated April 1993, (2) Combustor Cold Flow Model Report, dated August 28, 1992, (3) Coal Feed System Cold Flow Model Report, October 28, 1992. In this compilation, these three reports are included in one volume consisting of three parts, andmore » TRW proprietary information has been excluded.« less

Creativity as a Key Driver for Designing Context Sensitive Health Informatics.

PubMed

Zhou, Chunfang; Nøhr, Christian

2017-01-01

In order to face the increasing challenges of complexity and uncertainty in practice of health care, this paper aims to discuss how creativity can contribute to design new technologies in health informatics systems. It will firstly introduce the background highlighting creativity as a missing element in recent studies on context sensitive health informatics. Secondly, the concept of creativity and its relationship with activities of technology design will be discussed from a socio-culture perspective. This will be thirdly followed by understanding the roles of creativity in designing new health informatics technologies for meeting needs of high context sensitivity. Finally, a series of potential strategies will be suggested to improve creativity among technology designers working in healthcare industries. Briefly, this paper innovatively bridges two areas studies on creativity and context sensitive health informatics by issues of technology design that also indicates its important significances for future research.

Evaluation of a staged fuel combustor for turboprop engines

NASA Technical Reports Server (NTRS)

Verdouw, A. J.

1976-01-01

Proposed EPA emission regulations require emission reduction by 1979 for various gas turbine engine classes. Extensive combustion technology advancements are required to meet the proposed regulations. The T56 turboprop engine requires CO, UHC, and smoke reduction. A staged fuel combustor design was tested on a combustion rig to evaluate emission reduction potential in turboprop engines from fuel zoning. The can-type combustor has separately fueled-pilot and main combustion zones in series. The main zone fueling system was arranged for potential incorporation into the T56 with minor or no modifications to the basic engine. Three combustor variable geometry systems were incorporated to evaluate various airflow distributions. Emission results with fixed geometry operation met all proposed EPA regulations over the EPA LTO cycle. CO reduction was 82 percent, UHC reduction was 96 percent, and smoke reduction was 84 percent. NOx increased 14 percent over the LTO cycle. At high power, NOx reduction was 40 to 55 percent. This NOx reduction has potential application to stationary gas turbine powerplants which have different EPA regulations.

Utility gas turbine combustor viewing system: Volume 1, Conceptual design and initial field testing: Final report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Morey, W.W.

1988-12-01

This report summarizes the development and field testing of a combustor viewing probe (CVP) as a flame diagnostic monitor for utility gas turbine engines. The prototype system is capable of providing a visual record of combustor flame images, recording flame spectral data, analyzing image and spectral data, and diagnosing certain engine malfunctions. The system should provide useful diagnostic information to utility plant operators, and reduce maintenance costs. The field tests demonstrated the ability of the CVP to monitor combustor flame condition and to relate changes in the engine operation with variations in the flame signature. Engine light off, run upmore » to full speed, the addition of load, and the effect of water injection for NO/sub x/ control could easily be identified on the video monitor. The viewing probe was also valuable in identifying hard startups and shutdowns, as well as transient effects that can seriously harm the engine. 11 refs.« less

Lean, premixed, prevaporized fuel combustor conceptual design study

NASA Technical Reports Server (NTRS)

Fiorentino, A. J.; Greene, W.; Kim, J.

1979-01-01

Four combustor concepts, designed for the energy efficient engine, utilize variable geometry or other flow modulation techniques to control the equivalence ratio of the initial burning zone. Lean conditions are maintained at high power to control oxides of nitrogen while near stoichometric conditions are maintained at low power for low CO and THC emissions. Each concept was analyzed and ranked for its potential in meeting the goals of the program. Although the primary goal of the program is a low level of nitric oxide emissions at stratospheric cruise conditions, both the ground level EPA emission standards and combustor performance and operational requirements typical of advanced subsonic aircraft engines are retained as goals as well. Based on the analytical projections made, two of the concepts offer the potential of achieving the emission goals; however, the projected operational characteristics and reliability of any concept to perform satisfactorily over an entire aircraft flight envelope would require extensive experimental substantiation before engine adaptation can be considered.

Analytical fuel property effects, small combustors, phase 1

NASA Technical Reports Server (NTRS)

Cohen, J. D.

1983-01-01

The effects of nonstandard aviation fuels on a typical small gas turbine combustor was analyzed. The T700/CT7 engine family was chosen as being representative of the class of aircraft power plants desired. Fuel properties, as specified by NASA, are characterized by low hydrogen content and high aromatics levels. Higher than normal smoke output and flame radiation intensity for the current T700 combustor which serves as a baseline were anticipated. It is, therefore, predicted that out of specification smoke visibility and higher than normal shell temperatures will exist when using NASA ERBS fuels with a consequence of severe reduction in cyclic life. Three new designs are proposed to compensate for the deficiencies expected with the existing design. They have emerged as the best of the eight originally proposed redesigns or combinations thereof. After the five choices that were originally made by NASA on the basis of competing performance factors, General Electric narrowed the field to the three proposed.

SiC Recession Due to SiO2 Scale Volatility Under Combustor Conditions

NASA Technical Reports Server (NTRS)

Robinson, Raymond Craig

1997-01-01

One of today's most important and challenging technological problems is the development of advanced materials and processes required to design and build a fleet of supersonic High Speed Civil Transport (HSCT) airliners, a follow-up to the Concorde SST. The innovative combustor designs required for HSCT engines will need high-temperature materials with long-term environmental stability. Higher combustor liner temperatures than today's engines and the need for lightweight materials will require the use of advanced ceramic-matrix composites (CMC's) in hot-section components. The HSCT is just one example being used to demonstrate the need for such materials. This thesis evaluates silicon carbide (SiC) as a potential base material for HSCT and other similar applications. Key issues are the environmental durability for the materials of interest. One of the leading combustor design schemes leads to an environment which will contain both oxidizing and reducing gas mixtures. The concern is that these environments may affect the stability of the silica (SiO2) scale on which SiC depends for environmental protection. A unique High Pressure Burner Rig (HPBR) was developed to simulate the combustor conditions of future gas turbine engines, and a series of tests were conducted on commercially available SiC material. These tests are intended as a feasibility study for the use of these materials in applications such as the HSCT. Linear weight loss and surface recession of the SiC is observed as a result of SiO2 volatility for both fuel-lean and fuel-rich gas mixtures. These observations are compared and agree well with thermogravimetric analysis (TGA) experiments. A strong Arrhenius-type temperature dependence exists. In addition, the secondary dependencies of pressure and gas velocity are defined. As a result, a model is developed to enable extrapolation to points outside the experimental space of the burner rig, and in particular, to potential gas turbine engine conditions.

Laboratory studies of lean combustion

NASA Technical Reports Server (NTRS)

Sawyer, R. F.; Schefer, R. W.; Ganji, A. R.; Daily, J. W.; Pitz, R. W.; Oppenheim, A. K.; Angeli, J. W.

1977-01-01

The fundamental processes controlling lean combustion were observed for better understanding, with particular emphasis on the formation and measurement of gas-phase pollutants, the stability of the combustion process (blowout limits), methods of improving stability, and the application of probe and optical diagnostics for flow field characterization, temperature mapping, and composition measurements. The following areas of investigation are described in detail: (1) axisymmetric, opposed-reacting-jet-stabilized combustor studies; (2) stabilization through heat recirculation; (3) two dimensional combustor studies; and (4) spectroscopic methods. A departure from conventional combustor design to a premixed/prevaporized, lean combustion configuration is attractive for the control of oxides of nitrogen and smoke emissions, the promotion of uniform turbine inlet temperatures, and, possibly, the reduction of carbon monoxide and hydrocarbons at idle.

A First Look at the DGEN380 Engine Acoustic Data from a Core-Noise Perspective

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.

2015-01-01

This work is a first look at acoustic data acquired in the NASA Glenn Research Center Aero-Acoustic Propulsion Laboratory using the Price Induction DGEN380 small turbofan engine, with particular emphasis on broadband combustor (core) noise. Combustor noise is detected by using a two-signal source separation technique employing one engine-internal sensor and one semi-far-field microphone. Combustor noise is an important core-noise component and is likely to become a more prominent contributor to overall airport community noise due to turbofan design trends, expected aircraft configuration changes, and advances in fan-noise-mitigation techniques. This work was carried out under the NASA Fundamental Aeronautics Program, Fixed Wing Project, Quiet Performance Subproject

Core Noise - Increasing Importance

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.

2011-01-01

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 Subsonic Fixed Wing (SFW) Project. Sections of the presentation cover: the SFW system-level noise metrics for the 2015, 2020, and 2025 timeframes; turbofan design trends and their aeroacoustic implications; the emerging importance of core noise and its relevance to the SFW Reduced-Perceived-Noise Technical Challenge; and the current research activities in the core-noise area, with additional details given about the development of a high-fidelity combustor-noise prediction capability as well as activities supporting the development of improved reduced-order, physics-based models for combustor-noise prediction. The need for benchmark data for validation of high-fidelity and modeling work and the value of a potential future diagnostic facility for testing of core-noise-reduction concepts are indicated. The NASA Fundamental Aeronautics Program has the principal objective of overcoming today's national challenges in air transportation. The SFW Reduced-Perceived-Noise Technical Challenge aims to develop concepts and technologies to dramatically reduce the perceived aircraft noise outside of airport boundaries. This reduction of aircraft noise is critical to enabling the anticipated large increase in future air traffic. Noise generated in the jet engine core, by sources such as the compressor, combustor, and turbine, can be a significant contribution to the overall noise signature at low-power conditions, typical of approach flight. At high engine power during takeoff, jet and fan noise have traditionally dominated over core noise. However, current design trends and expected technological advances in engine-cycle design as well as noise-reduction methods are likely to reduce non-core noise even at engine-power points higher than approach. In addition, future low-emission combustor designs could increase the combustion-noise component. The trend towards high-power-density cores also means that the noise generated in the low-pressure turbine will likely increase. Consequently, the combined result from these emerging changes will be to elevate the overall importance of turbomachinery core noise, which will need to be addressed in order to meet future noise goals.

Time Pressure and Creativity in Industrial Design

ERIC Educational Resources Information Center

Hsiao, Shih-Wen; Wang, Ming-Feng; Chen, Chien-Wie

2017-01-01

Creativity is a critical aspect of competitiveness in all trades and professions. In the case of designers, creativity is of the utmost importance. Based on the perspective of industrial design, the relationship between creativity and time pressure was investigated in this study using control and experimental groups. In the first part of the…

Teaching Art and Design: Communicating Creative Practice through Embodied and Tacit Knowledge

ERIC Educational Resources Information Center

Budge, Kylie

2016-01-01

How do artists and designers teaching in universities communicate creative practice as they teach art/design? There is much discussion about the "mystery" of creativity, but little understanding of how teaching occurs in creative contexts. Understanding this topic better will develop greater knowledge within the academy of how art and…

High-Flow Jet Exit Rig Designed and Fabricated

NASA Technical Reports Server (NTRS)

Buehrle, Robert J.; Trimarchi, Paul A.

2003-01-01

The High-Flow Jet Exit Rig at the NASA Glenn Research Center is designed to test single flow jet nozzles and to measure the appropriate thrust and noise levels. The rig has been designed for the maximum hot condition of 16 lbm/sec of combustion air at 1960 R (maximum) and to produce a maximum thrust of 2000 lb. It was designed for cold flow of 29.1 lbm/sec of air at 530 R. In addition, it can test dual-flow nozzles (nozzles with bypass flow in addition to core flow) with independent control of each flow. The High- Flow Jet Exit Rig was successfully fabricated in late 2001 and is being readied for checkout tests. The rig will be installed in Glenn's Aeroacoustic Propulsion Laboratory. The High-Flow Jet Exit Rig consists of the following major components: a single component force balance, the natural-gas-fueled J-79 combustor assembly, the plenum and manifold assembly, an acoustic/instrumentation/seeding (A/I/S) section, a table, and the research nozzles. The rig will be unique in that it is designed to operate uncooled. The structure survives the 1960 R test condition because it uses carefully selected high temperature alloy materials such as Hastelloy-X. The lower plenum assembly was designed to operate at pressures to 450 psig at 1960 R, in accordance with the ASME B31.3 piping code. The natural gas-fueled combustor fires directly into the lower manifold. The hot air is directed through eight 1-1/2-in. supply pipes that supply the upper plenum. The flow is conditioned in the upper plenum prior to flowing to the research nozzle. The 1-1/2-in. supply lines are arranged in a U-shaped design to provide for a flexible piping system. The combustor assembly checkout was successfully conducted in Glenn's Engine Component Research Laboratory in the spring of 2001. The combustor is a low-smoke version of the J79 combustor used to power the F4 Phantom military aircraft. The natural gas-fueled combustor demonstrated high-efficiency combustion over a wide range of operating conditions. This wide operating envelope is required to support the testing of both single- and dual-flow nozzles. Key research goals include providing simultaneous, highly accurate acoustic, flow, and thrust measurements on jet nozzle models in realistic flight conditions, as well as providing scaleable acoustic results. The High-Flow Jet Exit Rig is a second-generation high-flow test rig. Improvements include cleaner flow with reduced levels of particulate, soot, and odor. Choked-flow metering is required with plus or minus 0.25-percent accuracy. Thrust measurements from 0 to 2000 lbf are required with plus or minus 0.25-percent accuracy. Improved acoustics will be achieved by minimizing noise through large pipe bend radii, lower internal flow velocities, and microdrilled choke plates with thousands of 0.040-in.- diameter holes.

Collaboration within Student Design Teams Participating in Architectural Design Competitions

ERIC Educational Resources Information Center

Erbil, Livanur; Dogan, Fehmi

2012-01-01

This paper investigates design collaboration with reference to convergent and divergent idea generation processes in architectural design teams entering a design competition. Study of design teams offer a unique opportunity to investigate how creativity is fostered through collaborative work. While views of creativity often relate creativity to…

Study on Product Innovative Design Process Driven by Ideal Solution

NASA Astrophysics Data System (ADS)

Zhang, Fuying; Lu, Ximei; Wang, Ping; Liu, Hui

Product innovative design in companies today relies heavily on individual members’ experience and creative ideation as well as their skills of integrating creativity and innovation tools with design methods agilely. Creative ideation and inventive ideas generation are two crucial stages in product innovative design process. Ideal solution is the desire final ideas for given problem, and the striving reaching target for product design. In this paper, a product innovative design process driven by ideal solution is proposed. This design process encourages designers to overcome their psychological inertia, to foster creativity in a systematic way for acquiring breakthrough creative and innovative solutions in a reducing sphere of solution-seeking, and results in effective product innovative design rapidly. A case study example is also presented to illustrate the effectiveness of the proposed design process.

Heavy hydrocarbon main injector technology program

NASA Technical Reports Server (NTRS)

Arbit, H. A.; Tuegel, L. M.; Dodd, F. E.

1991-01-01

The Heavy Hydrocarbon Main Injector Program was an analytical, design, and test program to demonstrate an injection concept applicable to an Isolated Combustion Compartment of a full-scale, high pressure, LOX/RP-1 engine. Several injector patterns were tested in a 3.4-in. combustor. Based on these results, features of the most promising injector design were incorporated into a 5.7-in. injector which was then hot-fire tested. In turn, a preliminary design of a 5-compartment 2D combustor was based on this pattern. Also the additional subscale injector testing and analysis was performed with an emphasis on improving analytical techniques and acoustic cavity design methodology. Several of the existing 3.5-in. diameter injectors were hot-fire tested with and without acoustic cavities for spontaneous and dynamic stability characteristics.

Start-up and Self-sustain Test of 500 W Ultra-Micro Gas Turbine Generator

NASA Astrophysics Data System (ADS)

Seo, Jeong Min; Park, Jun Young; Seog Choi, Bum

2013-12-01

This paper provides the performance test for start-up and self-sustaining of 500W ultra-micro gas turbine (UMGT) generator. Each component of UMGT, a centrifugal compressor, a radial turbine, an annular combustor and a shaft is already designed, manufactured and tested to meet design requirements in previous researches. However, they are not tested to work in an integrate system. Currently, integrated test unit with a compressor, a combustor and a turbine, is developed to find the proper condition of start-up and self-sustain. Ignition sequence depending on rotating speed is designed. Performance test for start-up and self-sustain is designed based on the ignition possible condition. An air impingement starter and a hot bulb inginer are applied. LPG is used as main fuel.

Making Space for the Act of Making: Creativity in the Engineering Design Classroom

ERIC Educational Resources Information Center

Lasky, Dorothea; Yoon, Susan A.

2011-01-01

Creativity continues to be an important goal for 21st century learning. However, teachers often have difficulties fostering creativity in their classrooms. Current creativity research suggests that the act of making can enhance the teaching of creativity. Hands-on engineering design lessons are ideal contexts for studying this effect. Through…

Three Dimensional CFD Analysis of the GTX Combustor

NASA Technical Reports Server (NTRS)

Steffen, C. J., Jr.; Bond, R. B.; Edwards, J. R.

2002-01-01

The annular combustor geometry of a combined-cycle engine has been analyzed with three-dimensional computational fluid dynamics. Both subsonic combustion and supersonic combustion flowfields have been simulated. The subsonic combustion analysis was executed in conjunction with a direct-connect test rig. Two cold-flow and one hot-flow results are presented. The simulations compare favorably with the test data for the two cold flow calculations; the hot-flow data was not yet available. The hot-flow simulation indicates that the conventional ejector-ramjet cycle would not provide adequate mixing at the conditions tested. The supersonic combustion ramjet flowfield was simulated with frozen chemistry model. A five-parameter test matrix was specified, according to statistical design-of-experiments theory. Twenty-seven separate simulations were used to assemble surrogate models for combustor mixing efficiency and total pressure recovery. ScramJet injector design parameters (injector angle, location, and fuel split) as well as mission variables (total fuel massflow and freestream Mach number) were included in the analysis. A promising injector design has been identified that provides good mixing characteristics with low total pressure losses. The surrogate models can be used to develop performance maps of different injector designs. Several complex three-way variable interactions appear within the dataset that are not adequately resolved with the current statistical analysis.

Art and Design Blogs: A Socially-Wise Approach to Creativity

ERIC Educational Resources Information Center

Budge, Kylie

2012-01-01

Many images of the "artist" or "designer" pervade the media and popular consciousness. Contemporary images of the artist and creativity that focus solely on the individual offer a very narrow depiction of the varying ways creativity occurs for artists and designers. These images do not capture the variety of creative processes and myriad ways…

Numerical study of innovative scramjet inlets coupled to combustors using hydrocarbon-air mixture

NASA Astrophysics Data System (ADS)

Malo-Molina, Faure Joel

The research objective is to use high-fidelity multi-physics Computational Fluid Dynamics (CFD) analysis to characterize 3-D scramjet flowfields in two novel streamline traced circular configurations without axisymmetric profiles. This work builds on a body of research conducted over the past several years. In addition, this research provides the modeling and simulation support, prior to ground (wind tunnel) and flight experiment programs. Two innovative inlets, Jaws and Scoop, are analyzed and compared to a Baseline inlet, a current state of the art rectangular inlet used as a baseline for on/off-design conditions. The flight trajectory conditions selected were Mach 6 and a dynamic pressure of 1,500 psf (71.82 kPa), corresponding to a static pressure of 43.7 psf (2.09 kPa) and temperature of 400.8 R° (222.67 C°). All inlets are designed for equal flight conditions, equal contraction ratios and exit cross-sectional areas, thus facilitating their comparison and integration to a common combustor design. Analysis of these hypersonic inlets was performed to investigate distortion effects downstream in common generic combustors. These combustors include a single cavity acting as flame holder and strategically positioned fuel injection ports. This research not only seeks to identify the most successful integrated scramjet inlet/combustor design, but also investigates the flow physics and quantifies the integrated performance impact of the two novel scramjet inlet designs. It contributes to the hypersonic air-breathing community by providing analysis and predictions on directly-coupled combustor numerical experiments for developing pioneering inlets or nozzles for scramjets. Several validations and verifications of General Propulsion Analysis Chemical-kinetic and Two-phase (GPACT), the CFD tool, were conducted throughout the research. In addition, this study uses 13 gaseous species and 20 reactions for an Ethylene/air finite-rate chemical model. The key conclusions of this research are: (1) Flow distortion in the innovative inlets is similar to some of the distortion in the Baseline inlet, despite design differences. In both innovative inlets, the resulting flowfield distortions were due to shock boundary layer interactions similar to those found in the Baseline. The Baseline and Jaws performance attributes are stronger than Scoop, but Jaws accomplishes this while eradicating the cowl lip interaction, and lessening the total drag and spillage penalties. (2) The innovative inlets work best on-design, whereas for off-design, the traditional inlet yields a higher performance. Although the innovative inlets' designs mitigated some of the issues encountered in traditional configurations, they underperform at off-design conditions. The strategy used in Jaws was less prone to interaction with the near wall flow, and yields lesser pressure losses and higher efficiency at on-design conditions compared to the others. In general, the overall values for Scoop seem lowest of all due to lesser entrainment. Its drag coefficient and thrust to mass capture ratios are higher than the Baseline configuration. (3) Early pressure losses and flow distortions actually aid downstream combustion in all cases. Although interactions captured by the viscous simulations for the on-design conditions increase losses in the inlets, they enhance turbulence in the isolator, favoring the mixing of air and fuel, and improving the overall factor of the system. Jaws inlet demonstrates the most valuable design with higher performance, but its factor later in the combustor drops relative to its rectangular counterpart. (4) A parametric study of the location and direction of injection is conducted to select the configuration for fuel penetration, mixing factor (factor) and other combustion qualities. Although the trends observed with and without chemical reactions are the same, the former yields roughly 10% higher mixing factor. Unlike at frozen conditions, when chemical reactions are considered, a high compression area was observed upstream of the cavity, not present when modeling Jaws. The upstream reactions from the cavity have a significant impact on the development of the shear layers and downstream development of the entire combustion. (5) Steady and unsteady simulations are conducted to characterize the ignition process, flame anchoring and flashback effects. This unsteadiness enlarges the circulation region in and around the cavity, allowing the reactions to propagate forward through the shear layer, and increases the mixing factor. In Scoop, these effects are exacerbated due to the thicker low energy profile surrounding the walls and most of the lower section of the combustor. In the steady assumptions, the forward reactions and their effects are positioned farthest upstream, closest to the combustor entrance. (6) Unsteady Reynolds Average Navier-Stokes (URANS) and Large Eddy Simulation (LES) modeling are compared to explore overall flow structure and for comparison of individual numerical methods. In URANS, the flashback effects are midway between the entrance and the step, whereas in LES, this effect is near the edge of the step in addition to yielding a higher combustion factor. Thus, the turbulence model and inflow assumptions can critically affect the total outcome of such devices.

Aircraft Engine Systems

NASA Technical Reports Server (NTRS)

Veres, Joseph

2001-01-01

This report outlines the detailed simulation of Aircraft Turbofan Engine. The objectives were to develop a detailed flow model of a full turbofan engine that runs on parallel workstation clusters overnight and to develop an integrated system of codes for combustor design and analysis to enable significant reduction in design time and cost. The model will initially simulate the 3-D flow in the primary flow path including the flow and chemistry in the combustor, and ultimately result in a multidisciplinary model of the engine. The overnight 3-D simulation capability of the primary flow path in a complete engine will enable significant reduction in the design and development time of gas turbine engines. In addition, the NPSS (Numerical Propulsion System Simulation) multidisciplinary integration and analysis are discussed.

National Combustion Code Parallel Performance Enhancements

NASA Technical Reports Server (NTRS)

Quealy, Angela; Benyo, Theresa (Technical Monitor)

2002-01-01

The National Combustion Code (NCC) is being developed by an industry-government team for the design and analysis of combustion systems. The unstructured grid, reacting flow code uses a distributed memory, message passing model for its parallel implementation. The focus of the present effort has been to improve the performance of the NCC code to meet combustor designer requirements for model accuracy and analysis turnaround time. Improving the performance of this code contributes significantly to the overall reduction in time and cost of the combustor design cycle. This report describes recent parallel processing modifications to NCC that have improved the parallel scalability of the code, enabling a two hour turnaround for a 1.3 million element fully reacting combustion simulation on an SGI Origin 2000.

Core/Combustor Noise - Research Overview

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.

2017-01-01

Contributions from the combustor to the overall propulsion noise of civilian transport aircraft are starting to become important due to turbofan design trends and advances in mitigation of other noise sources. Future propulsion systems for ultra-efficient commercial air vehicles are projected to be of increasingly higher bypass ratio from larger fans combined with much smaller cores, with ultra-clean burning fuel-flexible combustors. Unless effective noise-reduction strategies are developed, combustor noise is likely to become a prominent contributor to overall airport community noise in the future. This presentation gives a brief overview of the NASA outlook on pertinent issues and far-term research needs as well as current and planned research in the core/combustor-noise area. The research described herein is aligned with the NASA Ultra-Efficient Commercial Transport strategic thrust and is supported by the NASA Advanced Air Vehicle Program, Advanced Air Transport Technology Project, under the Aircraft Noise Reduction Subproject. The overarching goal of the Advanced Air Transport Technology (AATT) Project is to explore and develop technologies and concepts to revolutionize the energy efficiency and environmental compatibility of fixed wing transport aircrafts. These technological solutions are critical in reducing the impact of aviation on the environment even as this industry and the corresponding global transportation system continue to grow.

Hot section viewing system

NASA Technical Reports Server (NTRS)

Morey, W. W.

1983-01-01

The objective of the hot section viewing program is to develop a prototype optical system for viewing the interior of a gas turbine combustor during high temperature, high pressure operation in order to produce a visual record of some causes of premature hot section failures. The program began by identifying and analyzing system designs that would provide clearest images while being able to survive the hostile environment inside the combustion chamber. Different illumination methods and computer techniques for image enhancement and analysis were examined during a preliminary test phase. In the final phase of the program the prototype system was designed and fabricated and is currently being tested on a high pressure combustor rig.

H2O2 space shuttle APU

NASA Technical Reports Server (NTRS)

1975-01-01

A cryogenic H2-O2 auxiliary power unit (APU) was developed and successfully demonstrated. It has potential application as a minimum weight alternate to the space shuttle baseline APU because of its (1) low specific propellant consumption and (2) heat sink capabilities that reduce the amount of expendable evaporants. A reference system was designed with the necessary heat exchangers, combustor, turbine-gearbox, valves, and electronic controls to provide 400 shp to two aircraft hydraulic pumps. Development testing was carried out first on the combustor and control valves. This was followed by development of the control subsystem including the controller, the hydrogen and oxygen control valves, the combustor, and a turbine simulator. The complete APU system was hot tested for 10 hr with ambient and cryogenic propellants. Demonstrated at 95 percent of design power was 2.25 lb/hp-hr. At 10 percent design power, specific propellant consumption was 4 lb/hp-hr with space simulated exhaust and 5.2 lb/hp-hr with ambient exhaust. A 10 percent specific propellant consumption improvement is possible with some seal modifications. It was demonstrated that APU power levels could be changed by several hundred horsepower in less than 100 msec without exceeding allowable turbine inlet temperatures or turbine speed.

Fostering Creativity in Design Education: Using the Creative Product Analysis Matrix with Chinese Undergraduates in Macau

ERIC Educational Resources Information Center

Tsai, Kuan Chen

2016-01-01

The purpose of the present study is to explore to what extent the use of a more structured mode of assessing creative products--specifically, the CPAM--could beneficially influence design students' product creativity and creative processes. For this qualitative inquiry, following our CPAM-based intervention, students wrote reflective papers in…

Optimizing the CSP Tower Air Brayton Cycle System to Meet the SunShot Objectives - Final Technical Report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Bryner, Elliott; Brun, Klaus; Coogan, Shane

The objective of this project is to increase Concentrated Solar Power (CSP) tower air receiver and gas turbine temperature capabilities to 1,000ºC by the development of a novel gas turbine combustor, which can be integrated on a megawatt-scale gas turbine, such as the Solar Turbines Mercury 50™. No combustor technology currently available is compatible with the CSP application target inlet air temperature of 1,000°C. Autoignition and flashback at this temperature prevent the use of conventional lean pre-mix injectors that are currently employed to manage NOx emissions. Additional challenges are introduced by the variability of the high-temperature heat source provided bymore » the field of solar collectors, the heliostat in CSP plants. For optimum energy generation from the power turbine, the turbine rotor inlet temperature (TRIT) should remain constant. As a result of changing heat load provided to the solar collector from the heliostat, the amount of energy input required from the combustion system must be adjusted to compensate. A novel multi-bank lean micro-mix injector has been designed and built to address the challenges of high-temperature combustion found in CSP applications. The multi-bank arrangement of the micro-mix injector selectively injects fuel to meet the heat addition requirements to maintain constant TRIT with changing solar load. To validate the design, operation, and performance of the multi-bank lean micro-mix injector, a novel combustion test facility has been designed and built at Southwest Research Institute® (SwRI®) in San Antonio, TX. This facility, located in the Turbomachinery Research Facility, provides in excess of two kilograms per second of compressed air at nearly eight bar pressure. A two-megawatt electric heater raises the inlet temperature to 800°C while a secondary gas-fired heater extends the operational temperature range of the facility to 1,000°C. A combustor test rig connected to the heater has been designed and built to test the multi-bank lean micro-mix injector over the range of CSP operating conditions. The fuel is controlled and selectively delivered to the banks of the injector based on combustor inlet conditions that correspond to turbine operating points. The combustor rig is equipped with a data acquisition system and a suite of instrumentation for measuring temperature, pressure, and species concentration. This unique test facility has been built and commissioned and a prototype of the multi-bank lean micro-mix injector design has been tested. Operation of the combustor and injector has been demonstrated over the full range of CSP inlet conditions and for the range of turbine load conditions specified. The multi-bank operation of the injector has been proven to be an effective design for managing the variable flow rates of air and fuel due to changing inlet conditions from the solar field and turbine loads.« less

Creativity Processes of Students in the Design Studio

ERIC Educational Resources Information Center

Huber, Amy Mattingly; Leigh, Katharine E.; Tremblay, Kenneth R., Jr.

2012-01-01

The creative process is a multifaceted and dynamic path of thinking required to execute a project in design-based disciplines. The goal of this research was to test a model outlining the creative design process by investigating student experiences in a design project assignment. The study used an exploratory design to collect data from student…

Is There Creativity in Design? From a Perspective of School Design and Technology in Hong Kong

ERIC Educational Resources Information Center

Wong, Yi Lin; Siu, Kin Wai Michael

2012-01-01

As creativity is likely to become a crucial aspect of living in the future, it is important for educators to teach students to think creatively when solving constantly evolving and increasingly complex problems. Supported by the idea that creativity can be taught and learnt, elements of creativity are now embedded in secondary school education.…

Exploring the Relationships among Creativity, Engineering Knowledge, and Design Team Interaction on Senior Engineering Design Projects

ERIC Educational Resources Information Center

Ibrahim, Badaruddin

2012-01-01

In the 21st century, engineers are expected to be creative and work collaboratively in teams to solve or design new products. Research in the past has shown how creativity and good team communication, together with knowledge, can impact the outcomes in the organization. The purpose of this study was to explore the relationships among creativity,…

DART Core/Combustor-Noise Initial Test Results

NASA Technical Reports Server (NTRS)

Boyle, Devin K.; Henderson, Brenda S.; Hultgren, Lennart S.

2017-01-01

Contributions from the combustor to the overall propulsion noise of civilian transport aircraft are starting to become important due to turbofan design trends and advances in mitigation of other noise sources. Future propulsion systems for ultra-efficient commercial air vehicles are projected to be of increasingly higher bypass ratio from larger fans combined with much smaller cores, with ultra-clean burning fuel-flexible combustors. Unless effective noise-reduction strategies are developed, combustor noise is likely to become a prominent contributor to overall airport community noise in the future. The new NASA DGEN Aero0propulsion Research Turbofan (DART) is a cost-efficient testbed for the study of core-noise physics and mitigation. This presentation gives a brief description of the recently completed DART core combustor-noise baseline test in the NASA GRC Aero-Acoustic Propulsion Laboratory (AAPL). Acoustic data was simultaneously acquired using the AAPL overhead microphone array in the engine aft quadrant far field, a single midfield microphone, and two semi-infinite-tube unsteady pressure sensors at the core-nozzle exit. An initial assessment shows that the data is of high quality and compares well with results from a quick 2014 feasibility test. Combustor noise components of measured total-noise signatures were educed using a two-signal source-separation method an dare found to occur in the expected frequency range. The research described herein is aligned with the NASA Ultra-Efficient Commercial Transport strategic thrust and is supported by the NASA Advanced Air Vehicle Program, Advanced Air Transport Technology Project, under the Aircraft Noise Reduction Subproject.

Analytical and experimental investigations of the oblique detonation wave engine concept

NASA Technical Reports Server (NTRS)

Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc

1990-01-01

Wave combustors, which include the oblique detonation wave engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using Computational Fluid Dynamics (CFD) codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being performed in an arc heated hypersonic wind tunnel. Several fuel injection design were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Low NOx heavy fuel combustor concept program, phase 1

NASA Technical Reports Server (NTRS)

Cutrone, M. B.

1981-01-01

Combustion tests were completed with seven concepts, including three rich/lean concepts, three lean/lean concepts, and one catalytic combustor concept. Testing was conducted with ERBS petroleum distillate, petroleum residual, and SRC-II coal-derived liquid fuels over a range of operating conditions for the 12:1 pressure ratio General Electric MS7001E heavy-duty turbine. Blends of ERBS and SRC-II fuels were used to vary fuel properties over a wide range. In addition, pyridine was added to the ERBS and residual fuels to vary nitrogen level while holding other fuel properties constant. Test results indicate that low levels of NOx and fuel-bound nitrogen conversion can be achieved with the rich/lean combustor concepts for fuels with nitrogen contents up to 1.0% by weight. Multinozzle rich/lean Concept 2 demonstrated dry low Nox emissions within 10-15% of the EPA New Source Performance Standards goals for SRC-II fuel, with yields of approximately 15%, while meeting program goals for combustion efficiency, pressure drop, and exhaust gas temperature profile. Similar, if not superior, potential was demonstrated by Concept 3, which is a promising rich/lean combustor design.

Analytical and experimental investigations of the oblique detonation wave engine concept

NASA Technical Reports Server (NTRS)

Menees, Gene P.; Adelman, Henry G.; Cambier, Jean-Luc

1991-01-01

Wave combustors, which include the Oblique Detonation Wave Engine (ODWE), are attractive propulsion concepts for hypersonic flight. These engines utilize oblique shock or detonation waves to rapidly mix, ignite, and combust the air-fuel mixture in thin zones in the combustion chamber. Benefits of these combustion systems include shorter and lighter engines which will require less cooling and can provide thrust at higher Mach numbers than conventional scramjets. The wave combustor's ability to operate at lower combustor inlet pressures may allow the vehicle to operate at lower dynamic pressures which could lessen the heating loads on the airframe. The research program at NASA-Ames includes analytical studies of the ODWE combustor using CFD codes which fully couple finite rate chemistry with fluid dynamics. In addition, experimental proof-of-concept studies are being carried out in an arc heated hypersonic wind tunnel. Several fuel injection designs were studied analytically and experimentally. In-stream strut fuel injectors were chosen to provide good mixing with minimal stagnation pressure losses. Measurements of flow field properties behind the oblique wave are compared to analytical predictions.

Component testing of a ground based gas turbine steam cooled rich-burn primary zone combustor for emissions control of nitrogeneous fuels

NASA Technical Reports Server (NTRS)

Schultz, D. F.

1986-01-01

This effort summarizes the work performed on a steam cooled, rich-burn primary zone, variable geometry combustor designed for combustion of nitrogeneous fuels such as heavy oils or synthetic crude oils. The steam cooling was employed to determine its feasibility and assess its usefulness as part of a ground based gas turbine bottoming cycle. Variable combustor geometry was employed to demonstrate its ability to control primary and secondary zone equivalence ratios and overall pressure drop. Both concepts proved to be highly successful in achieving their desired objectives. The steam cooling reduced peak liner temperatures to less than 800 K. This low temperature offers the potential of both long life and reduced use of strategic materials for liner fabrication. These degrees of variable geometry were successfully employed to control air flow distribution within the combustor. A variable blade angle axial flow air swirler was used to control primary zone air flow, while the secondary and tertiary zone air flows were controlled by rotating bands which regulated air flow to the secondary zone quench holes and the dilutions holes respectively.

Downhole steam generator using low-pressure fuel and air supply

DOEpatents

Fox, R.L.

1981-01-07

For tertiary oil recovery, an apparatus for downhole steam generation is designed in which water is not injected directly onto the flame in the combustor, the combustion process is isolated from the reservoir pressure, the fuel and oxidant are supplied to the combustor at relatively low pressures, and the hot exhaust gases is prevented from entering the earth formation but is used to preheat the fuel and oxidant and water. The combustion process is isolated from the steam generation process. (DLC)

Dynamically balanced fuel nozzle and method of operation

DOEpatents

Richards, George A.; Janus, Michael C.; Robey, Edward H.

2000-01-01

An apparatus and method of operation designed to reduce undesirably high pressure oscillations in lean premix combustion systems burning hydrocarbon fuels are provided. Natural combustion and nozzle acoustics are employed to generate multiple fuel pockets which, when burned in the combustor, counteract the oscillations caused by variations in heat release in the combustor. A hybrid of active and passive control techniques, the apparatus and method eliminate combustion oscillations over a wide operating range, without the use of moving parts or electronics.

The Development of Environmental Barrier Coating Systems for SiC-SiC Ceramic Matrix Composites: Environment Effects on the Creep and Fatigue Resistance

NASA Technical Reports Server (NTRS)

Zhu, Dongming; Ghosn, Louis J.

2014-01-01

Topics covered include: Environmental barrier coating system development: needs, challenges and limitations; Advanced environmental barrier coating systems (EBCs) for CMC airfoils and combustors; NASA EBC systems and material system evolutions, Current turbine and combustor EBC coating emphases, Advanced development, processing, testing and modeling, EBC and EBC bond coats: recent advances; Design tool and life prediction of coated CMC components; Advanced CMC-EBC rig demonstrations; Summary and future directions.

Motivation for Creativity in Architectural Design and Engineering Design Students: Implications for Design Education

ERIC Educational Resources Information Center

Casakin, Hernan; Kreitler, Shulamith

2010-01-01

The investigation reported here dealt with the study of motivation for creativity. The goals were to assess motivation for creativity in architectural design and engineering design students based on the Cognitive Orientation theory which defines motivation as a function of a set of belief types, themes, and groupings identified as relevant for the…

Fuel-Air Mixing and Combustion in Scramjets

NASA Technical Reports Server (NTRS)

Drummond, J. P.; Diskin, Glenn S.; Cutler, A. D.

2002-01-01

Activities in the area of scramjet fuel-air mixing and combustion associated with the Research and Technology Organization Working Group on Technologies for Propelled Hypersonic Flight are described. Work discussed in this paper has centered on the design of two basic experiments for studying the mixing and combustion of fuel and air in a scramjet. Simulations were conducted to aid in the design of these experiments. The experimental models were then constructed, and data were collected in the laboratory. Comparison of the data from a coaxial jet mixing experiment and a supersonic combustor experiment with a combustor code were then made and described. This work was conducted by NATO to validate combustion codes currently employed in scramjet design and to aid in the development of improved turbulence and combustion models employed by the codes.

Operational procedure for computer program for design point characteristics of a compressed-air generator with through-flow combustor for V/STOL applications

NASA Technical Reports Server (NTRS)

Krebs, R. P.

1971-01-01

The computer program described in this report calculates the design-point characteristics of a compressed-air generator for use in V/STOL applications such as systems with a tip-turbine-driven lift fan. The program computes the dimensions and mass, as well as the thermodynamic performance of a model air generator configuration which involves a straight through-flow combustor. Physical and thermodynamic characteristics of the air generator components are also given. The program was written in FORTRAN IV language. Provision has been made so that the program will accept input values in either SI units or U.S. customary units. Each air generator design-point calculation requires about 1.5 seconds of 7094 computer time for execution.

The Importance of Design Thinking for Technological Literacy: A Phenomenological Perspective

ERIC Educational Resources Information Center

Wells, Alastair

2013-01-01

"We know that progress depends on discovery, inventions, creativity and design, but we have simply supposed that it happens anyway," de Bono (1999 p. 43). Technology education is ostensibly a foundation for future designers and creative thinking. However evidence of good design or creative thinking in outcomes displayed in school…

Clocked combustor can array

DOEpatents

Kim, Won-Wook; McMahan, Kevin Weston; Srinivasan, Shiva Kumar

2017-01-17

The present application provides a clocked combustor can array for coherence reduction in a gas turbine engine. The clocked combustor can array may include a number of combustor cans positioned in a circumferential array. A first set of the combustor cans may have a first orientation and a second set of the combustor cans may have a second orientation.

Coupling of Transport and Chemical Processes in Catalytic Combustion

NASA Technical Reports Server (NTRS)

Bracco, F. V.; Bruno, C.; Royce, B. S. H.; Santavicca, D. A.; Sinha, N.; Stein, Y.

1983-01-01

Catalytic combustors have demonstrated the ability to operate efficiently over a much wider range of fuel air ratios than are imposed by the flammability limits of conventional combustors. Extensive commercial use however needs the following: (1) the design of a catalyst with low ignition temperature and high temperature stability, (2) reducing fatigue due to thermal stresses during transient operation, and (3) the development of mathematical models that can be used as design optimization tools to isolate promising operating ranges for the numerous operating parameters. The current program of research involves the development of a two dimensional transient catalytic combustion model and the development of a new catalyst with low temperature light-off and high temperature stablity characteristics.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Rosfjord, T.J.

United Technologies Research Center has conducted a comprehensive investigation of the combustion characteristics of coal-water mixture (CWM) fuel for use in gas turbine combustors. Included in the program have been studies of the physical and chemical processes involved in CWM combustion. In particular, subrig test programs were performed to document the level of atomization achieved by candidate fuel nozzles and to evaluate the mixing/stability characteristics as evidenced by flow patterns in the combustor. Additionally, an extensive investigation of the pyrolysis behavior of coals used in the preparation of CWM fuels has been conducted for the high heat flux conditions experiencedmore » in a gas turbine combustor. These several activities were performed in preparation for tests in a combustor rig which simulated full-load gas turbine combustor conditions. Data were acquired to evaluate the ability of several configurations to achieve high levels of fuel burnout while controlling the conversion of fuel nitrogen to NOx. The bulk of the combustion tests were performed using one slurry, designated UCC-1, which was a 60% loaded CWM with a heating value of approximately 8800 Btu/lb. A limited test effort was conducted using a second fuel (UCC-2) which was chemically identical to the original CWM but contained coal possessing a larger particle size distribution. High levels of combustion efficiency were obtained using either UCC-1 or UCC-2 fuels. 9 refs., 53 figs., 7 tabs.« less

Review of modern low emissions combustion technologies for aero gas turbine engines

NASA Astrophysics Data System (ADS)

Liu, Yize; Sun, Xiaoxiao; Sethi, Vishal; Nalianda, Devaiah; Li, Yi-Guang; Wang, Lu

2017-10-01

Pollutant emissions from aircraft in the vicinity of airports and at altitude are of great public concern due to their impact on environment and human health. The legislations aimed at limiting aircraft emissions have become more stringent over the past few decades. This has resulted in an urgent need to develop low emissions combustors in order to meet legislative requirements and reduce the impact of civil aviation on the environment. This article provides a comprehensive review of low emissions combustion technologies for modern aero gas turbines. The review considers current high Technologies Readiness Level (TRL) technologies including Rich-Burn Quick-quench Lean-burn (RQL), Double Annular Combustor (DAC), Twin Annular Premixing Swirler combustors (TAPS), Lean Direct Injection (LDI). It further reviews some of the advanced technologies at lower TRL. These include NASA multi-point LDI, Lean Premixed Prevaporised (LPP), Axially Staged Combustors (ASC) and Variable Geometry Combustors (VGC). The focus of the review is placed on working principles, a review of the key technologies (includes the key technology features, methods of realising the technology, associated technology advantages and design challenges, progress in development), technology application and emissions mitigation potential. The article concludes the technology review by providing a technology evaluation matrix based on a number of combustion performance criteria including altitude relight auto-ignition flashback, combustion stability, combustion efficiency, pressure loss, size and weight, liner life and exit temperature distribution.

International Union of Theoretical and Applied Mechanics (IUTAM) Symposium on Aerothermodynamics in Combustors

NASA Astrophysics Data System (ADS)

Lee, Richard S.; Whitelaw, J. H.; Wung, T.-S.

1991-06-01

The subject of aerothermodynamics is playing an ever increasingly critical role in a variety of important industrial and technical problems in the design of combustors. In recent years, it has become the focus of attention among investigators from research laboratories and industries around the world resulting in a large number of meetings on its various aspects every year. However, most of these meetings deal with a certain problem area, for instance that of the global combustion of fuel droplets in a flow. Because of the inherent complexities involved in such flows, the analytical effort has been mostly confined to over-simplified and over-idealized flow systems while the experimental effort has been mostly directed towards global measurements of flows found in industrial applications. With the rapid and phenomenal developments of key research tools mostly in the last two decades, in particular those of modern digital computers, laser optics, and electronics, many of the previously unthinkable, rigorous investigations in real-life flows have gradually become feasible. It is against this background that this international conference on the aerothermodynamics in combustors is being held at this point in time. This symposium involves the presentation of papers concerned with flow and thermodynamic characteristics of combustors, with emphasis on gas-turbine combustors and including information relevant to rocket motors, internal combustion engines and furnaces.

The influence of cavity parameters on the combustion oscillation in a single-side expansion scramjet combustor

NASA Astrophysics Data System (ADS)

Ouyang, Hao; Liu, Weidong; Sun, Mingbo

2017-08-01

Cavity has been validated to be efficient flameholders for scramjet combustors, but the influence of its parameters on the combustion oscillation in scramjet combustor has barely been studied. In the present work, a series of experiments focusing on this issue have been carried out. The influence of flameholding cavity position, its length to depth ratio L/D and aft wall angle θ and number on ethylene combustion oscillation characteristics in scramjet combustor has been researched. The obtained experimental results show that, as the premixing distance between ethylene injector and flameholding cavity varies, the ethylene combustion flame will take on two distinct forms, small-amplitude high frequency fluctuation, and large-amplitude low frequency oscillation. The dominant frequency of the large-amplitude combustion oscillation is in inverse proportion to the pre-mixing distance. Moreover, the influence of cavity length to depth ratio and the aft wall angleθexists diversity when the flameholding cavity position is different and can be recognized as unnoticeable compared to the impact of the premixing distance. In addition, we also find that, when the premixing distance is identical and sufficient, increasing the number of tandem flameholding cavities can change the dominant frequency of combustion oscillation hardly, let alone avoid the combustion oscillation. It is believed that the present investigation will provide a useful reference for the design of the scramjet combustor.

Radial midframe baffle for can-annular combustor arrangement having tangentially oriented combustor cans

DOEpatents

Rodriguez, Jose L.

2015-09-15

A can-annular gas turbine engine combustion arrangement (10), including: a combustor can (12) comprising a combustor inlet (38) and a combustor outlet circumferentially and axially offset from the combustor inlet; an outer casing (24) defining a plenum (22) in which the combustor can is disposed; and baffles (70) configured to divide the plenum into radial sectors (72) and configured to inhibit circumferential motion of compressed air (16) within the plenum.

A Chaotic Intervention: Creativity and Peer Learning in Design Education

ERIC Educational Resources Information Center

Budge, Kylie; Beale, Claire; Lynas, Emma

2013-01-01

Peer feedback and critique is integral to the creative practice of studio-based textile designers. In a creative learning context, how do students perceive the role of peer feedback and critique? What conditions do students identify as being important to stimulating creativity in a collaborative peer feedback and critique-driven learning…

3D-CAD Effects on Creative Design Performance of Different Spatial Abilities Students

ERIC Educational Resources Information Center

Chang, Y.

2014-01-01

Students' creativity is an important focus globally and is interrelated with students' spatial abilities. Additionally, three-dimensional computer-assisted drawing (3D-CAD) overcomes barriers to spatial expression during the creative design process. Does 3D-CAD affect students' creative abilities? The purpose of this study was to explore the…

Secondary flow and heat transfer control in gas turbine inlet nozzle guide vanes

NASA Astrophysics Data System (ADS)

Burd, Steven Wayne

1998-12-01

Endwall heat transfer is a very serious problem in the inlet nozzle guide vane region of gas turbine engines. To resolve heat transfer concerns and provide the desired thermal protection, modern cooling flows for the vane endwalls tend to be excessive leading to lossy and inefficient designs. Coolant introduction is further complicated by the flow patterns along vane endwall surfaces. They are three-dimensional and dominated by strong, complex secondary flows. To achieve performance goals for next-generation engines, more aerodynamically efficient and advanced cooling concepts, including combustor bleed cooling, must be investigated. To this end, the overall performance characteristics of several combustor bleed flow designs are assessed in this experimental study. In particular, their contributions toward secondary flow control and component cooling are documented. Testing is performed in a large-scale, guide vane simulator comprised of three airfoils encased between one contoured and one flat endwall. Core flow is supplied to this simulator at an inlet chord Reynolds number of 350,000 and turbulence intensity of 9.5%. Combustor bleed cooling flow is injected through the contoured endwall via inclined slots. The slots vary in cross-sectional area, have equivalent slot widths, and are positioned with their leeward edges 10% of the axial chord ahead of the airfoil leading edges. Measurements with hot-wire anemometry characterize the inlet and exit flow fields of the cascade. Total and static pressure measurements document aerodynamic performance. Thermocouple measurements detail thermal fields and permit evaluation of surface adiabatic effectiveness. To elucidate the effects of bleed injection, data are compared to an experiment taken without bleed. The influence of bleed mass flow rate and slot geometry on the aerodynamic losses and thermal protection arc given. This study suggests that such combustor bleed flow cooling offers significant thermal protection without imposing aerodynamic penalties. Such performance is contrary to the performance of present vane cooling schemes. The results of this investigation support designs which incorporate combustor coolant injection upstream of the airfoil leading edges. To complement, a short exploratory study regarding the effects of surface roughness was also performed. Results indicate modified cooling performance and significantly higher aerodynamic losses with rough surfaces.

Advanced coal-fueled industrial cogeneration gas turbine system particle removal system development

DOE Office of Scientific and Technical Information (OSTI.GOV)

Stephenson, M.

1994-03-01

Solar Turbines developed a direct coal-fueled turbine system (DCFT) and tested each component in subscale facilities and the combustion system was tested at full-scale. The combustion system was comprised of a two-stage slagging combustor with an impact separator between the two combustors. Greater than 90 percent of the native ash in the coal was removed as liquid slag with this system. In the first combustor, coal water slurry mixture (CWM) was injected into a combustion chamber which was operated loan to suppress NO{sub x} formation. The slurry was introduced through four fuel injectors that created a toroidal vortex because ofmore » the combustor geometry and angle of orientation of the injectors. The liquid slag that was formed was directed downward toward an impaction plate made of a refractory material. Sixty to seventy percent of the coal-borne ash was collected in this fashion. An impact separator was used to remove additional slag that had escaped the primary combustor. The combined particulate collection efficiency from both combustors was above 95 percent. Unfortunately, a great deal of the original sulfur from the coal still remained in the gas stream and needed to be separated. To accomplish this, dolomite or hydrated lime were injected in the secondary combustor to react with the sulfur dioxide and form calcium sulfite and sulfates. This solution for the sulfur problem increased the dust concentrations to as much as 6000 ppmw. A downstream particulate control system was required, and one that could operate at 150 psia, 1850-1900{degrees}F and with low pressure drop. Solar designed and tested a particulate rejection system to remove essentially all particulate from the high temperature, high pressure gas stream. A thorough research and development program was aimed at identifying candidate technologies and testing them with Solar`s coal-fired system. This topical report summarizes these activities over a period beginning in 1987 and ending in 1992.« less

Scaling of Performance in Liquid Propellant Rocket Engine Combustion Devices

NASA Technical Reports Server (NTRS)

Hulka, James R.

2008-01-01

This paper discusses scaling of combustion and combustion performance in liquid propellant rocket engine combustion devices. In development of new combustors, comparisons are often made between predicted performance in a new combustor and measured performance in another combustor with different geometric and thermodynamic characteristics. Without careful interpretation of some key features, the comparison can be misinterpreted and erroneous information used in the design of the new device. This paper provides a review of this performance comparison, including a brief review of the initial liquid rocket scaling research conducted during the 1950s and 1960s, a review of the typical performance losses encountered and how they scale, a description of the typical scaling procedures used in development programs today, and finally a review of several historical development programs to see what insight they can bring to the questions at hand.

Experimental and computational study of the effect of shocks on film cooling effectiveness in scramjet combustors

NASA Technical Reports Server (NTRS)

Kamath, Pradeep S.; Holden, Michael S.; Mcclinton, Charles R.

1990-01-01

This paper presents results from a study conducted to investigate the effect of incident oblique shocks on the effectiveness of a coolant film at Mach numbers, typical of those expected in a scramjet combustor at Mach 15 to 20 flight. Computations with a parabolic code are in good agreement with the measured pressures and heat fluxes, after accounting for the influence of the shock upstream of its point of impingement on the plate, and the expansion from the trailing edge of the shock generator. The test data shows that, for the blowing rates tested, the film is rendered largely ineffective by the shock. Computations show that coolant blowing rates five to ten times those tested are required to protect against shock-induced heating. The implications of the results to scramjet combustor design are discussed.

Scramjet Combustor Characteristics at Hypervelocity Condition over Mach 10 Flight

NASA Astrophysics Data System (ADS)

Takahashi, M.; Komuro, T.; Sato, K.; Kodera, M.; Tanno, H.; Itoh, K.

2009-01-01

To investigate possibility of reduction of a scramjet combustor size without thrust performance loss, a two-dimensional constant-area combustor of a previous engine model was replaced with the one with 23% lower-height. With the application of the lower-height combustor, the pressure in the combustor becomes 50% higher and the combustor length for the optimal performance becomes 43% shorter than the original combustor. The combustion tests of the modified engine model were conducted using a large free-piston driven shock tunnel at flow conditions corresponding to the flight Mach number from 9 to 14. CFD was also applied to the engine internal flows. The results showed that the mixing and combustion heat release progress faster to the distance and the combustor performance similar to that of the previous engine was obtained with the modified engine. The reduction of the combustor size without the thrust performance loss is successfully achieved by applying the lower-height combustor.

Evaluation of advanced combustion concepts for dry NO sub x suppression with coal-derived, gaseous fuels

NASA Technical Reports Server (NTRS)

Beebe, K. W.; Symonds, R. A.; Notardonato, J. J.

1982-01-01

The emissions performance of a rich lean combustor (developed for liquid fuels) was determined for combustion of simulated coal gases ranging in heating value from 167 to 244 Btu/scf (7.0 to 10.3 MJ/NCM). The 244 Btu/scf gas is typical of the product gas from an oxygen blown gasifier, while the 167 Btu/scf gas is similar to that from an air blown gasifier. NOx performance of the rich lean combustor did not meet program goals with the 244 Btu/scf gas because of high thermal NOx, similar to levels expected from conventional lean burning combustors. The NOx emissions are attributed to inadequate fuel air mixing in the rich stage resulting from the design of the large central fuel nozzle delivering 71% of the total gas flow. NOx yield from ammonia injected into the fuel gas decreased rapidly with increasing ammonia level, and is projected to be less than 10% at NH3 levels of 0.5% or higher. NOx generation from NH3 is significant at ammonia concentrations significantly less than 0.5%. These levels may occur depending on fuel gas cleanup system design. CO emissions, combustion efficiency, smoke and other operational performance parameters were satisfactory. A test was completed with a catalytic combustor concept with petroleum distillate fuel. Reactor stage NOx emissions were low (1.4g NOx/kg fuel). CO emissions and combustion efficiency were satisfactory. Airflow split instabilities occurred which eventually led to test termination.

Instructional Design as Critical and Creative Thinking: A Journey through a Jamestown-Era Native American Village

ERIC Educational Resources Information Center

Baum, Liesl M.; Newbill, Phyllis Leary

2010-01-01

The role of critical and creative thinking has been debated within the field of instructional design. Through an instructional design and development project we have identified how critical and creative thinking are essential to the instructional design process. This paper highlights a recent project focused on a virtual Native American village…

Feedback control of combustion instabilities from within limit cycle oscillations using H∞ loop-shaping and the ν-gap metric

PubMed Central

Morgans, Aimee S.

2016-01-01

Combustion instabilities arise owing to a two-way coupling between acoustic waves and unsteady heat release. Oscillation amplitudes successively grow, until nonlinear effects cause saturation into limit cycle oscillations. Feedback control, in which an actuator modifies some combustor input in response to a sensor measurement, can suppress combustion instabilities. Linear feedback controllers are typically designed, using linear combustor models. However, when activated from within limit cycle, the linear model is invalid, and such controllers are not guaranteed to stabilize. This work develops a feedback control strategy guaranteed to stabilize from within limit cycle oscillations. A low-order model of a simple combustor, exhibiting the essential features of more complex systems, is presented. Linear plane acoustic wave modelling is combined with a weakly nonlinear describing function for the flame. The latter is determined numerically using a level set approach. Its implication is that the open-loop transfer function (OLTF) needed for controller design varies with oscillation level. The difference between the mean and the rest of the OLTFs is characterized using the ν-gap metric, providing the minimum required ‘robustness margin’ for an H∞ loop-shaping controller. Such controllers are designed and achieve stability both for linear fluctuations and from within limit cycle oscillations. PMID:27493558

A new unsteady mixing model to predict NO(x) production during rapid mixing in a dual-stage combustor

NASA Technical Reports Server (NTRS)

Menon, Suresh

1992-01-01

An advanced gas turbine engine to power supersonic transport aircraft is currently under study. In addition to high combustion efficiency requirements, environmental concerns have placed stringent restrictions on the pollutant emissions from these engines. A combustor design with the potential for minimizing pollutants such as NO(x) emissions is undergoing experimental evaluation. A major technical issue in the design of this combustor is how to rapidly mix the hot, fuel-rich primary zone product with the secondary diluent air to obtain a fuel-lean mixture for combustion in the second stage. Numerical predictions using steady-state methods cannot account for the unsteady phenomena in the mixing region. Therefore, to evaluate the effect of unsteady mixing and combustion processes, a novel unsteady mixing model is demonstrated here. This model has been used to study multispecies mixing as well as propane-air and hydrogen-air jet nonpremixed flames, and has been used to predict NO(x) production in the mixing region. Comparison with available experimental data show good agreement, thereby providing validation of the mixing model. With this demonstration, this mixing model is ready to be implemented in conjunction with steady-state prediction methods and provide an improved engineering design analysis tool.

Parametric Design of Injectors for LDI-3 Combustors

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Mongia, Hukam; Lee, Phil

2015-01-01

Application of a partially calibrated National Combustion Code (NCC) for providing guidance in the design of the 3rd generation of the Lean-Direct Injection (LDI) multi-element combustion configuration (LDI-3) is summarized. NCC was used to perform non-reacting and two-phase reacting flow computations on several LDI-3 injector configurations in a single-element and a five-element injector array. All computations were performed with a consistent approach for mesh-generation, turbulence, spray simulations, ignition and chemical kinetics-modeling. Both qualitative and quantitative assessment of the computed flowfield characteristics of the several design options led to selection of an optimal injector LDI- 3 design that met all the requirements including effective area, aerodynamics and fuel-air mixing criteria. Computed LDI-3 emissions (namely, NOx, CO and UHC) will be compared with the prior generation LDI- 2 combustor experimental data at relevant engine cycle conditions.

Fuel cell system with combustor-heated reformer

DOEpatents

Pettit, William Henry

2000-01-01

A fuel cell system including a fuel reformer heated by a catalytic combustor fired by anode effluent and/or fuel from a liquid fuel supply providing fuel for the fuel cell. The combustor includes a vaporizer section heated by the combustor exhaust gases for vaporizing the fuel before feeding it into the combustor. Cathode effluent is used as the principle oxidant for the combustor.

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

NASA Technical Reports Server (NTRS)

Shih, W.-P.; Lee, J. G.; Santavicca, D. A.

1994-01-01

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.

Predicting the Creativity of Design Majors Based on the Interaction of Diverse Personality Traits

ERIC Educational Resources Information Center

Chang, Chi-Cheng; Peng, Li-Pei; Lin, Ju-Sen; Liang, Chaoyun

2015-01-01

In this study, design majors were analysed to examine how diverse personality traits interact and influence student creativity. The study participants comprised 476 design majors. The results indicated that openness predicted the originality of creativity, whereas openness, conscientiousness and agreeableness predicted the usefulness of…

Designscholar: Examining Creative Thinking in an Online Learning Community for Interior Design Graduate Students

ERIC Educational Resources Information Center

Ransdell, Marlo Evelyn

2009-01-01

This study examined the creative thinking of interior design graduate students in an online learning community. This study considered potential changes in creative thinking (fluency, flexibility, originality, and elaboration) about design research resulting from peer-led online discussions. It further studied the learner characteristics of…

Searching Creativity: (N)On Place Design Workshop

ERIC Educational Resources Information Center

Önal, Gökçe Ketizmen

2017-01-01

This study is mainly about developing an approach for fostering creativity in design education through analyzing the interactions among creative dimensions resembling spatial and organizational pattern of folding as a technique and also by the help of cognitive action of designers: workshop participants. In order to make an assessment, a case…

Recognizing and Fostering Creativity in Technological Design Education

ERIC Educational Resources Information Center

Cropley, David; Cropley, Arthur

2010-01-01

The importance of creativity in technological design education is now clearly recognized, both in everyday understanding and also in formal curriculum guidelines. Design offers special opportunities for creativity because of the "openness" of problems (ill-defined problems, the existence of a variety of pathways to the solution, the absence of…

Creative Practices Embodied, Embedded, and Enacted in Architectural Settings: Toward an Ecological Model of Creativity.

PubMed

Malinin, Laura H

2015-01-01

Memoires by eminently creative people often describe architectural spaces and qualities they believe instrumental for their creativity. However, places designed to encourage creativity have had mixed results, with some found to decrease creative productivity for users. This may be due, in part, to lack of suitable empirical theory or model to guide design strategies. Relationships between creative cognition and features of the physical environment remain largely uninvestigated in the scientific literature, despite general agreement among researchers that human cognition is physically and socially situated. This paper investigates what role architectural settings may play in creative processes by examining documented first person and biographical accounts of creativity with respect to three central theories of situated cognition. First, the embodied thesis argues that cognition encompasses both the mind and the body. Second, the embedded thesis maintains that people exploit features of the physical and social environment to increase their cognitive capabilities. Third, the enaction thesis describes cognition as dependent upon a person's interactions with the world. Common themes inform three propositions, illustrated in a new theoretical framework describing relationships between people and their architectural settings with respect to different cognitive processes of creativity. The framework is intended as a starting point toward an ecological model of creativity, which may be used to guide future creative process research and architectural design strategies to support user creative productivity.

Creative Practices Embodied, Embedded, and Enacted in Architectural Settings: Toward an Ecological Model of Creativity

PubMed Central

Malinin, Laura H.

2016-01-01

Memoires by eminently creative people often describe architectural spaces and qualities they believe instrumental for their creativity. However, places designed to encourage creativity have had mixed results, with some found to decrease creative productivity for users. This may be due, in part, to lack of suitable empirical theory or model to guide design strategies. Relationships between creative cognition and features of the physical environment remain largely uninvestigated in the scientific literature, despite general agreement among researchers that human cognition is physically and socially situated. This paper investigates what role architectural settings may play in creative processes by examining documented first person and biographical accounts of creativity with respect to three central theories of situated cognition. First, the embodied thesis argues that cognition encompasses both the mind and the body. Second, the embedded thesis maintains that people exploit features of the physical and social environment to increase their cognitive capabilities. Third, the enaction thesis describes cognition as dependent upon a person’s interactions with the world. Common themes inform three propositions, illustrated in a new theoretical framework describing relationships between people and their architectural settings with respect to different cognitive processes of creativity. The framework is intended as a starting point toward an ecological model of creativity, which may be used to guide future creative process research and architectural design strategies to support user creative productivity. PMID:26779087

Fostering Creativity from an Emotional Perspective: Do Teachers Recognise and Handle Students' Emotions?

ERIC Educational Resources Information Center

Siu, Kin Wai; Wong, Yi Lin

2016-01-01

Emotions have a significant effect on the processes of designing and creative thinking. In an educational context, some emotions may even be detrimental to creativity. To further explore the link between creativity and emotion, a series of interviews were conducted with design and technology (D&T) teachers in Singapore, Hong Kong and Beijing…

Let's Think Creatively: Designing a High School Lesson on Metaphorical Creativity for English L2 Learners

ERIC Educational Resources Information Center

Wang, Hung-chun

2018-01-01

This article reports on a practitioner research study exploring how creative thinking activities can be designed and integrated into high school English classes. It delineates the process of developing a metaphorical creativity workshop for year 11 students in Taiwan and demonstrates the students' workshop experiences and learning outcomes.…

Energy Efficient Engine combustor test hardware detailed design report

NASA Technical Reports Server (NTRS)

Burrus, D. L.; Chahrour, C. A.; Foltz, H. L.; Sabla, P. E.; Seto, S. P.; Taylor, J. R.

1984-01-01

The Energy Efficient Engine (E3) Combustor Development effort was conducted as part of the overall NASA/GE E3 Program. This effort included the selection of an advanced double-annular combustion system design. The primary intent was to evolve a design which meets the stringent emissions and life goals of the E3 as well as all of the usual performance requirements of combustion systems for modern turbofan engines. Numerous detailed design studies were conducted to define the features of the combustion system design. Development test hardware was fabricated, and an extensive testing effort was undertaken to evaluate the combustion system subcomponents in order to verify and refine the design. Technology derived from this development effort will be incorporated into the engine combustion system hardware design. This advanced engine combustion system will then be evaluated in component testing to verify the design intent. What is evolving from this development effort is an advanced combustion system capable of satisfying all of the combustion system design objectives and requirements of the E3. Fuel nozzle, diffuser, starting, and emissions design studies are discussed.

One-Dimensional Spontaneous Raman Measurements Made in a Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

DeGroot, Wilhelmus A.; Hicks, Yolanda R.; Locke, Randy J.; Anderson, Robert C.

2001-01-01

The NASA Glenn Research Center and the aerospace industry are designing and testing low-emission combustor concepts to build the next generation of cleaner, more fuel efficient aircraft powerplants. These combustors will operate at much higher inlet temperatures and at pressures that are up to 3 to 5 times greater than combustors in the current fleet. From a test and analysis viewpoint, there is an increasing need for measurements from these combustors that are nonintrusive, simultaneous, multipoint, and more quantitative. Glenn researchers have developed several unique test facilities (refs. 1 and 2) that allow, for the first time, optical interrogation of combustor flow fields, including subcomponent performance, at pressures ranging from 1 to 60 bar (1 to 60 atm). Experiments conducted at Glenn are the first application of a visible laser-pumped, one-dimensional, spontaneous Raman-scattering technique to analyze the flow in a high-pressure, advanced-concept fuel injector at pressures thus far reaching 12 bar (12 atm). This technique offers a complementary method to the existing two- and three-dimensional imaging methods used, such as planar laser-induced fluorescence. Raman measurements benefit from the fact that the signal from each species is a linear function of its density, and the relative densities of all major species can be acquired simultaneously with good precision. The Raman method has the added potential to calibrate multidimensional measurements by providing an independent measurement of species number-densities at known points within the planar laser-induced fluorescence images. The visible Raman method is similar to an ultraviolet-Raman technique first tried in the same test facility (ref. 3). However, the visible method did not suffer from the ultraviolet technique's fuel-born polycyclic aromatic hydrocarbon fluorescence interferences.

Low Emissions RQL Flametube Combustor Component Test Results

NASA Technical Reports Server (NTRS)

Holdeman, James D.; Chang, Clarence T.

2001-01-01

This report describes and summarizes elements of the High Speed Research (HSR) Low Emissions Rich burn/Quick mix/Lean burn (RQL) flame tube combustor test program. This test program was performed at NASA Glenn Research Center circa 1992. The overall objective of this test program was to demonstrate and evaluate the capability of the RQL combustor concept for High Speed Civil Transport (HSCT) applications with the goal of achieving NOx emission index levels of 5 g/kg-fuel at representative HSCT supersonic cruise conditions. The specific objectives of the tests reported herein were to investigate component performance of the RQL combustor concept for use in the evolution of ultra-low NOx combustor design tools. Test results indicated that the RQL combustor emissions and performance at simulated supersonic cruise conditions were predominantly sensitive to the quick mixer subcomponent performance and not sensitive to fuel injector performance. Test results also indicated the mixing section configuration employing a single row of circular holes was the lowest NOx mixer tested probably due to the initial fast mixing characteristics of this mixing section. However, other quick mix orifice configurations such as the slanted slot mixer produced substantially lower levels of carbon monoxide emissions most likely due to the enhanced circumferential dispersion of the air addition. Test results also suggested that an optimum momentum-flux ratio exists for a given quick mix configuration. This would cause undesirable jet under- or over-penetration for test conditions with momentum-flux ratios below or above the optimum value. Tests conducted to assess the effect of quick mix flow area indicated that reduction in the quick mix flow area produced lower NOx emissions at reduced residence time, but this had no effect on NOx emissions measured at similar residence time for the configurations tested.

Rich burn combustor technology at Pratt and Whitney

NASA Technical Reports Server (NTRS)

Lohmann, Robert P.; Rosfjord, T. J.

1992-01-01

The topics covered include the following: near term objectives; rich burn quick quench combustor (RBQC); RBQC critical technology areas; cylindrical RBQQ combustor rig; modular RBQQ combustor; cylindrical rig objectives; quench zone mixing; noneffusive cooled liner; variable geometry requirements; and sector combustor rig.

Cultural Memory, an Asset for Design-Driven Innovation within the Creative Industries Sector: Lessons for Design Education

ERIC Educational Resources Information Center

Moalosi, Richie; Setlhatlhanyo, Keiphe Nani; Sealetsa, Oanthata Jester

2016-01-01

Culture is gaining recognition globally as an important driver of sustainable development in the creative economy. The significance of the role of design and culture with the creative industries is under-researched, especially from the new emerging economies perspective. Therefore, designers need a framework which will guide them on how they can…

Creative Minds Abroad: How Design Students Make Meaning of Their International Education Experiences

ERIC Educational Resources Information Center

Johnson, Rachel Sherman

2016-01-01

The purpose of this study is to explore the ways in which students majoring in a design discipline make meaning of their study abroad experiences in relation to their creativity and creative design work. Students and recent alumni from the College of Design (CDes) at the University of Minnesota-Twin Cities (UMTC) who had studied abroad formed the…

Design assessment of a 150 kWt CFBC Test Unit

DOE Office of Scientific and Technical Information (OSTI.GOV)

Batu, A.; Selcuk, N.; Kulah, G.

2010-04-15

For clean and efficient energy generation from coal, the most suitable technology known to date is 'Fluidized Bed Combustion' technology. Applications of circulating fluidized bed (CFB) combustion technology have been steadily increasing in both capacity and number over the past decade. Designs of these units have been based on the combustion tests carried out in pilot scale facilities to determine the combustion and desulfurization characteristics of coal and limestone reserves in CFB conditions. Similarly, utilization of Turkish lignites in CFB boilers necessitates adaptation of CFB combustion technology to these resources. However, the design of these test units are not basedmore » on firing coals with high ash, volatile matter and sulfur contents like Turkish lignites. For this purpose, a 150 kWt CFB combustor test unit is designed and constructed in Chemical Engineering Department of Middle East Technical University, based on the extensive experience acquired at the existing 0.3 MWt Bubbling Atmospheric Fluidized Bed Combustor (AFBC) Test Rig. Following the commissioning tests, a combustion test is carried out for investigation of combustion characteristics of Can lignite in CFB conditions and for assessment of the design of test unit. Comparison of the design outputs with experimental results reveals that most of the predictions and assumptions have acceptable agreement with the operating conditions. In conclusion, the performance of 150 kWt CFBC Test Unit is found to be satisfactory to be utilized for the long term research studies on combustion and desulfurization characteristics of indigenous lignite reserves in circulating fluidized bed combustors. (author)« less

Creative Thinking of Practical Engineering Students During a Design Project

NASA Astrophysics Data System (ADS)

Waks, Shlomo; Merdler, Moti

2003-01-01

Creativity in engineering design had become an economic necessity and not merely the privilege of unique individuals. The search for new, innovative and effective ideas in engineering design stands in center of daily creative performance. This search requires sensitivity to gaps of knowledge and information, and the ability to evoke numerous, different and unique ideas about engineering problems. The source of such information or knowledge can be either extrinsic-such as provided by an instructor or expert or intrinsic, which might involve transformation from one field or context to another. Furthermore, interaction with an exterior source as well as developing an inherent drive, have an impact on the motivation to perform creatively. This article, which is based on a study conducted among Israeli practical engineering students, deals with the variations in creative thinking during various stages of a design project and the relation between creative thinking and motivation factors.

Numerical analysis of the hot-gas-side and coolant-side heat transfer in liquid rocket engine combustors

NASA Technical Reports Server (NTRS)

Wang, Ten-See; Van, Luong

1992-01-01

The objective of this paper are to develop a multidisciplinary computational methodology to predict the hot-gas-side and coolant-side heat transfer and to use it in parametric studies to recommend optimized design of the coolant channels for a regeneratively cooled liquid rocket engine combustor. An integrated numerical model which incorporates CFD for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels, was developed. This integrated CFD/thermal model was validated by comparing predicted heat fluxes with those of hot-firing test and industrial design methods for a 40 k calorimeter thrust chamber and the Space Shuttle Main Engine Main Combustion Chamber. Parametric studies were performed for the Advanced Main Combustion Chamber to find a strategy for a proposed combustion chamber coolant channel design.

Sustainability as a Design Principle for Composition: Situational Creativity as a Habit of Mind

ERIC Educational Resources Information Center

Newcomb, Matthew

2012-01-01

Design is a rhetorical activity that requires creative thinking in response to difficult situations. That creative work ultimately builds new relationships and new contexts. Sustainable design can become an approach to composition that alters ways of thinking about writing situations, keeping ethical and contextual factors in focus, and…

The Need for Imagination and Creativity in Instructional Design

ERIC Educational Resources Information Center

Gibson, Pat

2013-01-01

The purpose of this article is to explore the need for imagination and creativity in adult education instructional design both online and face-to-face. It defines both imagination and creativity as well as provides an overview of the history of instructional design. It provides an examination of imagination and its application in educational…

Linear aerospike engine. [for reusable single-stage-to-orbit vehicle

NASA Technical Reports Server (NTRS)

Kirby, F. M.; Martinez, A.

1977-01-01

A description is presented of a dual-fuel modular split-combustor linear aerospike engine concept. The considered engine represents an approach to an integrated engine for a reusable single-stage-to-orbit (SSTO) vehicle. The engine burns two fuels (hydrogen and a hydrocarbon) with oxygen in separate combustors. Combustion gases expand on a linear aerospike nozzle. An engine preliminary design is discussed. Attention is given to the evaluation process for selecting the optimum number of modules or divisions of the engine, aspects of cooling and power cycle balance, and details of engine operation.

Experimental study of the operating characteristics of premixing-prevaporizing fuel/air mixing passages

NASA Technical Reports Server (NTRS)

Rohy, D. A.; Meier, J. G.

1983-01-01

Fuel spray and air flow characteristics were determined using nonintrusive (optical) measurement techniques in a fuel preparation duct. A very detailed data set was obtained at high pressures (to 10 atm) and temperatures (to 750 K). The data will be used to calibrate an analytical model which will facilitate the design of a lean premixed prevaporized combustor. This combustor has potential for achieving low pollutant emissions and low levels of flame radiation and pattern factors conductive to improved durability and performance for a variety of fuels.

14 CFR 25.1181 - Designated fire zones; regions included.

Code of Federal Regulations, 2012 CFR

2012-01-01

... engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain... Protection § 25.1181 Designated fire zones; regions included. (a) Designated fire zones are— (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete...

14 CFR 25.1181 - Designated fire zones; regions included.

Code of Federal Regulations, 2010 CFR

2010-01-01

... engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain... Protection § 25.1181 Designated fire zones; regions included. (a) Designated fire zones are— (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete...

14 CFR 25.1181 - Designated fire zones; regions included.

Code of Federal Regulations, 2013 CFR

2013-01-01

... engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain... Protection § 25.1181 Designated fire zones; regions included. (a) Designated fire zones are— (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete...

14 CFR 25.1181 - Designated fire zones; regions included.

Code of Federal Regulations, 2011 CFR

2011-01-01

... engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain... Protection § 25.1181 Designated fire zones; regions included. (a) Designated fire zones are— (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete...

14 CFR 25.1181 - Designated fire zones; regions included.

Code of Federal Regulations, 2014 CFR

2014-01-01

... engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain... Protection § 25.1181 Designated fire zones; regions included. (a) Designated fire zones are— (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete...

Creativity and Collaborative Learning and Teaching Strategies in the Design Disciplines

ERIC Educational Resources Information Center

Turnbull, Morag; Littlejohn, Allison; Allan, Malcolm

2010-01-01

Creativity can be described as the ability to generate new ideas and combine existing ideas in new ways to find novel solutions to problems. Creativity is enhanced by a free flow of knowledge and through social contact. On this basis, the authors argue that knowledge sharing is central to creativity in design and present preliminary evidence to…

The Effect of Using the Creative Reversal Act in Science Education on Middle School Students' Creativity Levels

ERIC Educational Resources Information Center

Karaca, Tulin; Koray, Ozlem

2017-01-01

Purpose: The purpose of this study is to examine the effects of the creative reversal act (CREACT) used in teaching ecosystems topics on the creativity levels of middle school students. Research Methods: The research was conducted using a quasi-experimental design, a quantitative research method, and a pretest-posttest control group design. The…

Perspectives on Pupil Creativity in Design and Technology in the Lower Secondary Curriculum in England

ERIC Educational Resources Information Center

Rutland, Marion; Barlex, David

2008-01-01

This paper is based on work carried out as part of a research study into the professional practices of secondary design and technology teachers in England. It focused on fostering creativity or teaching for creativity as defined by the Robinson Report (1999, All our futures: creativity, culture and education. London: Department for Education and…

Numerical Investigation of Fuel Distribution Effect on Flow and Temperature Field in a Heavy Duty Gas Turbine Combustor

NASA Astrophysics Data System (ADS)

Deng, Xiaowen; Xing, Li; Yin, Hong; Tian, Feng; Zhang, Qun

2018-03-01

Multiple-swirlers structure is commonly adopted for combustion design strategy in heavy duty gas turbine. The multiple-swirlers structure might shorten the flame brush length and reduce emissions. In engineering application, small amount of gas fuel is distributed for non-premixed combustion as a pilot flame while most fuel is supplied to main burner for premixed combustion. The effect of fuel distribution on the flow and temperature field related to the combustor performance is a significant issue. This paper investigates the fuel distribution effect on the combustor performance by adjusting the pilot/main burner fuel percentage. Five pilot fuel distribution schemes are considered including 3 %, 5 %, 7 %, 10 % and 13 %. Altogether five pilot fuel distribution schemes are computed and deliberately examined. The flow field and temperature field are compared, especially on the multiple-swirlers flow field. Computational results show that there is the optimum value for the base load of combustion condition. The pilot fuel percentage curve is calculated to optimize the combustion operation. Under the combustor structure and fuel distribution scheme, the combustion achieves high efficiency with acceptable OTDF and low NOX emission. Besides, the CO emission is also presented.

Parametric study of combustion oscillation in a single-side expansion scramjet combustor

NASA Astrophysics Data System (ADS)

Ouyang, Hao; Liu, Weidong; Sun, Mingbo

2016-10-01

As a promising candidate for future air-breathing systems, the viability and efficiency of scramjet propulsion is challenged by a variety of factors including the combustion oscillation in scramjet combustor. A series of comparative experiments focusing on the combustion oscillation issue has been carried out in the present work. The obtained experimental results show that as the global equivalence ratio increases, the combustion oscillation becomes more regular and frequent which is the most intensive in the vicinity of the fuel jet and the periodic combustion oscillation is more possible when the injectors and flame-holding cavity are mounted on the expansion-side wall. In order to avoid the combustion oscillation in scramjet combustor, distributed injection scheme is an effective method which can induce two parts interacting stable flame. In addition, the results reveal that the varying fuel including hydrogen, ethylene and kerosene with different chemical kinetics has a significant effect on the reaction process in scramjet combustor, which can result in stable combustion, periodic oscillation and failed ignition respectively on the same operating condition of this paper. We believe that the present work is helpful to the designing of scramjet propulsion device.

Modeling of high speed chemically reacting flow-fields

NASA Technical Reports Server (NTRS)

Drummond, J. P.; Carpenter, Mark H.; Kamath, H.

1989-01-01

The SPARK3D and SPARK3D-PNS computer programs were developed to model 3-D supersonic, chemically reacting flow-fields. The SPARK3D code is a full Navier-Stokes solver, and is suitable for use in scramjet combustors and other regions where recirculation may be present. The SPARK3D-PNS is a parabolized Navier-Stokes solver and provides an efficient means of calculating steady-state combustor far-fields and nozzles. Each code has a generalized chemistry package, making modeling of any chemically reacting flow possible. Research activities by the Langley group range from addressing fundamental theoretical issues to simulating problems of practical importance. Algorithmic development includes work on higher order and upwind spatial difference schemes. Direct numerical simulations employ these algorithms to address the fundamental issues of flow stability and transition, and the chemical reaction of supersonic mixing layers and jets. It is believed that this work will lend greater insight into phenomenological model development for simulating supersonic chemically reacting flows in practical combustors. Currently, the SPARK3D and SPARK3D-PNS codes are used to study problems of engineering interest, including various injector designs and 3-D combustor-nozzle configurations. Examples, which demonstrate the capabilities of each code are presented.

NASA's hypersonic propulsion program: History and direction

NASA Technical Reports Server (NTRS)

Wander, Steve

1992-01-01

Research into hypersonic propulsion; i.e., supersonic combustion, was seriously initiated at the Langley Research Center in the 1960's with the Hypersonic Research Engine (HRE) project. This project was designed to demonstrate supersonic combustion within the context of an engine module consisting of an inlet, combustor, and nozzle. In addition, the HRE utilized both subsonic and supersonic combustion (dual-mode) to demonstrate smooth operation over a Mach 4 to 7 speed range. The propulsion program thus concentrated on fundamental supersonic combustion studies and free jet propulsion tests for the three dimensional fixed geometry engine design to demonstrate inlet and combustor integration and installed performance potential. The developmental history of the program is presented. Additionally, the HRE program's effect on the current state of hypersonic propulsion is discussed.

Experimental clean combustor program, alternate fuels addendum, phase 2

NASA Technical Reports Server (NTRS)

Gleason, C. C.; Bahr, D. W.

1976-01-01

The characteristics of current and advanced low-emissions combustors when operated with special test fuels simulating broader range combustion properties of petroleum or coal derived fuels were studied. Five fuels were evaluated; conventional JP-5, conventional No. 2 Diesel, two different blends of Jet A and commercial aromatic mixtures - zylene bottoms and haphthalene charge stock, and a fuel derived from shale oil crude which was refined to Jet A specifications. Three CF6-50 engine size combustor types were evaluated; the standard production combustor, a radial/axial staged combustor, and a double annular combustor. Performance and pollutant emissons characteristics at idle and simulated takeoff conditions were evaluated in a full annular combustor rig. Altitude relight characteristics were evaluated in a 60 degree sector combustor rig. Carboning and flashback characteristics at simulated takeoff conditions were evaluated in a 12 degree sector combustor rig. For the five fuels tested, effects were moderate, but well defined.

Exhaust gas measurements in a propane fueled swirl stabilized combustor

NASA Technical Reports Server (NTRS)

Aanad, M. S.

1982-01-01

Exhaust gas temperature, velocity, and composition are measured and combustor efficiencies are calculated in a lean premixed swirl stabilized laboratory combustor. The radial profiles of the data between the co- and the counter swirl cases show significant differences. Co-swirl cases show evidence of poor turbulent mixing across the combustor in comparison to the counter-swirl cases. NO sub x levels are low in the combustor but substantial amounts of CO are present. Combustion efficiencies are low and surprisingly constant with varying outer swirl in contradiction to previous results under a slightly different inner swirl condition. This difference in the efficiency trends is expected to be a result of the high sensitivity of the combustor to changes in the inner swirl. Combustor operation is found to be the same for propane and methane fuels. A mechanism is proposed to explain the combustor operation and a few important characteristics determining combustor efficiency are identified.

Evaluation of a dual-chamber kerosene-heater combustion technology. Topical report, June-December 1985

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kardas, A.

1987-10-01

A kerosene heater equipped with a dual-chamber combustor was procured, tested, and technically evaluated to determine its applicability to natural gas combustion. The kerosene heater was found to have nitric oxide (NO), nitrogen dioxide (NO/sub 2/), and carbon monoxide (CO) emissions of 0.0)2, 0.006 and 0.02 lb/10/sup 6/ Btu input, respectively, much lower than those of blue-flame natural-gas combustors. A basic study was conducted to understand the interaction between kerosene combustion and the surrounding metal sleeves forming the dual chamber. Combustion characteristics of kerosene and natural gas were compared to formulate potential designs of low-emitting natural gas combustors. Three conceptsmore » were developed for low-emitting burners: an atmospheric burner to replace the kerosene wick in the dual chamber; the same concept with a powered vent; and a two-stage system equipped with a powered vent.« less

Model-Based Self-Tuning Multiscale Method for Combustion Control

NASA Technical Reports Server (NTRS)

Le, Dzu, K.; DeLaat, John C.; Chang, Clarence T.; Vrnak, Daniel R.

2006-01-01

A multi-scale representation of the combustor dynamics was used to create a self-tuning, scalable controller to suppress multiple instability modes in a liquid-fueled aero engine-derived combustor operating at engine-like conditions. Its self-tuning features designed to handle the uncertainties in the combustor dynamics and time-delays are essential for control performance and robustness. The controller was implemented to modulate a high-frequency fuel valve with feedback from dynamic pressure sensors. This scalable algorithm suppressed pressure oscillations of different instability modes by as much as 90 percent without the peak-splitting effect. The self-tuning logic guided the adjustment of controller parameters and converged quickly toward phase-lock for optimal suppression of the instabilities. The forced-response characteristics of the control model compare well with those of the test rig on both the frequency-domain and the time-domain.

Importance of inlet boundary conditions for numerical simulation of combustor flows

NASA Technical Reports Server (NTRS)

Sturgess, G. J.; Syed, S. A.; Mcmanus, K. R.

1983-01-01

Fluid dynamic computer codes for the mathematical simulation of problems in gas turbine engine combustion systems are required as design and diagnostic tools. To eventually achieve a performance standard with these codes of more than qualitative accuracy it is desirable to use benchmark experiments for validation studies. Typical of the fluid dynamic computer codes being developed for combustor simulations is the TEACH (Teaching Elliptic Axisymmetric Characteristics Heuristically) solution procedure. It is difficult to find suitable experiments which satisfy the present definition of benchmark quality. For the majority of the available experiments there is a lack of information concerning the boundary conditions. A standard TEACH-type numerical technique is applied to a number of test-case experiments. It is found that numerical simulations of gas turbine combustor-relevant flows can be sensitive to the plane at which the calculations start and the spatial distributions of inlet quantities for swirling flows.

Aerothermal modeling. Executive summary

NASA Technical Reports Server (NTRS)

Kenworthy, M. K.; Correa, S. M.; Burrus, D. L.

1983-01-01

One of the significant ways in which the performance level of aircraft turbine engines has been improved is by the use of advanced materials and cooling concepts that allow a significant increase in turbine inlet temperature level, with attendant thermodynamic cycle benefits. Further cycle improvements have been achieved with higher pressure ratio compressors. The higher turbine inlet temperatures and compressor pressure ratios with corresponding higher temperature cooling air has created a very hostile environment for the hot section components. To provide the technology needed to reduce the hot section maintenance costs, NASA has initiated the Hot Section Technology (HOST) program. One key element of this overall program is the Aerothermal Modeling Program. The overall objective of his program is to evolve and validate improved analysis methods for use in the design of aircraft turbine engine combustors. The use of such combustor analysis capabilities can be expected to provide significant improvement in the life and durability characteristics of both combustor and turbine components.

Effect of Blockage and Location on Mixing of Swirling Coaxial Jets in a Non-expanding Circular Confinement

NASA Astrophysics Data System (ADS)

Patel, V. K.; Singh, S. N.; Seshadri, V.

2013-06-01

A study is conducted to evolve an effective design concept to improve mixing in a combustor chamber to reduce the amount of intake air. The geometry used is that of a gas turbine combustor model. For simplicity, both the jets have been considered as air jets and effect of heat release and chemical reaction has not been modeled. Various contraction shapes and blockage have been investigated by placing them downstream at different locations with respect to inlet to obtain better mixing. A commercial CFD code `Fluent 6.3' which is based on finite volume method has been used to solve the flow in the combustor model. Validation is done with the experimental data available in literature using standard k-ω turbulence model. The study has shown that contraction and blockage at optimum location enhances the mixing process. Further, the effect of swirl in the jets has also investigated.

AST Combustion Workshop: Diagnostics Working Group Report

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Hicks, Yolanda R.; Hanson, Ronald K.

1996-01-01

A workshop was convened under NASA's Advanced Subsonics Technologies (AST) Program. Many of the principal combustion diagnosticians from industry, academia, and government laboratories were assembled in the Diagnostics/Testing Subsection of this workshop to discuss the requirements and obstacles to the successful implementation of advanced diagnostic techniques to the test environment of the proposed AST combustor. The participants, who represented the major relevant areas of advanced diagnostic methods currently applied to combustion and related fields, first established the anticipated AST combustor flowfield conditions. Critical flow parameters were then examined and prioritized as to their importance to combustor/fuel injector design and manufacture, environmental concerns, and computational interests. Diagnostic techniques were then evaluated in terms of current status, merits and obstacles for each flow parameter. All evaluations are presented in tabular form and recommendations are made on the best-suited diagnostic method to implement for each flow parameter in order of applicability and intrinsic value.

Mixing of Pure Air Jets with a Reacting Fuel-Rich Crossflow

NASA Technical Reports Server (NTRS)

Leong, M. Y.; Samuelsen, G. S.; Holdeman, J. D.

1997-01-01

Jets in a crossflow play an integral role in practical combustion systems such as can and annular gas turbine combustors in conventional systems, and the Rich-burn/Quick-mix/Lean-burn (RQL) combustor utilized in stationary applications and proposed for advanced subsonic and supersonic transports. The success of the RQL combustor rests with the performance of the quick-mixing section that bridges the rich and lean zones. The mixing of jet air with a rich crossflow to bring the reaction to completion in the lean zone must be performed rapidly and thoroughly in order to decrease the extent of near-stoichiometric fluid pocket formation. Fluid pockets at near-stoichiometric equivalence ratios are undesirable because the high temperatures attained accelerate pollutant formation kinetics associated with nitric oxide (NO). The present study develops a model experiment designed to reveal the processes that occur when jet air is introduced into hot effluent emanating from a fuel-rich reaction zone.

Design and evaluation of combustors for reducing aircraft engine pollution

NASA Technical Reports Server (NTRS)

Jones, R. E.; Grobman, J.

1973-01-01

Efforts in reducing exhaust emissions from turbine engines are reported. Various techniques employed and the results of testing are briefly described and referenced for detail. The experimental approaches taken to reduce oxides of nitrogen emissions include the use of: (1) multizone combustors incorporating reduced dwell times, (2) fuel-air premixing, (3) air atomization, (4) fuel prevaporization, and (5) gaseous fuel. Since emissions of unburned hydrocarbons and carbon monoxide are caused by poor combustion efficiency at engine idle, the studies of fuel staging in multizone combustors and air assist fuel nozzles have indicated that large reductions in these emissions can be achieved. Also, the effect of inlet-air humidity on oxides of nitrogen was studied as well as the very effective technique of direct water injection. The emission characteristics of natural gas and propane fuels were measured and compared with those of ASTM-Al kerosene fuel.

Design and evaluation of combustors for reducing aircraft engine pollution.

NASA Technical Reports Server (NTRS)

Jones, R. E.; Grobman, J.

1973-01-01

This report summarizes some of the NASA Lewis Research Center's recent efforts in reducing exhaust emissions from turbine engines. Various techniques employed and the results of testing are briefly described and referenced for detail. The experimental approaches taken to reduce oxides of nitrogen emissions include the use of: multizone combustors incorporating reduced dwell time, fuel-air premixing, air atomization, fuel prevaporization and gaseous fuel. Since emissions of unburned hydrocarbons and carbon monoxide are caused by poor combustion efficiency at engine idle, the studies of fuel staging in multizone combustors and air assist fuel nozzles have indicated that large reductions in these emissions can be achieved. Also, the effect of inlet-air humidity on oxides of nitrogen was studied as well as the very effective technique of direct water injection. The emission characteristics of natural gas and propane fuels were measured and compared with those of ASTM-Al kerosene fuel.

Principles for Designing Mathematical Tasks That Enhance Imitative and Creative Reasoning

ERIC Educational Resources Information Center

Lithner, Johan

2017-01-01

The design research programme learning by imitative and creative reasoning (LICR) studies whether, how and why tasks and teaching that enhance creative reasoning lead to a more productive struggle and more efficient learning than the common but inefficient task designs based on imitating given solution procedures. The purpose of this paper is to…

Professional Capabilities for Twenty-First Century Creative Careers: Lessons from Outstandingly Successful Australian Artists and Designers

ERIC Educational Resources Information Center

Bridgstock, Ruth

2013-01-01

Artists and designers are positioned at the centre of the twenty-first century creative economy. In order to recognise and make the most of the opportunities afforded by this new era, artists and designers still require the creativity, disciplinary depth of knowledge and technical skills traditionally possessed by professionals in these…

Modeling and Simulation of Swirl Stabilized Turbulent Non-Premixed Flames

NASA Astrophysics Data System (ADS)

Badillo-Rios, Salvador; Karagozian, Ann

2017-11-01

Flame stabilization is an important design criterion for many combustion chambers, especially at lean conditions and/or high power output, where insufficient stabilization can result in dangerous oscillations and noisy or damaged combustors. At high flow rates, swirling flow can offer a suitable stabilization mechanism, although understanding the dynamics of swirl-stabilized turbulent flames remains a significant challenge. Utilizing the General Equation and Mesh Solver (GEMS) code, which solves the Navier-Stokes equations along with the energy equation and five species equations, 2D axisymmetric and full 3D parametric studies and simulations are performed to guide the design and development of an experimental swirl combustor configuration and to study the effects of swirl on statistically stationary combustion. Results show that as the momentum of air is directed into the inner air inlet rather than the outer inlet of the swirl combustor, the central recirculating region becomes stronger and more unsteady, improving mixing and burning efficiency in that region. A high temperature region is found to occur as a result of burning of the trapped fuel from the central toroidal vortex. The effects of other parameters on flowfield and flame-stabilization dynamics are explored. Supported by ERC, Inc. (PS150006) and AFOSR (Dr. Chiping Li).

40 CFR 60.53b - Standards for municipal waste combustor operating practices.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Circulating fluidized bed combustor 100 4 Pulverized coal/refuse-derived fuel mixed fuel-fired combustor 150 4 Spreader stoker coal/refuse-derived fuel mixed fuel-fired combustor 150 24 a Measured at the combustor... activated carbon injection rate during dioxin/furan or mercury testing. [60 FR 65419, Dec. 19, 1995, as...

DOE Office of Scientific and Technical Information (OSTI.GOV)

Neocleous, C.C.; Esat, I.I.; Schizas, C.N.

The creativity phase is identified as an integral part of the design phase. The characteristics of creative persons which are relevant to designing artificial neural networks manifesting aspects of creativity, are identified. Based on these identifications, a general framework of artificial neural network characteristics to implement such a goal are proposed.

Computer Aided Creativity.

ERIC Educational Resources Information Center

Proctor, Tony

1988-01-01

Explores the conceptual components of a computer program designed to enhance creative thinking and reviews software that aims to stimulate creative thinking. Discusses BRAIN and ORACLE, programs intended to aid in creative problem solving. (JOW)

Combustor for a low-emissions gas turbine engine

DOEpatents

Glezer, Boris; Greenwood, Stuart A.; Dutta, Partha; Moon, Hee-Koo

2000-01-01

Many government entities regulated emission from gas turbine engines including CO. CO production is generally reduced when CO reacts with excess oxygen at elevated temperatures to form CO2. Many manufactures use film cooling of a combustor liner adjacent to a combustion zone to increase durability of the combustion liner. Film cooling quenches reactions of CO with excess oxygen to form CO2. Cooling the combustor liner on a cold side (backside) away from the combustion zone reduces quenching. Furthermore, placing a plurality of concavities on the cold side enhances the cooling of the combustor liner. Concavities result in very little pressure reduction such that air used to cool the combustor liner may also be used in the combustion zone. An expandable combustor housing maintains a predetermined distance between the combustor housing and combustor liner.

Rolling contact mounting arrangement for a ceramic combustor

DOEpatents

Boyd, G.L.; Shaffer, J.E.

1995-10-17

A combustor assembly having a preestablished rate of thermal expansion is mounted within a gas turbine engine housing having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the combustor assembly. The combustor assembly is constructed of a inlet end portion, a outlet end portion and a plurality of combustor ring segments positioned between the end portions. A mounting assembly is positioned between the combustor assembly and the gas turbine engine housing to allow for the difference in the rate of thermal expansion while maintaining axially compressive force on the combustor assembly to maintain contact between the separate components. 3 figs.

Rolling contact mounting arrangement for a ceramic combustor

DOEpatents

Boyd, Gary L.; Shaffer, James E.

1995-01-01

A combustor assembly having a preestablished rate of thermal expansion is mounted within a gas turbine engine housing having a preestablished rate of thermal expansion being greater than the preestablished rate of thermal expansion of the combustor assembly. The combustor assembly is constructed of a inlet end portion, a outlet end portion and a plurality of combustor ring segments positioned between the end portions. A mounting assembly is positioned between the combustor assembly and the gas turbine engine housing to allow for the difference in the rate of thermal expansion while maintaining axially compressive force on the combustor assembly to maintain contact between the separate components.

Non-Intrusive, Laser-Based Imaging of Jet-A Fuel Injection and Combustion Species in High Pressure, Subsonic Flows

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Hicks, Yolanda R.; Anderson, Robert C.; deGroot, Wilhelmus A.

2001-01-01

The emphasis of combustion research efforts at NASA Glenn Research Center (GRC) is on collaborating with industry to design and test gas-turbine combustors and subcomponents for both sub- and supersonic applications. These next-generation aircraft combustors are required to meet strict international environmental restrictions limiting emissions. To meet these goals, innovative combustor concepts require operation at temperatures and pressures far exceeding those of cur-rent designs. New and innovative diagnostic tools are necessary to characterize these flow streams since existing methods are inadequate. The combustion diagnostics team at GRC has implemented a suite of highly sensitive, nonintrusive optical imaging methods to diagnose the flowfields of these new engine concepts. By using optically accessible combustors and flametubes, imaging of fuel and intermediate combustion species via planar laser-induced fluorescence (PLIF) at realistic pressures are now possible. Direct imaging of the fuel injection process through both planar Mie scattering and PLIF methods is also performed. Additionally, a novel combination of planar fuel fluorescence imaging and computational analysis allows a 3-D examination of the flowfield, resulting in spatially and temporally resolved fuel/air volume distribution maps. These maps provide detailed insight into the fuel injection process at actual conditions, thereby greatly enhancing the evaluation of fuel injector performance and other combustion phenomena. Stable species such as CO2, O2, N2O. and hydrocarbons are also investigated by a newly demonstrated 1-D, spontaneous Raman spectroscopic method. This visible wavelength Raman technique allows the acquisition of quantitative. stable species concentration measurements from the flow.

Non-Intrusive, Laser-Based Imaging of Jet-A Fuel Injection and Combustion Species in High Pressure, Subsonic Flows

NASA Technical Reports Server (NTRS)

Locke, R. J.; Hicks, Y. R.; Anderson, R. C.; deGroot, W. A.

2000-01-01

The emphasis of combustion research efforts at NASA Glenn Research Center (GRC) is on collaborating with industry to design and test gas-turbine combustors and subcomponents for both sub- and supersonic applications. These next-generation aircraft combustors are required to meet strict international environmental restrictions limiting emissions. To meet these goals, innovative combustor concepts require operation at temperatures and pressures far exceeding those of current designs. New and innovative diagnostic tools are necessary to characterize these flow streams since existing methods are inadequate. The combustion diagnostics team at GRC has implemented a suite of highly sensitive, nonintrusive optical imaging methods to diagnose the flowfields of these new engine concepts. By using optically accessible combustors and flame-tubes, imaging of fuel and intermediate combustion species via planar laser-induced fluorescence (PLIF) at realistic pressures are now possible. Direct imaging of the fuel injection process through both planar Mie scattering and PLIF methods is also performed. Additionally, a novel combination of planar fuel fluorescence imaging and computational analysis allows a 3-D examination of the flowfield, resulting in spatially and temporally resolved fuel/air volume distribution maps. These maps provide detailed insight into the fuel injection process at actual conditions, thereby greatly enhancing the evaluation of fuel injector performance and other combustion phenomena. Stable species such as CO2, O2, N2, H2O, and hydrocarbons are also investigated by a newly demonstrated 1-D, spontaneous Raman spectroscopic method. This visible wavelength Raman technique allows the acquisition of quantitative, stable species concentration measurements from the flow.

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

NASA Technical Reports Server (NTRS)

Brankovic, Andreja; Ryder, Robert C., Jr.; Hendricks, Robert C.; Liu, Nan-Suey; Gallagher, John R.; Shouse, Dale T.; Roquemore, W. Melvyn; Cooper, Clayton S.; Burrus, David L.; Hendricks, John A.

2002-01-01

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.

Coherent anti-Stokes Raman scattering for quantitative temperature and concentration measurements in a high-pressure gas turbine combustor rig

NASA Astrophysics Data System (ADS)

Thariyan, Mathew Paul

Dual-pump coherent anti-Stokes Raman scattering (DP-CARS) temperature and major species (CO2/N2) concentration measurements have been performed in an optically-accessible high-pressure gas turbine combustor facility (GTCF) and for partially-premixed and non-premixed flames in a laminar counter-flow burner. A window assembly incorporating pairs of thin and thick fused silica windows on three sides was designed, fabricated, and assembled in the GTCF for advanced laser diagnostic studies. An injection-seeded optical parametric oscillator (OPO) was used as a narrowband pump laser source in the dual-pump CARS system. Large prisms on computer-controlled translation stages were used to direct the CARS beams either into the main optics leg for measurements in the GTCF or to a reference optics leg for measurements of the nonresonant CARS spectrum and for aligning the CARS system. Combusting flows were stabilized with liquid fuel injection only for the central injector of a 9-element lean direct injection (LDI) device developed at NASA Glenn Research Center. The combustor was operated using Jet A fuel at inlet air temperatures up to 725 K and combustor pressures up to 1.03 MPa. Single-shot DP-CARS spectra were analyzed using the Sandia CARSFT code in the batch operation mode to yield instantaneous temperature and CO2/N2 concentration ratio values. Spatial maps of mean and standard deviations of temperature and CO2/N2 concentrations were obtained in the high-pressure LDI flames by translating the CARS probe volume in axial and vertical directions inside the combustor rig. The mean temperature fields demonstrate the effect of the combustor conditions on the overall flame length and the average flame structure. The temperature relative standard deviation values indicate thermal fluctuations due to the presence of recirculation zones and/or flame brush fluctuations. The correlation between the temperature and relative CO 2 concentration data has been studied at various combustor conditions. The insight into the reacting flow structure provided by these measurements is discussed. Such measurements at conditions similar to those of aircraft gas turbine combustors are extremely useful for testing combustion models being used to predict performance of these systems.

Analytical fuel property effects--small combustors

NASA Technical Reports Server (NTRS)

Sutton, R. D.; Troth, D. L.; Miles, G. A.

1984-01-01

The consequences of using broad-property fuels in both conventional and advanced state-of-the-art small gas turbine combustors are assessed. Eight combustor concepts were selected for initial screening, of these, four final combustor concepts were chosen for further detailed analysis. These included the dual orifice injector baseline combustor (a current production 250-C30 engine combustor) two baseline airblast injected modifications, short and piloted prechamber combustors, and an advanced airblast injected, variable geometry air staged combustor. Final predictions employed the use of the STAC-I computer code. This quasi 2-D model includes real fuel properties, effects of injector type on atomization, detailed droplet dynamics, and multistep chemical kinetics. In general, fuel property effects on various combustor concepts can be classified as chemical or physical in nature. Predictions indicate that fuel chemistry has a significant effect on flame radiation, liner wall temperature, and smoke emission. Fuel physical properties that govern atomization quality and evaporation rates are predicted to affect ignition and lean-blowout limits, combustion efficiency, unburned hydrocarbon, and carbon monoxide emissions.

The Design and Implementation of a Physical Education Program to Promote Children's Creativity in the Early Years

ERIC Educational Resources Information Center

Zachopoulou, Evridiki; Trevlas, Efthimios; Konstadinidou, Elisavet

2006-01-01

The purpose of this study was to design and implement a physical education program to promote creativity in preschool children. The study was based on the following phases: (a) to design and formulate 20 physical education lessons in order to provide children with opportunities to develop their creative thinking through the use of movement…

Computational thermo-fluid dynamics contributions to advanced gas turbine engine design

NASA Technical Reports Server (NTRS)

Graham, R. W.; Adamczyk, J. J.; Rohlik, H. E.

1984-01-01

The design practices for the gas turbine are traced throughout history with particular emphasis on the calculational or analytical methods. Three principal components of the gas turbine engine will be considered: namely, the compressor, the combustor and the turbine.

National Combustion Code: Parallel Implementation and Performance

NASA Technical Reports Server (NTRS)

Quealy, A.; Ryder, R.; Norris, A.; Liu, N.-S.

2000-01-01

The National Combustion Code (NCC) is being developed by an industry-government team for the design and analysis of combustion systems. CORSAIR-CCD is the current baseline reacting flow solver for NCC. This is a parallel, unstructured grid code which uses a distributed memory, message passing model for its parallel implementation. The focus of the present effort has been to improve the performance of the NCC flow solver to meet combustor designer requirements for model accuracy and analysis turnaround time. Improving the performance of this code contributes significantly to the overall reduction in time and cost of the combustor design cycle. This paper describes the parallel implementation of the NCC flow solver and summarizes its current parallel performance on an SGI Origin 2000. Earlier parallel performance results on an IBM SP-2 are also included. The performance improvements which have enabled a turnaround of less than 15 hours for a 1.3 million element fully reacting combustion simulation are described.

Measuring Gifted Adolescents' Implicit Theories of Creativity

ERIC Educational Resources Information Center

Wickes, Katherine N. Saunders; Ward, Thomas B.

2006-01-01

This paper examines the structure of implicit theories of creativity in a sample of gifted adolescents and describes the development and use of the Creative Self Checklist and the Creative Individual Checklist, adjective checklists designed to assess endorsement of creativity-related personality and behavioral attributes. Findings indicate that…

Can Style Be Creative? An Exploratory Essay

ERIC Educational Resources Information Center

Clinton, Gregory

2016-01-01

This essay explores the nature of creativity of the practicing professional through the examination of the role of personal style in creative work, as well as how personality can affect and sustain creativity. Instructional designers, as practicing creatives, must balance the divergent and novel with the restraints of clients, projects, and…

Technique for predicting high-frequency stability characteristics of gaseous-propellant combustors

NASA Technical Reports Server (NTRS)

Priem, R. J.; Jefferson, Y. S. Y.

1973-01-01

A technique for predicting the stability characteristics of a gaseous-propellant rocket combustion system is developed based on a model that assumes coupling between the flow through the injector and the oscillating chamber pressure. The theoretical model uses a lumped parameter approach for the flow elements in the injection system plus wave dynamics in the combustion chamber. The injector flow oscillations are coupled to the chamber pressure oscillations with a delay time. Frequency and decay (or growth) rates are calculated for various combustor design and operating parameters to demonstrate the influence of various parameters on stability. Changes in oxidizer design parameters had a much larger influence on stability than a similar change in fuel parameters. A complete description of the computer program used to make these calculations is given in an appendix.

Adaptive Controls Method Demonstrated for the Active Suppression of Instabilities in Engine Combustors

NASA Technical Reports Server (NTRS)

Kopasakis, George

2004-01-01

An adaptive feedback control method was demonstrated that suppresses thermoacoustic instabilities in a liquid-fueled combustor of a type used in aircraft engines. Extensive research has been done to develop lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle to reduce the environmental impact of aerospace propulsion systems. However, these lean-burning combustors are susceptible to thermoacoustic instabilities (high-frequency pressure waves), which can fatigue combustor components and even the downstream turbine blades. This can significantly decrease the safe operating lives of the combustor and turbine. Thus, suppressing the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors under NASA's Propulsion and Power Program. This control methodology has been developed and tested in a partnership of the NASA Glenn Research Center, Pratt & Whitney, United Technologies Research Center, and the Georgia Institute of Technology. Initial combustor rig testing of the controls algorithm was completed during 2002. Subsequently, the test results were analyzed and improvements to the method were incorporated in 2003, which culminated in the final status of this controls algorithm. This control methodology is based on adaptive phase shifting. The combustor pressure oscillations are sensed and phase shifted, and a high-frequency fuel valve is actuated to put pressure oscillations into the combustor to cancel pressure oscillations produced by the instability.

"If I Was Going to Design a Chair, the Last Thing I Would Look at Is a Chair": Product Analysis and the Causes of Fixation in Students' Design Work 11-16 Years

ERIC Educational Resources Information Center

McLellan, Ros; Nicholl, Bill

2011-01-01

Creativity is acknowledged to be important for economic growth and as an everyday life-skill, however several influential reports have suggested that education could do more to harness creative talent. Creative cognition literature suggests the lack of creativity is at least partly the result of "fixation" (difficulty in generating novel…

A Pharmacy Elective Course on Creative Thinking, Innovation, and TED Talks.

PubMed

Cain, Jeff

2016-12-25

Objective. To implement and assess an elective course designed to enhance student creative thinking and presentation skills. Design. A two-credit elective course was developed that incorporated creative-thinking exercises, article discussions pertaining to creativity, TED Talk (TED Conferences, New York, NY) analyses, and presentation design and delivery exercises. Assessment. Assessment instruments included pre- and post-course Torrance Tests of Creative Thinking (TTCT). A scoring rubric was developed and used to evaluate a final mock TED Talk presented to faculty and students. Course evaluations were also used to assess student experiences in the course. Students' TTCT verbal creativity scores increased significantly ( p <0.05) during the course and their mock TED Talk mean scores (135±6.4) out of 150 were rated highly. Conclusion. The outcomes from this elective course confirmed that pharmacy students could develop and present an original "idea worth sharing" using the TED Talk format.

A Pharmacy Elective Course on Creative Thinking, Innovation, and TED Talks

PubMed Central

2016-01-01

Objective. To implement and assess an elective course designed to enhance student creative thinking and presentation skills. Design. A two-credit elective course was developed that incorporated creative-thinking exercises, article discussions pertaining to creativity, TED Talk (TED Conferences, New York, NY) analyses, and presentation design and delivery exercises. Assessment. Assessment instruments included pre- and post-course Torrance Tests of Creative Thinking (TTCT). A scoring rubric was developed and used to evaluate a final mock TED Talk presented to faculty and students. Course evaluations were also used to assess student experiences in the course. Students’ TTCT verbal creativity scores increased significantly (p<0.05) during the course and their mock TED Talk mean scores (135±6.4) out of 150 were rated highly. Conclusion. The outcomes from this elective course confirmed that pharmacy students could develop and present an original “idea worth sharing” using the TED Talk format. PMID:28179719

Multidisciplinary Analysis of a Hypersonic Engine

NASA Technical Reports Server (NTRS)

Suresh, Ambady; Stewart, Mark

2003-01-01

The objective is to develop high fidelity tools that can influence ISTAR design In particular, tools for coupling Fluid-Thermal-Structural simulations RBCC/TBCC designers carefully balance aerodynamic, thermal, weight, & structural considerations; consistent multidisciplinary solutions reveal details (at modest cost) At Scram mode design point, simulations give details of inlet & combustor performance, thermal loads, structural deflections.

The Structural Challenge: A Simple Design-Based Science Activity to Foster Creativity among Kinaesthetic Learners

ERIC Educational Resources Information Center

Amir, Nazir; Subramaniam, R.

2014-01-01

A suitable way for teachers to present science content and foster creativity in less academically inclined students is by getting them to engage in design-based science activities and guiding them along the way. This study illustrates how a design-and-make activity was carried out with the aim of getting students to showcase their creativity while…

Assessing Creativity: Drawing from the Experience of the UK's Creative Design Educators

ERIC Educational Resources Information Center

Penaluna, Andy; Penaluna, Kathryn

2009-01-01

Purpose: The aim of this paper is to respond to the acknowledged dearth of academic discourse on assessment strategies for entrepreneurship education. Using established approaches from design education as its fulcrum, it proposes a generic framework for assessment of "creativity" in an entrepreneurial context.…

Experimental study of the effects of flameholder geometry on emissions and performance of lean premixed combustors

NASA Technical Reports Server (NTRS)

Roffe, G.; Venkataramani, K. S.

1978-01-01

Emissions of NOx, CO, and unburned hydrocarbons (UHC) are reported for a lean premixed propane-air system at inlet conditions of 800K and 1MPa using twelve flameholder designs. The flameholders tested represent six design concepts with two values of blockage for each concept. Data were obtained at reference velocities of 35 m/s, 25 m/s and 20 m/s at combustor stations 10 cm and 30 cm downstream of the flameholders. Flameholder pressure drop was found to be a principal determinant of emissions performance. Designs producing larger pressure drops also produced less NOx, CO, and UHC emissions. The lean stability limit equivalence ratio was found to be approximately 0.35 for all designs. Flashback velocities (axial components in the flameholder passages) varied between 30 m/s and 40 m/s. A perforated plate flameholder was operated with a velocity as low as 23 m/s through the perforations at equivalence ratio 0.7 without producing flashback.

Innovative Adaptive Control Method Demonstrated for Active Suppression of Instabilities in Engine Combustors

NASA Technical Reports Server (NTRS)

Kopasakis, George

2005-01-01

This year, an improved adaptive-feedback control method was demonstrated that suppresses thermoacoustic instabilities in a liquid-fueled combustor of a type used in aircraft engines. Extensive research has been done to develop lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle to reduce the environmental impact of aerospace propulsion systems. However, these lean-burning combustors are susceptible to thermoacoustic instabilities (high-frequency pressure waves), which can fatigue combustor components and even downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppressing the thermoacoustic combustor instabilities is an enabling technology for meeting the low-emission goals of the NASA Ultra-Efficient Engine Technology (UEET) Project.

40 CFR 62.5870 - Identification of plan.

Code of Federal Regulations, 2010 CFR

2010-07-01

...) APPROVAL AND PROMULGATION OF STATE PLANS FOR DESIGNATED FACILITIES AND POLLUTANTS Minnesota Existing Large... combustors in Minnesota and should apply only to these sources. [63 FR 43083, Aug. 12, 1998] ...

40 CFR 62.5870 - Identification of plan.

Code of Federal Regulations, 2011 CFR

2011-07-01

...) APPROVAL AND PROMULGATION OF STATE PLANS FOR DESIGNATED FACILITIES AND POLLUTANTS Minnesota Existing Large... combustors in Minnesota and should apply only to these sources. [63 FR 43083, Aug. 12, 1998] ...

40 CFR 62.5870 - Identification of plan.

Code of Federal Regulations, 2014 CFR

2014-07-01

...) APPROVAL AND PROMULGATION OF STATE PLANS FOR DESIGNATED FACILITIES AND POLLUTANTS Minnesota Existing Large... combustors in Minnesota and should apply only to these sources. [63 FR 43083, Aug. 12, 1998] ...

40 CFR 62.5870 - Identification of plan.

Code of Federal Regulations, 2013 CFR

2013-07-01

...) APPROVAL AND PROMULGATION OF STATE PLANS FOR DESIGNATED FACILITIES AND POLLUTANTS Minnesota Existing Large... combustors in Minnesota and should apply only to these sources. [63 FR 43083, Aug. 12, 1998] ...

40 CFR 62.5870 - Identification of plan.

Code of Federal Regulations, 2012 CFR

2012-07-01

...) APPROVAL AND PROMULGATION OF STATE PLANS FOR DESIGNATED FACILITIES AND POLLUTANTS Minnesota Existing Large... combustors in Minnesota and should apply only to these sources. [63 FR 43083, Aug. 12, 1998] ...

Scaling study of the combustion performance of gas—gas rocket injectors

NASA Astrophysics Data System (ADS)

Wang, Xiao-Wei; Cai, Guo-Biao; Jin, Ping

2011-10-01

To obtain the key subelements that may influence the scaling of gas—gas injector combustor performance, the combustion performance subelements in a liquid propellant rocket engine combustor are initially analysed based on the results of a previous study on the scaling of a gas—gas combustion flowfield. Analysis indicates that inner wall friction loss and heat-flux loss are two key issues in gaining the scaling criterion of the combustion performance. The similarity conditions of the inner wall friction loss and heat-flux loss in a gas—gas combustion chamber are obtained by theoretical analyses. Then the theoretical scaling criterion was obtained for the combustion performance, but it proved to be impractical. The criterion conditions, the wall friction and the heat flux are further analysed in detail to obtain the specific engineering scaling criterion of the combustion performance. The results indicate that when the inner flowfields in the combustors are similar, the combustor wall shear stress will have similar distributions qualitatively and will be directly proportional to pc0.8dt-0.2 quantitatively. In addition, the combustion peformance will remain unchanged. Furthermore, multi-element injector chambers with different geometric sizes and at different pressures are numerically simulated and the wall shear stress and combustion efficiencies are solved and compared with each other. A multielement injector chamber is designed and hot-fire tested at several chamber pressures and the combustion performances are measured in a total of nine hot-fire tests. The numerical and experimental results verified the similarities among combustor wall shear stress and combustion performances at different chamber pressures and geometries, with the criterion applied.

Testing of DLR C/C-SiC and C/C for HIFiRE 8 Scramjet Combustor

NASA Technical Reports Server (NTRS)

Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael K.

2014-01-01

Ceramic Matrix Composites (CMCs) have been proposed for use as lightweight hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a carbon/carbon (C/C) and a carbon/carbon-silicon carbide (C/C-SiC) material fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for Hypersonic International Flight Research Experimentation (HIFiRE) 8, a joint Australia / Air Force Research Laboratory hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kilopascals. Flat panels of the DLR C/C and C/C-SiC materials were installed downstream of a hydrogen-fueled, dual-mode scramjet combustor and tested for several minutes at conditions simulating flight at Mach 5 and Mach 6. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used during some of the tests to reduce heat loss from the back surface and thus increase the heated surface temperature of the C/C-SiC panel approximately 177 C (350 F). The final C/C-SiC panel was tested for three cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

The impact of JP-4/JP-8 conversion on aircraft engine exhaust emissions. Interim technical report Jul 75--Feb 76

DOE Office of Scientific and Technical Information (OSTI.GOV)

Blazowski, W.S.

1976-05-01

The proposed conversion of predominant Air Force fuel usage from JP-4 to JP-8 has created the need to examine the dependence of engine pollutant emission on fuel type. Available data concerning the effect of fuel type on emissions has been reviewed. T56 single combustor testing has been undertaken to determine JP-4/JP-8 emission variations over a wide range of simulated engine cycle operating conditions at idle. In addition, a J85-5 engine was tested using JP-4 and JP-8. Results of the previous and new data collectively led to the following conclusions regarding conversion to JP-8: (a) HC and CO emission changes willmore » depend upon individual combustor design features, (b) no change to NOx emission will occur, and (c) an increase in smoke/particulate emissions will result. It is recommended that these findings be incorporated into air quality analytical models to define the overall impact of the proposed conversion. Further, it is recommended that combustor analytical models be employed to attempt prediction of the results described herein. Should these models be successful, analytical prediction of JP-8 emissions from other Air Force engine models may be substituted for more combustor rig or engine testing. (auth)« less

AST Critical Propulsion and Noise Reduction Technologies for Future Commercial Subsonic Engines Area of Interest 1.0: Reliable and Affordable Control Systems

NASA Technical Reports Server (NTRS)

Myers, William; Winter, Steve

2006-01-01

The General Electric Reliable and Affordable Controls effort under the NASA Advanced Subsonic Technology (AST) Program has designed, fabricated, and tested advanced controls hardware and software to reduce emissions and improve engine safety and reliability. The original effort consisted of four elements: 1) a Hydraulic Multiplexer; 2) Active Combustor Control; 3) a Variable Displacement Vane Pump (VDVP); and 4) Intelligent Engine Control. The VDVP and Intelligent Engine Control elements were cancelled due to funding constraints and are reported here only to the state they progressed. The Hydraulic Multiplexing element developed and tested a prototype which improves reliability by combining the functionality of up to 16 solenoids and servo-valves into one component with a single electrically powered force motor. The Active Combustor Control element developed intelligent staging and control strategies for low emission combustors. This included development and tests of a Controlled Pressure Fuel Nozzle for fuel sequencing, a Fuel Multiplexer for individual fuel cup metering, and model-based control logic. Both the Hydraulic Multiplexer and Controlled Pressure Fuel Nozzle system were cleared for engine test. The Fuel Multiplexer was cleared for combustor rig test which must be followed by an engine test to achieve full maturation.

Thermal analysis of cylindrical natural-gas steam reformer for 5 kW PEMFC

NASA Astrophysics Data System (ADS)

Jo, Taehyun; Han, Junhee; Koo, Bonchan; Lee, Dohyung

2016-11-01

The thermal characteristics of a natural-gas based cylindrical steam reformer coupled with a combustor are investigated for the use with a 5 kW polymer electrolyte membrane fuel cell. A reactor unit equipped with nickel-based catalysts was designed to activate the steam reforming reaction without the inclusion of high-temperature shift and low-temperature shift processes. Reactor temperature distribution and its overall thermal efficiency depend on various inlet conditions such as the equivalence ratio, the steam to carbon ratio (SCR), and the fuel distribution ratio (FDR) into the reactor and the combustor components. These experiments attempted to analyze the reformer's thermal and chemical properties through quantitative evaluation of product composition and heat exchange between the combustor and the reactor. FDR is critical factor in determining the overall performance as unbalanced fuel injection into the reactor and the combustor deteriorates overall thermal efficiency. Local temperature distribution also influences greatly on the fuel conversion rate and thermal efficiency. For the experiments, the operation conditions were set as SCR was in range of 2.5-4.0 and FDR was in 0.4-0.7 along with equivalence ratio of 0.9-1.1; optimum results were observed for FDR of 0.63 and SCR of 3.0 in the cylindrical steam reformer.

Performance of high mach number scramjets - Tunnel vs flight

NASA Astrophysics Data System (ADS)

Landsberg, Will O.; Wheatley, Vincent; Smart, Michael K.; Veeraragavan, Ananthanarayanan

2018-05-01

While typically analysed through ground-based impulse facilities, scramjets experience significant heating loads in flight, raising engine wall temperatures and the fuel used to cool them beyond standard laboratory conditions. Hence, the present work numerically compares an access-to-space scramjet's performance at both these conditions. The Mach 12 Rectangular-to-Elliptical Shape-Transitioning scramjet flow path is examined via three-dimensional and chemically reacting Reynolds-averaged Navier-Stokes solutions. Flight operation is modelled through 800 K and 1800 K inlet and combustor walls respectively, while fuel is injected at both inlet- and combustor-based stations at 1000 K stagnation temperature. Room temperature walls and fuel plena model shock tunnel conditions. Mixing and combustion performance indicates that while flight conditions promote rapid mixing, high combustor temperatures inhibit the completion of reaction pathways, with reactant dissociation reducing chemical heat release by 16%. However, the heated walls in flight ensured 28% less energy was absorbed by the walls. While inlet fuel injection promotes robust burning of combustor-injected fuel, premature ignition upon the inlet in flight suggests these injectors should be moved further downstream. Coupled with counteracting differences in heat release and loss to the walls, the optimal engine design for flight may differ considerably from that which gives the best performance in the tunnel.

Emission Reduction of Fuel-Staged Aircraft Engine Combustor Using an Additional Premixed Fuel Nozzle.

PubMed

Yamamoto, Takeshi; Shimodaira, Kazuo; Yoshida, Seiji; Kurosawa, Yoji

2013-03-01

The Japan Aerospace Exploration Agency (JAXA) is conducting research and development on aircraft engine technologies to reduce environmental impact for the Technology Development Project for Clean Engines (TechCLEAN). As a part of the project, combustion technologies have been developed with an aggressive target that is an 80% reduction over the NO x threshold of the International Civil Aviation Organization (ICAO) Committee on Aviation Environmental Protection (CAEP)/4 standard. A staged fuel nozzle with a pilot mixer and a main mixer was developed and tested using a single-sector combustor under the target engine's landing and takeoff (LTO) cycle conditions with a rated output of 40 kN and an overall pressure ratio of 25.8. The test results showed a 77% reduction over the CAEP/4 NO x standard. However, the reduction in smoke at thrust conditions higher than the 30% MTO condition and of CO emission at thrust conditions lower than the 85% MTO condition are necessary. In the present study, an additional fuel burner was designed and tested with the staged fuel nozzle in a single-sector combustor to control emissions. The test results show that the combustor enables an 82% reduction in NO x emissions relative to the ICAO CAEP/4 standard and a drastic reduction in smoke and CO emissions.

Combustion characteristics of hydrogen. Carbon monoxide based gaseous fuels

NASA Technical Reports Server (NTRS)

Notardonato, J. J.; White, D. J.; Kubasco, A. J.; Lecren, R. T.

1981-01-01

An experimental rig program was conducted with the objective of evaluating the combuston performance of a family of fuel gases based on a mixture of hydrogen and carbon monoxide. These gases, in addition to being members of a family, were also representative of those secondary fuels that could be produced from coal by various gasification schemes. In particular, simulated Winkler, Lurgi, and Blue-water low and medium energy content gases were used as fuels in the experimental combustor rig. The combustor used was originally designed as a low NOx rich-lean system for burning liquid fuels with high bound nitrogen levels. When used with the above gaseous fuels this combustor was operated in a lean-lean mode with ultra long residence times. The Blue-water gas was also operated in a rich-lean mode. The results of these tests indicate the possibility of the existence of an 'optimum' gas turbine hydrogen - carbon monoxide based secondary fuel. Such a fuel would exhibit NOx and high efficiency over the entire engine operating range. It would also have sufficient stability range to allow normal light-off and engine acceleration. Solar Turbines Incorporated would like to emphasize that the results presented here have been obtained with experimental rig combustors. The technologies generated could, however, be utilized in future commercial gas turbines.

Special Issue on Creativity at Work.

ERIC Educational Resources Information Center

Donnelly, Brian; And Others

1994-01-01

Special issue includes "Creativity at the Workplace" (Donnelly); "Creativity Revisited" (Iandoli); interviews with 16 people who work in or teach industrial engineering, software, and graphic design; "On Creativity and Schooling" (Coppola, Iandoli); and "End Notes: What I Learned" (Iandoli). (SK)

Steam reformer with catalytic combustor

DOEpatents

Voecks, Gerald E.

1990-03-20

A steam reformer is disclosed having an annular steam reforming catalyst bed formed by concentric cylinders and having a catalytic combustor located at the center of the innermost cylinder. Fuel is fed into the interior of the catalytic combustor and air is directed at the top of the combustor, creating a catalytic reaction which provides sufficient heat so as to maintain the catalytic reaction in the steam reforming catalyst bed. Alternatively, air is fed into the interior of the catalytic combustor and a fuel mixture is directed at the top. The catalytic combustor provides enhanced radiant and convective heat transfer to the reformer catalyst bed.

Small gas turbine combustor experimental study - Compliant metal/ceramic liner and performance evaluation

NASA Technical Reports Server (NTRS)

Acosta, W. A.; Norgren, C. T.

1986-01-01

Combustor research relating to the development of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently underway at NASA Lewis. As part of this combustor research, a basic reverse-flow combustor has been used to investigate advanced liner wall cooling techniques. Liner temperature, performance, and exhaust emissions of the experimental combustor utilizing compliant metal/ceramic liners were determined and compared with three previously reported combustors that featured: (1)splash film-cooled liner walls; (2) transpiration cooled liner walls; and (3) counter-flow film cooled panels.

Small gas turbine combustor experimental study: Compliant metal/ceramic liner and performance evaluation

NASA Technical Reports Server (NTRS)

Acosta, W. A.; Norgren, C. T.

1986-01-01

Combustor research relating to the development of fuel efficient small gas turbine engines capable of meeting future commercial and military aviation needs is currently underway at NASA Lewis. As part of this combustor research, a basic reverse-flow combustor has been used to investigate advanced liner wall cooling techniques. Liner temperature, performance, and exhaust emissions of the experimental combustor utilizing compliant metal/ceramic liners were determined and compared with three previously reported combustors that featured: (1) splash film-cooled liner walls; (2) transpiration cooled liner walls; and (3) counter-flow film cooled panels.

Computations of soot and and NO sub x emissions from gas turbine combustors

NASA Technical Reports Server (NTRS)

Srivatsa, S. K.

1982-01-01

An analytical program was conducted to compute the soot and NOx emissions from a combustor and the radiation heat transfer to the combustor walls. The program involved the formulation of an emission and radiation model and the incorporation of this model into the Garrett 3-D Combustor Perfomance Computer Program. Computations were performed for the idle, cruise, and take-off conditions of a JT8D can combustor. The predicted soot and NOx emissions and the radiation heat transfer to the combustor walls agree reasonably well with the limited experimental data available.

Energy Efficient Engine: Combustor component performance program

NASA Technical Reports Server (NTRS)

Dubiel, D. J.

1986-01-01

The results of the Combustor Component Performance analysis as developed under the Energy Efficient Engine (EEE) program are presented. This study was conducted to demonstrate the aerothermal and environmental goals established for the EEE program and to identify areas where refinements might be made to meet future combustor requirements. In this study, a full annular combustor test rig was used to establish emission levels and combustor performance for comparison with those indicated by the supporting technology program. In addition, a combustor sector test rig was employed to examine differences in emissions and liner temperatures obtained during the full annular performance and supporting technology tests.

Steam reformer with catalytic combustor

NASA Technical Reports Server (NTRS)

Voecks, Gerald E. (Inventor)

1990-01-01

A steam reformer is disclosed having an annular steam reforming catalyst bed formed by concentric cylinders and having a catalytic combustor located at the center of the innermost cylinder. Fuel is fed into the interior of the catalytic combustor and air is directed at the top of the combustor, creating a catalytic reaction which provides sufficient heat so as to maintain the catalytic reaction in the steam reforming catalyst bed. Alternatively, air is fed into the interior of the catalytic combustor and a fuel mixture is directed at the top. The catalytic combustor provides enhanced radiant and convective heat transfer to the reformer catalyst bed.

Downhole steam generator with improved preheating/cooling features

DOEpatents

Donaldson, A. Burl; Hoke, Donald E.; Mulac, Anthony J.

1983-01-01

An apparatus for downhole steam generation employing dual-stage preheaters for liquid fuel and for the water. A first heat exchange jacket for the fuel surrounds the fuel/oxidant mixing section of the combustor assembly downstream of the fuel nozzle and contacts the top of the combustor unit of the combustor assembly, thereby receiving heat directly from the combustion of the fuel/oxidant. A second stage heat exchange jacket surrounds an upper portion of the oxidant supply line adjacent the fuel nozzle receiving further heat from the compression heat which results from pressurization of the oxidant. The combustor unit includes an inner combustor sleeve whose inner wall defines the combustion zone. The inner combustor sleeve is surrounded by two concentric water channels, one defined by the space between the inner combustor sleeve and an intermediate sleeve, and the second defined by the space between the intermediate sleeve and an outer cylindrical housing. The channels are connected by an annular passage adjacent the top of the combustor assembly and the countercurrent nature of the water flow provides efficient cooling of the inner combustor sleeve. An annular water ejector with a plurality of nozzles is provided to direct water downwardly into the combustor unit at the boundary of the combustion zone and along the lower section of the intermediate sleeve.

Downhole steam generator with improved preheating/cooling features. [Patent application

DOEpatents

Donaldson, A.B.; Hoke, D.E.; Mulac, A.J.

1980-10-10

An apparatus is described for downhole steam generation employing dual-stage preheaters for liquid fuel and for the water. A first heat exchange jacket for the fuel surrounds the fuel/oxidant mixing section of the combustor assembly downstream of the fuel nozzle and contacts the top of the combustor unit of the combustor assembly, thereby receiving heat directly from the combustion of the fuel/oxidant. A second stage heat exchange jacket surrounds an upper portion of the oxidant supply line adjacent the fuel nozzle receiving further heat from the compression heat which results from pressurization of the oxidant. The combustor unit includes an inner combustor sleeve whose inner wall defines the combustion zone. The inner combustor sleeve is surrounded by two concentric water channels, one defined by the space between the inner combustor sleeve and an intermediate sleeve, and the second defined by the space between the intermediate sleeve and an outer cylindrical housing. The channels are connected by an annular passage adjacent the top of the combustor assembly and the countercurrent nature of the water flow provides efficient cooling of the inner combustor sleeve. An annular water ejector with a plurality of nozzles is provided to direct water downwardly into the combustor unit at the boundary of the combustion zone and along the lower section of the intermediate sleeve.

Developing Instrumentation for Assessing Creativity in Engineering Design

ERIC Educational Resources Information Center

Denson, Cameron D.; Buelin, Jennifer K.; Lammi, Matthew D.; D'Amico, Susan

2015-01-01

A perceived inability to assess creative attributes of students' work has often precluded creativity instruction in the classroom. The Consensual Assessment Technique (CAT) has shown promise in a variety of domains for its potential as a valid and reliable means of creativity assessment. Relying upon an operational definition of creativity and a…

Designing EEG Neurofeedback Procedures to Enhance Open-Ended versus Closed-Ended Creative Potentials

ERIC Educational Resources Information Center

Lin, Wei-Lun; Shih, Yi-Ling

2016-01-01

Recent empirical evidence demonstrated that open-ended creativity (which refers to creativity measures that require various and numerous responses, such as divergent thinking) correlated with alpha brain wave activation, whereas closed-ended creativity (which refers to creativity measures that ask for one final correct answer, such as insight…

Enhancing pre-service physics teachers' creative thinking skills through HOT lab design

NASA Astrophysics Data System (ADS)

Malik, Adam; Setiawan, Agus; Suhandi, Andi; Permanasari, Anna

2017-08-01

A research on the implementation of HOT (Higher Order Thinking) Laboratory has been carried out. This research is aimed to compare increasing of creative thinking skills of pre-service physics teachers who receive physics lesson with HOT Lab and with verification lab for the topic of electric circuit. This research used a quasi-experiment methods with control group pretest-posttest design. The subject of the research is 40 Physics Education pre-service physics teachers of UIN Sunan Gunung Djati Bandung. Research samples were selected by class random sampling technique. Data on pre-service physics teachers' creative thinking skills were collected using test of creative thinking skills in the form of essay. The results of the research reveal that average of N-gain of creative thinking skills are <0,69> for pre-service physics teachers who received lesson with HOT Lab design and <0,39> for pre-service physics teachers who received lesson with verification lab, respectively. Therefore, we conclude that application of HOT Lab design is more effective to increase creative thinking skills in the lesson of electric circuit.

Using a Moodle-Based Professional Development Program to Train Science Teachers to Teach for Creativity and its Effectiveness on their Teaching Practices

NASA Astrophysics Data System (ADS)

Al-Balushi, Sulaiman M.; Al-Abdali, Nasser S.

2015-08-01

This study describes a distance learning professional development program that we designed for the purpose of training science teachers to teach for creativity. The Moodle platform was used to host the training. To ensure that trainees would benefit from this distance learning program, we designed the instructional activities according to the Community of Inquiry framework, which consists of three main elements: cognitive presence, teaching presence and social presence. Nineteen science teachers in Oman engaged in the training, which lasted for 36 working days. To measure the effectiveness of the training program on science teachers' instructional practices related to teaching for creativity, we used a pre-post one-group quasi-experimental design. An observation form was used to assess and document participants' practices. Paired t test results showed that there was a statistically significant improvement in science teachers' practices related to teaching for creativity. During the implementation of the training program, we observed that cognitive presence and teaching presence were the two most successful elements of the program. The training program involved participants in different instructional activities which were designed to help them understand the role of creativity in science; a wide range of instructional techniques designed to nurture students' creativity was discussed. The program also provided participants with opportunities to relate their practices to teaching for creativity and to design and implement lesson plans geared toward teaching for creativity. However, the social presence element was not satisfying. Participants' virtual interactions with each other and their engagement in online discussion forums were limited. This paper provides some recommendations to overcome such pitfalls.

Creative inventive design and research

NASA Astrophysics Data System (ADS)

Kerley, James J.

1994-06-01

This paper is a summary of a course given at the Goddard Space Flight Center for graduate engineers entitled 'Creative Inventive Design and Research.' This course strikes at the heart of the problem as it describes the thinking process itself before it goes deeper into the design process as a structured method for performing creative design. Many problem examples and figures are presented in a form that should make clear to all students what this process is and how it can be used.

Creative inventive design and research

NASA Technical Reports Server (NTRS)

Kerley, James J.

1994-01-01

This paper is a summary of a course given at the Goddard Space Flight Center for graduate engineers entitled 'Creative Inventive Design and Research.' This course strikes at the heart of the problem as it describes the thinking process itself before it goes deeper into the design process as a structured method for performing creative design. Many problem examples and figures are presented in a form that should make clear to all students what this process is and how it can be used.

Liquid oxygen/liquid hydrogen auxiliary power system thruster investigation

NASA Technical Reports Server (NTRS)

Eberle, E. E.; Kusak, L.

1979-01-01

The design, fabrication, and demonstration of a 111 newton (25 lb) thrust, integrated auxiliary propulsion system (IAPS) thruster for use with LH2/LO2 propellants is described. Hydrogen was supplied at a temperature range of 22 to 33 K (40 to 60 R), and oxygen from 89 to 122 K (160 to 220 R). The thruster was designed to operate in both pulse mode and steady-state modes for vehicle attitude control, space maneuvering, and as an abort backup in the event of failure of the main propulsion system. A dual-sleeve, tri-axial injection system was designed that utilizes a primary injector/combustor where 100 percent of the oxygen and 8 percent of the hydrogen is introduced; a secondary injector/combustor where 45 percent of the hydrogen is introduced to mix with the primary combustor gases; and a boundary layer injector that uses the remaining 45 percent of the hydrogen to cool the thrust throat/nozzle design. Hot-fire evaluation of this thruster with a BLC injection distance of 2.79 cm (1.10 in.) indicated that a specific impulse value of 390 sec can be attained using a coated molybdenum thrust chamber. Pulse mode tests indicated that a chamber pressure buildup to 90 percent thrust can be achieved in a time on the order of 48 msec. Some problems were encountered in achieving ignition of each pulse during pulse trains. This was interpreted to indicate that a higher delivered spark energy level ( 100 mJ) would be required to maintain ignition reliability of the plasma torch ignition system under the extra 'cold' conditions resulting during pulsing.

Preliminary design of an alternate high-temperature turbine. A topical report for Phase II of the High-Temperature-Turbine Technology Program

DOE Office of Scientific and Technical Information (OSTI.GOV)

Strough, R.I.

The feasibility of designing a convectively air-cooled turbine to operate in the environment of a 3000/sup 0/F combustor exit temperature with maximum turbine airfoil metal temperatures held to 1500/sup 0/F was established. The United Technologies-Kraftwerk Union V84.3 gas turbine design was used as the basic configuration for the design of the 3000/sup 0/F turbine. Turbine cooling requirements were determined based on the use of the modified V84.3 type silo combustor with a pattern factor of 0.1. The convective air-cooling technology levels in terms of cooling effectiveness required to satisfy the airfoil cooling requirements were identified. Cooling schemes and fabrication technologiesmore » required are discussed. Turbine airfoil cooling technology levels required for the 3000/sup 0/F engine were selected. The performance of the 3000/sup 0/F convectively air-cooled gas turbine in simple and combined cycle was calculated. The 3000/sup 0/F gas turbine combined-cycle system provides an increase in power of 61% and a decrease in heat rate of 10% compared to a similar system with a combustor exit temperature of 2210/sup 0/F and the same airflow. The development of a successful 3000/sup 0/F convectively air-cooled turbine can be accomplished with a reasonable design and fabrication development effort on the cooled turbine airfoils. Use of the convectively air-cooled turbine provides the transfer of technology from extensive aircraft engines developed programs and operating experience to industrial gas turbines. It eliminates the requirement for large investments in alternate cooling techniques tailored specifically for industrial engines which offer no additional benefits.« less

Advanced Gas Turbine (AGT) technology development

NASA Technical Reports Server (NTRS)

1983-01-01

A 74.5 kW (100 hp) automotive gas turbine was evaluated. The engine structure, bearings, oil system, and electronics were demonstrated and no shaft dynamics or other vibration problem were encountered. Areas identified during the five tests are the scroll retention features, and transient thermal deflection of turbine backplates. Modifications were designed. Seroll retention is addressed by modifying the seal arrangement in front of the gasifier turbine assembly, which will increase the pressure load on the scroll in the forward direction and thereby increase the retention forces. the backplate thermal deflection is addressed by geometric changes and thermal insulation to reduce heat input. Combustor rig proof testing of two ceramic combustor assemblies was completed. The combustor was modified to incorporate slots and reduce sharp edges, which should reduce thermal stresses. The development work focused on techniques to sinter these barrier materials onto the ceramic rotors with successes for both material systems. Silicon carbide structural parts, including engine configuration gasifier rotors (ECRs), preliminary gasifier scroll parts, and gasifier and power turbine vanes are fabricated.

Parametric performance analysis of steam-injected gas turbine with a thermionic-energy-converter-lined combustor

NASA Technical Reports Server (NTRS)

Choo, Y. K.; Burns, R. K.

1982-01-01

The performance of steam-injected gas turbines having combustors lined with thermionic energy converters (STIG/TEC systems) was analyzed and compared with that of two baseline systems; a steam-injected gas turbine (without a TEC-lined combustor) and a conventional combined gas turbine/steam turbine cycle. Common gas turbine parameters were assumed for all of the systems. Two configurations of the STIG/TEC system were investigated. In both cases, steam produced in an exhaust-heat-recovery boiler cools the TEC collectors. It is then injected into the gas combustion stream and expanded through the gas turbine. The STIG/TEC system combines the advantage of gas turbine steam injection with the conversion of high-temperature combustion heat by TEC's. The addition of TEC's to the baseline steam-injected gas turbine improves both its efficiency and specific power. Depending on system configuration and design parameters, the STIG/TEC system can also achieve higher efficiency and specific power than the baseline combined cycle.

Full-Scale Turbofan Engine Noise-Source Separation Using a Four-Signal Method

NASA Technical Reports Server (NTRS)

Hultgren, Lennart S.; Arechiga, Rene O.

2016-01-01

Contributions from the combustor to the overall propulsion noise of civilian transport aircraft are starting to become important due to turbofan design trends and expected advances in mitigation of other noise sources. During on-ground, static-engine acoustic tests, combustor noise is generally sub-dominant to other engine noise sources because of the absence of in-flight effects. Consequently, noise-source separation techniques are needed to extract combustor-noise information from the total noise signature in order to further progress. A novel four-signal source-separation method is applied to data from a static, full-scale engine test and compared to previous methods. The new method is, in a sense, a combination of two- and three-signal techniques and represents an attempt to alleviate some of the weaknesses of each of those approaches. This work is supported by the NASA Advanced Air Vehicles Program, Advanced Air Transport Technology Project, Aircraft Noise Reduction Subproject and the NASA Glenn Faculty Fellowship Program.

A study on the Jordanian oil shale resources and utilization

NASA Astrophysics Data System (ADS)

Sakhrieh, Ahmad; Hamdan, Mohammed

2012-11-01

Jordan has significant oil shale deposits occurring in 26 known localities. Geological surveys indicate that the existing deposits underlie more than 60% of Jordan's territory. The resource consists of 40 to 70 billion tones of oil shale, which may be equivalent to more than 5 million tones of shale oil. Since the 1960s, Jordan has been investigating economical and environmental methods for utilizing oil shale. Due to its high organic content, is considered a suitable source of energy. This paper introduces a circulating fluidized bed combustor that simulates the behavior of full scale municipal oil shale combustors. The inside diameter of the combustor is 500 mm, the height is 3000 mm. The design of the CFB is presented. The main parameters which affect the combustion process are elucidated in the paper. The size of the laboratory scale fluidized bed reactor is 3 kW, which corresponds to a fuel-feeding rate of approximately 1.5 kg/h.

Emission calculations for a scramjet powered hypersonic transport

NASA Technical Reports Server (NTRS)

Lezberg, E. A.

1973-01-01

Calculations of exhaust emissions from a scramjet powered hypersonic transport burning hydrogen fuel were performed over a range of Mach numbers of 5 to 12 to provide input data for wake mixing calculations and forecasts of future levels of pollutants in the stratosphere. The calculations were performed utilizing a one-dimensional chemical kinetics computer program for the combustor and exhaust nozzle of a fixed geometry dual-mode scramjet engine. Inlet conditions to the combustor and engine size was based on a vehicle of 227,000 kg (500,000 lb) gross take of weight with engines sized for Mach 8 cruise. Nitric oxide emissions were very high for stoichiometric engine operation but for Mach 6 cruise at reduced equivalence ratio are in the range predicted for an advanced supersonic transport. Combustor designs which utilize fuel staging and rapid expansion to minimize residence time at high combustion temperatures were found to be effective in preventing nitric oxide formation from reaching equilibrium concentrations.

Effect of Flame Stabilizer Design on Performance and Exhaust Pollutants of a Two-Row Swirl-Can Combustor Operated to Near-Stoichiometric Conditions

NASA Technical Reports Server (NTRS)

Biaglow, James A.; Trout, Arthur M.

1977-01-01

Emissions and performance characteristics were determined for two full annulus modular combustors operated to near stoichiometric fuel air ratios. The tests were conducted to obtain stoichiometric data at inlet air temperatures from 756 to 894 K and to determine the effects of a flat plate circular flame stabilizer with upstream fuel injection and a contraswirl flame stabilizer with downstream fuel injection. Levels of unburned hydrocarbons were below 0.50 gram per kilogram of fuel for both combustors and thus there was no detectable difference in the two methods of fuel injection. The contraswirl flame stabilizer did not produce the level of mixing obtained with a flat plate circular flame stabilizer. It did produce higher levels of oxides of nitrogen, which peaked at a fuel air ratio of 0.037. For the flat plate circular flame stabilizer, oxides of nitrogen emission levels were still increasing with fuel air ratio to the maximum tested value of 0.045.

Low Emission Hydrogen Combustors for Gas Turbines Using Lean Direct Injection

NASA Technical Reports Server (NTRS)

Marek, C. John; Smith, Timothy D.; Kundu, Krishna

2005-01-01

One of the key technology challenges for the use of hydrogen in gas turbine engines is the performance of the combustion system, in particular the fuel injectors. To investigate the combustion performance of gaseous hydrogen fuel injectors flame tube combustor experiments were performed. Tests were conducted to measure the nitrogen oxide (NOx) emissions and combustion performance at inlet conditions of 600 to 1000 deg F, 60 to 200 pounds per square inch absolute (psia), and equivalence ratios up to 0.48. All the injectors were based on Lean Direct Injection (LDI) technology with multiple injection points and quick mixing. One challenge to hydrogen based premixing combustion systems is flashback since hydrogen has a reaction rate over seven times that of Jet-A. To reduce the risk, design mixing times were kept short and velocities high to minimize flashback. Five fuel injector designs were tested in 2.5 and 3.5-in. diameter flame tubes with non-vitiated heated air and gaseous hydrogen. Data is presented on measurements of NOx emissions and combustion efficiency for the hydrogen injectors at 1.0, 3.125, and 5.375 in. from the injector face. Results show that for some configurations, NOx emissions are comparable to that of state of the art Jet-A LDI combustor concepts.

Low-Emission Hydrogen Combustors for Gas Turbines Using Lean Direct Injection

NASA Technical Reports Server (NTRS)

Marek, C. John; Smith, Timothy D.; Kundu, Krishna

2007-01-01

One of the key technology challenges for the use of hydrogen in gas turbine engines is the performance of the combustion system, in particular the fuel injectors. To investigate the combustion performance of gaseous hydrogen fuel injectors flame tube combustor experiments were performed. Tests were conducted to measure the nitrogen oxide (NO(x)) emissions and combustion performance at inlet conditions of 588 to 811 K, 0.4 to 1.4 MPa, and equivalence ratios up to 0.48. All the injectors were based on Lean Direct Injection (LDI) technology with multiple injection points and quick mixing. One challenge to hydrogen-based premixing combustion systems is flashback since hydrogen has a reaction rate over 7 times that of Jet-A. To reduce the risk, design mixing times were kept short and velocities high to minimize flashback. Five fuel injector designs were tested in 6.35- and 8.9-cm-diameter flame tubes with non-vitiated heated air and gaseous hydrogen. Data is presented on measurements of NO(x) emissions and combustion efficiency for the hydrogen injectors at 2.540, 7.937, and 13.652 cm from the injector face. Results show that for some configurations, NO(x) emissions are comparable to that of state of the art Jet-A LDI combustor concepts.

Progress towards an Optimization Methodology for Combustion-Driven Portable Thermoelectric Power Generation Systems

DOE Office of Scientific and Technical Information (OSTI.GOV)

Krishnan, Shankar; Karri, Naveen K.; Gogna, Pawan K.

2012-03-13

Enormous military and commercial interests exist in developing quiet, lightweight, and compact thermoelectric (TE) power generation systems. This paper investigates design integration and analysis of an advanced TE power generation system implementing JP-8 fueled combustion and thermal recuperation. Design and development of a portable TE power system using a JP-8 combustor as a high temperature heat source and optimal process flows depend on efficient heat generation, transfer, and recovery within the system are explored. Design optimization of the system required considering the combustion system efficiency and TE conversion efficiency simultaneously. The combustor performance and TE sub-system performance were coupled directlymore » through exhaust temperatures, fuel and air mass flow rates, heat exchanger performance, subsequent hot-side temperatures, and cold-side cooling techniques and temperatures. Systematic investigation of this system relied on accurate thermodynamic modeling of complex, high-temperature combustion processes concomitantly with detailed thermoelectric converter thermal/mechanical modeling. To this end, this work reports on design integration of systemlevel process flow simulations using commercial software CHEMCADTM with in-house thermoelectric converter and module optimization, and heat exchanger analyses using COMSOLTM software. High-performance, high-temperature TE materials and segmented TE element designs are incorporated in coupled design analyses to achieve predicted TE subsystem level conversion efficiencies exceeding 10%. These TE advances are integrated with a high performance microtechnology combustion reactor based on recent advances at the Pacific Northwest National Laboratory (PNNL). Predictions from this coupled simulation established a basis for optimal selection of fuel and air flow rates, thermoelectric module design and operating conditions, and microtechnology heat-exchanger design criteria. This paper will discuss this simulation process that leads directly to system efficiency power maps defining potentially available optimal system operating conditions and regimes. This coupled simulation approach enables pathways for integrated use of high-performance combustor components, high performance TE devices, and microtechnologies to produce a compact, lightweight, combustion driven TE power system prototype that operates on common fuels.« less

The Distracting Effects of Music on the Cognitive Test Performance of Creative and Non-Creative Individuals

ERIC Educational Resources Information Center

Doyle, Maddie; Furnham, Adrian

2012-01-01

This study examined the effect of background music upon performance of creative and non-creative individuals on a reading comprehension task. In the presence of musical distraction and silence, 54 individuals (27 creative) carried out reading comprehension tasks in a repeated measures design. An interaction was predicted, such that musical…

System and method for controlling a combustor assembly

DOEpatents

York, William David; Ziminsky, Willy Steve; Johnson, Thomas Edward; Stevenson, Christian Xavier

2013-03-05

A system and method for controlling a combustor assembly are disclosed. The system includes a combustor assembly. The combustor assembly includes a combustor and a fuel nozzle assembly. The combustor includes a casing. The fuel nozzle assembly is positioned at least partially within the casing and includes a fuel nozzle. The fuel nozzle assembly further defines a head end. The system further includes a viewing device configured for capturing an image of at least a portion of the head end, and a processor communicatively coupled to the viewing device, the processor configured to compare the image to a standard image for the head end.

Learning Creativity in the Client-Agency Relationship

ERIC Educational Resources Information Center

Suh, Taewon; Jung, Jae C.; Smith, Bruce L.

2012-01-01

Purpose: This study aims to investigate creativity-related determinants of learning in the context of business-to-business services and client-agency relationships. Design/methodology/approach: The research model includes client encouragement, agency creativity, campaign creativity, and perceived performance. The study involved conducting a…

Computers and Creativity.

ERIC Educational Resources Information Center

Ten Dyke, Richard P.

1982-01-01

A traditional question is whether or not computers shall ever think like humans. This question is redirected to a discussion of whether computers shall ever be truly creative. Creativity is defined and a program is described that is designed to complete creatively a series problem in mathematics. (MP)

Convective Heat Transfer with and without Film Cooling in High Temperature, Fuel Rich and Lean Environments

NASA Astrophysics Data System (ADS)

Greiner, Nathan J.

Modern turbine engines require high turbine inlet temperatures and pressures to maximize thermal efficiency. Increasing the turbine inlet temperature drives higher heat loads on the turbine surfaces. In addition, increasing pressure ratio increases the turbine coolant temperature such that the ability to remove heat decreases. As a result, highly effective external film cooling is required to reduce the heat transfer to turbine surfaces. Testing of film cooling on engine hardware at engine temperatures and pressures can be exceedingly difficult and expensive. Thus, modern studies of film cooling are often performed at near ambient conditions. However, these studies are missing an important aspect in their characterization of film cooling effectiveness. Namely, they do not model effect of thermal property variations that occur within the boundary and film cooling layers at engine conditions. Also, turbine surfaces can experience significant radiative heat transfer that is not trivial to estimate analytically. The present research first computationally examines the effect of large temperature variations on a turbulent boundary layer. Subsequently, a method to model the effect of large temperature variations within a turbulent boundary layer in an environment coupled with significant radiative heat transfer is proposed and experimentally validated. Next, a method to scale turbine cooling from ambient to engine conditions via non-dimensional matching is developed computationally and the experimentally validated at combustion temperatures. Increasing engine efficiency and thrust to weight ratio demands have driven increased combustor fuel-air ratios. Increased fuel-air ratios increase the possibility of unburned fuel species entering the turbine. Alternatively, advanced ultra-compact combustor designs have been proposed to decrease combustor length, increase thrust, or generate power for directed energy weapons. However, the ultra-compact combustor design requires a film cooled vane within the combustor. In both these environments, the unburned fuel in the core flow encounters the oxidizer rich film cooling stream, combusts, and can locally heat the turbine surface rather than the intended cooling of the surface. Accordingly, a method to quantify film cooling performance in a fuel rich environment is prescribed. Finally, a method to film cool in a fuel rich environment is experimentally demonstrated.

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

NASA Technical Reports Server (NTRS)

Childs, J. Howard; McCafferty, Richard J.

1948-01-01

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.

Experimental Clean Combustor Program (ECCP), phase 3. [commercial aircraft turbofan engine tests with double annular combustor

NASA Technical Reports Server (NTRS)

Gleason, C. C.; Bahr, D. W.

1979-01-01

A double annular advanced technology combustor with low pollutant emission levels was evaluated in a series of CF6-50 engine tests. Engine lightoff was readily obtained and no difficulties were encountered with combustor staging. Engine acceleration and deceleration were smooth, responsive and essentially the same as those obtainable with the CF6-50 combustor. The emission reductions obtained in carbon monoxide, hydrocarbons, and nitrogen oxide levels were 55, 95, and 30 percent, respectively, at an idle power setting of 3.3 percent of takeoff power on an EPA parameter basis. Acceptable smoke levels were also obtained. The exit temperature distribution of the combustor was found to be its major performance deficiency. In all other important combustion system performance aspects, the combustor was found to be generally satisfactory.

Vortex combustor for low NOX emissions when burning lean premixed high hydrogen content fuel

DOEpatents

Steele, Robert C; Edmonds, Ryan G; Williams, Joseph T; Baldwin, Stephen P

2012-11-20

A trapped vortex combustor. The trapped vortex combustor is configured for receiving a lean premixed gaseous fuel and oxidant stream, where the fuel includes hydrogen gas. The trapped vortex combustor is configured to receive the lean premixed fuel and oxidant stream at a velocity which significantly exceeds combustion flame speed in a selected lean premixed fuel and oxidant mixture. The combustor is configured to operate at relatively high bulk fluid velocities while maintaining stable combustion, and low NOx emissions. The combustor is useful in gas turbines in a process of burning synfuels, as it offers the opportunity to avoid use of diluent gas to reduce combustion temperatures. The combustor also offers the possibility of avoiding the use of selected catalytic reaction units for removal of oxides of nitrogen from combustion gases exiting a gas turbine.

Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel

DOEpatents

Steele, Robert C [Woodinville, WA; Edmonds, Ryan G [Renton, WA; Williams, Joseph T [Kirkland, WA; Baldwin, Stephen P [Winchester, MA

2009-10-20

A trapped vortex combustor. The trapped vortex combustor is configured for receiving a lean premixed gaseous fuel and oxidant stream, where the fuel includes hydrogen gas. The trapped vortex combustor is configured to receive the lean premixed fuel and oxidant stream at a velocity which significantly exceeds combustion flame speed in a selected lean premixed fuel and oxidant mixture. The combustor is configured to operate at relatively high bulk fluid velocities while maintaining stable combustion, and low NOx emissions. The combustor is useful in gas turbines in a process of burning synfuels, as it offers the opportunity to avoid use of diluent gas to reduce combustion temperatures. The combustor also offers the possibility of avoiding the use of selected catalytic reaction units for removal of oxides of nitrogen from combustion gases exiting a gas turbine.

Multifuel evaluation of rich/quench/lean combustor

NASA Technical Reports Server (NTRS)

Novick, A. S.; Troth, D. L.; Notardonato, J.

1982-01-01

Test results on the RQL low NO(x) industrial gas turbine engine are reported. The air-staged combustor comprises an initial rich burning zone, followed by a quench zone, and a lean reaction and dilution zone. The combustor was tested as part of the DoE/NASA program to define the technology for developing a durable, low-emission gas turbine combustor capable of operation with minimally processed petroleum residual, synthetic, or low/mid-heating value gaseous fuels. The properties of three liquid and two gaseous fuels burned in the combustor trials are detailed. The combustor featured air staging, variable geometry, and generative/convective cooling. The lean/rich mixtures could be varied in zones simultaneously or separately while maintaining a specified pressure drop. Low NO(x) and smoke emissions were produced with each fuel burned, while high combustor efficiencies were obtained.

78 FR 21871 - Approval and Promulgation of State Plans for Designated Facilities and Pollutants: Connecticut...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-04-12

... promulgated on May 10, 2006. Connecticut DEEP's State Plan is for implementing and enforcing provisions at... and Promulgation of State Plans for Designated Facilities and Pollutants: Connecticut; 111(d)/129 Revised State Plan for Large and Small Municipal Waste Combustors AGENCY: Environmental Protection Agency...

Effects of Spaced Training in Creative Imagination and Delayed Posttesting on Originality.

ERIC Educational Resources Information Center

Masten, William G.; Hairston, Susan A.

While there is increasing evidence that creativity can be improved through training, the spacing of the training has not been studied. This study assessed the effect of spaced training on the use of creative imagination in 110 undergraduate students. The research design was a randomized delayed posttest-only design. The independent variable was…

The Role of Flow Experience and CAD Tools in Facilitating Creative Behaviours for Architecture Design Students

ERIC Educational Resources Information Center

Dawoud, Husameddin M.; Al-Samarraie, Hosam; Zaqout, Fahed

2015-01-01

This study examined the role of flow experience in intellectual activity with an emphasis on the relationship between flow experience and creative behaviour in design using CAD. The study used confluence and psychometric approaches because of their unique abilities to depict a clear image of creative behaviour. A cross-sectional study…

Fuel Spray Diagnostics

NASA Technical Reports Server (NTRS)

Humenik, F. M.; Bosque, M. A.

1983-01-01

Fundamental experimental data base for turbulent flow mixing models is provided and better prediction of the more complex turbulent chemical reacting flows. Analytical application to combustor design is provided and a better fundamental understanding of the combustion process.

Aerothermal modeling program, phase 2. Element B: Flow interaction experiment

NASA Technical Reports Server (NTRS)

Nikjooy, M.; Mongia, H. C.; Murthy, S. N. B.; Sullivan, J. P.

1986-01-01

The design process was improved and the efficiency, life, and maintenance costs of the turbine engine hot section was enhanced. Recently, there has been much emphasis on the need for improved numerical codes for the design of efficient combustors. For the development of improved computational codes, there is a need for an experimentally obtained data base to be used at test cases for the accuracy of the computations. The purpose of Element-B is to establish a benchmark quality velocity and scalar measurements of the flow interaction of circular jets with swirling flow typical of that in the dome region of annular combustor. In addition to the detailed experimental effort, extensive computations of the swirling flows are to be compared with the measurements for the purpose of assessing the accuracy of current and advanced turbulence and scalar transport models.

Computational models for the analysis/design of hypersonic scramjet components. I - Combustor and nozzle models

NASA Technical Reports Server (NTRS)

Dash, S. M.; Sinha, N.; Wolf, D. E.; York, B. J.

1986-01-01

An overview of computational models developed for the complete, design-oriented analysis of a scramjet propulsion system is provided. The modular approach taken involves the use of different PNS models to analyze the individual propulsion system components. The external compression and internal inlet flowfields are analyzed by the SCRAMP and SCRINT components discussed in Part II of this paper. The combustor is analyzed by the SCORCH code which is based upon SPLITP PNS pressure-split methodology formulated by Dash and Sinha. The nozzle is analyzed by the SCHNOZ code which is based upon SCIPVIS PNS shock-capturing methodology formulated by Dash and Wolf. The current status of these models, previous developments leading to this status, and, progress towards future hybrid and 3D versions are discussed in this paper.

Combustor and combustor screech mitigation methods

DOEpatents

Kim, Kwanwoo; Johnson, Thomas Edward; Uhm, Jong Ho; Kraemer, Gilbert Otto

2014-05-27

The present application provides for a combustor for use with a gas turbine engine. The combustor may include a cap member and a number of fuel nozzles extending through the cap member. One or more of the fuel nozzles may be provided in a non-flush position with respect to the cap member.

Rigorous Mathematical Thinking Approach to Enhance Students’ Mathematical Creative and Critical Thinking Abilities

NASA Astrophysics Data System (ADS)

Hidayat, D.; Nurlaelah, E.; Dahlan, J. A.

2017-09-01

The ability of mathematical creative and critical thinking are two abilities that need to be developed in the learning of mathematics. Therefore, efforts need to be made in the design of learning that is capable of developing both capabilities. The purpose of this research is to examine the mathematical creative and critical thinking ability of students who get rigorous mathematical thinking (RMT) approach and students who get expository approach. This research was quasi experiment with control group pretest-posttest design. The population were all of students grade 11th in one of the senior high school in Bandung. The result showed that: the achievement of mathematical creative and critical thinking abilities of student who obtain RMT is better than students who obtain expository approach. The use of Psychological tools and mediation with criteria of intentionality, reciprocity, and mediated of meaning on RMT helps students in developing condition in critical and creative processes. This achievement contributes to the development of integrated learning design on students’ critical and creative thinking processes.

The Industrial Property Rights Education in Collaboration with the Creative Product Design Education

NASA Astrophysics Data System (ADS)

Tokoro, Tetsuro; Habuchi, Hitoe; Chonan, Isao

Recently, the Advanced Courses of Electronic System Engineering and Architecture and Civil Engineering of Gifu National College of Technology have introduced a creative subject, “Creative Engineering Practice”. In this subject, students study intellectual property rights. More specifically, they learn and practice industrial proprietary rights, procedures for obtaining a patent right, how to use Industrial Property Digital Library and so forth, along with the practice of creative product design. The industrial property rights education in collaboration with the creative product design education has been carried out by the cooperation of Japan Patent Office, Japan Institute of Invention and Innovation and a patent attorney. Through the instruction of the cooperative members, great educative results have been obtained. In this paper, we will describe the contents of the subject together with its items to pursue an upward spiral of progress.

Controlled pilot oxidizer for a gas turbine combustor

DOEpatents

Laster, Walter R.; Bandaru, Ramarao V.

2010-07-13

A combustor (22) for a gas turbine (10) includes a main burner oxidizer flow path (34) delivering a first portion (32) of an oxidizer flow (e.g., 16) to a main burner (28) of the combustor and a pilot oxidizer flow path (38) delivering a second portion (36) of the oxidizer flow to a pilot (30) of the combustor. The combustor also includes a flow controller (42) disposed in the pilot oxidizer flow path for controlling an amount of the second portion delivered to the pilot.

Teaching Practices in ipad-Classrooms: Alignment of Didactical Designs, Mobile Devices and Creativity

ERIC Educational Resources Information Center

Jahnke, Isa

2013-01-01

Creativity is socially constructed and is not an objective fact at all. How do teachers perceive students' creativity and how can they foster students' creative learning? From two case studies, one in higher education and a second on iPad-classrooms in schools, the paper reflects on didactical concepts for creativity using mobile devices.…

NOx reduction in catalytically stabilized thermal burners. Annual report, pril 1, 1988-March 31, 1989

DOE Office of Scientific and Technical Information (OSTI.GOV)

Pfefferle, L.D.

1989-09-01

Catalytically stabilized combustors can be designed to combine the high reaction rates of thermal combustors with low-NOx emissions. The objectives of the research are to understand why the CST burner has inherently low-NOx emissions and whether preexisting NOx can be reduced in-situ in the post-flame zone of a CST burner. Initial results indicate that reduced NOx emissions are, at least for some operating conditions, due to more than just the ability to stabilize combustion at low temperatures. The next phase of the investigation will focus on isothermal flow-tube kinetics studies to isolate catalytic and thermal effects.

Testing and Characterization of CMC Combustor Liners

NASA Technical Reports Server (NTRS)

Robinson, R. Craig; Verrilli, Michael J.

2003-01-01

Multiple combustor liner applications, both segmented and fully annular designs, have been configured for exposure in NASA's High Pressure Burner Rig (HPBR). The segmented liners were attached to the rig structure with SiC/SiC fasteners and exposed to simulated gas turbine conditions for nearly 200 hours. Test conditions included pressures of 6 atm., gas velocity of 42 m/s, and gas temperatures near 1450 C. The temperatures of both the cooled and combustion flow sides of the liners were measured using optical and contact measurement techniques. Minor weight loss was observed, but the liners remained structural sound, although damage was noted in some fasteners.

The perceived importance and the presence of creative potential in the health professional's work environment.

PubMed

Lukersmith, Sue; Burgess-Limerick, Robin

2013-01-01

The value of creative employees to an organisation's growth and innovative development, productivity, quality and sustainability is well established. This study examined the perceived relationship between creativity and work environment factors of 361 practicing health professionals, and whether these factors were present (realised) in their work environment. Job design (challenges, team work, task rotation, autonomy) and leadership (coaching supervisor, time for thinking, creative goals, recognition and incentives for creative ideas and results) were perceived as the most important factors for stimulating creativity. There was room for improvement of these in the work environment. Many aspects of the physical work environment were less important. Public health sector employers and organisations should adopt sustainable strategies which target the important work environment factors to support employee creativity and so enhance service quality, productivity, performance and growth. Implications of the results for ergonomists and workplace managers are discussed with a participatory ergonomics approach recommended. Creative employees are important to an organisation's innovation, productivity and sustainability. The survey identified health professionals perceive a need to improve job design and leadership factors at work to enhance and support employee creativity. There are implications for organisations and ergonomists to investigate the creative potential of work environments.

Experimental clean combustor program, phase 3

NASA Technical Reports Server (NTRS)

Roberts, R.; Fiorentino, A.; Greene, W.

1977-01-01

A two-stage vortex burning and mixing combustor and associated fuel system components were successfully tested at steady state and transient operating conditions. The combustor exceeded the program goals for all three emissions species, with oxides of nitrogen 10 percent below the goal, carbon monoxide 26 percent below the goal, and total unburned hydrocarbons 75 percent below the goal. Relative to the JT9D-7 combustor, the oxides of nitrogen were reduced by 58 percent, carbon monoxide emissions were reduced by 69 percent, and total unburned hydrocarbons were reduced by 9 percent. The combustor efficiency and exit temperature profiles were comparable to those of production combustor. Acceleration and starting characteristics were deficient relative to the production engine.

Turbine combustor with fuel nozzles having inner and outer fuel circuits

DOEpatents

Uhm, Jong Ho; Johnson, Thomas Edward; Kim, Kwanwoo

2013-12-24

A combustor cap assembly for a turbine engine includes a combustor cap and a plurality of fuel nozzles mounted on the combustor cap. One or more of the fuel nozzles would include two separate fuel circuits which are individually controllable. The combustor cap assembly would be controlled so that individual fuel circuits of the fuel nozzles are operated or deliberately shut off to provide for physical separation between the flow of fuel delivered by adjacent fuel nozzles and/or so that adjacent fuel nozzles operate at different pressure differentials. Operating a combustor cap assembly in this fashion helps to reduce or eliminate the generation of undesirable and potentially harmful noise.

Lean-rich axial stage combustion in a can-annular gas turbine engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Laster, Walter R.; Szedlacsek, Peter

2016-06-14

An apparatus and method for lean/rich combustion in a gas turbine engine (10), which includes a combustor (12), a transition (14) and a combustor extender (16) that is positioned between the combustor (12) and the transition (14) to connect the combustor (12) to the transition (14). Openings (18) are formed along an outer surface (20) of the combustor extender (16). The gas turbine (10) also includes a fuel manifold (28) to extend along the outer surface (20) of the combustor extender (16), with fuel nozzles (30) to align with the respective openings (18). A method (200) for axial stage combustionmore » in the gas turbine engine (10) is also presented.« less

Method of making an aero-derivative gas turbine engine

DOE Office of Scientific and Technical Information (OSTI.GOV)

Wiebe, David J.

A method of making an aero-derivative gas turbine engine (100) is provided. A combustor outer casing (68) is removed from an existing aero gas turbine engine (60). An annular combustor (84) is removed from the existing aero gas turbine engine. A first row of turbine vanes (38) is removed from the existing aero gas turbine engine. A can annular combustor assembly (122) is installed within the existing aero gas turbine engine. The can annular combustor assembly is configured to accelerate and orient combustion gasses directly onto a first row of turbine blades of the existing aero gas turbine engine. Amore » can annular combustor assembly outer casing (108) is installed to produce the aero-derivative gas turbine engine (100). The can annular combustor assembly is installed within an axial span (85) of the existing aero gas turbine engine vacated by the annular combustor and the first row of turbine vanes.« less

40 CFR 60.53a - Standard for municipal waste combustor organics.

Code of Federal Regulations, 2014 CFR

2014-07-01

... Municipal Waste Combustors for Which Construction Is Commenced After December 20, 1989 and On or Before September 20, 1994 § 60.53a Standard for municipal waste combustor organics. (a) [Reserved] (b) On and after... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standard for municipal waste combustor...

40 CFR 60.55b - Standards for municipal waste combustor fugitive ash emissions.

Code of Federal Regulations, 2014 CFR

2014-07-01

... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... municipal waste combustor fugitive ash emissions. (a) On and after the date on which the initial performance...

40 CFR 60.55b - Standards for municipal waste combustor fugitive ash emissions.

Code of Federal Regulations, 2010 CFR

2010-07-01

... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standards for municipal waste combustor... municipal waste combustor fugitive ash emissions. (a) On and after the date on which the initial performance...

40 CFR 60.54a - Standard for municipal waste combustor acid gases.

Code of Federal Regulations, 2014 CFR

2014-07-01

... for Municipal Waste Combustors for Which Construction Is Commenced After December 20, 1989 and On or Before September 20, 1994 § 60.54a Standard for municipal waste combustor acid gases. (a)-(b) [Reserved... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standard for municipal waste combustor...

40 CFR 60.54b - Standards for municipal waste combustor operator training and certification.

Code of Federal Regulations, 2014 CFR

2014-07-01

... Standards of Performance for Large Municipal Waste Combustors for Which Construction is Commenced After... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... Standards for municipal waste combustor operator training and certification. (a) No later than the date 6...

40 CFR 60.53b - Standards for municipal waste combustor operating practices.

Code of Federal Regulations, 2014 CFR

2014-07-01

... Performance for Large Municipal Waste Combustors for Which Construction is Commenced After September 20, 1994... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standards for municipal waste combustor... municipal waste combustor operating practices. (a) On and after the date on which the initial performance...

Advanced composite combustor structural concepts program

NASA Technical Reports Server (NTRS)

Sattar, M. A.; Lohmann, R. P.

1984-01-01

An analytical study was conducted to assess the feasibility of and benefits derived from the use of high temperature composite materials in aircraft turbine engine combustor liners. The study included a survey and screening of the properties of three candidate composite materials including tungsten reinforced superalloys, carbon-carbon and silicon carbide (SiC) fibers reinforcing a ceramic matrix of lithium aluminosilicate (LAS). The SiC-LAS material was selected as offering the greatest near term potential primarily on the basis of high temperature capability. A limited experimental investigation was conducted to quantify some of the more critical mechanical properties of the SiC-LAS composite having a multidirection 0/45/-45/90 deg fiber orientation favored for the combustor linear application. Rigorous cyclic thermal tests demonstrated that SiC-LAS was extremely resistant to the thermal fatigue mechanisms that usually limit the life of metallic combustor liners. A thermal design study led to the definition of a composite liner concept that incorporated film cooled SiC-LAS shingles mounted on a Hastelloy X shell. With coolant fluxes consistent with the most advanced metallic liner technology, the calculated hot surface temperatures of the shingles were within the apparent near term capability of the material. Structural analyses indicated that the stresses in the composite panels were low, primarily because of the low coefficient of expansion of the material and it was concluded that the dominant failure mode of the liner would be an as yet unidentified deterioration of the composite from prolonged exposure to high temperature. An economic study, based on a medium thrust size commercial aircraft engine, indicated that the SiC-LAS combustor liner would weigh 22.8N (11.27 lb) less and cost less to manufacture than advanced metallic liner concepts intended for use in the late 1980's.

Thinking inside the Box Constrained Creativity and New Technology

ERIC Educational Resources Information Center

Gill, Bradley

2012-01-01

Thinking outside of the box is a standard cliche for creativity. Yet an awareness of the boxed nature of new media can empower young students to think creatively about design. This article is a reflection of one teacher who led a group of young students in a series of lessons based on basic design principles related to technology. It is based on…

Active Control of Combustor Instability Shown to Help Lower Emissions

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Chang, Clarence T.

2002-01-01

In a quest 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, or 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 safe operating life of the combustor and turbine. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Aerospace Propulsion and Power Base Research and Technology Program, the NASA Glenn Research Center, in partnership with Pratt & Whitney and United Technologies Research Center, is developing technologies for the active control of combustion instabilities. With active combustion control, the fuel is pulsed to put pressure oscillations into the system. This cancels out the pressure oscillations being produced by the instabilities. Thus, the engine can have lower pollutant emissions and long life.The use of active combustion instability control to reduce thermo-acoustic-driven combustor pressure oscillations was demonstrated on a single-nozzle combustor rig at United Technologies. This rig has many of the complexities of a real engine combustor (i.e., an actual fuel nozzle and swirler, dilution cooling, etc.). Control was demonstrated through modeling, developing, and testing a fuel-delivery system able to the 280-Hz instability frequency. The preceding figure shows the capability of this system to provide high-frequency fuel modulations. Because of the high-shear contrarotating airflow in the fuel injector, there was some concern that the fuel pulses would be attenuated to the point where they would not be effective for control. Testing in the combustor rig showed that open-loop pulsing of the fuel was, in fact, able to effectively modulate the combustor pressure. To suppress the combustor pressure oscillations due to thermoacoustic instabilities, it is desirable to time the injection of the fuel so that it interferes with the instability. A closed-loop control scheme was developed that uses combustion pressure feedback and a phase-shifting controller to time the fuel-injection pulses. Some suppression of the pressure oscillations at the 280-Hz instability frequency was demonstrated (see the next figure). However, the overall peak-to- peak pressure oscillations in the combustor were only mildly reduced. Improvements to control hardware and control methods are being continued to gain improved closed-loop reduction of the pressure oscillations.pulse the fuel at

Fostering and Assessing Creativity in Technology Education

ERIC Educational Resources Information Center

Buelin-Biesecker, Jennifer Katherine

2012-01-01

This study compared the creative outcomes in student work resulting from two pedagogical approaches to creative problem solving activities. A secondary goal was to validate the Consensual Assessment Technique (CAT) as a means of assessing creativity. Linear models for problem solving and design processes serve as the current paradigm in classroom…

Analysis of Factors Influencing Creative Personality of Elementary School Students

ERIC Educational Resources Information Center

Park, Jongman; Kim, Minkee; Jang, Shinho

2017-01-01

This quantitative research examined factors that affect elementary students' creativity and how those factors correlate. Aiming to identify significant factors that affect creativity and to clarify the relationship between these factors by path analysis, this research was designed to be a stepping stone for creativity enhancement studies. Data…

Learning to Relate: Interweaving Creative Approaches in Group Counseling with Adolescents

ERIC Educational Resources Information Center

Skudrzyk, Bogusia; Zera, David Aloyzy; McMahon, George; Schmidt, Robert; Boyne, Joey; Spannaus, Rudy L.

2009-01-01

This article describes three creative arts exercises designed to promote belonging, connectedness, and wellness among adolescents with and without disabilities. The scenarios are multifaceted and based on utilization of creative visual and performing arts. Creative outlets may help groups be more effective and meaningful for participants,…

Mandalas: A Simple Project to Explore Creativity

ERIC Educational Resources Information Center

Holbrook, Robert L., Jr.; Comer, Debra R.

2017-01-01

Creativity is more important than ever in today's knowledge-based economy. Although many students doubt their own creativity, very few exercises are designed to help them access this ability. We believe that self-expression and self-reflection are important for understanding personal creative ability. Jung introduced the mandala to promote these…

Linear Test Bed. Volume 2: Test Bed No. 2. [linear aerospike test bed for thrust vector control

NASA Technical Reports Server (NTRS)

1974-01-01

Test bed No. 2 consists of 10 combustors welded in banks of 5 to 2 symmetrical tubular nozzle assemblies, an upper stationary thrust frame, a lower thrust frame which can be hinged, a power package, a triaxial combustion wave ignition system, a pneumatic control system, pneumatically actuated propellant valves, a purge and drain system, and an electrical control system. The power package consists of the Mark 29-F fuel turbopump, the Mark 29-0 oxidizer turbopump, a gas generator assembly, and propellant ducting. The system, designated as a linear aerospike system, was designed to demonstrate the feasibility of the concept and to explore technology related to thrust vector control, thrust vector optimization, improved sequencing and control, and advanced ignition systems. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure at an engine mixture ratio of 5.5. With 10 combustors, the sea level thrust is 95,000 pounds.

Evaluation of fuel preparation systems for lean premixing-prevaporizing combustors

NASA Technical Reports Server (NTRS)

Dodds, W. J.; Ekstedt, E. E.

1985-01-01

A series of experiments was carried out in order to produce design data for a premixing prevaporizing fuel-air mixture preparation system for aircraft gas turbine engine combustors. The fuel-air mixture uniformity of four different system design concepts was evaluated over a range of conditions representing the cruise operation of a modern commercial turbofan engine. Operating conditions including pressure, temperature, fuel-to-air ratio, and velocity, exhibited no clear effect on mixture uniformity of systems using pressure-atomizing fuel nozzles and large-scale mixing devices. However, the performance of systems using atomizing fuel nozzles and large-scale mixing devices was found to be sensitive to operating conditions. Variations in system design variables were also evaluated and correlated. Mixing uniformity was found to improve with system length, pressure drop, and the number of fuel injection points per unit area. A premixing system capable of providing mixing uniformity to within 15 percent over a typical range of cruise operating conditions is demonstrated.

The Role of Healthy Social Interaction and Communications in Provoking Creativity in the Design Studio

ERIC Educational Resources Information Center

Sidawi, Bhzad

2016-01-01

Design is a social phenomenon and researchers suggest that negotiations and communications between designers are essential to initiate creativity. Within the design studio environment, a number of factors affect the healthy social interaction and design negotiations, such as the teaching style of tutors and the culture that governs a design studio…

Applying MacKinnon's 4Ps to Foster Creative Thinking and Creative Behaviours in Kindergarten Children

ERIC Educational Resources Information Center

Riga, Vassiliki; Chronopoulou, Elena

2014-01-01

The purpose of this study was to identify certain strategies and conditions that should be used by teachers in kindergarten so as to foster creative thinking and creative behaviours to children. We used a quasi-experimental research design for 6 months in a public kindergarten in a suburban area of Greece, and we developed a creative music and…

Low Emissions RQL Flametube Combustor Test Results

NASA Technical Reports Server (NTRS)

Chang, Clarence T.; Holdeman, James D.

2001-01-01

The overall objective of this test program was to demonstrate and evaluate the capability of the Rich-burn/Quick-mix/Lean-burn (RQL) combustor concept for HSR applications. This test program was in support of the Pratt & Whitney and GE Aircraft Engines HSR low-NOx Combustor Program. Collaborative programs with Parker Hannifin Corporation and Textron Fuel Systems resulted in the development and testing of the high-flow low-NOx rich-burn zone fuel-to-air ratio research fuel nozzles used in this test program. Based on the results obtained in this test program, several conclusions can be made: (1) The RQL tests gave low NOx and CO emissions results at conditions corresponding to HSR cruise. (2) The Textron fuel nozzle design with optimal multiple partitioning of fuel and air circuits shows potential of providing an acceptable uniform local fuel-rich region in the rich burner. (3) For the parameters studied in this test series, the tests have shown T3 is the dominant factor in the NOx formation for RQL combustors. As T3 increases from 600 to 1100 F, EI(NOx) increases approximately three fold. (4) Factors which appear to have secondary influence on NOx formation are P4, T4, infinity(sub rb), V(sub ref,ov). (5) Low smoke numbers were measured for infinity(sub rb) of 2.0 at P4 of 120 psia.

Testing of DLR C/C-SiC for HIFiRE 8 Scramjet Combustor

NASA Technical Reports Server (NTRS)

Glass, David E.; Capriotti, Diego P.; Reimer, Thomas; Kutemeyer, Marius; Smart, Michael

2013-01-01

Ceramic Matrix Composites (CMCs) have been proposed for hot structures in scramjet combustors. Previous studies have calculated significant weight savings by utilizing CMCs (active and passive) versus actively cooled metallic scramjet structures. Both a C/C and a C/C-SiC material system fabricated by DLR (Stuttgart, Germany) are being considered for use in a passively cooled combustor design for HIFiRE 8, a joint Australia / AFRL hypersonic flight program, expected to fly at Mach 7 for approximately 30 sec, at a dynamic pressure of 55 kPa. Flat panels of the DLR C/C and the C/C-SiC were tested in the NASA Langley Direct Connect Rig (DCR) at Mach 5 and Mach 6 enthalpy for several minutes. Gaseous hydrogen fuel was used to fuel the scramjet combustor. The test panels were instrumented with embedded Type K and Type S thermocouples. Zirconia felt insulation was used in some of the tests to increase the surface temperature of the C/C-SiC panel for approximately 350degF. The final C/C-SiC panel was tested for 3 cycles totaling over 135 sec at Mach 6 enthalpy. Slightly more erosion was observed on the C/C panel than the C/C-SiC panels, but both material systems demonstrated acceptable recession performance for the HIFiRE 8 flight.

40 CFR 62.14103 - Emission limits for municipal waste combustor metals, acid gases, organics, and nitrogen oxides.

Code of Federal Regulations, 2010 CFR

2010-07-01

... combustor metals, acid gases, organics, and nitrogen oxides. 62.14103 Section 62.14103 Protection of... combustor metals, acid gases, organics, and nitrogen oxides. (a) The emission limits for municipal waste combustor metals are specified in paragraphs (a)(1) through (a)(3) of this section. (1) The owner or...

40 CFR 60.52a - Standard for municipal waste combustor metals.

Code of Federal Regulations, 2014 CFR

2014-07-01

... Municipal Waste Combustors for Which Construction Is Commenced After December 20, 1989 and On or Before September 20, 1994 § 60.52a Standard for municipal waste combustor metals. (a) On and after the date on... 40 Protection of Environment 7 2014-07-01 2014-07-01 false Standard for municipal waste combustor...

40 CFR Table 3 to Subpart Cb of... - Municipal Waste Combustor Operating Guidelines

Code of Federal Regulations, 2010 CFR

2010-07-01

... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Municipal Waste Combustor Operating... and Compliance Times for Large Municipal Waste Combustors That are Constructed on or Before September 20, 1994 Pt. 60, Subpt. Cb, Table 3 Table 3 to Subpart Cb of Part 60—Municipal Waste Combustor...

System and method for reducing combustion dynamics in a combustor

DOEpatents

Uhm, Jong Ho; Ziminsky, Willy Steve; Johnson, Thomas Edward; Srinivasan, Shiva; York, William David

2016-11-29

A system for reducing combustion dynamics in a combustor includes an end cap that extends radially across the combustor and includes an upstream surface axially separated from a downstream surface. A combustion chamber is downstream of the end cap, and tubes extend from the upstream surface through the downstream surface. Each tube provides fluid communication through the end cap to the combustion chamber. The system further includes means for reducing combustion dynamics in the combustor. A method for reducing combustion dynamics in a combustor includes flowing a working fluid through tubes that extend axially through an end cap that extends radially across the combustor and obstructing at least a portion of the working fluid flowing through a first set of the tubes.

The 3-D CFD modeling of gas turbine combustor-integral bleed flow interaction

NASA Technical Reports Server (NTRS)

Chen, D. Y.; Reynolds, R. S.

1993-01-01

An advanced 3-D Computational Fluid Dynamics (CFD) model was developed to analyze the flow interaction between a gas turbine combustor and an integral bleed plenum. In this model, the elliptic governing equations of continuity, momentum and the k-e turbulence model were solved on a boundary-fitted, curvilinear, orthogonal grid system. The model was first validated against test data from public literature and then applied to a gas turbine combustor with integral bleed. The model predictions agreed well with data from combustor rig testing. The model predictions also indicated strong flow interaction between the combustor and the integral bleed. Integral bleed flow distribution was found to have a great effect on the pressure distribution around the gas turbine combustor.

Gas turbine combustor

NASA Technical Reports Server (NTRS)

Burd, Steven W. (Inventor); Cheung, Albert K. (Inventor); Dempsey, Dae K. (Inventor); Hoke, James B. (Inventor); Kramer, Stephen K. (Inventor); Ols, John T. (Inventor); Smith, Reid Dyer Curtis (Inventor); Sowa, William A. (Inventor)

2011-01-01

A gas turbine engine has a combustor module including an annular combustor having a liner assembly that defines an annular combustion chamber having a length, L. The liner assembly includes a radially inner liner, a radially outer liner that circumscribes the inner liner, and a bulkhead, having a height, H1, which extends between the respective forward ends of the inner liner and the outer liner. The combustor has an exit height, H3, at the respective aft ends of the inner liner and the outer liner interior. The annular combustor has a ratio H1/H3 having a value less than or equal to 1.7. The annular combustor may also have a ration L/H3 having a value less than or equal to 6.0.

Apparatus and method for gas turbine active combustion control system

NASA Technical Reports Server (NTRS)

Knobloch, Aaron (Inventor); Mancini, Alfred Albert (Inventor); Myers, William J. (Inventor); Fortin, Jeffrey B. (Inventor); Umeh, Chukwueloka (Inventor); Kammer, Leonardo C. (Inventor); Shah, Minesh (Inventor)

2011-01-01

An Active Combustion Control System and method provides for monitoring combustor pressure and modulating fuel to a gas turbine combustor to prevent combustion dynamics and/or flame extinguishments. The system includes an actuator, wherein the actuator periodically injects pulsed fuel into the combustor. The apparatus also includes a sensor connected to the combustion chamber down stream from an inlet, where the sensor generates a signal detecting the pressure oscillations in the combustor. The apparatus controls the actuator in response to the sensor. The apparatus prompts the actuator to periodically inject pulsed fuel into the combustor at a predetermined sympathetic frequency and magnitude, thereby controlling the amplitude of the pressure oscillations in the combustor by modulating the natural oscillations.

Supporting Creative Responses in Design Education--The Development and Application of the Graphic Design Composition Method

ERIC Educational Resources Information Center

Lu, Hui-Ping; Chen, Jun-Hong; Lee, Chang-Franw

2016-01-01

Inspiration is the primary element of good design. Designers, however, also risk not being able to find inspiration. Novice designers commonly find themselves to be depressed during the conceptual design phase when they fail to find inspiration and the information to be creative. Accordingly, under the graphic design parameter, we have developed…

Off-Design Analysis of a High Bypass Turbofan Using a Pulsed Detonation Combustor

DTIC Science & Technology

2010-03-01

Engine Off-Design Results.............................................................39 Code Verification and Operating Limit ...38 4.4. Maximum Operating Limit Baseline and Hybrid Engine ......................................... 41 4.5. Throttle...that an isentropic expansion process takes place followed by a heat rejection to close the cycle. The derivation for the solutions for the Chapman

Peacocks, Picasso, and parental investment: The effects of romantic motives on creativity.

PubMed

Griskevicius, Vladas; Cialdini, Robert B; Kenrick, Douglas T

2006-07-01

Four experiments explored the effects of mating motivation on creativity. Even without other incentives to be creative, romantic motives enhanced creativity on subjective and objective measures. For men, any cue designed to activate a short-term or a long-term mating goal increased creative displays; however, women displayed more creativity only when primed to attract a high-quality long-term mate. These creative boosts were unrelated to increased effort on creative tasks or to changes in mood or arousal. Furthermore, results were unaffected by the application of monetary incentives for creativity. These findings align with the view that creative displays in both sexes may be linked to sexual selection, qualified by unique exigencies of human parental investment. Copyright 2006 APA, all rights reserved.

Creativity: Creativity in Complex Military Systems

DTIC Science & Technology

2017-05-25

generation later in the problem-solving process. The design process is an alternative problem-solving framework individuals or groups use to orient...no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control ...the potential of their formations. 15. SUBJECT TERMS Creativity, Divergent Thinking, Design , Systems Thinking, Operational Art 16. SECURITY

Designing Creative Inter-Disciplinary Science and Art Interventions in Schools: The Case of Write a Science Opera (WASO)

ERIC Educational Resources Information Center

Ben-Horin, Oded; Chappell, Kerry A.; Halstead, Jill; Espeland, Magne

2017-01-01

The goal of this qualitative study is to provide theoretical knowledge and design principles for a creative educational environment characterized by simultaneous study and exploration of science or math, and the arts: Write a Science Opera (WASO). To do so, we used a theory of creativity in education which links collaborative co-creation in…

An Analysis of Creative Process Learning in Computer Game Activities through Player Experiences

ERIC Educational Resources Information Center

Inchamnan, Wilawan

2016-01-01

This research investigates the extent to which creative processes can be fostered through computer gaming. It focuses on creative components in games that have been specifically designed for educational purposes: Digital Game Based Learning (DGBL). A behavior analysis for measuring the creative potential of computer game activities and learning…

An Interdisciplinary Design Studio: How Can Art and Engineering Collaborate to Increase Students' Creativity?

ERIC Educational Resources Information Center

Costantino, Tracie; Kellam, Nadia; Cramond, Bonnie; Crowder, Isabelle

2010-01-01

Creativity often has been associated with the arts, although creativity also is essential for innovative discoveries and applications in science and engineering. In this article, a pilot study is presented about an investigation concerning how creativity is fostered in an art education course in conjunction with an undergraduate engineering…

"Pedagogical Creativity" as a Framework for Music Teacher Education

ERIC Educational Resources Information Center

Abramo, Joseph Michael; Reynolds, Amy

2015-01-01

Creativity research has a long history in music education, including the development of theories and strategies to foster the music creativity of students of all ages and levels. Underexplored is how teacher education programs can cultivate pre- and in-service teachers' abilities to develop their educational creativity when designing curricula and…

Validation of Predictive Relationship of Creative Problem-Solving Attrubutes with Math Creativity

ERIC Educational Resources Information Center

Pham, Linh Hung

2014-01-01

This study was designed to investigate the predictive relationships of creative problem-solving attributes, which comprise divergent thinking, convergent thinking, motivation, general and domain knowledge and skills, and environment, with mathematical creativity of sixth grade students in Thai Nguyen City, Viet Nam. The study also aims to revise…

NOX CONTROL TECHNOLOGIES APPLICABLE TO MUNICIPAL WASTE COMBUSTION

EPA Science Inventory

The report documents the key design and operating parameters, commercial status, demonstrated performance, and cost of three technologies available for reducing nitrogen oxide (NOx) emissions from municipal waste combustors (MWCs), and identifies technology research and developme...

Rich catalytic injection

DOEpatents

Veninger, Albert [Coventry, CT

2008-12-30

A gas turbine engine includes a compressor, a rich catalytic injector, a combustor, and a turbine. The rich catalytic injector includes a rich catalytic device, a mixing zone, and an injection assembly. The injection assembly provides an interface between the mixing zone and the combustor. The injection assembly can inject diffusion fuel into the combustor, provides flame aerodynamic stabilization in the combustor, and may include an ignition device.

An Experimental Investigation of Self-Excited Combustion Dynamics in a Single Element Lean Direct Injection (LDI) Combustor

NASA Astrophysics Data System (ADS)

Gejji, Rohan M.

The management of combustion dynamics in gas turbine combustors has become more challenging as strict NOx/CO emission standards have led to engine operation in a narrow, lean regime. While premixed or partially premixed combustor configurations such as the Lean Premixed Pre-vaporized (LPP), Rich Quench Lean burn (RQL), and Lean Direct Injection (LDI) have shown a potential for reduced NOx emissions, they promote a coupling between acoustics, hydrodynamics and combustion that can lead to combustion instabilities. These couplings can be quite complex, and their detailed understanding is a pre-requisite to any engine development program and for the development of predictive capability for combustion instabilities through high-fidelity models. The overarching goal of this project is to assess the capability of high-fidelity simulation to predict combustion dynamics in low-emissions gas turbine combustors. A prototypical lean-direct-inject combustor was designed in a modular configuration so that a suitable geometry could be found by test. The combustor comprised a variable length air plenum and combustion chamber, air swirler, and fuel nozzle located inside a subsonic venturi. The venturi cross section and the fuel nozzle were consistent with previous studies. Test pressure was 1 MPa and variables included geometry and acoustic resonance, inlet temperatures, equivalence ratio, and type of liquid fuel. High-frequency pressure measurements in a well-instrumented metal chamber yielded frequencies and mode shapes as a function of inlet air temperature, equivalence ratio, fuel nozzle placement, and combustor acoustic resonances. The parametric survey was a significant effort, with over 105 tests on eight geometric configurations. A good dataset was obtained that could be used for both operating-point-dependent quantitative comparisons, and testing the ability of the simulation to predict more global trends. Results showed a very strong dependence of instability amplitude on the geometric configuration of the combustor, i.e., its acoustic resonance characteristics, with measured pressure fluctuation amplitudes ranged from 5 kPa (0.5% of mean pressure) to 200 kPa ( 20% of mean pressure) depending on combustor geometry. The stability behavior also showed a consistent and pronounced dependence on equivalence ratio and inlet air temperature. Instability amplitude increased with higher equivalence ratio and with lower inlet air temperature. A pronounced effect of fuel nozzle location on the combustion dynamics was also observed. Combustion instabilities with the fuel nozzle at the throat of the venturi throat were stronger than in the configuration with fuel nozzle 2.6 mm upstream of the nozzle. A second set of dynamics data was based on high-response-rate laser-based combustion diagnostics using an optically accessible combustor section. High-frequency measurements of OH*-chemiluminescence and OH-PLIF and velocity fields using PIV were obtained at a relatively stable, low equivalence ratio case and a less stable case at higher equivalence ratio. PIV measurements were performed at 5 kHz for non-reacting flow but glare from the cylindrical quartz chamber limited the field of view to a small region in the combustor. Quantitative and qualitative comparisons were made for five different combinations of geometry and operating condition that yielded discriminating stability behavior in the experiment with simulations that were carried out concurrently. Comparisons were made on the basis of trends and pressure mode data as well as with OH-PLIF measurements for the baseline geometry at equivalence ratios of 0.44 and 0.6. Overall, the ability of the simulation to match experimental data and trends was encouraging. Dynamic Mode Decomposition (DMD) analysis was performed on two sets of computations - a global 2-step chemistry mechanism and an 18-step chemistry mechanism - and the OH-PLIF images to allow comparison of dynamic patterns of heat release and OH distribution in the combustion zone. The DMD analysis was able to identify similar dominant unstable modes in the combustor. Recommendations for future work are based on the continued requirement for quantitative and spatio-temporally resolved data for direct comparison with computational efforts to develop predictive capabilities for combustion instabilities at relevant operating conditions. Discriminating instability behavior for the prototypical combustor demonstrated in this study is critical for any robust validation effort Unit physics based scaling of the current effort to multi-element combustors along with improvement in diagnostic techniques and analysis efforts are recommended for advancement in understanding of the complex physics in the multi-phase, three dimensional and turbulent combustion processes in the LDI combustor.

Combustion Dynamics in Multi-Nozzle Combustors Operating on High-Hydrogen Fuels

DOE Office of Scientific and Technical Information (OSTI.GOV)

Santavicca, Dom; Lieuwen, Tim

Actual gas turbine combustors for power generation applications employ multi-nozzle combustor configurations. Researchers at Penn State and Georgia Tech have extended previous work on the flame response in single-nozzle combustors to the more realistic case of multi-nozzle combustors. Research at Georgia Tech has shown that asymmetry of both the flow field and the acoustic forcing can have a significant effect on flame response and that such behavior is important in multi-flame configurations. As a result, the structure of the flame and its response to forcing is three-dimensional. Research at Penn State has led to the development of a three-dimensional chemiluminescencemore » flame imaging technique that can be used to characterize the unforced (steady) and forced (unsteady) flame structure of multi-nozzle combustors. Important aspects of the flame response in multi-nozzle combustors which are being studied include flame-flame and flame-wall interactions. Research at Penn State using the recently developed three-dimensional flame imaging technique has shown that spatial variations in local flame confinement must be accounted for to accurately predict global flame response in a multi-nozzle can combustor.« less

O and temperature in a hydrocarbon-fueled scramjet combustor

NASA Astrophysics Data System (ADS)

Goldenstein, C. S.; Schultz, I. A.; Spearrin, R. M.; Jeffries, J. B.; Hanson, R. K.

2014-09-01

The design and demonstration of a two-color tunable diode laser sensor for measurements of temperature and H2O in an ethylene-fueled model scramjet combustor are presented. This sensor probes multiple H2O transitions in the fundamental vibration bands near 2.5 μm that are up to 20 times stronger than those used by previous near-infrared H2O sensors. In addition, two design measures enabled high-fidelity measurements in the nonuniform flow field. (1) A recently developed calibration-free scanned-wavelength-modulation spectroscopy spectral-fitting strategy was used to infer the integrated absorbance of each transition without a priori knowledge of the absorption lineshape and (2) transitions with strengths that scale near-linearly with temperature were used to accurately determine the H2O column density and the H2O-weighted path-averaged temperature from the integrated absorbance of two transitions.

Material requirements for the High Speed Civil Transport

NASA Technical Reports Server (NTRS)

Stephens, Joseph R.; Hecht, Ralph J.; Johnson, Andrew M.

1993-01-01

Under NASA-sponsored High Speed Research (HSR) programs, the materials and processing requirements have been identified for overcoming the environmental and economic barriers of the next generation High Speed Civil Transport (HSCT) propulsion system. The long (2 to 5 hours) supersonic cruise portion of the HSCT cycle will place additional durability requirements on all hot section engine components. Low emissions combustor designs will require high temperature ceramic matrix composite liners to meet an emission goal of less than 5g NO(x) per Kg fuel burned. Large axisymmetric and two-dimensional exhaust nozzle designs are now under development to meet or exceed FAR 36 Stage III noise requirements, and will require lightweight, high temperature metallic, intermetallic, and ceramic matrix composites to reduce nozzle weight and meet structural and acoustic component performance goals. This paper describes and discusses the turbomachinery, combustor, and exhaust nozzle requirements of the High Speed Civil Transport propulsion system.

Aviation fuel property effects on altitude relight

NASA Technical Reports Server (NTRS)

Venkataramani, K.

1987-01-01

The major objective of this experimental program was to investigate the effects of fuel property variation on altitude relight characteristics. Four fuels with widely varying volatility properties (JP-4, Jet A, a blend of Jet A and 2040 Solvent, and Diesel 2) were tested in a five-swirl-cup-sector combustor at inlet temperatures and flows representative of windmilling conditions of turbofan engines. The effects of fuel physical properties on atomization were eliminated by using four sets of pressure-atomizing nozzles designed to give the same spray Sauter mean diameter (50 + or - 10 micron) for each fuel at the same design fuel flow. A second series of tests was run with a set of air-blast nozzles. With comparable atomization levels, fuel volatility assumes only a secondary role for first-swirl-cup lightoff and complete blowout. Full propagation first-cup blowout were independent of fuel volatility and depended only on the combustor operating conditions.

Reduced Order Modeling of Combustion Instability in a Gas Turbine Model Combustor

NASA Astrophysics Data System (ADS)

Arnold-Medabalimi, Nicholas; Huang, Cheng; Duraisamy, Karthik

2017-11-01

Hydrocarbon fuel based propulsion systems are expected to remain relevant in aerospace vehicles for the foreseeable future. Design of these devices is complicated by combustion instabilities. The capability to model and predict these effects at reduced computational cost is a requirement for both design and control of these devices. This work focuses on computational studies on a dual swirl model gas turbine combustor in the context of reduced order model development. Full fidelity simulations are performed utilizing URANS and Hybrid RANS-LES with finite rate chemistry. Following this, data decomposition techniques are used to extract a reduced basis representation of the unsteady flow field. These bases are first used to identify sensor locations to guide experimental interrogations and controller feedback. Following this, initial results on developing a control-oriented reduced order model (ROM) will be presented. The capability of the ROM will be further assessed based on different operating conditions and geometric configurations.

When Theater Comes to Engineering Design: Oh How Creative They Can Be.

PubMed

Pfeiffer, Ferris M; Bauer, Rachel E; Borgelt, Steve; Burgoyne, Suzanne; Grant, Sheila; Hunt, Heather K; Pardoe, Jennie J; Schmidt, David C

2017-07-01

The creative process is fun, complex, and sometimes frustrating, but it is critical to the future of our nation and progress in science, technology, engineering, mathematics (STEM), as well as other fields. Thus, we set out to see if implementing methods of active learning typical to the theater department could impact the creativity of senior capstone design students in the bioengineering (BE) department. Senior bioengineering capstone design students were allowed to self-select into groups. Prior to the beginning of coursework, all students completed a validated survey measuring engineering design self-efficacy. The control and experimental groups both received standard instruction, but in addition the experimental group received 1 h per week of creativity training developed by a theater professor. Following the semester, the students again completed the self-efficacy survey. The surveys were examined to identify differences in the initial and final self-efficacy in the experimental and control groups over the course of the semester. An analysis of variance was used to compare the experimental and control groups with p < 0.05 considered significant. Students in the experimental group reported more than a twofold (4.8 (C) versus 10.9 (E)) increase of confidence. Additionally, students in the experimental group were more motivated and less anxious when engaging in engineering design following the semester of creativity instruction. The results of this pilot study indicate that there is a significant potential to improve engineering students' creative self-efficacy through the implementation of a "curriculum of creativity" which is developed using theater methods.

Real-Time Control of Lean Blowout in a Turbine Engine for Minimizing No(x) Emissions

NASA Technical Reports Server (NTRS)

Zinn, Ben

2004-01-01

This report describes research on the development and demonstration of a controlled combustor operates with minimal NO, emissions, thus meeting one of NASA s UEET program goals. NO(x) emissions have been successfully minimized by operating a premixed, lean burning combustor (modeling a lean prevaporized, premixed LPP combustor) safely near its lean blowout (LBO) limit over a range of operating conditions. This was accomplished by integrating the combustor with an LBO precursor sensor and closed-loop, rule-based control system that allowed the combustor to operate far closer to the point of LBO than an uncontrolled combustor would be allowed to in a current engine. Since leaner operation generally leads to lower NO, emissions, engine NO, was reduced without loss of safety.

Gas turbine combustor transition

DOEpatents

Coslow, Billy Joe; Whidden, Graydon Lane

1999-01-01

A method of converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit.

Gas turbine combustor transition

DOEpatents

Coslow, B.J.; Whidden, G.L.

1999-05-25

A method is described for converting a steam cooled transition to an air cooled transition in a gas turbine having a compressor in fluid communication with a combustor, a turbine section in fluid communication with the combustor, the transition disposed in a combustor shell and having a cooling circuit connecting a steam outlet and a steam inlet and wherein hot gas flows from the combustor through the transition and to the turbine section, includes forming an air outlet in the transition in fluid communication with the cooling circuit and providing for an air inlet in the transition in fluid communication with the cooling circuit. 7 figs.

Creative Mosaics, Art Education: 6683.14.

ERIC Educational Resources Information Center

Marinaccio, Louis M.

This quinmester guide describes a laboratory course designed to offer opportunities for experimentation in mosaic design by exploring uses of various techniques and materials. The course, which is offered to 7th through 12th grade pre-vocational students, emphasizes the use of design in relation to creative composition. Students learn to design,…

Foundational Performance Analyses of Pressure Gain Combustion Thermodynamic Benefits for Gas Turbines

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Kaemming, Thomas A.

2012-01-01

A methodology is described whereby the work extracted by a turbine exposed to the fundamentally nonuniform flowfield from a representative pressure gain combustor (PGC) may be assessed. The method uses an idealized constant volume cycle, often referred to as an Atkinson or Humphrey cycle, to model the PGC. Output from this model is used as input to a scalable turbine efficiency function (i.e., a map), which in turn allows for the calculation of useful work throughout the cycle. Integration over the entire cycle yields mass-averaged work extraction. The unsteady turbine work extraction is compared to steady work extraction calculations based on various averaging techniques for characterizing the combustor exit pressure and temperature. It is found that averages associated with momentum flux (as opposed to entropy or kinetic energy) provide the best match. This result suggests that momentum-based averaging is the most appropriate figure-of-merit to use as a PGC performance metric. Using the mass-averaged work extraction methodology, it is also found that the design turbine pressure ratio for maximum work extraction is significantly higher than that for a turbine fed by a constant pressure combustor with similar inlet conditions and equivalence ratio. Limited results are presented whereby the constant volume cycle is replaced by output from a detonation-based PGC simulation. The results in terms of averaging techniques and design pressure ratio are similar.

Melt Infiltrated Ceramic Matrix Composites for Shrouds and Combustor Liners of Advanced Industrial Gas Turbines

DOE Office of Scientific and Technical Information (OSTI.GOV)

Gregory Corman; Krishan Luthra; Jill Jonkowski

2011-01-07

This report covers work performed under the Advanced Materials for Advanced Industrial Gas Turbines (AMAIGT) program by GE Global Research and its collaborators from 2000 through 2010. A first stage shroud for a 7FA-class gas turbine engine utilizing HiPerComp{reg_sign}* ceramic matrix composite (CMC) material was developed. The design, fabrication, rig testing and engine testing of this shroud system are described. Through two field engine tests, the latter of which is still in progress at a Jacksonville Electric Authority generating station, the robustness of the CMC material and the shroud system in general were demonstrated, with shrouds having accumulated nearly 7,000more » hours of field engine testing at the conclusion of the program. During the latter test the engine performance benefits from utilizing CMC shrouds were verified. Similar development of a CMC combustor liner design for a 7FA-class engine is also described. The feasibility of using the HiPerComp{reg_sign} CMC material for combustor liner applications was demonstrated in a Solar Turbines Ceramic Stationary Gas Turbine (CSGT) engine test where the liner performed without incident for 12,822 hours. The deposition processes for applying environmental barrier coatings to the CMC components were also developed, and the performance of the coatings in the rig and engine tests is described.« less

The Creative Web.

ERIC Educational Resources Information Center

Yudess, Jo

2003-01-01

This article lists the Web sites of 12 international not-for-profit creativity associations designed to trigger more creative thought and research possibilities. Along with Web addresses, the entries include telephone contact information and a brief description of the organization. (CR)

P and W propulsion systems studies results/status

NASA Technical Reports Server (NTRS)

Smith, Martin G., Jr.; Champagne, George A.

1992-01-01

The topics covered include the following: Pratt and Whitney (P&W) propulsion systems studies - NASA funded efforts to date; P&W engine concepts; P&W combustor focus - rich burn quick quench (RBQQ) concept; mixer ejector nozzle concept - large flow entrainment reduces jet noise; technology impact on NO(x) emissions - mature RBQQ combustor reduces NO(x) up to 85 percent; technology impact on sideline noise characteristics of Mach 2.4 turbine bypass engines (TBE's) - 600 lb/sec airflow size; technology impact on takeoff gross weight (TOGW) - provides up to 12 percent TOGW reduction; HSCT quiet engine concepts; TBE inlet valve/ejector nozzle concept schematic; mixed flow turbofan study; and exhaust nozzle conceptual design.