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At least 433 records · Page 24

Advanced Low Emissions Subsonic Combustor Study

Recent advances in commercial and military aircraft gas turbines have yielded significant improvements in fuel efficiency and thrust-to-weight ratio, due in large part to increased combustor operating pressures and temperatures. However, the higher operating conditions have increased the emission of oxides of nitrogen (NOx), which is a pollutant with adverse impact on the atmosphere and environment. Since commercial and military aircraft are the only important direct source of NOx emissions at high altitudes, there is a growing consensus that considerably more stringent limits on NOx emissions will be required in the future for all aircraft. In fact, the regulatory communities have recently agreed to reduce NOx limits by 20 percent from current requirements effective in 1996. Further reductions at low altitude, together with introduction of limits on NOx at altitude, are virtual certainties. In addition, the U.S. Government recently conducted hearings on the introduction of federal fees on the local emission of pollutants from all sources, including aircraft. While no action was taken regarding aircraft in this instance, the threat of future action clearly remains. In these times of intense and growing international competition, the U.S. le-ad in aerospace can only be maintained through a clear technological dominance that leads to a product line of maximum value to the global airline customer. Development of a very low NOx combustor will be essential to meet the future needs of both the commercial and military transport markets, if additional economic burdens and/or operational restrictions are to be avoided. In this report, Pratt & Whitney (P&W) presents the study results with the following specific objectives: Development of low-emissions combustor technologies for advances engines that will enter into service circa 2005, while producing a goal of 70 percent lower NOx emissions, compared to 1996 regulatory levels. Identification of solution approaches to barriers to the productization and economic viability of the low-emissions technologies. Preparation of these technologies to facilitate an annular rig high-pressure demonstration.

Smith, Reid↗

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

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.

Locke, Randy J.↗

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

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.

Vardakas, M. A.↗

Numerical Simulation of Dual-Mode Scramjet Combustors

Results of a numerical investigation of a three-dimensional dual-mode scramjet isolator-combustor flow-field are presented. Specifically, the effect of wall cooling on upstream interaction and flow-structure is examined for a case assuming jet-to-jet symmetry within the combustor. Comparisons are made with available experimental wall pressures. The full half-duct for the isolator-combustor is then modeled in order to study the influence of side-walls. Large scale three-dimensionality is observed in the flow with massive separation forward on the side-walls of the duct. A brief review of convergence-acceleration techniques useful in dual-mode simulations is presented, followed by recommendations regarding the development of a reliable and unambiguous experimental data base for guiding CFD code assessments in this area.

Rodriguez, C. G.↗

Augmentation of Stagnation Region Heat Transfer Due to Turbulence From a DLN Can Combustor

Heat transfer measurements have been made in the stagnation region of a flat plate with a circular leading edge. Electrically heated aluminum strips placed symmetrically about the leading edge stagnation region were used to measure spanwise averaged heat transfer coefficients. The maximum Reynolds number obtained, based on leading edge diameter, was about 100,000. The model was immersed in the flow field downstream of an approximately half scale model of a can-type combustor from a low NO(x), ground based power-generating turbine. The tests were conducted with room temperature air; no fuel was added. Room air flowed into the combustor through six vane type fuel/air swirlers. The combustor can contained no dilution holes. The fuel/air swirlers all swirled the incoming airflow in a counter clockwise direction (facing downstream). A 5-hole probe flow field survey in the plane of the model stagnation point showed the flow was one big vortex with flow angles up to 36' at the outer edges of the rectangular test section. Hot wire measurements showed test section flow had very high levels of turbulence, around 28.5 percent, and had a relatively large axial-length scale-to-leading edge diameter ratio of 0.5. X-wire measurements showed the turbulence to be nearly isotropic. Stagnation heat transfer augmentation over laminar levels was around 77 percent and was about 14 percent higher than predicted by a previously developed correlation for isotropic grid generated turbulence.

VanFossen, G. James↗

Augmentation of Stagnation Region Heat Transfer Due to Turbulence from a DLN Can Combustor

Heat transfer measurements have been made in the stagnation region of a flat plate with a circular leading edge. Electrically heated aluminum strips placed symmetrically about the leading edge stagnation region were used to measure spanwise-averaged heat transfer coefficients. The maximum Reynolds number obtained, based on leading edge diameter, was about 100,000. The model was immersed in the flow field downstream of an approximately half-scale model of a can-type combustor from a low NO(x), ground-based power-generating turbine. The tests were conducted with room temperature air; no fuel was added. Room air flowed into the combustor through six vane-type fuel/air swirlers. The combustor can contained no dilution holes. The fuel/air swirlers all swirled the incoming airflow in a counterclockwise direction (facing downstream). A five-hole probe flow field survey in the plane of the model stagnation point showed the flow was one big vortex with flow angles up to 36 deg at the outer edges of the rectangular test section. Hot-wire measurements showed test section flow had very high levels of turbulence, around 28.5%, and had a relatively large axial-length scale-to-leading edge diameter ratio of 0.5. X-wire measurements showed the turbulence to be nearly isotropic. Stagnation heat transfer augmentation over laminar levels was around 77% and was about 14% higher than predicted by a previously developed correlation for isotropic grid-generated turbulence.

VanFossen, G. James↗

Review of the PDWA Concept for Combustion Enhancement in a Supersonic Air-Breathing Combustor Environment

This paper reviews the design of the Pulsed Detonation Wave Augmentor (PDWA) concept and the preliminary computational fluid dynamics studies that supported it. The PDWA relies on the rapid generation of detonation waves in a small tube, which are then injected into the supersonic stream of the main combustor. The blast waves thus generated are used to stimulate the mixing and combustion inside the main combustor. The mixing enhancement relies on various forms of the baroclinic interaction, where misaligned pressure and density gradients combine to produce vortical flow. By using unsteady shock waves, the concept also uses the Richtmyer-Meshkov effect to further increase the rate of mixing. By carefully designing the respective configurations of the combustor and the detonation tubes, one can also increase the penetration of the fuel into the supersonic air stream. The unsteady shocks produce lower stagnation pressure losses than steady shocks. Combustion enhancement can also be obtained through the transient shock-heating of the fuel-air interface, and the lowering of the ignition delay in these regions. The numerical simulations identify these processes, and show which configurations give the best results. Engineering considerations are also presented, and discuss the feasibility of the concept. Of primary importance are the enhancements in performance, the design simplicity, the minimization of the power, cost, and weight, and the methods to achieve very rapid cycling.

Canbier, Jean-Luc↗

Cars Thermometry in a Supersonic Combustor for CFD Code Validation

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

Cutler, A. D.↗

Low Emissions RQL Flametube Combustor Test Results

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.

Chang, Clarence T.↗

Dual-Pump CARS Thermometry and Species Concentration Measurements in a Supersonic Combustor

The dual-pump coherent anti-Stokes Raman spectroscopy (CARS) method was used to measure temperature and the absolute mole fractions of N2, O2 and H2 in a supersonic combustor. Experiments were conducted in NASA Langley Research Center's Direct Connect Supersonic Combustion Test Facility. In this facility, hydrogen and air bum to increase the enthalpy of the test gas; O2 is then added to simulate air. This gas is expanded through a Mach 2 nozzle and into a combustor model consisting of a short constant-area section followed by a small rearward facing step and another constant area section. At the end of this straight section H2 fuel is then injected at Mach 2 and at 30 deg. angle with respect to the freestream. One wall of the duct then expands at a 3 deg. angle for over 1 meter. The ensuing combustion is monitored optically through ports in the side of the combustor. CARS measurements were performed at the nozzle exit and at four different planes downstream fuel injection. Maps were obtained of the mean temperature, as well as quantitative N2 and O2 and qualitative H2 mean mole fraction fields. Correlations between fluctuations of the different measured parameters are presented for one of the planes of data.

OByrne, Sean↗

Enhancement of Mixing and Reaction in High-speed Combustor Flowfields

Research has been underway for a number of years, both in the United States and abroad, to develop advanced aerospace propulsion systems for use late in this century and beyond. One program is now underway at the NASA Langley Research Research has been underway for a number of years, both in the United States and abroad, to develop advanced aerospace propulsion systems for use late in this century and beyond. One program is now underway at the NASA Langley Research Center to develop a hydrogen-fueled supersonic combustion ramjet (scramjet) that is capable of propelling a vehicle at hypersonic speeds in the atmosphere. A part of that research has been directed toward the optimization of the scramjet combustor and, in particular, the efficiency of fuel-air mixing and reaction taking place in the engine. In the high-speed vehicle configurations currently being considered, achieving a high combustor efficiency becomes particularly difficult. With increasing combustor Mach number, the degree of fuel-air mixing that can be achieved through natural convective and diffusive processes is reduced leading to an overall decrease in combustion efficiency and thrust. Compressible shear layers and jets provide a good model for studying the physical processes occurring in high speed mixing and combustion in a scramjet. Mixing layers are characterized by large scale eddies that form due to the high shear that is present between the fuel and the air streams. These eddies entrain the fuel and air into the mixing region between the fluids leading to increased surface area and locally steep concentration gradients. Molecular diffusion then occurs across the strained interfaces. In an early study of high-speed mixing, Brown and Roshko show that the spreading rate of a supersonic mixing layer decreases with increasing Mach number, exhibiting a factor of three decrease in spread rate as compared with an incompressible mixing layer with the same density ratio. They conclude that the reduced spread rate is primarily due to compressibility. Papamoschou and Roshko also observe that the spreading rate of compressible mixing layers is significantly reduced over that of incompressible layers. To characterize the structure of the flow quantitatively, they define a convective Mach number. The reduction in mixing layer spreading rate is shown in these experiments to correlate well with increasing convective Mach number.

Drummond, J. Philip↗

Numerical Investigation of Dual-Mode Scramjet Combustor with Large Upstream Interaction

Dual-mode scramjet combustor configuration with significant upstream interaction is investigated numerically, The possibility of scaling the domain to accelerate the convergence and reduce the computational time is explored. The supersonic combustor configuration was selected to provide an understanding of key features of upstream interaction and to identify physical and numerical issues relating to modeling of dual-mode configurations. The numerical analysis was performed with vitiated air at freestream Math number of 2.5 using hydrogen as the sonic injectant. Results are presented for two-dimensional models and a three-dimensional jet-to-jet symmetric geometry. Comparisons are made with experimental results. Two-dimensional and three-dimensional results show substantial oblique shock train reaching upstream of the fuel injectors. Flow characteristics slow numerical convergence, while the upstream interaction slowly increases with further iterations. As the flow field develops, the symmetric assumption breaks down. A large separation zone develops and extends further upstream of the step. This asymmetric flow structure is not seen in the experimental data. Results obtained using a sub-scale domain (both two-dimensional and three-dimensional) qualitatively recover the flow physics obtained from full-scale simulations. All results show that numerical modeling using a scaled geometry provides good agreement with full-scale numerical results and experimental results for this configuration. This study supports the argument that numerical scaling is useful in simulating dual-mode scramjet combustor flowfields and could provide an excellent convergence acceleration technique for dual-mode simulations.

Mohieldin, T. O.↗

National Combustion Code: A Multidisciplinary Combustor Design System

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.

Stubbs, Robert M.↗

Critical Propulsion Components: Combustor - Volume 2

Several studies have concluded that a supersonic aircraft, if environmentally acceptable and economically viable, could successfully compete in the 21st century marketplace. However, before industry can commit to what is estimated as a 15 to 20 billion dollar investment, several barrier issues must be resolved. In an effort to address these barrier issues, NASA and Industry teamed to form the High-Speed Research (HSR) program. As part of this program, the Critical Propulsion Components (CPC) element was created and assigned the task of developing those propulsion component technologies necessary to: (1) reduce cruise emissions by a factor of 10 and (2) meet the ever-increasing airport noise restrictions with an economically viable propulsion system. The CPC-identified critical components were ultra-low emission combustors, low-noise/high-performance exhaust nozzles, low-noise fans, and stable/high-performance inlets. Propulsion cycle studies (coordinated with NASA Langley Research Center sponsored airplane studies) were conducted throughout this CPC program to help evaluate candidate components and select the best concepts for the more complex and larger scale research efforts. The propulsion cycle and components ultimately selected were a mixed-flow turbofan (MFTF) engine employing a lean, premixed, prevaporized (LPP) combustor coupled to a two-dimensional mixed compression inlet and a two-dimensional mixer/ejector nozzle. Due to the large amount of material presented in this report, it was prepared in four volumes; Volume 1: Summary, Introduction, and Team. Propulsion System Studies, Volume 2: Combustor, Volume 3: Exhaust Nozzle, and Volume 4: Inlet and Fan/Inlet Acoustic Team.

Source record↗

Electric Plasma Arc-Lamp Combustor Liner Durability Test System Developed

Silicon carbide matrix composites are candidate materials for high-temperature combustor liners. Because through-the-thickness thermal gradients are the primary cause of stress on combustor liners, a unique test facility was developed at the NASA Glenn Research Center at Lewis Field to simulate in-service pure thermal stress distributions in fiber reinforced silicon carbide cylinders. It was developed initially under Phase II of the High- Speed Research Program. This test stand can accommodate 8-in.-long test cylinders that have outer diameters of 4 in. and a wall thickness of about 0.08 to 0.12 in. One cylinder at a time is loaded vertically into the test stand. Water-cooled plates enclose the open ends of the cylinder and provide cooling. Load plates on the exterior side of the water-cooled plates provide support and compression loads. To evaluate a combustor liner material s potential performance, researchers induce thermal gradients with an axisymmetric, direct-current, electric arc within the cylinder while refrigerated air at a rate of 1.5 lb/sec impinges on the outside surface of the cylinder. The achievable through-the-thickness thermal gradient is predicted to be in excess of 200 C. The 8-in. long, 0.5-in.-diameter plasma arc emits full spectrum visible light; radiant intensity exceeds 300 Watts per square centimeters to produce temperatures in excess of 1500 C on materials with emissivity near unity. Because the system does not rely upon the combustion of fuels to achieve the related thermal conditions, ancillary environmental reactions with the sample are eliminated. The system incorporates a standard mechanical test frame, which can impose constant as well cyclical axial stresses up to 2200 lb upon the test piece. Silicon-carbide-fiber reinforced silicon carbide matrix composite cylinders were instrumented with thin-film thermocouples to obtain through-the-thickness thermal flux measurements. Inside wall temperatures reached 1200 C with only 250 A of current. One of the special features of this configuration is the creation of hoop stress states within the cylinder, which up this point have not been obtainable in planar coupon tests. This facility will allow various operational modes, including accelerated tests of thermal transients simulating the effects of repeated engine ignition as well as prescribed thermal and mechanical histories to simulate various duty cycle profiles. Tests can now be performed on thermal-barrier-coated metallic liners and ceramic composite liners that require a combination of high heat flux and controlled mechanical stresses.

Goldsby, Jon C.↗

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

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.

Hicks, Yolanda R.↗

Combustor and Vane Features and Components Tested in a Gas Turbine Environment

The use of ceramic matrix composites (CMCs) as combustor liners and turbine vanes provides the potential of improving next-generation turbine engine performance, through lower emissions and higher cycle efficiency, relative to today s use of superalloy hot-section components. For example, the introduction of film-cooling air in metal combustor liners has led to higher levels of nitrogen oxide (NOx) emissions from the combustion process. An environmental barrier coated (EBC) siliconcarbide- fiber-reinforced silicon carbide matrix (SiC/SiC) composite is a new material system that can operate at higher temperatures, significantly reducing the film-cooling requirements and enabling lower NOx production. Evaluating components and subcomponents fabricated from these advanced CMCs under gas turbine conditions is paramount to demonstrating that the material system can perform as required in the complex thermal stress and environmentally aggressive engine environment. To date, only limited testing has been conducted on CMC combustor and turbine concepts and subelements of this type throughout the industry. As part of the Ultra-Efficient Engine Technology (UEET) Program, the High Pressure Burner Rig (HPBR) at the NASA Glenn Research Center was selected to demonstrate coupon, subcomponent feature, and component testing because it can economically provide the temperatures, pressures, velocities, and combustion gas compositions that closely simulate the engine environments. The results have proven the HPBR to be a highly versatile test rig amenable to multiple test specimen configurations essential to coupon and component testing.

Roinson, R. Craig↗

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

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.

Davoudzadeh, Farhad↗