Numerical Simulation in a Supercirtical CFB Boiler

NASA Astrophysics Data System (ADS)

Zhang, Yanjun; Gaol, Xiang; Luo, Zhongyang; Jiang, Xiaoguo

The dimension of the hot circulation loop of the supercritical CFB boiler is large, and there are many unknowns and challenges that should be identified and resolved during the development. In order to realize a reasonable and reliable design of the hot circulation loop, numerical simulation of gas-solid flow in a supercritical CFB boiler was conducted by using FLUENT software. The working condition of hot circulation loop flow field, gas-solid flow affected by three unsymmetrical cyclones, air distribution and pressure drop in furnace were analyzed. The simulation results showed that the general arrangement of the 600MWe supercritical CFB boiler is reasonable.

CFD Modelling of Particle Mixtures in a 2D CFB

NASA Astrophysics Data System (ADS)

Seppälä, M.; Kallio, S.

The capability of Fluent 6.2.16 to simulate particle mixtures in a laboratory scale 2D circulating fluidized bed (CFB) unit has been tested. In the simulations, the solids were described as one or two particle phases. The loading ratio of small to large particles, particle diameters and the gas inflow velocity were varied. The 40 cm wide and 3 m high 2D CFB was modeled using a grid with 31080 cells. The outflow of particles at the top of the CFB was monitored and emanated particles were fed back to the riser through a return duct. The paper presents the segregation patterns of the particle phases obtained from the simulations. When the fraction of large particles was 50% or larger, large particles segregated, as expected, to the wall regions and to the bottom part of the riser. However, when the fraction of large particles was 10%, an excess of large particles was found in the upper half of the riser. The explanation for this unexpected phenomenon was found in the distribution of the large particles between the slow clusters and the faster moving lean suspension.

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

Novel CFB Boiler Technology with Reconstruction of its Fluidization State

NASA Astrophysics Data System (ADS)

Yang, H. R.; Zhang, H.; Lu, J. F.; Lfu, Q.; Wu, Y. X.; Yuet, G. X.; Su, J.; Fu, Z. P.

Compared with a conventional pulverized coal fired boiler, the combustion efficiency of a CFB boiler is lower while the self-consumed service power is 1-2% higher. The solution of these problems is the key research topic for researchers and manufacturers of CFB boilers. Based on the State Specification Design Theory of CFB boilers, Tsinghua University proposed a novel CFB technology by reconstruction of the fluidization state in the furnace by adjusting the bed inventory and bed quality. Theoretical analyses show that there is an optimal bed pressure drop, around which the boiler operation can achieve the maximal combustion efficiency and with significant reduction of the wear of the heating surface and fan power consumption. The proposed novel process was implemented in a 75t/h CFB boiler. The results of field tests on this boiler validated the theoretical analyses.

A comparison of fuzzy logic and cluster renewal approaches for heat transfer modeling in a 1296 t/h CFB boiler with low level of flue gas recirculation

NASA Astrophysics Data System (ADS)

Błaszczuk, Artur; Krzywański, Jarosław

2017-03-01

The interrelation between fuzzy logic and cluster renewal approaches for heat transfer modeling in a circulating fluidized bed (CFB) has been established based on a local furnace data. The furnace data have been measured in a 1296 t/h CFB boiler with low level of flue gas recirculation. In the present study, the bed temperature and suspension density were treated as experimental variables along the furnace height. The measured bed temperature and suspension density were varied in the range of 1131-1156 K and 1.93-6.32 kg/m3, respectively. Using the heat transfer coefficient for commercial CFB combustor, two empirical heat transfer correlation were developed in terms of important operating parameters including bed temperature and also suspension density. The fuzzy logic results were found to be in good agreement with the corresponding experimental heat transfer data obtained based on cluster renewal approach. The predicted bed-to-wall heat transfer coefficient covered a range of 109-241 W/(m2K) and 111-240 W/(m2K), for fuzzy logic and cluster renewal approach respectively. The divergence in calculated heat flux recovery along the furnace height between fuzzy logic and cluster renewal approach did not exceeded ±2%.

MP-Pic simulation of CFB riser with EMMS-based drag model

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

Li, F.; Song, F.; Benyahia, S.

2012-01-01

MP-PIC (multi-phase particle in cell) method combined with the EMMS (energy minimization multi- scale) drag force model was implemented with the open source program MFIX to simulate the gas–solid flows in CFB (circulatingfluidizedbed) risers. Calculated solid flux by the EMMS drag agrees well with the experimental value; while the traditional homogeneous drag over-predicts this value. EMMS drag force model can also predict the macro-and meso-scale structures. Quantitative comparison of the results by the EMMS drag force model and the experimental measurements show high accuracy of the model. The effects of the number of particles per parcel and wall conditions onmore » the simulation results have also been investigated in the paper. This work proved that MP-PIC combined with the EMMS drag model can successfully simulate the fluidized flows in CFB risers and it serves as a candidate to realize real-time simulation of industrial processes in the future.« less

Effect of particle size distribution on the hydrodynamics of dense CFB risers

NASA Astrophysics Data System (ADS)

Bakshi, Akhilesh; Khanna, Samir; Venuturumilli, Raj; Altantzis, Christos; Ghoniem, Ahmed

2015-11-01

Circulating Fluidized Beds (CFB) are favorable in the energy and chemical industries, due to their high efficiency. While accurate hydrodynamic modeling is essential for optimizing performance, most CFB riser simulations are performed assuming equally-sized solid particles, owing to limited computational resources. Even though this approach yields reasonable predictions, it neglects commonly observed experimental findings suggesting the strong effect of particle size distribution (psd) on the hydrodynamics and chemical conversion. Thus, this study is focused on the inclusion of discrete particle sizes to represent the psd and its effect on fluidization via 2D numerical simulations. The particle sizes and corresponding mass fluxes are obtained using experimental data in dense CFB riser while the modeling framework is described in Bakshi et al 2015. Simulations are conducted at two scales: (a) fine grid to resolve heterogeneous structures and (b) coarse grid using EMMS sub-grid modifications. Using suitable metrics which capture bed dynamics, this study provides insights into segregation and mixing of particles as well as highlights need for improved sub-grid models.

Improvement of CFD Methods for Modeling Full Scale Circulating Fluidized Bed Combustion Systems

NASA Astrophysics Data System (ADS)

Shah, Srujal; Klajny, Marcin; Myöhänen, Kari; Hyppänen, Timo

With the currently available methods of computational fluid dynamics (CFD), the task of simulating full scale circulating fluidized bed combustors is very challenging. In order to simulate the complex fluidization process, the size of calculation cells should be small and the calculation should be transient with small time step size. For full scale systems, these requirements lead to very large meshes and very long calculation times, so that the simulation in practice is difficult. This study investigates the requirements of cell size and the time step size for accurate simulations, and the filtering effects caused by coarser mesh and longer time step. A modeling study of a full scale CFB furnace is presented and the model results are compared with experimental data.

Carbon attrition during the circulating fluidized bed combustion of a packaging-derived fuel

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

Mastellone, M.L.; Arena, U.

1999-05-01

Cylindrical pellets of a market-available packaging-derived fuel, obtained from a mono-material collection of polyethylene terephthalate (PET) bottles, were batchwise fed to a laboratory scale circulating fluidized bed (CFB) combustor. The apparatus, whose riser was 41 mm ID and 4 m high, was operated under both inert and oxidizing conditions to establish the relative importance of purely mechanical attrition and combustion-assisted attrition in generating carbon fines. Silica sand particles of two size distributions were used as inert materials. For each run, carbon load and carbon particle size distribution in the riser and rates of attrited carbon fines escaping the combustor weremore » determined as a function of time. A parallel investigation was carried out with a bubbling fluidized bed (BFB) combustor to point out peculiarities of attrition in CFB combustors. After devolatilization, PET pellets generated fragile aggregates of char and sand, which easily crumbled, leading to single particles, partially covered by a carbon-rich layer. The injected fixed carbon was therefore present in the bed in three phases: an A-phase, made of aggregates of sand and char, an S-phase, made of individual carbon-covered sand particles and an F-phase, made of carbon fines, abraded by the surfaces of the A- and S-phases. The effects of the size of inert material on the different forms under which fixed carbon was present in the bed and on the rate of escape of attrited carbon fines from the combustor were investigated. Features of carbon attrition in CFB and BFB combustors are discussed.« less

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.

Full-scale results for TAM limestone injection

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

Baer, S.

1996-12-31

Information is outlined on the use of thermally active marble (TAM) sorbents in boilers. Data are presented on: the comparison of TAM to limestone; NOVACON process development history; CFB test history; CFB pilot scale test; full-scale CFB trial; August, 1996 CFB demonstration; Foster Wheeler Mount Carmel sorbent feed rate comparison and Ca:S comparison; unburned carbon is ash; and advantages and savings in CFB boilers.

Optimization of Computational Performance and Accuracy in 3-D Transient CFD Model for CFB Hydrodynamics Predictions

NASA Astrophysics Data System (ADS)

Rampidis, I.; Nikolopoulos, A.; Koukouzas, N.; Grammelis, P.; Kakaras, E.

2007-09-01

This work aims to present a pure 3-D CFD model, accurate and efficient, for the simulation of a pilot scale CFB hydrodynamics. The accuracy of the model was investigated as a function of the numerical parameters, in order to derive an optimum model setup with respect to computational cost. The necessity of the in depth examination of hydrodynamics emerges by the trend to scale up CFBCs. This scale up brings forward numerous design problems and uncertainties, which can be successfully elucidated by CFD techniques. Deriving guidelines for setting a computational efficient model is important as the scale of the CFBs grows fast, while computational power is limited. However, the optimum efficiency matter has not been investigated thoroughly in the literature as authors were more concerned for their models accuracy and validity. The objective of this work is to investigate the parameters that influence the efficiency and accuracy of CFB computational fluid dynamics models, find the optimum set of these parameters and thus establish this technique as a competitive method for the simulation and design of industrial, large scale beds, where the computational cost is otherwise prohibitive. During the tests that were performed in this work, the influence of turbulence modeling approach, time and space density and discretization schemes were investigated on a 1.2 MWth CFB test rig. Using Fourier analysis dominant frequencies were extracted in order to estimate the adequate time period for the averaging of all instantaneous values. The compliance with the experimental measurements was very good. The basic differences between the predictions that arose from the various model setups were pointed out and analyzed. The results showed that a model with high order space discretization schemes when applied on a coarse grid and averaging of the instantaneous scalar values for a 20 sec period, adequately described the transient hydrodynamic behaviour of a pilot CFB while the

Integration of concentrated solar power (CSP) and circulating fluidized bed (CFB) power plants - final results of the COMBO-CFB project

NASA Astrophysics Data System (ADS)

Suojanen, Suvi; Hakkarainen, Elina; Kettunen, Ari; Kapela, Jukka; Paldanius, Juha; Tuononen, Minttu; Selek, Istvan; Kovács, Jenö; Tähtinen, Matti

2017-06-01

Hybridization of solar energy together with another energy source is an option to provide heat and power reliably on demand. Hybridization allows decreasing combustion related fuel consumption and emissions, assuring stable grid connection and cutting costs of concentrated solar power technology due to shared power production equipment. The research project "Integration of Concentrated Solar Power (CSP) and Circulating Fluidized Bed (CFB) Power Plants" (COMBO-CFB) has been carried out to investigate the technical possibilities and limitations of the concept. The main focus was on the effect of CSP integration on combustion dynamics and on the joint power cycle, and on the interactions of subsystems. The research provides new valuable experimental data and knowhow about dynamic behaviour of CFB combustion under boundary conditions of the hybrid system. Limiting factors for maximum solar share in different hybridization schemes and suggestions for enhancing the performance of the hybrid system are derived.

Experimental Research on the Dense CFB's Riser and the Simulation Based on the EMMS Model

NASA Astrophysics Data System (ADS)

Wang, X. Y.; Wang, S. D.; Fan, B. G.; Liao, L. L.; Jiang, F.; Xu, X.; Wu, X. Z.; Xiao, Y. H.

2010-03-01

The flow structure in the CFB (circulating fluidized bed) riser has been investigated. Experimental studies were performed in a cold square section unit with 270 mm×270 mm×10 m. Since the drag force model based on homogeneous two-phase flow such as the Gidaspow drag model could not depict the heterogeneous structures of the gas-solid flow, the structure-dependent energy-minimization multi-scale (EMMS) model based on the heterogenerity was applied in the paper and a revised drag force model based on the EMMS model was proposed. A 2D two-fluid model was used to simulate a bench-scale square cross-section riser of a cold CFB. The typical core-annulus structure and the back-mixing near the wall of the riser were observed and the assembly and fragmentation processes of clusters were captured. By comparing with the Gidaspow drag model, the results obtained by the revised drag model based on EMMS shows better consistency with the experimental data. The model can also depict the difference from the two exit configurations. This study once again proves the key role of drag force in CFD (Computational Fluid Dynamics) simulation and also shows the availability of the revised drag model to describe the gas-solid flow in CFB risers.

Study of Gas Solid Flow Characteristics in Cyclone Inlet Ducts of A300Mwe CFB Boiler

NASA Astrophysics Data System (ADS)

Tang, J. Y.; Lu, X. F.; Lai, J.; Liu, H. Z.

Gas solid flow characteristics in cyclone's inlet duct of a 300MW CFB boiler were studied in a cold circulating fluidized bed (CFB) experimental setup according to a 410t/h CFB boiler with a scale of 10∶1. Tracer particles were adopted in the experiment and their motion trajectories in the two kinds of cyclone's inlet ducts were photographed by a high-speed camera. By analyzing the motion trajectories of tracer particles, acceleration performance of particle phases in the two inlet ducts was obtained. Results indicate that the acceleration performance of particles in the long inlet duct is better than that in the short inlet duct, but the pressure drop of the long inlet duct is higher. Meanwhile, under the same operating conditions, both the separation efficiency and the pressure drop of the cyclone are higher when the cyclone is connected with the long inlet duct. Figs 11, Tabs 4 and refs 10.

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.

Low NO sub x heavy fuel combustor concept program

NASA Technical Reports Server (NTRS)

Russell, P.; Beal, G.; Hinton, B.

1981-01-01

A gas turbine technology program to improve and optimize the staged rich lean low NOx combustor concept is described. Subscale combustor tests to develop the design information for optimization of the fuel preparation, rich burn, quick air quench, and lean burn steps of the combustion process were run. The program provides information for the design of high pressure full scale gas turbine combustors capable of providing environmentally clean combustion of minimally of minimally processed and synthetic fuels. It is concluded that liquid fuel atomization and mixing, rich zone stoichiometry, rich zone liner cooling, rich zone residence time, and quench zone stoichiometry are important considerations in the design and scale up of the rich lean combustor.

CFD Modeling of a CFB Riser Using Improved Inlet Boundary Conditions

NASA Astrophysics Data System (ADS)

Peng, B. T.; Zhang, C.; Zhu, J. X.; Qi, X. B.

2010-03-01

A computational fluid dynamics (CFD) model based on Eulerian-Eulerian approach coupled with granular kinetics theory was adopted to investigate the hydrodynamics and flow structures in a circulating fluidized bed (CFB) riser column. A new approach to specify the inlet boundary conditions was proposed in this study to simulate gas-solids flow in CFB risers more accurately. Simulation results were compared with the experimental data, and good agreement between the numerical results and experimental data was observed under different operating conditions, which indicates the effectiveness and accuracy of the CFD model with the proposed inlet boundary conditions. The results also illustrate a clear core annulus structure in the CFB riser under all operating conditions both experimentally and numerically.

Basic design studies for a 600 MWe CFB boiler (270b, 2 x 600 C)

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

Bursi, J.M.; Lafanechere, L.; Jestin, L.

Commercial CFB boilers are currently available in the 300 MWe equivalent range for use with international coal. Retrofitting of Provence 4 with a 250 MWe CFB boiler was an important step in CFB development. In light of the results obtained from two large French units--Emile Huchet 4 (125 MWe) and Provence 4 (250 MWe)--this paper focuses on the main technical points which are currently being studied in relation to the basic design of a 600 MWe CFB boiler, a project that has been undertaken by EDF. The general aim of this project is to demonstrate the competitiveness of a CFBmore » boiler compared with a PF boiler. The main areas of focus in the design of this large CFB boiler with advanced steam conditions are described. These points are subjected to particular analysis from a design standpoint. The objective is to prepare the precise specifications needed to ensure a product which is optimized in terms of quality/cost or service/cost. Due to the present lack of theoretical understanding of the refined and complex two-phase flow, design is a challenge which has to be based on reliable and comprehensive data obtained from large plants in commercial operation. This will ensure that the advantages of CFB which arise from the hydrodynamics within the circulation loop are maintained. The major goals of maintaining good particle residence time and concentration in the furnace are described. Misunderstanding of CFB furnace bottom conditions is also pointed out, with cost reduction and better NO{sub x} capture certainly among the major new targets in relation to bottom furnace design. General problems associated with the heat exchanger arrangement, principally those linked to high steam conditions and, especially, the vaporization system, are discussed. Once again, comparison with PF in this area showed that CFB boilers appear more competitive. Finally, the main area in which there is a need for sharing of CFB experience among CFB users is pointed out.« less

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

Chaos in an imperfectly premixed model combustor.

PubMed

Kabiraj, Lipika; Saurabh, Aditya; Karimi, Nader; Sailor, Anna; Mastorakos, Epaminondas; Dowling, Ann P; Paschereit, Christian O

2015-02-01

This article reports nonlinear bifurcations observed in a laboratory scale, turbulent combustor operating under imperfectly premixed mode with global equivalence ratio as the control parameter. The results indicate that the dynamics of thermoacoustic instability correspond to quasi-periodic bifurcation to low-dimensional, deterministic chaos, a route that is common to a variety of dissipative nonlinear systems. The results support the recent identification of bifurcation scenarios in a laminar premixed flame combustor (Kabiraj et al., Chaos: Interdiscip. J. Nonlinear Sci. 22, 023129 (2012)) and extend the observation to a practically relevant combustor configuration.

Small Gas Turbine Combustor Primary Zone Study

NASA Technical Reports Server (NTRS)

Sullivan, R. E.; Young, E. R.; Miles, G. A.; Williams, J. R.

1983-01-01

A development process is described which consists of design, fabrication, and preliminary test evaluations of three approaches to internal aerodynamic primary zone flow patterns: (1) conventional double vortex swirl stabilization; (2) reverse flow swirl stabilization; and (3) large single vortex flow system. Each concept incorporates special design features aimed at extending the performance capability of the small engine combustor. Since inherent geometry of these combustors result in small combustion zone height and high surface area to volume ratio, design features focus on internal aerodynamics, fuel placement, and advanced cooling. The combustors are evaluated on a full scale annular combustor rig. A correlation of the primary zone performance with the overall performance is accomplished using three intrusion type gas sampling probes located at the exit of the primary zone section. Empirical and numerical methods are used for designing and predicting the performance of the three combustor concepts and their subsequent modifications. The calibration of analytical procedures with actual test results permits an updating of the analytical design techniques applicable to small reverse flow annular combustors.

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

Development of a small-scale power system with meso-scale vortex combustor and thermo-electric device

NASA Astrophysics Data System (ADS)

Shimokuri, D.; Hara, T.; Matsumoto, R.

2015-10-01

A small-scale vortex combustion power system has been developed using a thermo-electric device (TED). The system consisted of a heat medium, TED, and cooling plates. A vortex combustion chamber (7 mm inner diameter and 27 mm long) was fabricated inside the heat medium (40  ×  40  ×  20 mm and 52 g of duralumin). It was found that a stable propane/air flame could be established in the narrow 7 mm channel even for the large heat input conditions of 213 ~ 355 W. With a couple of TEDs, the maximum of 8.1 W (9.8 V  ×  0.83 A) could be successfully obtained for 355 W heat input, which corresponded to the energy conversion rate of 2.4%. The results of the gas and the combustor wall temperature measurements showed that the heat transfer from the burned gas to combustor wall was significantly enhanced by the vortex flow, which contributed to the relatively high efficiency energy conversion on the vortex combustion power system.

Analysis of CFB, a cytokinin-responsive gene of Arabidopsis thaliana encoding a novel F-box protein regulating sterol biosynthesis.

PubMed

Brenner, Wolfram G; Leuendorf, Jan Erik; Cortleven, Anne; Martin, Laetitia B B; Schaller, Hubert; Schmülling, Thomas

2017-05-17

Protein degradation by the ubiquitin-26S proteasome pathway is important for the regulation of cellular processes, but the function of most F-box proteins relevant to substrate recognition is unknown. We describe the analysis of the gene Cytokinin-induced F-box encoding (CFB, AT3G44326), identified in a meta-analysis of cytokinin-related transcriptome studies as one of the most robust cytokinin response genes. F-box domain-dependent interaction with the E3 ubiquitin ligase complex component ASK1 classifies CFB as a functional F-box protein. Apart from F-box and transmembrane domains, CFB contains no known functional domains. CFB is expressed in all plant tissues, predominantly in root tissue. A ProCFB:GFP-GUS fusion gene showed strongest expression in the lateral root cap and during lateral root formation. CFB-GFP fusion proteins were mainly localized in the nucleus and the cytosol but also at the plasma membrane. cfb mutants had no discernible phenotype, but CFB overexpressing plants showed several defects, such as a white upper inflorescence stem, similar to the hypomorphic cycloartenol synthase mutant cas1-1. Both CFB overexpressing plants and cas1-1 mutants accumulated the CAS1 substrate 2,3-oxidosqualene in the white stem tissue, the latter even more after cytokinin treatment, indicating impairment of CAS1 function. This suggests that CFB may link cytokinin and the sterol biosynthesis pathway. © The Author 2017. Published by Oxford University Press on behalf of the Society for Experimental Biology.

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.

Development of Flexi-Burn™ CFB Power Plant to Meet the Challenge of Climate Change

NASA Astrophysics Data System (ADS)

Hackt, Horst; Fant, Zhen; Seltzert, Andrew; Hotta, Arto; Erikssoni, Timo; Sippu, Ossi

Carbon-dioxide capture and storage (CCS) offers the potential for major reductions in carbon- dioxide emissions of fossil fuel-based power generation in the fairly short term, and oxyfuel combustion is one of the identified CCS technology options. Foster Wheeler (FW) is working on reduction of carbon-dioxide with its integrated Flexi-Burn™ CFB technology. The proven high efficiency circulating fluidized-bed (CFB) technology, when coupled with air separation units and carbon purification units, offers a solution for carbon dioxide reduction both in re-powering and in greenfield power plants. CFB technology has the advantages over pulverized coal technology of a more uniform furnace heat flux, increased fuel flexibility and offers the opportunity to further reduce carbon dioxide emissions by co-firing coal with bio-fuels. Development and design of an integrated Flexi-Bum™ CFB steam generator and balance of plant system was conducted for both air mode and oxyfuel mode. Through proper configuration and design, the same steam generator can be switched from air mode to oxyfuel mode without the need for unit shutdown for modifications. The Flexi-Burn™ CFB system incorporates features to maximize plant efficiency and power output when operating in the oxy-firing mode through firing more fuel in the same boiler.

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.

Postcombustion and its influences in 135 MWe CFB boilers

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

Shaohua Li; Hairui Yang; Hai Zhang

2009-09-15

In the cyclone of a circulating fluidized bed (CFB) boiler, a noticeable increment of flue gas temperature, caused by combustion of combustible gas and unburnt carbon content, is often found. Such phenomenon is defined as post combustion, and it could introduce overheating of reheated and superheated steam and extra heat loss of exhaust flue gas. In this paper, mathematical modeling and field measurements on post combustion in 135MWe commercial CFB boilers were conducted. A novel one-dimensional combustion model taking post combustion into account was developed. With this model, the overall combustion performance, including size distribution of various ashes, temperature profile,more » and carbon content profiles along the furnace height, heat release fraction in the cyclone and furnace were predicted. Field measurements were conducted by sampling gas and solid at different positions in the boiler under different loads. The measured data and corresponding model-calculated results were compared. Both prediction and field measurements showed post combustion introduced a temperature increment of flue gas in the cyclone of the 135MWe CFB boiler in the range of 20-50{sup o}C when a low-volatile bituminous coal was fired. Although it had little influence on ash size distribution, post combustion had a remarkable influence on the carbon content profile and temperature profile in the furnace. Moreover, it introduced about 4-7% heat release in the cyclone over the total heat release in the boiler. This fraction slightly increased with total air flow rate and boiler load. Model calculations were also conducted on other two 135MWe CFB boilers burning lignite and anthracite coal, respectively. The results confirmed that post combustion was sensitive to coal type and became more severe as the volatile content of the coal decreased. 15 refs., 11 figs., 4 tabs.« less

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.

Application of CFB technology for large power generating units and CO{sub 2} capture

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

Ryabov, G. A., E-mail: georgy.ryabov@gmail.com; Folomeev, O. M.; Sankin, D. A.

2010-07-15

Data on the development of the circulating fluidized bed (CFB) technology for combustion of fuels in large power generating units are examined. The problems with raising the steam parameters and unit power of boilers with a circulating fluidized bed are examined. With the boiler system at the 460 MW unit at Lagisza (Poland) as an example, the feasibility of raising the efficiency of units with CFB boilers through deep recovery of the heat of the effluent gases and reducing expenditure for in-house needs is demonstrated. Comparative estimates of the capital and operating costs of 225 and 330 MW units aremore » used to determine the conditions for optimum use of CFB boilers in the engineering renovation of thermal power plants in Russia. New areas for the application of CFB technology in CO{sub 2} capture are analyzed in connection with the problem of reducing greenhouse gas emissions.« less

Feasible experimental study on the utilization of a 300 MW CFB boiler desulfurizating bottom ash for construction applications

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

Lu, X.F.; Amano, R.S.

2006-12-15

CFB boiler ash cannot be used as a cement replacement in concrete due to its unacceptably high sulfur content. The disposal in landfills has been the most common means of handling ash in circulating fluidized bed boiler power plants. However for a 300 MW CFB boiler power plant, there will be 600,000 tons of ash discharged per year and will result in great volumes and disposal cost of ash byproduct. It was very necessary to solve the utilization of CFB ash and to decrease the disposal cost of CFB ash. The feasible experimental study results on the utilization of themore » bottom ashes of a 300 MW CFB boiler in Baima power plant in China were reported in this paper. The bottom ashes used for test came from the discharged bottom ashes in a 100 MW CFB boiler in which the anthracite and limestone designed for the 300 MW CFB project was burned. The results of this study showed that the bottom ash could be used for cementitious material, road concrete, and road base material. The masonry cements, road concrete with 30 MPa compressive strength and 4.0 MPa flexural strength, and the road base material used for base courses of the expressway, the main road and the minor lane were all prepared with milled CFB bottom ashes in the lab. The better methods of utilization of the bottom ashes were discussed in this paper.« less

Static and transient performance prediction for CFB boilers using a Bayesian-Gaussian Neural Network

NASA Astrophysics Data System (ADS)

Ye, Haiwen; Ni, Weidou

1997-06-01

A Bayesian-Gaussian Neural Network (BGNN) is put forward in this paper to predict the static and transient performance of Circulating Fluidized Bed (CFB) boilers. The advantages of this network over Back-Propagation Neural Networks (BPNNs), easier determination of topology, simpler and time saving in training process as well as self-organizing ability, make this network more practical in on-line performance prediction for complicated processes. Simulation shows that this network is comparable to the BPNNs in predicting the performance of CFB boilers. Good and practical on-line performance predictions are essential for operation guide and model predictive control of CFB boilers, which are under research by the authors.

Characterization of Centrifugally-Loaded Flame Migration for Ultra-Compact Combustors

DTIC Science & Technology

2011-10-01

11 T04 combustor exit temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11 Q b combustor heat addition...11 Q ab afterburner heat addition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11...the mass flow rates, with heat addition, lead to reaching a specific g-load. In addition to varying g-load, a larger scale UCC will require a

Clean catalytic combustor program

NASA Technical Reports Server (NTRS)

Ekstedt, E. E.; Lyon, T. F.; Sabla, P. E.; Dodds, W. J.

1983-01-01

A combustor program was conducted to evolve and to identify the technology needed for, and to establish the credibility of, using combustors with catalytic reactors in modern high-pressure-ratio aircraft turbine engines. Two selected catalytic combustor concepts were designed, fabricated, and evaluated. The combustors were sized for use in the NASA/General Electric Energy Efficient Engine (E3). One of the combustor designs was a basic parallel-staged double-annular combustor. The second design was also a parallel-staged combustor but employed reverse flow cannular catalytic reactors. Subcomponent tests of fuel injection systems and of catalytic reactors for use in the combustion system were also conducted. Very low-level pollutant emissions and excellent combustor performance were achieved. However, it was obvious from these tests that extensive development of fuel/air preparation systems and considerable advancement in the steady-state operating temperature capability of catalytic reactor materials will be required prior to the consideration of catalytic combustion systems for use in high-pressure-ratio aircraft turbine engines.

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.

Numerical Simulation of Dual-Mode Scramjet Combustors

NASA Technical Reports Server (NTRS)

Rodriguez, C. G.; Riggins, D. W.; Bittner, R. D.

2000-01-01

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.

Alternate-Fueled Combustor-Sector Performance. Parts A and B; (A) Combustor Performance; (B) Combustor Emissions

NASA Technical Reports Server (NTRS)

Shouse, D. T.; Hendricks, R. C.; Lynch, A.; Frayne, C. W.; Stutrud, J. S.; Corporan, E.; Hankins, T.

2012-01-01

Alternate aviation fuels for military or commercial use are required to satisfy MIL-DTL-83133F(2008) or ASTM D 7566 (2010) standards, respectively, and are classified as "drop-in" fuel replacements. To satisfy legacy issues, blends to 50% alternate fuel with petroleum fuels are certified individually on the basis of processing and assumed to be feedstock agnostic. Adherence to alternate fuels and fuel blends requires "smart fueling systems" or advanced fuel-flexible systems, including combustors and engines, without significant sacrifice in performance or emissions requirements. This paper provides preliminary performance (Part A) and emissions and particulates (Part B) combustor sector data. The data are for nominal inlet conditions at 225 psia and 800 F (1.551 MPa and 700 K), for synthetic-paraffinic-kerosene- (SPK-) type (Fisher-Tropsch (FT)) fuel and blends with JP-8+100 relative to JP-8+100 as baseline fueling. Assessments are made of the change in combustor efficiency, wall temperatures, emissions, and luminosity with SPK of 0%, 50%, and 100% fueling composition at 3% combustor pressure drop. The performance results (Part A) indicate no quantifiable differences in combustor efficiency, a general trend to lower liner and higher core flow temperatures with increased FT fuel blends. In general, emissions data (Part B) show little differences, but with percent increase in FT-SPK-type fueling, particulate emissions and wall temperatures are less than with baseline JP-8. High-speed photography illustrates both luminosity and combustor dynamic flame characteristics.

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.

Flame Stability Limit and Exhaust Emissions of Low Calorific Fuel Combustion in Turbulent Diffusion Combustor for a Small-Scale Fuel Cell

NASA Astrophysics Data System (ADS)

Koseki, Hidenori

This paper describes an investigation conducted on flame stability and exhaust emissions from a turbulent diffusion combustor, fueled with low-calorific gas, for a small-scale fuel cell. It is important to maintain flame stability in the combustor, even under lean fuel conditions, and to suppress CO emission in the exhaust gas. An imitation off-gas, in which hydrogen and methane were diluted by adding nitrogen, with Wobbe indices ranging from ca. 4400-8700, corresponding to the fuel utility ratio of 90%-60%in the fuel cell, was supplied to the combustor, and the blow-off limits, CO, and NOx emissions were experimentally investigated. The results show that the blow-off excess air ratios increases with an increasing Wobbe index and with decreasing fuel input to the combustor, and that they are proportional to the hydrogen concentration in the fuel to the power of 0.5-1.0. In addition, it was found that the Damköhler numbers at blow-off limits decreased with decreasing fuel input and with increasing Wobbe indices, and that the product of (SS / V·M)A[H2][O2]0.5 was constant at blow-off limits. Furthermore, NOx emissions from the combustor were low, less than 20ppmV (O2=0%), it was also found that the apparent activation energy of NOx emission derived from Arrhenius plots was almost equal to that of prompt NO in the combustion of imitation off-gas.

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.

Computational Simulation of Acoustic Modes in Rocket Combustors

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Merkle, C. L.; Sankaran, V.; Ellis, M.

2004-01-01

A combination of computational fluid dynamic analysis and analytical solutions is being used to characterize the dominant modes in liquid rocket engines in conjunction with laboratory experiments. The analytical solutions are based on simplified geometries and flow conditions and are used for careful validation of the numerical formulation. The validated computational model is then extended to realistic geometries and flow conditions to test the effects of various parameters on chamber modes, to guide and interpret companion laboratory experiments in simplified combustors, and to scale the measurements to engine operating conditions. In turn, the experiments are used to validate and improve the model. The present paper gives an overview of the numerical and analytical techniques along with comparisons illustrating the accuracy of the computations as a function of grid resolution. A representative parametric study of the effect of combustor mean flow Mach number and combustor aspect ratio on the chamber modes is then presented for both transverse and longitudinal modes. The results show that higher mean flow Mach numbers drive the modes to lower frequencies. Estimates of transverse wave mechanics in a high aspect ratio combustor are then contrasted with longitudinal modes in a long and narrow combustor to provide understanding of potential experimental simulations.

Computational fluid dynamics assessment: Volume 2, Isothermal simulations of the METC bench-scale coal-water slurry combustor: Final report

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

Celik, I.; Chattree, M.

1988-09-01

The isothermal turbulent, swirling flow inside the METC pressurized bench-scale combustor has been simulated using ISOPCGC-2. The effects of the swirl numbers, the momentum ratio of the primary to secondary streams, the annular wall thickness, and the quarl angle on the flow and mixing patterns have been investigated. The results that with the present configuration of the combustor, an annular recirculation zone is present up to secondary swirl number of four. A central (on axis) recirculation zone can be obtained by increasing the momentum of the secondary stream by decreasing the annular area at the reactor inlet. The mixing ofmore » the primary (fuel carrier) air with the secondary air improves only slightly due to swirl unless a central recirculation zone is present. Good mixing is achieved in the quarl region when a central recirculation zone is present. A preliminary investigation of the influence of placing flow regulators inside the the combustor shows that they influence the flow field significantly and that there is a potential of obtaining optimum flow conditions using these flow regulators. 58 refs., 47 figs., 12 tabs.« less

Low NO(x) Combustor Development

NASA Technical Reports Server (NTRS)

Kastl, J. A.; Herberling, P. V.; Matulaitis, J. M.

2005-01-01

The goal of these efforts was the development of an ultra-low emissions, lean-burn combustor for the High Speed Civil Transport. The HSCT Mach 2.4 FLADE C1 Cycle was selected as the baseline engine cycle. A preliminary compilation of performance requirements for the HSCT combustor system was developed. The emissions goals of the program, baseline engine cycle, and standard combustor performance requirements were considered in developing the compilation of performance requirements. Seven combustor system designs were developed. The development of these system designs was facilitated by the use of spreadsheet-type models which predicted performance of the combustor systems over the entire flight envelope of the HSCT. A chemical kinetic model was developed for an LPP combustor and employed to study NO(x) formation kinetics, and CO burnout. These predictions helped to define the combustor residence time. Five fuel-air mixer concepts were analyzed for use in the combustor system designs. One of the seven system designs, one using the Swirl-Jet and Cyclone Swirler fuel-air mixers, was selected for a preliminary mechanical design study.

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.

Experimental clean combustor program, phase 1

NASA Technical Reports Server (NTRS)

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

1975-01-01

Full annular versions of advanced combustor designs, sized to fit within the CF6-50 engine, were defined, manufactured, and tested at high pressure conditions. Configurations were screened, and significant reductions in CO, HC, and NOx emissions levels were achieved with two of these advanced combustor design concepts. Emissions and performance data at a typical AST cruise condition were also obtained along with combustor noise data as a part of an addendum to the basic program. The two promising combustor design approaches evolved in these efforts were the Double Annular Combustor and the Radial/Axial Combustor. With versions of these two basic combustor designs, CO and HC emissions levels at or near the target levels were obtained. Although the low target NOx emissions level was not obtained with these two advanced combustor designs, significant reductions were relative to the NOx levels of current technology combustors. Smoke emission levels below the target value were obtained.

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.

Unstructured LES of Reacting Multiphase Flows in Realistic Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Ham, Frank; Apte, Sourabh; Iaccarino, Gianluca; Wu, Xiao-Hua; Herrmann, Marcus; Constantinescu, George; Mahesh, Krishnan; Moin, Parviz

2003-01-01

As part of the Accelerated Strategic Computing Initiative (ASCI) program, an accurate and robust simulation tool is being developed to perform high-fidelity LES studies of multiphase, multiscale turbulent reacting flows in aircraft gas turbine combustor configurations using hybrid unstructured grids. In the combustor, pressurized gas from the upstream compressor is reacted with atomized liquid fuel to produce the combustion products that drive the downstream turbine. The Large Eddy Simulation (LES) approach is used to simulate the combustor because of its demonstrated superiority over RANS in predicting turbulent mixing, which is central to combustion. This paper summarizes the accomplishments of the combustor group over the past year, concentrating mainly on the two major milestones achieved this year: 1) Large scale simulation: A major rewrite and redesign of the flagship unstructured LES code has allowed the group to perform large eddy simulations of the complete combustor geometry (all 18 injectors) with over 100 million control volumes; 2) Multi-physics simulation in complex geometry: The first multi-physics simulations including fuel spray breakup, coalescence, evaporation, and combustion are now being performed in a single periodic sector (1/18th) of an actual Pratt & Whitney combustor geometry.

Study of CFB Simulation Model with Coincidence at Multi-Working Condition

NASA Astrophysics Data System (ADS)

Wang, Z.; He, F.; Yang, Z. W.; Li, Z.; Ni, W. D.

A circulating fluidized bed (CFB) two-stage simulation model was developed. To realize the model results coincident with the design value or real operation value at specified multi-working conditions and with capability of real-time calculation, only the main key processes were taken into account and the dominant factors were further abstracted out of these key processes. The simulation results showed a sound accordance at multi-working conditions, and confirmed the advantage of the two-stage model over the original single-stage simulation model. The combustion-support effect of secondary air was investigated using the two-stage model. This model provides a solid platform for investigating the pant-leg structured CFB furnace, which is now under design for a supercritical power plant.

Heat Transfer Coefficient Distribution in the Furnace of a 300MWe CFB Boiler

NASA Astrophysics Data System (ADS)

Zhang, P.; Lu, J. F.; Yang, H. R.; Zhang, J. S.; Zhang, H.; Yue, G. X.

Properly understanding and calculating the distributions of heat flux and heat transfer coefficient (α) in the furnace is important in designing a circulating fluidized bed (CFB) boiler, especially with supercritical parameters. Experimental study on the heat transfer in a commercial 300MWe CFB boiler was conducted. The α from the bed to the water wall was measured by the finite element method (FEM), at five different heights. The influence of suspension density and bed temperature on α was analyzed. It was found that the pressure difference between the inlet and exit of the three cyclones, and the chamber pressure of the corresponding loop seal were not equal. The results indicated the suspension solid density was non-uniform in the cross section at a certain height. Consequently, the distributions of heat flux and α in the horizontal plane in the furnace was non-uniform. The furnace can divided into three sections according to the arrangement of the platen superheaters hanging in the upper CFB furnace. In each section, the heat flux near the center showed increasing trend.

Dual-Mode Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J (Inventor); Dippold, Vance F (Inventor)

2013-01-01

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

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.

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.

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.

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.

Numerical exploration of mixing and combustion in ethylene fueled scramjet combustor

NASA Astrophysics Data System (ADS)

Dharavath, Malsur; Manna, P.; Chakraborty, Debasis

2015-12-01

Numerical simulations are performed for full scale scramjet combustor of a hypersonic airbreathing vehicle with ethylene fuel at ground test conditions corresponding to flight Mach number, altitude and stagnation enthalpy of 6.0, 30 km and 1.61 MJ/kg respectively. Three dimensional RANS equations are solved along with species transport equations and SST-kω turbulence model using Commercial CFD software CFX-11. Both nonreacting (with fuel injection) and reacting flow simulations [using a single step global reaction of ethylene-air with combined combustion model (CCM)] are carried out. The computational methodology is first validated against experimental results available in the literature and the performance parameters of full scale combustor in terms of thrust, combustion efficiency and total pressure loss are estimated from the simulation results. Parametric studies are conducted to study the effect of fuel equivalence ratio on the mixing and combustion behavior of the combustor.

Application of Multivariable Model Predictive Advanced Control for a 2×310T/H CFB Boiler Unit

NASA Astrophysics Data System (ADS)

Weijie, Zhao; Zongllao, Dai; Rong, Gou; Wengan, Gong

When a CFB boiler is in automatic control, there are strong interactions between various process variables and inverse response characteristics of bed temperature control target. Conventional Pill control strategy cannot deliver satisfactory control demand. Kalman wave filter technology is used to establish a non-linear combustion model, based on the CFB combustion characteristics of bed fuel inventory, heating values, bed lime inventory and consumption. CFB advanced combustion control utilizes multivariable model predictive control technology to optimize primary and secondary air flow, bed temperature, air flow, fuel flow and heat flux. In addition to providing advanced combustion control to 2×310t/h CFB+1×100MW extraction condensing turbine generator unit, the control also provides load allocation optimization and advanced control for main steam pressure, combustion and temperature. After the successful implementation, under 10% load change, main steam pressure varied less than ±0.07MPa, temperature less than ±1°C, bed temperature less than ±4°C, and air flow (O2) less than ±0.4%.

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

Prediction of Agglomeration, Fouling, and Corrosion Tendency of Fuels in CFB Co-Combustion

NASA Astrophysics Data System (ADS)

Barišć, Vesna; Zabetta, Edgardo Coda; Sarkki, Juha

Prediction of agglomeration, fouling, and corrosion tendency of fuels is essential to the design of any CFB boiler. During the years, tools have been successfully developed at Foster Wheeler to help with such predictions for the most commercial fuels. However, changes in fuel market and the ever-growing demand for co-combustion capabilities pose a continuous need for development. This paper presents results from recently upgraded models used at Foster Wheeler to predict agglomeration, fouling, and corrosion tendency of a variety of fuels and mixtures. The models, subject of this paper, are semi-empirical computer tools that combine the theoretical basics of agglomeration/fouling/corrosion phenomena with empirical correlations. Correlations are derived from Foster Wheeler's experience in fluidized beds, including nearly 10,000 fuel samples and over 1,000 tests in about 150 CFB units. In these models, fuels are evaluated based on their classification, their chemical and physical properties by standard analyses (proximate, ultimate, fuel ash composition, etc.;.) alongside with Foster Wheeler own characterization methods. Mixtures are then evaluated taking into account the component fuels. This paper presents the predictive capabilities of the agglomeration/fouling/corrosion probability models for selected fuels and mixtures fired in full-scale. The selected fuels include coals and different types of biomass. The models are capable to predict the behavior of most fuels and mixtures, but also offer possibilities for further improvements.

HSCT Sector Combustor Evaluations for Demonstration Engine

NASA Technical Reports Server (NTRS)

Greenfield, Stuart; Heberling, Paul; Kastl, John; Matulaitis, John; Huff, Cynthia

2004-01-01

In LET Task 10, critical development issues of the HSCT lean-burn low emissions combustor were addressed with a range of engineering tools. Laser diagnostics and CFD analysis were applied to develop a clearer understanding of the fuel-air premixing process and premixed combustion. Subcomponent tests evaluated the emissions and operability performance of the fuel-air premixers. Sector combustor tests evaluated the performance of the integrated combustor system. A 3-cup sector was designed and procured for laser diagnostics studies at NASA Glenn. The results of these efforts supported the earlier selection of the Cyclone Swirler as the pilot stage premixer and the IMFH (Integrated Mixer Flame Holder) tube as the main stage premixer of the LPP combustor. In the combustor system preliminary design subtask, initial efforts to transform the sector combustor design into a practical subscale engine combustor met with significant challenges. Concerns about the durability of a stepped combustor dome and the need for a removable fuel injection system resulted in the invention and refinement of the MRA (Multistage Radial Axial) combustor system in 1994. The MRA combustor was selected for the HSR Phase II LPP subscale combustor testing in the CPC Program.

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.

Coexistence of alkaline-carbonatite complexes and high-MgO CFB in the Paranà-Etendeka province: Insights on plume-lithosphere interactions in the Gondwana realm

NASA Astrophysics Data System (ADS)

Natali, Claudio; Beccaluva, Luigi; Bianchini, Gianluca; Siena, Franca

2018-01-01

A careful review of petrological and geochemical data on the Paranà-Etendeka igneous province is reported, with particular attention being devoted to the relationships between high-MgO CFB (tholeiitic basalts-picrites) and nearly coeval alkaline-carbonatite complexes linked to the same extensional tectonics on a regional scale. At 135-130 Ma, the tectonomagmatic activity was focused in Etendeka, the centre of the restored province, and characterised by an exclusive occurrence of the hottest and deepest high-MgO CFB (potential temperature Tp up to 1590 °C and pressure up to 5 GPa) possessing the same Sr-Nd-Pb isotopic composition of the "Gough" geochemical component, a marker of the initial Tristan plume activity. Etendeka high-MgO CFB thus represent the most genuine proxies of sublithospheric melts generated at the plume axis and are relatively unaffected by lithospheric contamination. Nearly coeval (133-128 Ma) alkaline‑carbonatite complexes cluster around the extensional structures of the Ponta Grossa Arch (e.g., Jacupiranga and Juquia in Brazil) and the Damara Belt (e.g., Erongo, Okurusu, Okenyenya and Paresis in Namibia), both of which intersect the early track of the south Atlantic opening. Compared to high-MgO CFB, alkaline magmas display distinctive isotopic signatures and an incompatible element distribution consistent with their generation from lithospheric mantle sources, which were variably metasomatised (veined?) by amphibole and phlogopite. Metasomes of alkaline mantle sources have a HIMU affinity and are dominated by amphibole in Namibia, whereas they display EM1 tendency and a more relevant role of phlogopite in Brazil, which implies important lithospheric differences at a regional scale. The tectonomagmatic features of Paranà-Etendeka -also shared by other Gondwana LIPs, such as Deccan and Karoo- can be reconciled by a generalized model where a hot plume impinging on a relatively thick lithosphere caused, in the axial zone, the contemporaneous

Variable volume combustor

DOEpatents

Ostebee, Heath Michael; Ziminsky, Willy Steve; Johnson, Thomas Edward; Keener, Christopher Paul

2017-01-17

The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a linear actuator so as to maneuver the micro-mixer fuel nozzles axially along the liner.

Fuel cell system combustor

DOEpatents

Pettit, William Henry

2001-01-01

A fuel cell system including a fuel reformer heated by a catalytic combustor fired by anode and cathode effluents. The combustor includes a turbulator section at its input end for intimately mixing the anode and cathode effluents before they contact the combustors primary catalyst bed. The turbulator comprises at least one porous bed of mixing media that provides a tortuous path therethrough for creating turbulent flow and intimate mixing of the anode and cathode effluents therein.

Flow process in combustors

NASA Technical Reports Server (NTRS)

Gouldin, F. C.

1982-01-01

Fluid mechanical effects on combustion processes in steady flow combustors, especially gas turbine combustors were investigated. Flow features of most interest were vorticity, especially swirl, and turbulence. Theoretical analyses, numerical calculations, and experiments were performed. The theoretical and numerical work focused on noncombusting flows, while the experimental work consisted of both reacting and nonreacting flow studies. An experimental data set, e.g., velocity, temperature and composition, was developed for a swirl flow combustor for use by combustion modelers for development and validation work.

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.

Characterization of Combustion and Emission of Several Kinds of Herbaceous Biomass Pellets in a Circulating Fluidized Bed Combustor

NASA Astrophysics Data System (ADS)

Li, S. Y.; Teng, H. P.; Jiao, W. H.; Shang, L. L.; Lu, Q. G.

Characterizations of combustion and emission of four kinds of herbaceous biomass pellets were investigated in a 0.15 MWt circulating fluidized bed. Corn stalk, wheat stalk, cotton stalk and king grass, which are typical herbaceous biomass in China, were chosen for this study. Temperature profile, emission in flue gas and agglomeration were studied by changing the combustion temperature between 750°C and 880°C. The combustion efficiencies are in the range from 97.4% to 99.4%, which are relatively high due to the homogeneous temperature profiles and good circulating fluidization of bed material. Suitable combustion temperatures for the different herbaceous biomass are mainly depended on the emission and bed agglomeration. SO2 and HCl concentrations in flue gas are in direct proportion to the sulfur and chlorine contents of the herbaceous biomass. Agglomeration at the cyclone leg and the loop seal is the main reason for defluidization in the CFB combustor.

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

NASA Technical Reports Server (NTRS)

Mohieldin, T. O.; Tiwari, S. N.; Reubush, David E. (Technical Monitor)

2004-01-01

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.

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 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

Low NOx heavy fuel combustor concept program. Phase 1: Combustion technology generation

NASA Astrophysics Data System (ADS)

Lew, H. G.; Carl, D. R.; Vermes, G.; Dezubay, E. A.; Schwab, J. A.; Prothroe, D.

1981-10-01

The viability of low emission nitrogen oxide (NOx) gas turbine combustors for industrial and utility application. Thirteen different concepts were evolved and most were tested. Acceptable performance was demonstrated for four of the combustors using ERBS fuel and ultralow NOx emissions were obtained for lean catalytic combustion. Residual oil and coal derived liquids containing fuel bound nitrogen (FBN) were also used at test fuels, and it was shown that staged rich/lean combustion was effective in minimizing the conversion of FBN to NOx. The rich/lean concept was tested with both modular and integral combustors. While the ceramic lined modular configuration produced the best results, the advantages of the all metal integral burners make them candidates for future development. An example of scaling the laboratory sized combustor to a 100 MW size engine is included in the report as are recommendations for future work.

Low NOx heavy fuel combustor concept program. Phase 1: Combustion technology generation

NASA Technical Reports Server (NTRS)

Lew, H. G.; Carl, D. R.; Vermes, G.; Dezubay, E. A.; Schwab, J. A.; Prothroe, D.

1981-01-01

The viability of low emission nitrogen oxide (NOx) gas turbine combustors for industrial and utility application. Thirteen different concepts were evolved and most were tested. Acceptable performance was demonstrated for four of the combustors using ERBS fuel and ultralow NOx emissions were obtained for lean catalytic combustion. Residual oil and coal derived liquids containing fuel bound nitrogen (FBN) were also used at test fuels, and it was shown that staged rich/lean combustion was effective in minimizing the conversion of FBN to NOx. The rich/lean concept was tested with both modular and integral combustors. While the ceramic lined modular configuration produced the best results, the advantages of the all metal integral burners make them candidates for future development. An example of scaling the laboratory sized combustor to a 100 MW size engine is included in the report as are recommendations for future work.

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.

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

NASA Astrophysics Data System (ADS)

Bobba, Mohan K.

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.

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.

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.

Alternate-Fueled Combustor-Sector Performance: Part A: Combustor Performance Part B: Combustor Emissions

NASA Technical Reports Server (NTRS)

Shouse, D. T.; Neuroth, C.; Henricks, R. C.; Lynch, A.; Frayne, C.; Stutrud, J. S.; Corporan, E.; Hankins, T.

2010-01-01

Alternate aviation fuels for military or commercial use are required to satisfy MIL-DTL-83133F(2008) or ASTM D 7566 (2010) standards, respectively, and are classified as drop-in fuel replacements. To satisfy legacy issues, blends to 50% alternate fuel with petroleum fuels are certified individually on the basis of feedstock. Adherence to alternate fuels and fuel blends requires smart fueling systems or advanced fuel-flexible systems, including combustors and engines without significant sacrifice in performance or emissions requirements. This paper provides preliminary performance (Part A) and emissions and particulates (Part B) combustor sector data for synthetic-parafinic-kerosene- (SPK-) type fuel and blends with JP-8+100 relative to JP-8+100 as baseline fueling.

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.

Effect of Spray Cone Angle on Flame Stability in an Annular Gas Turbine Combustor

NASA Astrophysics Data System (ADS)

Mishra, R. K.; Kumar, S. Kishore; Chandel, Sunil

2016-04-01

Effect of fuel spray cone angle in an aerogas turbine combustor has been studied using computational fluid dynamics (CFD) and full-scale combustor testing. For CFD analysis, a 22.5° sector of an annular combustor is modeled and the governing equations are solved using the eddy dissipation combustion model in ANSYS CFX computational package. The analysis has been carried out at 125 kPa and 303 K inlet conditions for spray cone angles from 60° to 140°. The lean blowout limits are established by studying the behavior of combustion zone during transient engine operation from an initial steady-state condition. The computational study has been followed by testing the practical full-scale annular combustor in an aerothermal test facility. The experimental result is in a good agreement with the computational predictions. The lean blowout fuel-air ratio increases as the spray cone angle is decreased at constant operating pressure and temperature. At higher spray cone angle, the flame and high-temperature zone moves upstream close to atomizer face and a uniform flame is sustained over a wide region causing better flame stability.

Integrated CFD modeling of gas turbine combustors

NASA Technical Reports Server (NTRS)

Fuller, E. J.; Smith, C. E.

1993-01-01

3D, curvilinear, multi-domain CFD analysis is becoming a valuable tool in gas turbine combustor design. Used as a supplement to experimental testing. CFD analysis can provide improved understanding of combustor aerodynamics and used to qualitatively assess new combustor designs. This paper discusses recent advancements in CFD combustor methodology, including the timely integration of the design (i.e. CAD) and analysis (i.e. CFD) processes. Allied Signal's F124 combustor was analyzed at maximum power conditions. The assumption of turbulence levels at the nozzle/swirler inlet was shown to be very important in the prediction of combustor exit temperatures. Predicted exit temperatures were compared to experimental rake data, and good overall agreement was seen. Exit radial temperature profiles were well predicted, while the predicted pattern factor was 25 percent higher than the harmonic-averaged experimental pattern factor.

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.

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.

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.

Applications of tribology to determine attrition by wear of particulate solids in CFB systems

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

Bayham, Samuel C.; Breault, Ronald; Monazam, Esmail

In recent years, much attention has been focused on the development of novel technologies for carbon capture and chemicals production that utilize a circulating fluidized bed (CFB) configuration; examples include chemical looping combustion and circulation of temperature swing adsorbents in a CFB configuration for CO 2 capture. A major uncertainty in determining the economic feasibility of these technologies is the required solids makeup rate, which, among other factors, is due to impact and wear attrition at various locations, including standpipes, cyclones, and the gas jets in fluid beds. While correlations have been developed that estimate the attrition rates at thesemore » areas, these correlations are dependent on constants that are uncertain without extensive experiment in the corresponding unit operation. Thus, it is difficult to determine the attrition rate a priori without performing extensive experiments on the materials or scaling up entirely. In this work, the authors outline a methodology for predictive attrition based on fundamental material properties from fields of tribology—specifically, the study of wear—to the knowledge of forces and sliding distances determined from hydrodynamic models to develop basic attrition models for novel CFB systems. The equations are derived for the standpipe and cyclone, which are common components found in CFBs, and the cyclone equation is compared to experimental data of attrition in the literature. The cyclone equation derived in this work results in an abrasion rate based on (1) material properties such as particle density and hardness, (2) inlet velocity, and (3) cyclone geometry. According to this equation, increasing the diameter of the cyclone and the solids inlet velocity tends to increase the rate of abrasion of the catalyst, while decreasing the hardness increases the abrasion rate. The functionality of the increasing attrition rate with velocity increase implies that increasing the efficiency of the

Applications of tribology to determine attrition by wear of particulate solids in CFB systems

DOE PAGES

Bayham, Samuel C.; Breault, Ronald; Monazam, Esmail

2016-11-03

In recent years, much attention has been focused on the development of novel technologies for carbon capture and chemicals production that utilize a circulating fluidized bed (CFB) configuration; examples include chemical looping combustion and circulation of temperature swing adsorbents in a CFB configuration for CO 2 capture. A major uncertainty in determining the economic feasibility of these technologies is the required solids makeup rate, which, among other factors, is due to impact and wear attrition at various locations, including standpipes, cyclones, and the gas jets in fluid beds. While correlations have been developed that estimate the attrition rates at thesemore » areas, these correlations are dependent on constants that are uncertain without extensive experiment in the corresponding unit operation. Thus, it is difficult to determine the attrition rate a priori without performing extensive experiments on the materials or scaling up entirely. In this work, the authors outline a methodology for predictive attrition based on fundamental material properties from fields of tribology—specifically, the study of wear—to the knowledge of forces and sliding distances determined from hydrodynamic models to develop basic attrition models for novel CFB systems. The equations are derived for the standpipe and cyclone, which are common components found in CFBs, and the cyclone equation is compared to experimental data of attrition in the literature. The cyclone equation derived in this work results in an abrasion rate based on (1) material properties such as particle density and hardness, (2) inlet velocity, and (3) cyclone geometry. According to this equation, increasing the diameter of the cyclone and the solids inlet velocity tends to increase the rate of abrasion of the catalyst, while decreasing the hardness increases the abrasion rate. The functionality of the increasing attrition rate with velocity increase implies that increasing the efficiency of the

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.

Staged cascade fluidized bed combustor

DOEpatents

Cannon, Joseph N.; De Lucia, David E.; Jackson, William M.; Porter, James H.

1984-01-01

A fluid bed combustor comprising a plurality of fluidized bed stages interconnected by downcomers providing controlled solids transfer from stage to stage. Each stage is formed from a number of heat transfer tubes carried by a multiapertured web which passes fluidizing air to upper stages. The combustor cross section is tapered inwardly from the middle towards the top and bottom ends. Sorbent materials, as well as non-volatile solid fuels, are added to the top stages of the combustor, and volatile solid fuels are added at an intermediate stage.

Lean Blow-out Studies in a Swirl Stabilized Annular Gas Turbine Combustor

NASA Astrophysics Data System (ADS)

Mishra, R. K.; Kishore Kumar, S.; Chandel, Sunil

2015-05-01

Lean blow out characteristics in a swirl stabilized aero gas turbine combustor have been studied using computational fluid dynamics. For CFD analysis, a 22.5° sector of an annular combustor is modeled using unstructured tetrahedral meshes comprising 1.2 × 106 elements. The governing equations are solved using the eddy dissipation combustion model in CFX. The primary combustion zone is analyzed by considering it as a well stirred reactor. The analysis has been carried out for different operating conditions of the reactants entering into the control volume. The results are treated as the base-line or reference values. Combustion lean blow-out limits are further characterized studying the behavior of combustion zone during transient engine operation. The validity of the computational study has been established by experimental study on a full-scale annular combustor in an air flow test facility that is capable of simulating different conditions at combustor inlet. The experimental result is in a good agreement with the analytical predictions. Upon increasing the combustor mass flow, the lean blow out limit increases, i.e., the blow out occurs at higher fuel-air ratios. In addition, when the operating pressure decreases, the lean blow out limit increases, i.e., blow out occurs at higher fuel-air ratios.

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.

EFFECT OF SOOT AND COPPER COMBUSTOR DEPOSITS ON DIOXIN EMISSIONS

EPA Science Inventory

An experimental study was conducted to investigate the effects of residual soot and copper combustor deposits on the formation of polychlorinated dibenzo-p-dioxins (PCDDs) and polychlorinated dibenzofurans (PCDFs) during the combustion of a chlorinated waste. In a bench-scale set...

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.

Alternate-Fueled Combustor-Sector Performance

NASA Technical Reports Server (NTRS)

Thomas, Anna E.; Saxena, Nikita T.; Shouse, Dale T.; Neuroth, Craig; Hendricks, Robert C.; Lynch, Amy; Frayne, Charles W.; Stutrud, Jeffrey S.; Corporan, Edwin; Hankins, Terry

2013-01-01

In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MIL-DTL-83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as "drop-in" fuel replacements. This report provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fischer-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 degF (533 K), 125 psia (0.86 MPa) at 625 degF (603 K), 175 psia (1.21 MPa) at 725 degF (658 K), and 225 psia (1.55 MPa) at 790 degF (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3, 4, and 5 percent combustor pressure drop (DP) for fuel:air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade and vane lives.

Alternate-Fueled Combustor-Sector Performance

NASA Technical Reports Server (NTRS)

Thomas, Anna E.; Saxena, Nikita T.; Shouse, Dale T.; Neuroth, Craig; Hendricks, Robert C.; Lynch, Amy; Frayne, Charles W.; Stutrud, Jeffrey S.; Corporan, Edwin; Hankins, Terry

2012-01-01

In order to realize alternative fueling for military and commercial use, the industry has set forth guidelines that must be met by each fuel. These aviation fueling requirements are outlined in MILDTL- 83133F(2008) or ASTM D 7566 Annex (2011) standards, and are classified as drop-in fuel replacements. This paper provides combustor performance data for synthetic-paraffinic-kerosene- (SPK-) type (Fisher-Tropsch (FT)) fuel and blends with JP-8+100, relative to JP-8+100 as baseline fueling. Data were taken at various nominal inlet conditions: 75 psia (0.52 MPa) at 500 F (533 K), 125 psia (0.86 MPa) at 625 F (603 K), 175 psia (1.21 MPa) at 725 F (658 K), and 225 psia (1.55 MPa) at 790 F (694 K). Combustor performance analysis assessments were made for the change in flame temperatures, combustor efficiency, wall temperatures, and exhaust plane temperatures at 3%, 4%, and 5% combustor pressure drop (% delta P) for fuel: air ratios (F/A) ranging from 0.010 to 0.025. Significant general trends show lower liner temperatures and higher flame and combustor outlet temperatures with increases in FT fueling relative to JP-8+100 fueling. The latter affects both turbine efficiency and blade/vane life.

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.

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.

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

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, 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.

Computational Study of Combustor-Turbine Interactions

NASA Technical Reports Server (NTRS)

Miki, Kenji; Liou, Meng-Sing

2017-01-01

The Open National Combustion Code (OpenNCC) is applied to the simulation of a realisticcombustor configuration (Energy Efficient Engine (E3)) in order to investigate the unsteady flow fields inside the combustor and around the first stage stator of a high pressure turbine (HPT). We consider one-twelfth (24 degrees) of the full annular E3 combustor with three different geometries of the combustor exit: one without the vane, and two others with the vane set at different relative positions in relation to the fuel nozzle (clocking). Although it is common to take the exit flow profiles obtained by separately simulating the combustor and then feed it as the inflow profile when modeling the HPT, our studies show that the unsteady flow fields are influenced by the presence of the vane as well as clocking. More importantly, the characteristics (e.g., distribution and strength) of the high temperature spots (i.e., hot-streaks) appearing on the vane significantly alters. This indicates the importance of simultaneously modeling both the combustor and the HPT to understand the mechanics of the unsteady formulation of hot-streaks.

Ejector-Enhanced, Pulsed, Pressure-Gain Combustor

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Dougherty, Kevin T.

2009-01-01

An experimental combination of an off-the-shelf valved pulsejet combustor and an aerodynamically optimized ejector has shown promise as a prototype of improved combustors for gas turbine engines. Despite their name, the constant pressure combustors heretofore used in gas turbine engines exhibit typical pressure losses ranging from 4 to 8 percent of the total pressures delivered by upstream compressors. In contrast, the present ejector-enhanced pulsejet combustor exhibits a pressure rise of about 3.5 percent at overall enthalpy and temperature ratios compatible with those of modern turbomachines. The modest pressure rise translates to a comparable increase in overall engine efficiency and, consequently, a comparable decrease in specific fuel consumption. The ejector-enhanced pulsejet combustor may also offer potential for reducing the emission of harmful exhaust compounds by making it practical to employ a low-loss rich-burn/quench/lean-burn sequence. Like all prior concepts for pressure-gain combustion, the present concept involves an approximation of constant-volume combustion, which is inherently unsteady (in this case, more specifically, cyclic). The consequent unsteadiness in combustor exit flow is generally regarded as detrimental to the performance of downstream turbomachinery. Among other adverse effects, this unsteadiness tends to detract from the thermodynamic benefits of pressure gain. Therefore, it is desirable in any intermittent combustion process to minimize unsteadiness in the exhaust path.

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.

Ceramic combustor mounting

DOEpatents

Hoffman, Melvin G.; Janneck, Frank W.

1982-01-01

A combustor for a gas turbine engine includes a metal engine block including a wall portion defining a housing for a combustor having ceramic liner components. A ceramic outlet duct is supported by a compliant seal on the metal block and a reaction chamber liner is stacked thereon and partly closed at one end by a ceramic bypass swirl plate which is spring loaded by a plurality of circumferentially spaced, spring loaded guide rods and wherein each of the guide rods has one end thereof directed exteriorly of a metal cover plate on the engine block to react against externally located biasing springs cooled by ambient air and wherein the rod spring support arrangement maintains the stacked ceramic components together so that a normal force is maintained on the seal between the outlet duct and the engine block under all operating conditions. The support arrangement also is operative to accommodate a substantial difference in thermal expansion between the ceramic liner components of the combustor and the metal material of the engine block.

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.

Critical Propulsion Components. Volume 2; Combustor

NASA Technical Reports Server (NTRS)

2005-01-01

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.

Computational Study of Combustor-Turbine Interactions

NASA Technical Reports Server (NTRS)

Miki, Kenji; Liou, Meng-Sing

2017-01-01

The Open National Combustion Code (OpenNCC) is applied to the simulation of a realisticcombustor configuration [Energy Efficient Engine (E(exp. 3))] in order to investigate the unsteady flow fields inside the combustor and around the first stage stator of a high pressure turbine (HPT). We consider one-twelfth (24 degrees) of the full annular E(exp. 3) combustor with three different geometries of the combustor exit: one without the vane, and two others with the vane set at different relative positions in relation to the fuel nozzle (clocking). Although it is common to take the exit flow profiles obtained by separately simulating the combustor and then feed it as the inflow profile when modeling the HPT, our studies show that the unsteady flow fields are influenced by the presence of the vane as well as clocking. More importantly, the characteristics (e.g., distribution and strength) of the high temperature spots (i.e., hot-streaks) appearing on the vane significantly alters. This indicates the importance of simultaneously modeling both the combustor and the HPT to understand the mechanics of the unsteady formulation of hot-streaks.

Combustor oscillating pressure stabilization and method

DOEpatents

Gemmen, R.S.; Richards, G.A.; Yip, M.T.J.; Robey, E.H.; Cully, S.R.; Addis, R.E.

1998-08-11

High dynamic pressure oscillations in hydrocarbon-fueled combustors typically occur when the transport time of the fuel to the flame front is at some fraction of the acoustic period. These oscillations are reduced to acceptably lower levels by restructuring or repositioning the flame front in the combustor to increase the transport time. A pilot flame front located upstream of the oscillating flame and pulsed at a selected frequency and duration effectively restructures and repositions the oscillating flame in the combustor to alter the oscillation-causing transport time. 7 figs.

Combustor liner durability analysis

NASA Technical Reports Server (NTRS)

Moreno, V.

1981-01-01

An 18 month combustor liner durability analysis program was conducted to evaluate the use of advanced three dimensional transient heat transfer and nonlinear stress-strain analyses for modeling the cyclic thermomechanical response of a simulated combustor liner specimen. Cyclic life prediction technology for creep/fatigue interaction is evaluated for a variety of state-of-the-art tools for crack initiation and propagation. The sensitivity of the initiation models to a change in the operating conditions is also assessed.

Systems Characterization of Combustor Instabilities With Controls Design Emphasis

NASA Technical Reports Server (NTRS)

Kopasakis, George

2004-01-01

This effort performed test data analysis in order to characterize the general behavior of combustor instabilities with emphasis on controls design. The analysis is performed on data obtained from two configurations of a laboratory combustor rig and from a developmental aero-engine combustor. The study has characterized several dynamic behaviors associated with combustor instabilities. These are: frequency and phase randomness, amplitude modulations, net random phase walks, random noise, exponential growth and intra-harmonic couplings. Finally, the very cause of combustor instabilities was explored and it could be attributed to a more general source-load type impedance interaction that includes the thermo-acoustic coupling. Performing these characterizations on different combustors allows for more accurate identification of the cause of these phenomena and their effect on instability.

Computation of losses in a scramjet combustor

NASA Technical Reports Server (NTRS)

Kamath, Pradeep S.; Mcclinton, Charles R.

1992-01-01

The losses in a conceptual scramjet combustor at flight Mach numbers of 8, 10, 12, 16 and 20 are computed. These losses are extracted from three-dimensional parabolized Navier-Stokes solutions of the turbulent, reacting combustor flow field. A combustor performance index was defined based on the rationale that an efficient scramjet combustor should add heat to the fluid in such a manner as to maximize the stream thrust at the combustor exit while minimizing the losses. This index showed a decrease of more than 40 percent as the flight Mach number increased from 8 to 20, indicative of a drop in the thrust-producing potential of the scramjet at the upper end of the speed regime studied. A breakdown of the losses showed that dissipation, nonequilibrium chemistry and heat diffusion contributed roughly 15 percent, 35 percent, and 50 percent to the irreversible increase in entropy at Mach 8 and 22 percent, 13 and 65 percent at Mach 20.

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.

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.

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.

Dual-Mode Free-Jet Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Dippold, Vance F., III; Yungster, Shaye

2017-01-01

The dual-mode free-jet combustor concept, pictured in figure 1, is described. It was introduced in 2010 as a wide- operating-range propulsion device using a novel supersonic free-jet combustion process. The unique feature of the free-jet combustor pictured in figure 1a, is supersonic combustion in an unconfined free-jet that traverses a larger subsonic combustion chamber to a variable nozzle. During this mode of operation, the propulsive stream is not in contact with the combustor walls, and equilibrates to the combustion chamber pressure. To a first order, thermodynamic efficiency is similar to that of a traditional scramjet under the assumption of constant-pressure combustion. Qualitatively, a number of possible benefits to this approach are obvious.

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.

Lagrangian coherent structures during combustion instability in a premixed-flame backward-step combustor.

PubMed

Sampath, Ramgopal; Mathur, Manikandan; Chakravarthy, Satyanarayanan R

2016-12-01

This paper quantitatively examines the occurrence of large-scale coherent structures in the flow field during combustion instability in comparison with the flow-combustion-acoustic system when it is stable. For this purpose, the features in the recirculation zone of the confined flow past a backward-facing step are studied in terms of Lagrangian coherent structures. The experiments are conducted at a Reynolds number of 18600 and an equivalence ratio of 0.9 of the premixed fuel-air mixture for two combustor lengths, the long duct corresponding to instability and the short one to the stable case. Simultaneous measurements of the velocity field using time-resolved particle image velocimetry and the CH^{*} chemiluminescence of the flame along with pressure time traces are obtained. The extracted ridges of the finite-time Lyapunov exponent (FTLE) fields delineate dynamically distinct regions of the flow field. The presence of large-scale vortical structures and their modulation over different time instants are well captured by the FTLE ridges for the long combustor where high-amplitude acoustic oscillations are self-excited. In contrast, small-scale vortices signifying Kelvin-Helmholtz instability are observed in the short duct case. Saddle-type flow features are found to separate the distinct flow structures for both combustor lengths. The FTLE ridges are found to align with the flame boundaries in the upstream regions, whereas farther downstream, the alignment is weaker due to dilatation of the flow by the flame's heat release. Specifically, the FTLE ridges encompass the flame curl-up for both the combustor lengths, and thus act as the surrogate flame boundaries. The flame is found to propagate upstream from an earlier vortex roll-up to a newer one along the backward-time FTLE ridge connecting the two structures.

Combustor with non-circular head end

DOEpatents

Kim, Won -Wook; McMahan, Kevin Weston

2015-09-29

The present application provides a combustor for use with a gas turbine engine. The combustor may include a head end with a non-circular configuration, a number of fuel nozzles positioned about the head end, and a transition piece extending downstream of the head end.

Gas turbine topping combustor

DOEpatents

Beer, J.; Dowdy, T.E.; Bachovchin, D.M.

1997-06-10

A combustor is described for burning a mixture of fuel and air in a rich combustion zone, in which the fuel bound nitrogen in converted to molecular nitrogen. The fuel rich combustion is followed by lean combustion. The products of combustion from the lean combustion are rapidly quenched so as to convert the fuel bound nitrogen to molecular nitrogen without forming NOx. The combustor has an air radial swirler that directs the air radially inward while swirling it in the circumferential direction and a radial fuel swirler that directs the fuel radially outward while swirling it in the same circumferential direction, thereby promoting vigorous mixing of the fuel and air. The air inlet has a variable flow area that is responsive to variations in the heating value of the fuel, which may be a coal-derived fuel gas. A diverging passage in the combustor in front of a bluff body causes the fuel/air mixture to recirculate with the rich combustion zone. 14 figs.

Gas turbine topping combustor

DOEpatents

Beer, Janos; Dowdy, Thomas E.; Bachovchin, Dennis M.

1997-01-01

A combustor for burning a mixture of fuel and air in a rich combustion zone, in which the fuel bound nitrogen in converted to molecular nitrogen. The fuel rich combustion is followed by lean combustion. The products of combustion from the lean combustion are rapidly quenched so as to convert the fuel bound nitrogen to molecular nitrogen without forming NOx. The combustor has an air radial swirler that directs the air radially inward while swirling it in the circumferential direction and a radial fuel swirler that directs the fuel radially outward while swirling it in the same circumferential direction, thereby promoting vigorous mixing of the fuel and air. The air inlet has a variable flow area that is responsive to variations in the heating value of the fuel, which may be a coal-derived fuel gas. A diverging passage in the combustor in front of a bluff body causes the fuel/air mixture to recirculate with the rich combustion zone.

High-temperature durability considerations for HSCT combustor

NASA Technical Reports Server (NTRS)

Jacobson, Nathan S.

1992-01-01

The novel combustor designs for the High Speed Civil Transport will require high temperature materials with long term environmental stability. Higher liner temperatures than in conventional combustors and the need for reduced weight necessitates the use of advanced ceramic matrix composites. The combustor environment is defined at the current state of design, the major degradation routes are discussed for each candidate ceramic material, and where possible, the maximum use temperatures are defined for these candidate ceramics.

Experimental clean combustor program, phase 2

NASA Technical Reports Server (NTRS)

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

1976-01-01

The alternate fuels investigation objective was to experimentally determine the impacts, if any, on exhaust emissions, performance, and durability characteristics of the hybrid and vorbix low pollution combustor concepts when operated on test fuels which simulate composition and property changes which might result from future broadened aviation turbine fuel specifications or use of synthetically derived crude feedstocks. Results of the program indicate a significant increase in CO and small NOX increase in emissions at idle for both combustor concepts, and an increase in THC for the vorbix concept. Minimal impact was observed on gaseous emissions at high power. The vorbix concept exhibited significant increase in exhaust smoke with increasing fuel aromatic content. Altitude stability was not affected for the vorbix combustor, but was substantially reduced for the hybrid concept. Severe carbon deposition was observed in both combustors following limited endurance testing with No. 2 home heat fuel. Liner temperature levels were insensitive to variations in aromatic content over the range of conditions investigated.

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.

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

Combustor technology for future small gas turbine aircraft

NASA Technical Reports Server (NTRS)

Lyons, Valerie J.; Niedzwiecki, Richard W.

1993-01-01

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

Energy efficient engine combustor test hardware detailed design report

NASA Technical Reports Server (NTRS)

Zeisser, M. H.; Greene, W.; Dubiel, D. J.

1982-01-01

The combustor for the Energy Efficient Engine is an annular, two-zone component. As designed, it either meets or exceeds all program goals for performance, safety, durability, and emissions, with the exception of oxides of nitrogen. When compared to the configuration investigated under the NASA-sponsored Experimental Clean Combustor Program, which was used as a basis for design, the Energy Efficient Engine combustor component has several technology advancements. The prediffuser section is designed with short, strutless, curved-walls to provide a uniform inlet airflow profile. Emissions control is achieved by a two-zone combustor that utilizes two types of fuel injectors to improve fuel atomization for more complete combustion. The combustor liners are a segmented configuration to meet the durability requirements at the high combustor operating pressures and temperatures. Liner cooling is accomplished with a counter-parallel FINWALL technique, which provides more effective heat transfer with less coolant.

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.

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.

Experimental investigation on heat transfer and frictional characteristics of vertical upward rifled tube in supercritical CFB boiler

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

Yang, Dong; Pan, Jie; Zhu, Xiaojing

2011-02-15

Water wall design is a key issue for supercritical Circulating Fluidized Bed (CFB) boiler. On account of the good heat transfer performance, rifled tube is applied in the water wall design of a 600 MW supercritical CFB boiler in China. In order to investigate the heat transfer and frictional characteristics of the rifled tube with vertical upward flow, an in-depth experiment was conducted in the range of pressure from 12 to 30 MPa, mass flux from 230 to 1200 kg/(m{sup 2} s), and inner wall heat flux from 130 to 720 kW/m{sup 2}. The wall temperature distribution and pressure dropmore » in the rifled tube were obtained in the experiment. The normal, enhanced and deteriorated heat transfer characteristics were also captured. In this paper, the effects of pressure, inner wall heat flux and mass flux on heat transfer characteristics are analyzed, the heat transfer mechanism and the frictional resistance performance are discussed, and the corresponding empirical correlations are presented. The experimental results show that the rifled tube can effectively prevent the occurrence of departure from nucleate boiling (DNB) and keep the tube wall temperature in a permissible range under the operating condition of supercritical CFB boiler. (author)« less

Multifuel evaluation of rich/quench/lean combustor

NASA Technical Reports Server (NTRS)

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

1982-01-01

The fuel flexible combustor technology was developed for application to the Model 570-K industrial gas turbine engine. The technology, to achieve emission goals, emphasizes dry NOx reduction methods. Due to the high levels of fuel-bound nitrogen (FBN), control of NOx can be effected through a staged combustor with a rich initial combustion zone. A rich/quench/lean variable geometry combustor utilizes the technology presented to achieve low NOx from alternate fuels containing FBN. The results focus on emissions and durability for multifuel operation.

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.

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.

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.

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

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

Process for Operating a Dual-Mode Combustor

NASA Technical Reports Server (NTRS)

Trefny, Charles J. (Inventor); Dippold, Vance F. (Inventor)

2017-01-01

A new dual-mode ramjet combustor used for operation over a wide flight Mach number range is described. Subsonic combustion mode is usable to lower flight Mach numbers than current dual-mode scramjets. High speed mode is characterized by supersonic combustion in a free-jet that traverses the subsonic combustion chamber to a variable nozzle throat. Although a variable combustor exit aperture is required, the need for fuel staging to accommodate the combustion process is eliminated. Local heating from shock-boundary-layer interactions on combustor walls is also eliminated.

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.

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.

Low NOx Fuel Flexible Combustor Integration Project Overview

NASA Technical Reports Server (NTRS)

Walton, Joanne C.; Chang, Clarence T.; Lee, Chi-Ming; Kramer, Stephen

2015-01-01

The Integrated Technology Demonstration (ITD) 40A Low NOx Fuel Flexible Combustor Integration development is being conducted as part of the NASA Environmentally Responsible Aviation (ERA) Project. Phase 2 of this effort began in 2012 and will end in 2015. This document describes the ERA goals, how the fuel flexible combustor integration development fulfills the ERA combustor goals, and outlines the work to be conducted during project execution.

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.

Combustor assembly in a gas turbine engine

DOEpatents

Wiebe, David J; Fox, Timothy A

2013-02-19

A combustor assembly in a gas turbine engine. The combustor assembly includes a combustor device coupled to a main engine casing, a first fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner disposed radially inwardly from the flow sleeve. The first fuel injection system provides fuel that is ignited with the pressurized air creating first working gases. The intermediate duct is disposed between the liner and the transition duct and defines a path for the first working gases to flow from the liner to the transition duct. An intermediate duct inlet portion is associated with a liner outlet and allows movement between the intermediate duct and the liner. An intermediate duct outlet portion is associated with a transition duct inlet section and allows movement between the intermediate duct and the transition duct.

NASA Project Develops Next-Generation Low-Emissions Combustor Technologies

NASA Technical Reports Server (NTRS)

Lee, Chi-Ming; Chang, Clarence T.; Herbon, John T.; Kramer, Stephen K.

2013-01-01

NASA's Environmentally Responsible Aviation (ERA) Project is working with industry to develop the fuel flexible combustor technologies for a new generation of low-emissions engine targeted for the 2020 timeframe. These new combustors will reduce nitrogen oxide (NOx) emissions to half of current state-of-the-art (SOA) combustors, while simultaneously reducing noise and fuel burn. The purpose of the low NOx fuel-flexible combustor research is to advance the Technology Readiness Level (TRL) and Integration Readiness Level (IRL) of a low NOx, fuel flexible combustor to the point where it can be integrated in the next generation of aircraft. To reduce project risk and optimize research benefit NASA chose to found two Phase 1 contracts. The first Phase 1 contracts went to engine manufactures and were awarded to: General Electric Company, and Pratt & Whitney Company. The second Phase 1 contracts went to fuel injector manufactures Goodrich Corporation, Parker Hannifin Corporation, and Woodward Fuel System Technology. In 2012, two sector combustors were tested at NASA's ASCR. The results indicated 75% NOx emission reduction below the 2004 CAEP/6 regulation level.

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

Advanced liner-cooling techniques for gas turbine combustors

NASA Technical Reports Server (NTRS)

Norgren, C. T.; Riddlebaugh, S. M.

1985-01-01

Component research for advanced small gas turbine engines is currently underway at the NASA Lewis Research Center. As part of this program, a basic reverse-flow combustor geometry was being maintained while different advanced liner wall cooling techniques were investigated. Performance and liner cooling effectiveness of the experimental combustor configuration featuring counter-flow film-cooled panels is presented and compared with two previously reported combustors featuring: splash film-cooled liner walls; and transpiration cooled liner walls (Lamilloy).

Combustor air flow control method for fuel cell apparatus

DOEpatents

Clingerman, Bruce J.; Mowery, Kenneth D.; Ripley, Eugene V.

2001-01-01

A method for controlling the heat output of a combustor in a fuel cell apparatus to a fuel processor where the combustor has dual air inlet streams including atmospheric air and fuel cell cathode effluent containing oxygen depleted air. In all operating modes, an enthalpy balance is provided by regulating the quantity of the air flow stream to the combustor to support fuel cell processor heat requirements. A control provides a quick fast forward change in an air valve orifice cross section in response to a calculated predetermined air flow, the molar constituents of the air stream to the combustor, the pressure drop across the air valve, and a look up table of the orifice cross sectional area and valve steps. A feedback loop fine tunes any error between the measured air flow to the combustor and the predetermined air flow.

Numerical analysis of flow instability in the water wall of a supercritical CFB boiler with annular furnace

NASA Astrophysics Data System (ADS)

Xie, Beibei; Yang, Dong; Xie, Haiyan; Nie, Xin; Liu, Wanyu

2016-08-01

In order to expand the study on flow instability of supercritical circulating fluidized bed (CFB) boiler, a new numerical computational model considering the heat storage of the tube wall metal was presented in this paper. The lumped parameter method was proposed for wall temperature calculation and the single channel model was adopted for the analysis of flow instability. Based on the time-domain method, a new numerical computational program suitable for the analysis of flow instability in the water wall of supercritical CFB boiler with annular furnace was established. To verify the code, calculation results were respectively compared with data of commercial software. According to the comparisons, the new code was proved to be reasonable and accurate for practical engineering application in analysis of flow instability. Based on the new program, the flow instability of supercritical CFB boiler with annular furnace was simulated by time-domain method. When 1.2 times heat load disturbance was applied on the loop, results showed that the inlet flow rate, outlet flow rate and wall temperature fluctuated with time eventually remained at constant values, suggesting that the hydrodynamic flow was stable. The results also showed that in the case of considering the heat storage, the flow in the water wall is easier to return to stable state than without considering heat storage.

Low NO.sub.x combustor

DOEpatents

Taylor, Jack R.

1987-01-01

A combustor having an annular first stage, a generally cylindrically-shaped second stage, and an annular conduit communicably connecting the first and second stages. The conduit has a relatively small annular height and a large number of quench holes in the walls thereof such that quench air injected into the conduit through the quench holes will mix rapidly with, or quench, the combustion gases flowing through the conduit. The rapid quenching reduces the amount of NO.sub.x produced in the combustor.

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.

Pulse combustor with controllable oscillations

DOEpatents

Richards, George A.; Welter, Michael J.; Morris, Gary J.

1992-01-01

A pulse combustor having thermally induced pulse combustion in a continuously flowing system is described. The pulse combustor is fitted with at lease one elongated ceramic body which significantly increases the heat transfer area in the combustion chamber of the combustor. The ceramic body or bodies possess sufficient mass and heat capacity to ignite the fuel-air charge once the ceramic body or bodies are heated by conventional spark plug initiated combustion so as to provide repetitive ignition and combustion of sequentially introduced fuel-air charges without the assistance of the spark plug and the rapid quenching of the flame after each ignition in a controlled manner so as to provide a selective control over the oscillation frequency and amplitude. Additional control over the heat transfer in the combustion chamber is provided by employing heat exchange mechanisms for selectively heating or cooling the elongated ceramic body or bodies and/or the walls of the combustion chamber.

Combustor with multistage internal vortices

DOEpatents

Shang, Jer Y.; Harrington, Richard E.

1989-01-01

A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard area to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard.

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.

Test results of low NO(x) catalytic combustors for gas turbines

NASA Astrophysics Data System (ADS)

Ozawa, Y.; Hirano, J.; Sato, M.; Saiga, M.; Watanabe, S.

1994-07-01

Catalytic combustion is an ultralow NO(x) combustion method, so it is expected that this method will be applied to a gas turbine combustor. However, it is difficult to develop a catalytic combustor because catalytic reliability at high temperature is still insufficient. To overcome this difficulty, we designed a catalytic combust gas at a combustion temperature of 1300 C while keeping the catalytic temperature below 1000 C. After performing preliminary tests using LPG, we designed two types of combustor for natural gas with a capacity equivalent to one combustor used in a 20 MW class multican-type gas turbine. Combustion tests were conducted at atmospheric pressure using natural gas. As a result, it was confirmed that a combustor in which catalytic combustor segments were arranged alternately with premixing nozzles could achieve low NO(x) and high combustion efficiency in the range from 1000 C to 1300 C of the combustor exit gas temperature.

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.

Soot loading in a generic gas turbine combustor

NASA Technical Reports Server (NTRS)

Eckerle, W. A.; Rosfjord, T. J.

1987-01-01

Variation in soot loading along the centerline of a generic gas turbine combustor was experimentally investigated. The 12.7-cm dia burner consisted of six sheet-metal louvers. Soot loading along the burner length was quantified by acquiring measurements first at the exit of the full-length combustor and then at upstream stations by sequential removal of liner louvers to shorten the burner length. Alteration of the flow field approaching removed louvers, maintaining a constant liner pressure drop. Burner exhaust flow was sampled at the burner centerline to determine soot mass concentration and smoke number. Characteristic particle size and number density, transmissivity of the exhaust flow, and local radiation from luminous soot particles in the exhaust flow were determined by optical techniques. Four test fuels were burned at three fuel-air ratios to determine fuel chemical property and flow temperature influences. Data were acquired at two combustor pressures. Particulate concentration data indicated a strong oxidation mechanism in the combustor secondary zone, though the oxidation was significantly affected by flow temperature. Soot production was directly related to fuel smoke point. Less soot production and lower secondary-zone oxidation rates were observed at reduced combustor pressure.

Advanced Low NO Sub X Combustors for Supersonic High-Altitude Aircraft Gas Turbines

NASA Technical Reports Server (NTRS)

Roberts, P. B.; White, D. J.; Shekleton, J. R.

1975-01-01

A test rig program was conducted with the objective of evaluating and minimizing the exhaust emissions, in particular NO sub x, of three advanced aircraft combustor concepts at a simulated, high altitude cruise condition. The three combustor designs, all members of the lean reaction, premixed family, are the Jet Induced Circulation (JIC) combustor, the Vortex Air Blast (VAB) combustor, and a catalytic combustor. They were rig tested in the form of reverse flow can combustors in the 0.127 m. (5.0 in.) size range. Various configuration modifications were applied to each of the initial JIC and VAB combustor model designs in an effort to reduce the emissions levels. The VAB combustor demonstrated a NO sub x level of 1.1 gm NO2/kg fuel with essentially 100% combustion efficiency at the simulated cruise combustor condition of 50.7 N/sq cm (5 atm), 833 K (1500 R) inlet pressure and temperature respectively and 1778 K (3200 R) outlet temperature on Jet-A1 fuel. Early tests on the catalytic combustor were unsuccessful due to a catalyst deposition problem and were discontinued in favor of the JIC and VAB tests. In addition emissions data were obtained on the JIC and VAB combustors at low combustor inlet pressure and temperatures that indicate the potential performance at engine off-design conditions.

Combustor with fuel preparation chambers

NASA Technical Reports Server (NTRS)

Zelina, Joseph (Inventor); Myers, Geoffrey D. (Inventor); Srinivasan, Ram (Inventor); Reynolds, Robert S. (Inventor)

2001-01-01

An annular combustor having fuel preparation chambers mounted in the dome of the combustor. The fuel preparation chamber comprises an annular wall extending axially from an inlet to an exit that defines a mixing chamber. Mounted to the inlet are an air swirler and a fuel atomizer. The air swirler provides swirled air to the mixing chamber while the atomizer provides a fuel spray. On the downstream side of the exit, the fuel preparation chamber has an inwardly extending conical wall that compresses the swirling mixture of fuel and air exiting the mixing chamber.

Methanol tailgas combustor control method

DOEpatents

Hart-Predmore, David J.; Pettit, William H.

2002-01-01

A method for controlling the power and temperature and fuel source of a combustor in a fuel cell apparatus to supply heat to a fuel processor where the combustor has dual fuel inlet streams including a first fuel stream, and a second fuel stream of anode effluent from the fuel cell and reformate from the fuel processor. In all operating modes, an enthalpy balance is determined by regulating the amount of the first and/or second fuel streams and the quantity of the first air flow stream to support fuel processor power requirements.

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...

Isolator-combustor interaction in a dual-mode scramjet engine

NASA Technical Reports Server (NTRS)

Pratt, David T.; Heiser, William H.

1993-01-01

A constant-area diffuser, or 'isolator', is required in both the ramjet and scramjet operating regimes of a dual-mode engine configuration in order to prevent unstarts due to pressure feedback from the combustor. Because the nature of the combustor-isolator interaction is different in the two operational modes, however, attention is presently given to the use of thermal vs kinetic energy coordinates for these interaction processes' visualization. The results of the analysis thus conducted indicate that the isolator requires severe flow separation at combustor entry, and that its entropy-generating characteristics are more severe than an equivalent oblique shock. A constant-area diffuser is only marginally able to contain the equivalent normal shock required for subsonic combustor entry.

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.

Low NOx, Lean Direct Wall Injection Combustor Concept Developed

NASA Technical Reports Server (NTRS)

Tacina, Robert R.; Wey, Changlie; Choi, Kyung J.

2003-01-01

The low-emissions combustor development at the NASA Glenn Research Center is directed toward advanced high-pressure aircraft gas turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low-power conditions. Low-NOx combustors can be classified into rich burn and lean burn concepts. Lean burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) combustors. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibly of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone and, thus, does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, the key is good atomization and mixing of the fuel quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP.

Low NO/x/ heavy fuel combustor program

NASA Technical Reports Server (NTRS)

Lister, E.; Niedzwiecki, R. W.; Nichols, L.

1980-01-01

The paper deals with the 'Low NO/x/ Heavy Fuel Combustor Program'. Main program objectives are to generate and demonstrate the technology required to develop durable gas turbine combustors for utility and industrial applications, which are capable of sustained, environmentally acceptable operation with minimally processed petroleum residual fuels. The program will focus on 'dry' reductions of oxides of nitrogen (NO/x/), improved combustor durability and satisfactory combustion of minimally processed petroleum residual fuels. Other technology advancements sought include: fuel flexibility for operation with petroleum distillates, blends of petroleum distillates and residual fuels, and synfuels (fuel oils derived from coal or shale); acceptable exhaust emissions of carbon monoxide, unburned hydrocarbons, sulfur oxides and smoke; and retrofit capability to existing engines.

Low NO(x) heavy fuel combustor program

NASA Technical Reports Server (NTRS)

Lister, E.; Niedzwiecki, R. W.; Nichols, L.

1979-01-01

The 'low nitrogen oxides heavy fuel combustor' program is described. Main program objectives are to generate and demonstrate the technology required to develop durable gas turbine combustors for utility and industrial applications, which are capable of sustained, environmentally acceptable operation with minimally processed petroleum residual fuels. The program will focus on 'dry' reductions of oxides of nitrogen, improved combustor durability, and satisfactory combustion of minimally processed petroleum residual fuels. Other technology advancements sought include: fuel flexibility for operation with petroleum distillates, blends of petroleum distillates and residual fuels, and synfuels (fuel oils derived from coal or shale); acceptable exhaust emissions of carbon monoxide, unburned hydrocarbons, sulfur oxides and smoke; and retrofit capability to existing engines.

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.

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.

Combustor with multistage internal vortices

DOEpatents

Shang, Jer Yu; Harrington, R.E.

1987-05-01

A fluidized bed combustor is provided with a multistage arrangement of vortex generators in the freeboard area. The vortex generators are provided by nozzle means which extend into the interior of the freeboard for forming vortices within the freeboard areas to enhance the combustion of particulate material entrained in product gases ascending into the freeboard from the fluidized bed. Each of the nozzles are radially inwardly spaced from the combustor walls defining the freeboard to provide for the formation of an essentially vortex-free, vertically extending annulus about the vortices whereby the particulate material centrifuged from the vortices against the inner walls of the combustor is returned through the annulus to the fluidized bed. By adjusting the vortex pattern within the freeboard, a significant portion of the full cross-sectional area of the freeboard except for the peripheral annulus can be contacted with the turbulent vortical flow for removing the particulate material from the gaseous products and also for enhancing the combustion thereof within the freeboard. 2 figs.

Variable volume combustor with a conical liner support

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

Johnson, Thomas Edward; McConnaughhay, Johnie Franklin; Keener, Chrisophter Paul

The present application provides a variable volume combustor for use with a gas turbine engine. The variable volume combustor may include a liner, a number of micro-mixer fuel nozzles positioned within the liner, and a conical liner support supporting the liner.

NOx results from two combustors tested on medium BTU coal gas

NASA Technical Reports Server (NTRS)

Sherlock, T. P.; Carl, D. E.; Vermes, G.; Schwab, J.; Notardonato, J. J.

1982-01-01

The results of tests of two combustor configurations using coal gas from a 25 ton/day fluidized bed coal gasifier are reported. The trials were run with a ceramic-lined, staged rich/lean burner and an integral, all metal multiannular swirl burner (MASB) using a range of temperatures and pressures representative of industrial turbine inlet conditions. A lean mixture was examined at 104, 197, and 254 Btu/Scf, yielding NO(x) emissions of 5, 20, and 70 ppmv, respectively. The MASB was employed only with a gas rated at 220-270 Btu/Scf, producing 80 ppmv NO(x) at rated engine conditions. The results are concluded to be transferrable to current machines. Further tests on the effects of gas composition, the scaling of combustors to utility size, and the development of improved wall cooling techniques and variable geometry are indicated.

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.

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%.

Thermally-Choked Combustor Technology

NASA Technical Reports Server (NTRS)

Knuth, William H.; Gloyer, P.; Goodman, J.; Litchford, R. J.

1993-01-01

A program is underway to demonstrate the practical feasibility of thermally-choked combustor technology with particular emphasis on rocket propulsion applications. Rather than induce subsonic to supersonic flow transition in a geometric throat, the goal is to create a thermal throat by adding combustion heat in a diverging nozzle. Such a device would have certain advantages over conventional flow accelerators assuming that the pressure loss due to heat addition does not severely curtail propulsive efficiency. As an aid to evaluation, a generalized one-dimensional compressible flow analysis tool was constructed. Simplified calculations indicate that the process is fluid dynamically and thermodynamically feasible. Experimental work is also being carried out in an attempt to develop, assuming an array of practical issues are surmountable, a practical bench-scale demonstrator using high flame speed H2/O2 combustibles.

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

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.

Hydrogen Fuel Capability Added to Combustor Flametube Rig

NASA Technical Reports Server (NTRS)

Frankenfield, Bruce J.

2003-01-01

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

Experimental evaluation of combustor concepts for burning broad property fuels

NASA Technical Reports Server (NTRS)

Kasper, J. M.; Ekstedt, E. E.; Dodds, W. J.; Shayeson, M. W.

1980-01-01

A baseline CF6-50 combustor and three advanced combustor designs were evaluated to determine the effects of combustor design on operational characteristics using broad property fuels. Three fuels were used in each test: Jet A, a broad property 13% hydrogen fuel, and a 12% hydrogen fuel blend. Testing was performed in a sector rig at true cruise and simulated takeoff conditions for the CF6-50 engine cycle. The advanced combustors (all double annular, lean dome designs) generally exhibited lower metal temperatures, exhaust emissions, and carbon buildup than the baseline CF6-50 combustor. The sensitivities of emissions and metal temperatures to fuel hydrogen content were also generally lower for the advanced designs. The most promising advanced design used premixing tubes in the main stage. This design was chosen for additional testing in which fuel/air ratio, reference velocity, and fuel flow split were varied.

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.

Flame structure and stabilization in miniature liquid film combustors

NASA Astrophysics Data System (ADS)

Pham, Trinh Kim

Liquid-fueled miniature combustion systems can be promising portable power devices when high specific power and long operation duration are required. A uniquely viable fueling option for small scale combustion is to introduce the liquid fuel as a film on the combustor walls. As one example of such systems, this dissertation characterizes 1-cm-diameter tubular combustors fed by liquid fuel films, and seeks to identify the mechanisms by which flames are stabilized within them. Early experimental work demonstrates that flame behavior is dependent upon steadiness in fuel and air injection and in geometric symmetry and uniformity. Significant discoveries in later work include the impact of direct strain on the flame by the airflow, the fact that no local recirculation zone appears to exist for stabilization as was previously believed, and that the film thickness, uniformity, and location directly affect the flame's characteristics and stability. A gradient in film thickness is required for stable operation, and this requirement may explain why the combustor maintains overall rich conditions. Initial numerical simulations of two-dimensional cold and reacting flows in a simplified model of the combustor yields flame shape and flow field results that do not match experiments in the burning case, therefore suggesting that local turbulence in the fuel injection region provides the necessary degree of mixing. A three-dimensional model of the combustor is needed if reacting flows are to be simulated accurately. It was also found that thermal conduction from the chamber exit to the chamber base plays an important role in fuel vaporization and the stability of the flame. Consequently, flames cannot be sustained in quartz and other transparent but thermally insulating materials for the selected geometry, so observation of the flame's entire structure cannot be accomplished without either the addition of other flameholding elements or the employment of a more thermally conductive

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.

Advanced Combustor in the Four Burner Area

NASA Image and Video Library

1966-03-21

Engineer Frank Kutina and a National Aeronautics and Space Administration (NASA) mechanic examine the setup of an advanced combustor rig inside one of the test cells at the Lewis Research Center’s Four Burner Area in the Engine Research Building. Kutina, of the Research Operations Branch, served as go-between for the researchers and the mechanics. He helped develop the test configurations and get the hardware installed. At the time of this photograph, Lewis Center Director Abe Silverstein had just established the Airbreathing Engine Division to address the new propulsion of the 1960s. After nearly a decade of focusing almost exclusively on space, NASA Lewis began tackling issues relating to the new turbofan engine, noise reduction, energy efficiency, supersonic transport, and the never-ending quest for higher performance levels with smaller and more lightweight engines. The Airbreathing Engine Division’s Combustion Branch was dedicated to the study and mitigation of the high temperatures and pressures found in advanced combustor designs. These high temperatures and pressures could destroy engine components. The Lewis investigation included film cooling, diffuser flow, and jet mixing. Components were tested in smaller test cells, but a full-scale augmenting burner rig, seen here, was tested extensively in the Four Burner Area test cell.

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.

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.

Rapid-quench axially staged combustor

DOEpatents

Feitelberg, Alan S.; Schmidt, Mark Christopher; Goebel, Steven George

1999-01-01

A combustor cooperating with a compressor in driving a gas turbine includes a cylindrical outer combustor casing. A combustion liner, having an upstream rich section, a quench section and a downstream lean section, is disposed within the outer combustor casing defining a combustion chamber having at least a core quench region and an outer quench region. A first plurality of quench holes are disposed within the liner at the quench section having a first diameter to provide cooling jet penetration to the core region of the quench section of the combustion chamber. A second plurality of quench holes are disposed within the liner at the quench section having a second diameter to provide cooling jet penetration to the outer region of the quench section of the combustion chamber. In an alternative embodiment, the combustion chamber quench section further includes at least one middle region and at least a third plurality of quench holes disposed within the liner at the quench section having a third diameter to provide cooling jet penetration to at least one middle region of the quench section of the combustion chamber.

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 60.52a - 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...

Performance of a small annular turbojet combustor designed for low cost

NASA Technical Reports Server (NTRS)

Fear, J. S.

1972-01-01

Performance investigations were conducted on a combustor utilizing several cost-reducing innovations and designed for use in a low-cost 4448-N thrust turbojet engine for commercial light aircraft. Low-cost features included simple, air-atomizing fuel injectors; combustor liners of perforated sheet; and the use of inexpensive type 304 stainless-steel material. Combustion efficiencies at the cruise and sea-level-takeoff design points were approximately 97 and 98 percent, respectively. The combustor isothermal pressure loss was 6.3 percent at the cruise-condition diffuser inlet Mach number of 0.34. The combustor exit temperature pattern factor was less than 0.24 at both the cruise and sea-level-takeoff design points. The combustor exit average radial temperature profiles at all conditions were in very good agreement with the design profile.

Exhaust gas emissions of a vortex breakdown stabilized combustor

NASA Technical Reports Server (NTRS)

Yetter, R. A.; Gouldin, F. C.

1976-01-01

Exhaust gas emission data are described for a swirl stabilized continuous combustor. The combustor consists of confined concentric jets with premixed fuel and air in the inner jet and air in the outer jet. Swirl may be induced in both inner and outer jets with the sense of rotation in the same or opposite directions (co-swirl and counter-swirl). The combustor limits NO emissions by lean operation without sacrificing CO and unburned hydrocarbon emission performance, when commercial-grade methane and air fired at one atmosphere without preheat are used. Relative swirl direction and magnitude are found to have significant effects on exhaust gas concentrations, exit temperatures, and combustor efficiencies. Counter-swirl gives a large recirculation zone, a short luminous combustion zone, and large slip velocities in the interjet shear layer. For maximum counter-swirl conditions, the efficiency is low.

Combustor and method for purging a combustor

DOEpatents

Berry, Jonathan Dwight; Hughes, Michael John

2015-06-09

A combustor includes an end cap. The end cap includes a first surface and a second surface downstream from the first surface, a shroud that circumferentially surrounds at least a portion of the first and second surfaces, a plate that extends radially within the shroud, a plurality of tubes that extend through the plate and the first and second surfaces, and a first purge port that extends through one or more of the plurality of tubes, wherein the purge port is axially aligned with the plate.

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

Idealized gas turbine combustor for performance research and validation of large eddy simulations.

PubMed

Williams, Timothy C; Schefer, Robert W; Oefelein, Joseph C; Shaddix, Christopher R

2007-03-01

This paper details the design of a premixed, swirl-stabilized combustor that was designed and built for the express purpose of obtaining validation-quality data for the development of large eddy simulations (LES) of gas turbine combustors. The combustor features nonambiguous boundary conditions, a geometrically simple design that retains the essential fluid dynamics and thermochemical processes that occur in actual gas turbine combustors, and unrestrictive access for laser and optical diagnostic measurements. After discussing the design detail, a preliminary investigation of the performance and operating envelope of the combustor is presented. With the combustor operating on premixed methane/air, both the equivalence ratio and the inlet velocity were systematically varied and the flame structure was recorded via digital photography. Interesting lean flame blowout and resonance characteristics were observed. In addition, the combustor exhibited a large region of stable, acoustically clean combustion that is suitable for preliminary validation of LES models.

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.

Micro-combustor for gas turbine engine

DOEpatents

Martin, Scott M.

2010-11-30

An improved gas turbine combustor (20) including a basket (26) and a multiplicity of micro openings (29) arrayed across an inlet wall (27) for passage of a fuel/air mixture for ignition within the combustor. The openings preferably have a diameter on the order of the quenching diameter; i.e. the port diameter for which the flame is self-extinguishing, which is a function of the fuel mixture, temperature and pressure. The basket may have a curved rectangular shape that approximates the shape of the curved rectangular shape of the intake manifolds of the turbine.

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.

Parametric study of flame radiation characteristics of a tubular-can combustor

NASA Technical Reports Server (NTRS)

Humenik, F. M.; Claus, R. W.; Neely, G. M.

1983-01-01

A series of combustor tests were conducted with a tubular-can combustor to study flame radiation characteristics and effects with parametric variations in combustor operating conditions. Two alternate combustor assemblies using a different fuel nozzle were compared. Spectral and total radiation detectors were positioned at three stations along the length of the combustor can. Data were obtained for a range of pressures from 0.34 to 2.07 MPa (50 to 300 psia), inlet temperatures from 533 to 700K (500 to 800 F), for Jet A (13.9 deg hydrogen) and ERBS (12.9% hydrogen) fuels, and with fuel-air ratios nominally from 0.008 to 0.021. Spectral radiation data, total radiant heat flux data, and liner temperature data are presented to illustrate the flame radiation characteristics and effects in the primary, secondary, and tertiary combustion zones.

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.

Combustor assembly in a gas turbine engine

DOEpatents

Wiebe, David J; Fox, Timothy A

2015-04-28

A combustor assembly in a gas turbine engine includes a combustor device, a fuel injection system, a transition duct, and an intermediate duct. The combustor device includes a flow sleeve for receiving pressurized air and a liner surrounded by the flow sleeve. The fuel injection system provides fuel to be mixed with the pressurized air and ignited in the liner to create combustion products. The intermediate duct is disposed between the liner and the transition duct so as to define a path for the combustion products to flow from the liner to the transition duct. The intermediate duct is associated with the liner such that movement may occur therebetween, and the intermediate duct is associated with the transition duct such that movement may occur therebetween. The flow sleeve includes structure that defines an axial stop for limiting axial movement of the intermediate duct.

Diesel engine catalytic combustor system. [aircraft engines

NASA Technical Reports Server (NTRS)

Ream, L. W. (Inventor)

1984-01-01

A low compression turbocharged diesel engine is provided in which the turbocharger can be operated independently of the engine to power auxiliary equipment. Fuel and air are burned in a catalytic combustor to drive the turbine wheel of turbine section which is initially caused to rotate by starter motor. By opening a flapper value, compressed air from the blower section is directed to catalytic combustor when it is heated and expanded, serving to drive the turbine wheel and also to heat the catalytic element. To start, engine valve is closed, combustion is terminated in catalytic combustor, and the valve is then opened to utilize air from the blower for the air driven motor. When the engine starts, the constituents in its exhaust gas react in the catalytic element and the heat generated provides additional energy for the turbine section.

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.

Variable volume combustor with nested fuel manifold system

DOEpatents

McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward; Ostebee, Heath Michael

2016-09-13

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles, a fuel manifold system in communication with the micro-mixer fuel nozzles to deliver a flow of fuel thereto, and a linear actuator to maneuver the micro-mixer fuel nozzles and the fuel manifold system.

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

NASA Technical Reports Server (NTRS)

VanFossen, G. James; Bunker, Ronald S.

2000-01-01

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.

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

Properties of Fuels Employed in a Gas Turbine Combustor Program.

DTIC Science & Technology

1983-09-01

potence nateonale PROPERTIES OF FUELS EMPLOYED IN A GAS TURBINE COMBUSTOR PROGRAM by .J.R. Coleman and L.D. Gallop JAN 1O t84’ La.I DEFENCE ROSOARCH...ESTABLISHMENT OTTAWA T~INCAMNTE M4 1-05 - ottwa , National Dibense3 Detence nationale PROPERTIES OF FUELS EMPLOYED IN A GAS TURBINE COMBUSTOR PROGRAM by...made of the physical and chemical properties of sixteen fuels employed in an aircraft gas turbine combustor programme. Several of these are specification

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 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...

High-temperature combustor liner tests in structural component response test facility

NASA Technical Reports Server (NTRS)

Moorhead, Paul E.

1988-01-01

Jet engine combustor liners were tested in the structural component response facility at NASA Lewis. In this facility combustor liners were thermally cycled to simulate a flight envelope of takeoff, cruise, and return to idle. Temperatures were measured with both thermocouples and an infrared thermal imaging system. A conventional stacked-ring louvered combustor liner developed a crack at 1603 cycles. This test was discontinued after 1728 cycles because of distortion of the liner. A segmented or float wall combustor liner tested at the same heat flux showed no significant change after 1600 cycles. Changes are being made in the facility to allow higher temperatures.

Combustor Simulation

NASA Technical Reports Server (NTRS)

Norris, Andrew

2003-01-01

The goal was to perform 3D simulation of GE90 combustor, as part of full turbofan engine simulation. Requirements of high fidelity as well as fast turn-around time require massively parallel code. National Combustion Code (NCC) was chosen for this task as supports up to 999 processors and includes state-of-the-art combustion models. Also required is ability to take inlet conditions from compressor code and give exit conditions to turbine code.

Catalytic combustor for integrated gasification combined cycle power plant

DOEpatents

Bachovchin, Dennis M [Mauldin, SC; Lippert, Thomas E [Murrysville, PA

2008-12-16

A gasification power plant 10 includes a compressor 32 producing a compressed air flow 36, an air separation unit 22 producing a nitrogen flow 44, a gasifier 14 producing a primary fuel flow 28 and a secondary fuel source 60 providing a secondary fuel flow 62 The plant also includes a catalytic combustor 12 combining the nitrogen flow and a combustor portion 38 of the compressed air flow to form a diluted air flow 39 and combining at least one of the primary fuel flow and secondary fuel flow and a mixer portion 78 of the diluted air flow to produce a combustible mixture 80. A catalytic element 64 of the combustor 12 separately receives the combustible mixture and a backside cooling portion 84 of the diluted air flow and allows the mixture and the heated flow to produce a hot combustion gas 46 provided to a turbine 48. When fueled with the secondary fuel flow, nitrogen is not combined with the combustor portion.

Combustion efficiency of a premixed continuous flow combustor

NASA Technical Reports Server (NTRS)

Anand, M. S.; Gouldin, F. C.

1985-01-01

Exhaust gas temperature, velocity, and composition measurements at various radial locations at the combustor exit are presented for a swirling-flow continuous combustor of a confined concentric jet configuration operating on premixed propane or methane and air. The main objective of the study is to determine the effect of fuel substitution and of changes in outer flow swirl conditions on the combustor performance. It is found that there is no difference in observed properties for propane and methane firing; the use of either of the fuels results in nearly the same exit temperature and velocity profiles and the same efficiency for a given operating condition. A mechanism for combustion is proposed which explains qualitatively the changes in efficiency and pollutant emissions observed with changing swirl.

Method for operating a combustor in a fuel cell system

DOEpatents

Chalfant, Robert W.; Clingerman, Bruce J.

2002-01-01

A method of operating a combustor to heat a fuel processor in a fuel cell system, in which the fuel processor generates a hydrogen-rich stream a portion of which is consumed in a fuel cell stack and a portion of which is discharged from the fuel cell stack and supplied to the combustor, and wherein first and second streams are supplied to the combustor, the first stream being a hydrocarbon fuel stream and the second stream consisting of said hydrogen-rich stream, the method comprising the steps of monitoring the temperature of the fuel processor; regulating the quantity of the first stream to the combustor according to the temperature of the fuel processor; and comparing said quantity of said first stream to a predetermined value or range of predetermined values.

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...

Increasing Operational Stability in Low NOX GT Combustor Using Fuel Rich Concentric Pilot Combustor

NASA Astrophysics Data System (ADS)

Levy, Yeshayahou; Erenburg, Vladimir; Sherbaum, Valery; Ovcharenko, Vitali; Rosentsvit, Leonid; Chudnovsky, Boris; Herszage, Amiel; Talanker, Alexander

2012-03-01

Lean combustion is a method in which combustion takes place under low equivalence ratio and relatively low combustion temperatures. As such, it has the potential to lower the effect of the relatively high activation energy nitrogen-oxygen reactions which are responsible for substantial NOX formation during combustion processes. However, lowering temperature reduces the reaction rate and deteriorates combustion stability. The objective of the present study is to reduce the lower equivalence ratio limit of the stable combustion operational boundary in lean Gas Turbine (GT) combustors while still maintaining combustion stability. A lean premixed gaseous combustor was equipped with a surrounding concentric pilot flame operating under rich conditions, thus generating a hot stream of combustion products with significant amount of reactive radicals. The main combustor's fuel-air composition was varied from stoichiometric to lean mixtures. The pilot's mixture composition was also varied by changing the air flow rate, within a limited rich mixtures range. The pilot fuel flow rate was always lower than five percent of the total fuel supply at the specific stage of the experiments.

Large eddy simulation of premixed and non-premixed combustion in a Stagnation Point Reverse Flow combustor

NASA Astrophysics Data System (ADS)

Undapalli, Satish

A new combustor referred to as Stagnation Point Reverse Flow (SPRF) combustor has been developed at Georgia Tech to meet the increasingly stringent emission regulations. The combustor incorporates a novel design to meet the conflicting requirements of low pollution and high stability in both premixed and non-premixed modes. The objective of this thesis work is to perform Large Eddy Simulations (LES) on this lab-scale combustor and elucidate the underlying physics that has resulted in its excellent performance. To achieve this, numerical simulations have been performed in both the premixed and non-premixed combustion modes, and velocity field, species field, entrainment characteristics, flame structure, emissions, and mixing characteristics have been analyzed. Simulations have been carried out first for a non-reactive case to resolve relevant fluid mechanics without heat release by the computational grid. The computed mean and RMS quantities in the non-reacting case compared well with the experimental data. Next, the simulations were extended for the premixed reactive case by employing different sub-grid scale combustion chemistry closures: Eddy Break Up (EBU), Artificially Thickened Flame (TF) and Linear Eddy Mixing (LEM) models. Results from the EBU and TF models exhibit reasonable agreement with the experimental velocity field. However, the computed thermal and species fields have noticeable discrepancies. Only LEM with LES (LEMLES), which is an advanced scalar approach, has been able to accurately predict both the velocity and species fields. Scalar mixing plays an important role in combustion, and this is solved directly at the sub-grid scales in LEM. As a result, LEM accurately predicts the scalar fields. Due to the two way coupling between the super-grid and sub-grid quantities, the velocity predictions also compare very well with the experiments. In other approaches, the sub-grid effects have been either modeled using conventional approaches (EBU) or need

Orbit transfer rocket engine technology program enhanced heat transfer combustor technology

NASA Technical Reports Server (NTRS)

Brown, William S.

1991-01-01

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

Variable volume combustor with pre-nozzle fuel injection system

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

Keener, Christopher Paul; Johnson, Thomas Edward; McConnaughhay, Johnie Franklin

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of fuel nozzles, a pre-nozzle fuel injection system supporting the fuel nozzles, and a linear actuator to maneuver the fuel nozzles and the pre-nozzle fuel injection system.

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.

Coaxial Dump Ramjet Combustor Combustion Instabilities. Part I. Parametric Test Data.

DTIC Science & Technology

1981-07-01

AD-AIII 355 COAXIAL DUP RA8.? COMBUSTOR COMBUSTION INSTABILITIES I/~ PART I PARAUER1C. 1111 AIR FORCE WRIONT AERONUTICAL LAOS WRIOIII-PATTERSOll...MICROCOPY RESOLUTION TEST CHART NATIONAL BUREAU OF STANOAROS - 193- A AFWAL-TR-81 -2047 Part 1 COAXIAL DUMP RAMJET COMBUSTOR COMBUSTION INSTABILITIES PART...COMBUSTOR Interim Report for Period COMBUSTION INSTABILITIES February 1979 - March 1980 Part I - Parametric Test Data S. PERFORMING ORG. REPORT NUMBER 7

Variable volume combustor with an air bypass system

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

Johnson, Thomas Edward; Ziminsky, Willy Steve; Ostebee, Heath Michael

The present application provides a combustor for use with flow of fuel and a flow of air in a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles positioned within a liner and an air bypass system position about the liner. The air bypass system variably allows a bypass portion of the flow of air to bypass the micro-mixer fuel nozzles.

Combustor nozzle for a fuel-flexible combustion system

DOEpatents

Haynes, Joel Meier [Niskayuna, NY; Mosbacher, David Matthew [Cohoes, NY; Janssen, Jonathan Sebastian [Troy, NY; Iyer, Venkatraman Ananthakrishnan [Mason, OH

2011-03-22

A combustor nozzle is provided. The combustor nozzle includes a first fuel system configured to introduce a syngas fuel into a combustion chamber to enable lean premixed combustion within the combustion chamber and a second fuel system configured to introduce the syngas fuel, or a hydrocarbon fuel, or diluents, or combinations thereof into the combustion chamber to enable diffusion combustion within the combustion chamber.

Investigation of soot and carbon formation in small gas turbine combustors

NASA Technical Reports Server (NTRS)

Rosfjord, T. J.

1982-01-01

An investigation of hardware configurations which attempt to minimize carbon and soot-production without sacrificing performance in small gas turbine combustors was conducted. Four fuel injectors, employing either airblast atomization, pressure atomization, or fuel vaporization techniques were combined with nozzle air swirlers and injector sheaths. Eight configurations were screened at sea-level takeoff and idle test conditions. Selected configurations were focused upon in an attempt to quantify the influence of combustor pressure, inlet temperature, primary zone operation, and combustor loading on soot and carbon formation. Cycle tests were also performed. It was found that smoke emission levels depended on the combustor fluid mechanics, the atomization quality of the injector and the fuel hydrogen content.

Mount assembly for porous transition panel at annular combustor outlet

NASA Technical Reports Server (NTRS)

Sweeney, Ralph B. (Inventor); Verdouw, Albert J. (Inventor)

1980-01-01

A gas turbine engine combustor assembly of annular configuration has outer and inner walls made up of a plurality of axially extending multi-layered porous metal panels joined together at butt joints therebetween and each outer and inner wall including a transition panel of porous metal defining a combustor assembly outlet supported by a combustor mount assembly including a stiffener ring having a side undercut thereon fit over a transition panel end face; and wherein an annular weld joins the ring to the end face to transmit exhaust heat from the end face to the stiffener ring for dissipation from the combustor; a combustor pilot member is located in axially spaced, surrounding relationship to the end face and connector means support the stiffener ring in free floating relationship with the pilot member to compensate for both radial and axial thermal expansion of the transition panel; and said connector means includes a radial gap for maintaining a controlled flow of coolant from outside of the transition panel into cooling relationship with the stiffener ring and said weld to further cool the end face against excessive heat build-up therein during flow of hot gas exhaust through said outlet.

Combustor Operability and Performance Verification for HIFiRE Flight 2

NASA Technical Reports Server (NTRS)

Storch, Andrea M.; Bynum, Michael; Liu, Jiwen; Gruber, Mark

2011-01-01

As part of the Hypersonic International Flight Research Experimentation (HIFiRE) Direct-Connect Rig (HDCR) test and analysis activity, three-dimensional computational fluid dynamics (CFD) simulations were performed using two Reynolds-Averaged Navier Stokes solvers. Measurements obtained from ground testing in the NASA Langley Arc-Heated Scramjet Test Facility (AHSTF) were used to specify inflow conditions for the simulations and combustor data from four representative tests were used as benchmarks. Test cases at simulated flight enthalpies of Mach 5.84, 6.5, 7.5, and 8.0 were analyzed. Modeling parameters (e.g., turbulent Schmidt number and compressibility treatment) were tuned such that the CFD results closely matched the experimental results. The tuned modeling parameters were used to establish a standard practice in HIFiRE combustor analysis. Combustor performance and operating mode were examined and were found to meet or exceed the objectives of the HIFiRE Flight 2 experiment. In addition, the calibrated CFD tools were then applied to make predictions of combustor operation and performance for the flight configuration and to aid in understanding the impacts of ground and flight uncertainties on combustor operation.

Impact of the addition of chicken litter on mercury speciation and emissions from coal combustion in a laboratory-scale fluidized bed combustor

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

Songgeng Li; Shuang Deng; Andy Wu

Co-combustion of chicken litter with coal was performed in a laboratory-scale fluidized bed combustor to investigate the effect of chicken litter addition on the partitioning behavior of mercury. Gaseous total and elemental mercury concentrations in the flue gas were measured online, and ash was analyzed for particle-bound mercury along with other elemental and surface properties. The mercury mass balance was between 85 and 105%. The experimental results show that co-combustion of chicken litter decreases the amount of elemental and total mercury in the gas phase. Mercury content in fly ash increases with an increasing chicken litter share. 22 refs., 6more » figs., 5 tabs.« less

Characteristics of a trapped-vortex (TV) combustor

NASA Technical Reports Server (NTRS)

Hsu, K.-Y.; Gross, L. P.; Trump, D. D.; Roquemore, W. M.

1994-01-01

The characteristics of a Trapped-Vortex (TV) combustor are presented. A vortex is trapped in the cavity established between two disks mounted in tandem. Fuel and air are injected directly into the cavity in such a way as to increase the vortex strength. Some air from the annular flow is also entrained into the recirculation zone of the vortex. Lean blow-out limits of the combustor are determined for a wide range of annular air flow rates. These data indicate that the lean blow-out limits are considerably lower for the TV combustor than for flames stabilized using swirl or bluff-bodies. The pressure loss through the annular duct is also low, being less than 2% for the flow conditions in this study. The instantaneous shape of the recirculation zone of the trapped vortex is measured using a two-color PIV technique. Temperature profiles obtained with CARS indicate a well mixed recirculation zone and demonstrate the impact of primary air injection on the local equivalence ratio.

Computational Analysis of Dynamic SPK(S8)-JP8 Fueled Combustor-Sector Performance

NASA Technical Reports Server (NTRS)

Ryder, R.; Hendricks, Roberts C.; Huber, M. L.; Shouse, D. T.

2010-01-01

Civil and military flight tests using blends of synthetic and biomass fueling with jet fuel up to 50:50 are currently considered as "drop-in" fuels. They are fully compatible with aircraft performance, emissions and fueling systems, yet the design and operations of such fueling systems and combustors must be capable of running fuels from a range of feedstock sources. This paper provides Smart Combustor or Fuel Flexible Combustor designers with computational tools, preliminary performance, emissions and particulates combustor sector data. The baseline fuel is kerosene-JP-8+100 (military) or Jet A (civil). Results for synthetic paraffinic kerosene (SPK) fuel blends show little change with respect to baseline performance, yet do show lower emissions. The evolution of a validated combustor design procedure is fundamental to the development of dynamic fueling of combustor systems for gas turbine engines that comply with multiple feedstock sources satisfying both new and legacy systems.

Combustor liner construction

NASA Technical Reports Server (NTRS)

Craig, H. M.; Wagner, W. B.; Strock, W. J. (Inventor)

1983-01-01

A combustor liner is fabricated from a plurality of individual segments each containing counter/parallel Finwall material and are arranged circumferentially and axially to define the combustion zone. Each segment is supported by a hook and ring construction to an opened lattice frame with sufficient tolerance between the hook and ring to permit thermal expansion with a minimum of induced stresses.

Variable volume combustor with aerodynamic support struts

DOEpatents

Ostebee, Heath Michael; Johnson, Thomas Edward; Stewart, Jason Thurman; Keener, Christopher Paul

2017-03-07

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and providing the flow of fuel therethrough. The support struts may include an aerodynamic contoured shape so as to distribute evenly a flow of air to the micro-mixer fuel nozzles.

Using the NASA GRC Sectored-One-Dimensional Combustor Simulation

NASA Technical Reports Server (NTRS)

Paxson, Daniel E.; Mehta, Vishal R.

2014-01-01

The document is a user manual for the NASA GRC Sectored-One-Dimensional (S-1-D) Combustor Simulation. It consists of three sections. The first is a very brief outline of the mathematical and numerical background of the code along with a description of the non-dimensional variables on which it operates. The second section describes how to run the code and includes an explanation of the input file. The input file contains the parameters necessary to establish an operating point as well as the associated boundary conditions (i.e. how it is fed and terminated) of a geometrically configured combustor. It also describes the code output. The third section describes the configuration process and utilizes a specific example combustor to do so. Configuration consists of geometrically describing the combustor (section lengths, axial locations, and cross sectional areas) and locating the fuel injection point and flame region. Configuration requires modifying the source code and recompiling. As such, an executable utility is included with the code which will guide the requisite modifications and insure that they are done correctly.

Transient/structural analysis of a combustor under explosive loads

NASA Technical Reports Server (NTRS)

Gregory, Peyton B.; Holland, Anne D.

1992-01-01

The 8-Foot High Temperature Tunnel (HTT) at NASA Langley Research Center is a combustion-driven blow-down wind tunnel. A major potential failure mode that was considered during the combustor redesign was the possibility of a deflagration and/or detonation in the combustor. If a main burner flame-out were to occur, then unburned fuel gases could accumulate and, if reignited, an explosion could occur. An analysis has been performed to determine the safe operating limits of the combustor under transient explosive loads. The failure criteria was defined and the failure mechanisms were determined for both peak pressures and differential pressure loadings. An overview of the gas dynamics analysis was given. A finite element model was constructed to evaluate 13 transient load cases. The sensitivity of the structure to the frequency content of the transient loading was assessed. In addition, two closed form dynamic analyses were conducted to verify the finite element analysis. It was determined that the differential pressure load or thrust load was the critical load mechanism and that the nozzle is the weak link in the combustor system.

Coanda injection system for axially staged low emission combustors

DOEpatents

Evulet, Andrei Tristan [Clifton Park, NY; Varatharajan, Balachandar [Cincinnati, OH; Kraemer, Gilbert Otto [Greer, SC; ElKady, Ahmed Mostafa [Niskayuna, NY; Lacy, Benjamin Paul [Greer, SC

2012-05-15

The low emission combustor includes a combustor housing defining a combustion chamber having a plurality of combustion zones. A liner sleeve is disposed in the combustion housing with a gap formed between the liner sleeve and the combustor housing. A secondary nozzle is disposed along a centerline of the combustion chamber and configured to inject a first fluid comprising air, at least one diluent, fuel, or combinations thereof to a downstream side of a first combustion zone among the plurality of combustion zones. A plurality of primary fuel nozzles is disposed proximate to an upstream side of the combustion chamber and located around the secondary nozzle and configured to inject a second fluid comprising air and fuel to an upstream side of the first combustion zone. The combustor also includes a plurality of tertiary coanda nozzles. Each tertiary coanda nozzle is coupled to a respective dilution hole. The tertiary coanda nozzles are configured to inject a third fluid comprising air, at least one other diluent, fuel, or combinations thereof to one or more remaining combustion zones among the plurality of combustion zones.

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.

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.

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

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.

Fluidized bed combustor modeling

NASA Technical Reports Server (NTRS)

Horio, M.; Rengarajan, P.; Krishnan, R.; Wen, C. Y.

1977-01-01

A general mathematical model for the prediction of performance of a fluidized bed coal combustor (FBC) is developed. The basic elements of the model consist of: (1) hydrodynamics of gas and solids in the combustor; (2) description of gas and solids contacting pattern; (3) kinetics of combustion; and (4) absorption of SO2 by limestone in the bed. The model is capable of calculating the combustion efficiency, axial bed temperature profile, carbon hold-up in the bed, oxygen and SO2 concentrations in the bubble and emulsion phases, sulfur retention efficiency and particulate carry over by elutriation. The effects of bed geometry, excess air, location of heat transfer coils in the bed, calcium to sulfur ratio in the feeds, etc. are examined. The calculated results are compared with experimental data. Agreement between the calculated results and the observed data are satisfactory in most cases. Recommendations to enhance the accuracy of prediction of the model are suggested.

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.

Parametric Modeling Investigation of a Radially-Staged Low-Emission Aviation Combustor

NASA Technical Reports Server (NTRS)

Heath, Christopher M.

2016-01-01

Aviation gas-turbine combustion demands high efficiency, wide operability and minimal trace gas emissions. Performance critical design parameters include injector geometry, combustor layout, fuel-air mixing and engine cycle conditions. The present investigation explores these factors and their impact on a radially staged low-emission aviation combustor sized for a next-generation 24,000-lbf-thrust engine. By coupling multi-fidelity computational tools, a design exploration was performed using a parameterized annular combustor sector at projected 100% takeoff power conditions. Design objectives included nitrogen oxide emission indices and overall combustor pressure loss. From the design space, an optimal configuration was selected and simulated at 7.1, 30 and 85% part-power operation, corresponding to landing-takeoff cycle idle, approach and climb segments. All results were obtained by solution of the steady-state Reynolds-averaged Navier-Stokes equations. Species concentrations were solved directly using a reduced 19-step reaction mechanism for Jet-A. Turbulence closure was obtained using a nonlinear K-epsilon model. This research demonstrates revolutionary combustor design exploration enabled by multi-fidelity physics-based simulation.

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...

Prediction of soot and thermal radiation in a model gas turbine combustor burning kerosene fuel spray at different swirl levels

NASA Astrophysics Data System (ADS)

Ghose, Prakash; Patra, Jitendra; Datta, Amitava; Mukhopadhyay, Achintya

2016-05-01

Combustion of kerosene fuel spray has been numerically simulated in a laboratory scale combustor geometry to predict soot and the effects of thermal radiation at different swirl levels of primary air flow. The two-phase motion in the combustor is simulated using an Eulerian-Lagragian formulation considering the stochastic separated flow model. The Favre-averaged governing equations are solved for the gas phase with the turbulent quantities simulated by realisable k-ɛ model. The injection of the fuel is considered through a pressure swirl atomiser and the combustion is simulated by a laminar flamelet model with detailed kinetics of kerosene combustion. Soot formation in the flame is predicted using an empirical model with the model parameters adjusted for kerosene fuel. Contributions of gas phase and soot towards thermal radiation have been considered to predict the incident heat flux on the combustor wall and fuel injector. Swirl in the primary flow significantly influences the flow and flame structures in the combustor. The stronger recirculation at high swirl draws more air into the flame region, reduces the flame length and peak flame temperature and also brings the soot laden zone closer to the inlet plane. As a result, the radiative heat flux on the peripheral wall decreases at high swirl and also shifts closer to the inlet plane. However, increased swirl increases the combustor wall temperature due to radial spreading of the flame. The high incident radiative heat flux and the high surface temperature make the fuel injector a critical item in the combustor. The injector peak temperature increases with the increase in swirl flow mainly because the flame is located closer to the inlet plane. On the other hand, a more uniform temperature distribution in the exhaust gas can be attained at the combustor exit at high swirl condition.

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.

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.

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.

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.

Experimental investigation of supersonic combustion in a strut-cavity based combustor

NASA Astrophysics Data System (ADS)

Sathiyamoorthy, K.; Danish, Tahzeeb Hassan; Srinivas, J.; Manjunath, P.

2018-07-01

Supersonic combustion was experimentally investigated in a strut-cavity based scramjet combustor with kerosene and pilot hydrogen as fuels. Strut-cavity is the space between two tandem struts in streamwise direction. The occurrence of cavity induced pressure oscillations in the strut-cavity was confirmed through cold flow experiments. The dominant modes of pressure oscillations were strongly influenced by the cavity aspect ratio. A ventilated rear wall (VRW), which is a new passive control device, was adopted in the strut-cavity. The strut-cavity with the VRW attenuated pressure oscillations better than the 'ramp rear wall' configuration. A scramjet combustor was realized with two strut-cavities in tandem for mixing enhancement and a strut-cavity with the VRW for flame stabilization. The combustor was tested at the following inlet conditions: total pressure of 4.89 bar, total temperature of 1517 K, and Mach number of 2. Supersonic combustion was observed. Steep increase in static pressure in the region of the strut-cavity with the VRW indicated that the flame was stabilized. The combustor was operated at a wide range of equivalence ratios (0.3-0.7) without inlet interactions. The total pressure at the combustor exit plane indicated that the flow was uniform, except at the central region. The total pressure loss and combustion efficiency of the combustor were evaluated for various equivalence ratios.

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

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...

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.

Experimental investigations on active cooling thermal protection structure of hydrocarbon-fueled scramjet combustor in arc heated facility

NASA Astrophysics Data System (ADS)

Jianqiang, Tu; Jinlong, Peng; Xianning, Yang; Lianzhong, Chen

2016-10-01

The active cooling thermal protection technology is the efficient method to resolve the long-duration work and reusable problems of hydrocarbon-fueled scramjet combustor, where worst thermo-mechanical loads occur. The fuel is passed through coolant channels adjacent to the heated surfaces to absorb heat from the heating exchanger panels, prior to injection into the combustor. The heating exchanger both cooled down the wall temperature of the combustor wall and heats and cracks the hydrocarbon fuel inside the panel to permit an easier combustion and satisfying combustion efficiency. The subscale active cooling metallic panels, with dimensions of 100×100 mm and different coolant channel sizes, have been tested under typical combustion thermal environment produced by arc heated Turbulent Flow Duct (TFD). The heat exchange ability of different coolant channel sizes has been obtained. The big-scale active cooling metallic panel, with dimensions of 100 × 750 mm and the coolant channel sizes of better heating exchange performance, has been made and tested in the big-scale arc heated TFD facility. The test results show that the local superheated ablation is easy to happen for the cooling fuel assigned asymmetrically in the bigscale active cooling metallic panel, and the cooling fuel rate can reduce 8%˜10% after spraying the Thermal Barrier Coating (TBC) in the heating surface.

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.

On the modelling of scalar and mass transport in combustor flows

NASA Technical Reports Server (NTRS)

Nikjooy, M.; So, R. M. C.

1989-01-01

Results are presented of a numerical study of swirling and nonswirling combustor flows with and without density variations. Constant-density arguments are used to justify closure assumptions invoked for the transport equations for turbulent momentum and scalar fluxes, which are written in terms of density-weighted variables. Comparisons are carried out with measurements obtained from three different axisymmetric model combustor experiments covering recirculating flow, swirling flow, and variable-density swirling flow inside the model combustors. Results show that the Reynolds stress/flux models do a credible job of predicting constant-density swirling and nonswirling combustor flows with passive scalar transport. However, their improvements over algebraic stress/flux models are marginal. The extension of the constant-density models to variable-density flow calculations shows that the models are equally valid for such flows.

Two and three-dimensional prediffuser combustor studies with air-water mixture

NASA Technical Reports Server (NTRS)

Laing, Peter; Ehresman, C. M.; Murthy, S. N. B.

1993-01-01

Two- and three-dimensional gas turbine prediffuser-combustor sectors were experimentally studied under a number of mixture and flow conditions in a tunnel operating with a two-phase, air-liquid film-droplet mixture. It is concluded that water vaporization in the combustor causes changes in both local gas temperature and state of vitiation and reduces reaction rates. Substantial accumulation of water and water vapor takes place in pocket over the combustor volume, even when the air-water mixture is steady in time. The accuracy of determining combustor performance changes increases with a better knowledge of the state of the air-water mixture in the primary zone. To establish flame-out conditions it is considered to be necessary to combine the prediction of detailed flowfield and chemical activity with that of flame stability and motion characteristics.

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.

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

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.

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.

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.

Development of a combustor liner composed of ceramic matrix composite (CMC)

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

Nishio, K.; Igashira, K.I.; Take, K.

The Research Institute of Advanced Materials Gas-Generator (AMG), which is a joint effort by the Japan Key Technology Center and 14 firms in Japan, has, since fiscal year 1992, been conducting technological studies on an innovative gas generator that will use 20% less fuel, weight 50% less, and emit 70% less NO{sub x} than the conventional gas generator through the use of advanced materials. Within this project, there is an R and D program for applying ceramic matrix composite (CMC) liners to the combustor, which is a major component of the gas generator. In the course of R and D,more » continuous SiC fiber-reinforced SiC composite (SiC{sup F}/SiC) was selected as the most suitable CMD for the combustor liner because of its thermal stability and formability. An evaluation of the applicability of the SiC{sup F}/SiC composite to the combustor liner on the basis of an evaluation of its mechanical properties and stress analysis of a SiC{sup F}/SiC combustor liner was carried out, and trial SiC{sup F}/SiC combustor liners, the largest of which was 500-mm in diameter, were fabricated by the filament winding and PIP (polymer impregnation and pyrolysis) method. Using a SiC{sup F}/SiC liner built to the actual dimensions, a noncooling combustion test was carried out and even when the gas temperature was raised to 1873K at outlet of the liner, no damage was observed after the test. Through their studies, the authors have confirmed the applicability of the selected SiC{sup F}/SiC composite as a combustor liner. In this paper, the authors describe the present state of the R and D of a CMC combustor liner.« less

Variable residence time vortex combustor

DOEpatents

Melconian, Jerry O.

1987-01-01

A variable residence time vortex combustor including a primary combustion chamber for containing a combustion vortex, and a plurality of louvres peripherally disposed about the primary combustion chamber and longitudinally distributed along its primary axis. The louvres are inclined to impel air about the primary combustion chamber to cool its interior surfaces and to impel air inwardly to assist in driving the combustion vortex in a first rotational direction and to feed combustion in the primary combustion chamber. The vortex combustor also includes a second combustion chamber having a secondary zone and a narrowed waist region in the primary combustion chamber interconnecting the output of the primary combustion chamber with the secondary zone for passing only lower density particles and trapping higher density particles in the combustion vortex in the primary combustion chamber for substantial combustion.

Systems and methods for detection of blowout precursors in combustors

DOEpatents

Lieuwen, Tim C.; Nair, Suraj

2006-08-15

The present invention comprises systems and methods for detecting flame blowout precursors in combustors. The blowout precursor detection system comprises a combustor, a pressure measuring device, and blowout precursor detection unit. A combustion controller may also be used to control combustor parameters. The methods of the present invention comprise receiving pressure data measured by an acoustic pressure measuring device, performing one or a combination of spectral analysis, statistical analysis, and wavelet analysis on received pressure data, and determining the existence of a blowout precursor based on such analyses. The spectral analysis, statistical analysis, and wavelet analysis further comprise their respective sub-methods to determine the existence of blowout precursors.

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.

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.

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

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Paxson, Daniel E.

2008-01-01

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

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...

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.

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.

Multi-dimensional computer simulation of MHD combustor hydrodynamics

NASA Astrophysics Data System (ADS)

Berry, G. F.; Chang, S. L.; Lottes, S. A.; Rimkus, W. A.

1991-04-01

Argonne National Laboratory is investigating the nonreacting jet gas mixing patterns in an MHD second stage combustor by using a 2-D multiphase hydrodynamics computer program and a 3-D single phase hydrodynamics computer program. The computer simulations are intended to enhance the understanding of flow and mixing patterns in the combustor, which in turn may lead to improvement of the downstream MHD channel performance. A 2-D steady state computer model, based on mass and momentum conservation laws for multiple gas species, is used to simulate the hydrodynamics of the combustor in which a jet of oxidizer is injected into an unconfined cross stream gas flow. A 3-D code is used to examine the effects of the side walls and the distributed jet flows on the non-reacting jet gas mixing patterns. The code solves the conservation equations of mass, momentum, and energy, and a transport equation of a turbulence parameter and allows permeable surfaces to be specified for any computational cell.

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.

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

NASA Technical Reports Server (NTRS)

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

1979-01-01

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

A study of air breathing rockets. 3: Supersonic mode combustors

NASA Astrophysics Data System (ADS)

Masuya, G.; Chinzel, N.; Kudo, K.; Murakami, A.; Komuro, T.; Ishii, S.

An experimental study was made on supersonic mode combustors of an air breathing rocket engine. Supersonic streams of room-temperature air and hot fuel-rich rocket exhaust were coaxially mixed and burned in a concially diverging duct of 2 deg half-angle. The effect of air inlet Mach number and excess air ratio was investigated. Axial wall pressure distribution was measured to calculate one dimensional change of Mach number and stagnation temperature. Calculated results showed that supersonic combustion occurred in the duct. At the exit of the duct, gas sampling and Pitot pressure measurement was made, from which radial distributions of various properties were deduced. The distribution of mass fraction of elements from rocket exhaust showed poor mixing performance in the supersonic mode combustors compared with the previously investigated cylindrical subsonic mode combustors. Secondary combustion efficiency correlated well with the centerline mixing parameter, but not with Annushkin's non-dimensional combustor length. No major effect of air inlet Mach number or excess air ratio was seen within the range of conditions under which the experiment was conducted.

Radiant heat transfer from flames in a single tubular turbojet combustor / Leonard Topper

NASA Technical Reports Server (NTRS)

Topper, Leonard

1952-01-01

An experimental investigation of thermal radiation from the flame of a single tubular turbojet-engine combustor to the combustor liner is presented. The effects of combustor inlet-air pressure, air mass flow, and fuel-air ratio on the radiant intensity and the temperature and emissivity of the flame are reported. The total radiation of the "luminous" flames (containing incandescent soot particles) was much greater (4 to 21 times) than the "nonluminous" molecular radiation. The intensity of radiation from the flame increased rapidly with an increase in combustor inlet-air pressure; it was affected to a lesser degree by variations in fuel-air ratio and air mass flow.

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.

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.

Performance and Pollution Measurements of Two-Row Swirl-Can Combustor Having 72 Modules

NASA Technical Reports Server (NTRS)

Biaglow, James A.; Trout, Arthur M.

1975-01-01

A test program was conducted to evaluate the performance and gaseous-pollutant levels of an experimental full-annulus 72-module swirl-can combustor. A comparison of data with those for a 120-module swirl-can combustor showed no significant difference in performance or levels of gaseous pollutants. Oxides of nitrogen were correlated for the 72- and 120-swirl-can combustors by using a previously developed parameter.

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

NASA Technical Reports Server (NTRS)

Hendricks, R. C.; Shouse, D. T.; Roquernore, W. M.; Burrus, D. L.; Duncan, B. S.; Ryder, R. C.; Brankovic, A.; Liu, N.-S.; Gallagher, J. R.; Hendricks, J. A.

2004-01-01

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

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

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.

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.

A Unique, Optically Accessible Flame Tube Facility for Lean Combustor Studies

NASA Technical Reports Server (NTRS)

Hicks, Yolanda R.; Locke, Randy J.; Wey, Chowen C.; Bianco, Jean

1995-01-01

A facility that allows interrogation of combusting flows by advanced diagnostic methods and instrumentation has been developed at the NASA Lewis Research Center. An optically accessible flame tube combustor is described which has high temperature, pressure, and air flow capabilities. The windows in the combustor measure 3.8 cm axially by 5.1 cm radially, providing 67% optical access to the 7.6 cm x 7.6 cm cross section flow chamber. Advanced gas analysis instrumentation is available through a gas chromatography/mass spectrometer system (GC/MS), which has on-line capability for heavy hydrocarbon measurement with resolution to the parts per billion level. The instrumentation allows one to study combusting flows and combustor subcomponents, such as fuel injectors and air swirlers. Planar Laser Induced Fluorescence (PLIF) can measure unstable combustion species, which cannot be obtained with traditional gas sampling. This type of data is especially useful to combustion modellers. The optical access allows measurements to have high spatial and temporal resolution. GC/MS data and PLIF images of OH- are presented from experiments using a lean direct injection (LDI) combustor burning Jet-A fuel at inlet temperatures ranging from 810 K to 866 K, combustor pressures up to 1380 kPa, and equivalence ratios from 0.41 to 0.59.

Effects of Burning Alternative Fuel in a 5-Cup Combustor Sector

NASA Technical Reports Server (NTRS)

Tacina, K. M.; Chang, C. T.; Lee, C.-M.; He, Z.; Herbon, J.

2015-01-01

A goal of NASA's Environmentally Responsible Aviation (ERA) program is to develop a combustor that will reduce the NOx emissions and that can burn both standard and alternative fuels. To meet this goal, NASA partnered with General Electric Aviation to develop a 5-cup combustor sector; this sector was tested in NASA Glenn's Advanced Subsonic Combustion Rig (ASCR). To verify that the combustor sector was fuel-flexible, it was tested with a 50-50 blend of JP-8 and a biofuel made from the camelina sativa plant. Results from this test were compared to results from tests where the fuel was neat JP-8. Testing was done at three combustor inlet conditions: cruise, 30% power, and 7% power. When compared to burning JP-8, burning the 50-50 blend did not significantly affect emissions of NOx, CO, or total hydrocarbons. Furthermore, it did not significantly affect the magnitude and frequency of the dynamic pressure fluctuations.

Nondestructive evaluation of ceramic matrix composite combustor components.

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

Sun, J. G.; Verrilli, M. J.; Stephan, R.

Combustor liners fabricated from a SiC/SiC composite were nondestructively interrogated before and after combustion rig testing. The combustor liners were inspected by X-ray, ultrasonic and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured liners, and from coupons machined from the components after combustion exposure. Thermography indications were found to correlate with reduced material properties obtained after rig testing. Microstructural examination of the SiC/SiC liners revealed the thermography indications to be delaminations and damaged fiber tows.

Variable volume combustor with center hub fuel staging

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

Ostebee, Heath Michael; McConnaughhay, Johnie Franklin; Stewart, Jason Thurman

The present application and the resultant patent provide a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a center hub for providing the flow of fuel therethrough. The center hub may include a first supply circuit for a first micro-mixer fuel nozzle and a second supply circuit for a second micro-mixer fuel nozzle.

Serial cooling of a combustor for a gas turbine engine

DOEpatents

Abreu, Mario E.; Kielczyk, Janusz J.

2001-01-01

A combustor for a gas turbine engine uses compressed air to cool a combustor liner and uses at least a portion of the same compressed air for combustion air. A flow diverting mechanism regulates compressed air flow entering a combustion air plenum feeding combustion air to a plurality of fuel nozzles. The flow diverting mechanism adjusts combustion air according to engine loading.

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.

Aerodynamic effect of combustor inlet-air pressure on fuel jet atomization

NASA Technical Reports Server (NTRS)

Ingebo, R. D.

1984-01-01

Mean drop diameters were measured with a recently developed scanning radiometer in a study of the atomization of liquid jets injected cross stream in high velocity and high pressure airflows. At constant inlet air pressure, reciprocal mean drop diameter, was correlated with airflow mass velocity. Over a combustor inlet-air pressure range of 1 to 21 atmospheres, the ratio of orifice to mean drop diameter, D(O)/D(M), was correlated with the product of Weber and Reynolds number, WeRe, and with the molecular scale momentum transfer ratio of gravitational to inertial forces.

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

NASA Technical Reports Server (NTRS)

Hendricks, Robert C.; Shouse, D. T.; Roquemore, W. M.; Burrus, D. L.; Duncan, B. S.; Ryder, R. C.; Brankovic, A.; Liu, N.-S.; Gallagher, J. R.; Hendricks, J. A.

2001-01-01

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

Experimental Investigation of Reacting Flow Characteristics in a Dual-Mode Scramjet Combustor

NASA Astrophysics Data System (ADS)

Shi, Deyong; Song, Wenyan; Ye, Jingfeng; Tao, Bo; Wang, Yanhua

2016-06-01

In this work, a hydrogen-fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS), respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y-direction, i.e. the transverse direction.

Fuel supply device for supplying fuel to an engine combustor

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

Lindsay, M.H.; Kerr, W.B.

1990-05-29

This patent describes a variable flow rate fuel supply device for supplying fuel to an engine combustor. It comprises: fuel metering means having a fuel valve means for controlling the flow rate of fuel to the combustor; piston means for dividing a first cooling fluid chamber from a second cooling fluid chamber; coupling means for coupling the piston means to the fuel valve means; and cooling fluid supply means in communication with the first and second cooling fluid chamber for producing a first pressure differential across the piston means for actuating the fuel valve means in a first direction, andmore » for producing a second pressure differential across the piston means for actuating the valve means in a second direction opposite the first direction, to control the flow rate of the fuel through the fuel metering means and into the engine combustor; and means for positioning the fuel metering means within the second cooling air chamber enabling the cooling air supply means to both cool the fuel metering means and control the fuel supply rate of fuel supplied by the fuel metering means to the combustor.« less

CFD analysis of jet mixing in low NOx flametube combustors

NASA Technical Reports Server (NTRS)

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

1991-01-01

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor was identified as a potential gas turbine combustor concept to reduce NO(x) emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NO(x) levels, cylindrical flametube versions of RQL combustors are being tested at NASA Lewis Research Center. A critical technology needed for the RQL combustor is a method of quickly mixing by-pass combustion air with rich-burn gases. Jet mixing in a cylindrical quick-mix section was numerically analyzed. The quick-mix configuration was five inches in diameter and employed twelve radial-inflow slots. The numerical analyses were performed with an advanced, validated 3-D Computational Fluid Dynamics (CFD) code named REFLEQS. Parametric variation of jet-to-mainstream momentum flux ratio (J) and slot aspect ratio was investigated. Both non-reacting and reacting analyses were performed. Results showed mixing and NO(x) emissions to be highly sensitive to J and slot aspect ratio. Lowest NO(x) emissions occurred when the dilution jet penetrated to approximately mid-radius. The viability of using 3-D CFD analyses for optimizing jet mixing was demonstrated.

Gas turbine engine combustor can with trapped vortex cavity

DOEpatents

Burrus, David Louis; Joshi, Narendra Digamber; Haynes, Joel Meier; Feitelberg, Alan S.

2005-10-04

A gas turbine engine combustor can downstream of a pre-mixer has a pre-mixer flowpath therein and circumferentially spaced apart swirling vanes disposed across the pre-mixer flowpath. A primary fuel injector is positioned for injecting fuel into the pre-mixer flowpath. A combustion chamber surrounded by an annular combustor liner disposed in supply flow communication with the pre-mixer. An annular trapped dual vortex cavity located at an upstream end of the combustor liner is defined between an annular aft wall, an annular forward wall, and a circular radially outer wall formed therebetween. A cavity opening at a radially inner end of the cavity is spaced apart from the radially outer wall. Air injection first holes are disposed through the forward wall and air injection second holes are disposed through the aft wall. Fuel injection holes are disposed through at least one of the forward and aft walls.

Sectoral combustor for burning low-BTU fuel gas

DOEpatents

Vogt, Robert L.

1980-01-01

A high-temperature combustor for burning low-BTU coal gas in a gas turbine is disclosed. The combustor includes several separately removable combustion chambers each having an annular sectoral cross section and a double-walled construction permitting separation of stresses due to pressure forces and stresses due to thermal effects. Arrangements are described for air-cooling each combustion chamber using countercurrent convective cooling flow between an outer shell wall and an inner liner wall and using film cooling flow through liner panel grooves and along the inner liner wall surface, and for admitting all coolant flow to the gas path within the inner liner wall. Also described are systems for supplying coal gas, combustion air, and dilution air to the combustion zone, and a liquid fuel nozzle for use during low-load operation. The disclosed combustor is fully air-cooled, requires no transition section to interface with a turbine nozzle, and is operable at firing temperatures of up to 3000.degree. F. or within approximately 300.degree. F. of the adiabatic stoichiometric limit of the coal gas used as fuel.

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.

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.

40 CFR 60.52a - Standard for municipal waste combustor metals.

Code of Federal Regulations, 2011 CFR

2011-07-01

... 40 Protection of Environment 6 2011-07-01 2011-07-01 false Standard for municipal waste combustor metals. 60.52a Section 60.52a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR... September 20, 1994 § 60.52a Standard for municipal waste combustor metals. (a) On and after the date on...

40 CFR 60.52a - Standard for municipal waste combustor metals.

Code of Federal Regulations, 2010 CFR

2010-07-01

... 40 Protection of Environment 6 2010-07-01 2010-07-01 false Standard for municipal waste combustor metals. 60.52a Section 60.52a Protection of Environment ENVIRONMENTAL PROTECTION AGENCY (CONTINUED) AIR... September 20, 1994 § 60.52a Standard for municipal waste combustor metals. (a) On and after the date on...

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.

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.

Simultaneous high-speed schlieren and OH chemiluminescence imaging in a hybrid rocket combustor at elevated pressures

NASA Astrophysics Data System (ADS)

Miller, Victor; Jens, Elizabeth T.; Mechentel, Flora S.; Cantwell, Brian J.; Stanford Propulsion; Space Exploration Group Team

2014-11-01

In this work, we present observations of the overall features and dynamics of flow and combustion in a slab-type hybrid rocket combustor. Tests were conducted in the recently upgraded Stanford Combustion Visualization Facility, a hybrid rocket combustor test platform capable of generating constant mass-flux flows of oxygen. High-speed (3 kHz) schlieren and OH chemiluminescence imaging were used to visualize the flow. We present imaging results for the combustion of two different fuel grains, a classic, low regression rate polymethyl methacrylate (PMMA), and a high regression rate paraffin, and all tests were conducted in gaseous oxygen. Each fuel grain was tested at multiple free-stream pressures at constant oxidizer mass flux (40 kg/m2s). The resulting image sequences suggest that aspects of the dynamics and scaling of the system depend strongly on both pressure and type of fuel.

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 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.

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

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.

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.

Effect of Fuel Injection and Mixing Characteristics on Pulse-Combustor Performance at High-Pressure

NASA Technical Reports Server (NTRS)

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

2014-01-01

Recent calculations of pulse-combustors operating at high-pressure conditions produced pressure gains significantly lower than those observed experimentally and computationally at atmospheric conditions. The factors limiting the pressure-gain at high-pressure conditions are identified, and the effects of fuel injection and air mixing characteristics on performance are investigated. New pulse-combustor configurations were developed, and the results show that by suitable changes to the combustor geometry, fuel injection scheme and valve dynamics the performance of the pulse-combustor operating at high-pressure conditions can be increased to levels comparable to those observed at atmospheric conditions. In addition, the new configurations can significantly reduce the levels of NOx emissions. One particular configuration resulted in extremely low levels of NO, producing an emission index much less than one, although at a lower pressure-gain. Calculations at representative cruise conditions demonstrated that pulse-combustors can achieve a high level of performance at such conditions.

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.

A chemical reactor network for oxides of nitrogen emission prediction in gas turbine combustor

NASA Astrophysics Data System (ADS)

Hao, Nguyen Thanh

2014-06-01

This study presents the use of a new chemical reactor network (CRN) model and non-uniform injectors to predict the NOx emission pollutant in gas turbine combustor. The CRN uses information from Computational Fluid Dynamics (CFD) combustion analysis with two injectors of CH4-air mixture. The injectors of CH4-air mixture have different lean equivalence ratio, and they control fuel flow to stabilize combustion and adjust combustor's equivalence ratio. Non-uniform injector is applied to improve the burning process of the turbine combustor. The results of the new CRN for NOx prediction in the gas turbine combustor show very good agreement with the experimental data from Korea Electric Power Research Institute.

Stagnation point reverse flow combustor for a combustion system

NASA Technical Reports Server (NTRS)

Zinn, Ben T. (Inventor); Neumeier, Yedidia (Inventor); Seitzman, Jerry M. (Inventor); Jagoda, Jechiel (Inventor); Hashmonay, Ben-Ami (Inventor)

2007-01-01

A combustor assembly includes a combustor vessel having a wall, a proximate end defining an opening and a closed distal end opposite said proximate end. A manifold is carried by the proximate end. The manifold defines a combustion products exit. The combustion products exit being axially aligned with a portion of the closed distal end. A plurality of combustible reactant ports is carried by the manifold for directing combustible reactants into the combustion vessel from the region of the proximate end towards the closed distal end.

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.

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

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.

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.

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.

Fuel-Flexible Gas Turbine Combustor Flametube Facility

NASA Technical Reports Server (NTRS)

Little, James E.; Nemets, Stephen A.; Tornabene, Robert T.; Smith, Timothy D.; Frankenfield, Bruce J.; Manning, Stephen D.; Thompson, William K.

2004-01-01

Facility modifications have been completed to an existing combustor flametube facility to enable testing with gaseous hydrogen propellants at the NASA Glenn Research Center. The purpose of the facility is to test a variety of fuel nozzle and flameholder hardware configurations for use in aircraft combustors. Facility capabilities have been expanded to include testing with gaseous hydrogen, along with the existing hydrocarbon-based jet fuel. Modifications have also been made to the facility air supply to provide heated air up to 350 psig, 1100 F, and 3.0 lbm/s. The facility can accommodate a wide variety of flametube and fuel nozzle configurations. Emissions and performance data are obtained via a variety of gas sample probe configurations and emissions measurement equipment.

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.

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

Wedge edge ceramic combustor tile

DOEpatents

Shaffer, J.E.; Holsapple, A.C.

1997-06-10

A multipiece combustor has a portion thereof being made of a plurality of ceramic segments. Each of the plurality of ceramic segments have an outer surface and an inner surface. Each of the plurality of ceramic segments have a generally cylindrical configuration and including a plurality of joints. The joints define joint portions, a first portion defining a surface being skewed to the outer surface and the inner surface. The joint portions have a second portion defining a surface being skewed to the outer surface and the inner surface. The joint portions further include a shoulder formed intermediate the first portion and the second portion. The joints provide a sealing interlocking joint between corresponding ones of the plurality of ceramic segments. Thus, the multipiece combustor having the plurality of ceramic segment with the plurality of joints reduces the physical size of the individual components and the degradation of the surface of the ceramic components in a tensile stress zone is generally eliminated reducing the possibility of catastrophic failures. 7 figs.

Wedge edge ceramic combustor tile

DOEpatents

Shaffer, James E.; Holsapple, Allan C.

1997-01-01

A multipiece combustor has a portion thereof being made of a plurality of ceramic segments. Each of the plurality of ceramic segments have an outer surface and an inner surface. Each of the plurality of ceramic segments have a generally cylindrical configuration and including a plurality of joints. The joints define joint portions, a first portion defining a surface being skewed to the outer surface and the inner surface. The joint portions have a second portion defining a surface being skewed to the outer surface and the inner surface. The joint portions further include a shoulder formed intermediate the first portion and the second portion. The joints provide a sealing interlocking joint between corresponding ones of the plurality of ceramic segments. Thus, the multipiece combustor having the plurality of ceramic segment with the plurality of joints reduces the physical size of the individual components and the degradation of the surface of the ceramic components in a tensile stress zone is generally eliminated reducing the possibility of catastrophic failures.

Stably operating pulse combustor and method

DOEpatents

Zinn, Ben T.; Reiner, David

1990-01-01

A pulse combustor apparatus adapted to burn either a liquid fuel or a pulverized solid fuel within a preselected volume of the combustion chamber. The combustion process is substantially restricted to an optimum combustion zone in order to attain effective pulse combustion operation.

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.

Resolution of the buoyancy in the 8-foot high temperature tunnel combustor

NASA Technical Reports Server (NTRS)

Loney, Norman W.

1995-01-01

Currently, the 8-Foot High Temperature Tunnel (8-Ft. HTT) combustor produces a good profile at only one point (2000 psia and 3650 R with oxygen enrichment). Air is enriched with oxygen (liquid) so that the combustor product gas will contain the volumetric amount of oxygen normally found in air. The oxygen enriched air has a large fraction that is not reacted and flows through the outer periphery of the fuel injector. This ring of cold air in addition to the relatively cold walls of the combustor set up buoyancy forces that produce a segregation of relatively cool gases at the bottom of the combustor exit. The basic problem is to produce a test gas that has uniform properties at all combustor conditions. The combustor temperature may be as high as 3700 R or as low as 2000 R. Combustor pressures can be as high as 3500 psia (no oxygen enrichment) and as low as 600 psia. The segregation is most severe with oxygen enriched air, since its temperature is lower and its density is high. The combustor is lined with nickel 201 and can be operated at about 1600 R maximum. A global mixing process is desired that produces an acceptable profile of temperature, species, and velocity at the exit of the combustor. The ultimate goal is a temperature profile with about 100 R variance and about 2 percent variance in oxygen. The exit total temperature must not be lowered significantly by the mixing apparatus or mechanisms employed. If immersed bodies are used, they must also be kept very hot. All combustor wall modifications must be able to survive the heat and structural conditions of the varied operating conditions. Our approach to resolving this issue is being conducted in three stages: (1) Consider mixing exclusively, (2) Resolve the heat transfer concerns resulting from the chosen mixing strategy, and (3) Solve the material and structural problems resulting from stages (1) and (2). Since the 8-Ft. HTT is unavailable for experimentation, the study is conducted exclusively with

Aerodynamic effect of combustor inlet-air pressure on fuel jet atomization

NASA Technical Reports Server (NTRS)

Ingebo, R. D.

1984-01-01

Mean drop diameters were measured with a recently developed scanning radiometer in a study of the atomization of liquid jets injected cross stream in high velocity and high pressure airflows. At constant inlet air pressure, reciprocal mean drop diameter was correlated with airflow mass velocity. Over a combustor inlet-air pressure range of 1 to 21 atmospheres, the ratio of orifice to mean drop diameter, D(O)/D(M), was correlated with the product of Weber and Reynolds number, WeRe, and with the molecular scale momentum transfer ratio of gravitational to inertial forces. Previously announced in STAR as N84-22910

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.

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.

CFD analysis of jet mixing in low NO(x) flametube combustors

NASA Technical Reports Server (NTRS)

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

1991-01-01

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor has been identified as a potential gas turbine combustor concept to reduce NO(x) emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NO(x) levels, cylindrical flametube versions of RQL combustors are being tested at NASA Lewis Research Center. A critical technology needed for the RQL combustor is a method of quickly mixing by-pass combustion air with rich-burn gases. Jet mixing in a cylindrical quick-mix section was numerically analyzed. The quick-mix configuration was five inches in diameter and employed twelve radial-inflow slots. The numerical analyses were performed with an advanced, validated 3D Computational Fluid Dynamics (CFD) code named REFLEQS. Parametric variation of jet-to-mainstream momentum flux ratio (J) and slot aspect ratio was investigated. Both non-reacting and reacting analyses were performed. Results showed mixing and NO(x) emissions to be highly sensitive to J and slot aspect ratio. Lowest NO(x) emissions occurred when the dilution jet penetrated to approximately mid-radius. The viability of using 3D CFD analyses for optimizing jet mixing was demonstrated.

An Experimental Study of Swirling Flows as Applied to Annular Combustors

NASA Technical Reports Server (NTRS)

Seal, Michael Damian, II

1997-01-01

This thesis presents an experimental study of swirling flows with direct applications to gas turbine combustors. Two separate flowfields were investigated: a round, swirling jet and a non-combusting annular combustor model. These studies were intended to allow both a further understanding of the behavior of general swirling flow characteristics, such as the recirculation zone, as well as to provide a base for the development of computational models. In order to determine the characteristics of swirling flows the concentration fields of a round, swirling jet were analyzed for varying amount of swirl. The experimental method used was a light scattering concentration measurement technique known as marker nephelometry. Results indicated the formation of a zone of recirculating fluid for swirl ratios (rotational speed x jet radius over mass average axial velocity) above a certain critical value. The size of this recirculation zone, as well as the spread angle of the jet, was found to increase with increase in the amount of applied swirl. The annular combustor model flowfield simulated the cold-flow characteristics of typical current annular combustors: swirl, recirculation, primary air cross jets and high levels of turbulence. The measurements in the combustor model made by the Laser Doppler Velocimetry technique, allowed the evaluation of the mean and rms velocities in the three coordinate directions, one Reynold's shear stress component and the turbulence kinetic energy: The primary cross jets were found to have a very strong effect on both the mean and turbulence flowfields. These cross jets, along with a large step change in area and wall jet inlet flow pattern, reduced the overall swirl in the test section to negligible levels. The formation of the strong recirculation zone is due mainly to the cross jets and the large step change in area. The cross jets were also found to drive a four-celled vortex-type motion (parallel to the combustor longitudinal axis) near the

Combustor cap having non-round outlets for mixing tubes

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

Hughes, Michael John; Boardman, Gregory Allen; McConnaughhay, Johnie Franklin

2016-12-27

A system includes a a combustor cap configured to be coupled to a plurality of mixing tubes of a multi-tube fuel nozzle, wherein each mixing tube of the plurality of mixing tubes is configured to mix air and fuel to form an air-fuel mixture. The combustor cap includes multiple nozzles integrated within the combustor cap. Each nozzle of the multiple nozzles is coupled to a respective mixing tube of the multiple mixing tubes. In addition, each nozzle of the multiple nozzles includes a first end and a second end. The first end is coupled to the respective mixing tube ofmore » the multiple mixing tubes. The second end defines a non-round outlet for the air-fuel mixture. Each nozzle of the multiple nozzles includes an inner surface having first and second portions, the first portion radially diverges along an axial direction from the first end to the second end, and the second portion radially converges along the axial direction from the first end to the second end.« less

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.

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.

NASA Lewis Research Center's Preheated Combustor and Materials Test Facility

NASA Technical Reports Server (NTRS)

Nemets, Steve A.; Ehlers, Robert C.; Parrott, Edith

1995-01-01

The Preheated Combustor and Materials Test Facility (PCMTF) in the Engine Research Building (ERB) at the NASA Lewis Research Center is one of two unique combustor facilities that provide a nonvitiated air supply to two test stands, where the air can be used for research combustor testing and high-temperature materials testing. Stand A is used as a research combustor stand, whereas stand B is used for cyclic and survivability tests of aerospace materials at high temperatures. Both stands can accommodate in-house and private industry research programs. The PCMTF is capable of providing up to 30 lb/s (pps) of nonvitiated, 450 psig combustion air at temperatures ranging from 850 to 1150 g F. A 5000 gal tank located outdoors adjacent to the test facility can provide jet fuel at a pressure of 900 psig and a flow rate of 11 gal/min (gpm). Gaseous hydrogen from a 70,000 cu ft (CF) tuber is also available as a fuel. Approximately 500 gpm of cooling water cools the research hardware and exhaust gases. Such cooling is necessary because the air stream reaches temperatures as high as 3000 deg F. The PCMTF provides industry and Government with a facility for studying the combustion process and for obtaining valuable test information on advanced materials. This report describes the facility's support systems and unique capabilities.

A cross-ethnic survey of CFB and SLC44A4, Indian ulcerative colitis GWAS hits, underscores their potential role in disease susceptibility

PubMed Central

Gupta, Aditi; Juyal, Garima; Sood, Ajit; Midha, Vandana; Yamazaki, Keiko; Vich Vila, Arnau; Esaki, Motohiro; Matsui, Toshiyuki; Takahashi, Atsushi; Kubo, Michiaki; Weersma, Rinse K; Thelma, B K

2017-01-01

The first ever genome-wide association study (GWAS) of ulcerative colitis in genetically distinct north Indian population identified two novel genes namely CFB and SLC44A4. Considering their biological relevance, we investigated allelic/genetic heterogeneity in these genes among ulcerative colitis cohorts of north Indian, Japanese and Dutch origin using high-density ImmunoChip case–control genotype data. Comparative linkage disequilibrium profiling and test of association were performed. Of the 28 CFB SNPs, similar strength of association was observed for rs4151657 (novel ulcerative colitis GWAS SNP) in north Indians (P=1.73 × 10−10) and Japanese (P=2.02 × 10−12) but not in the Dutch. Further, a three-marker haplotype was shared between north Indians and Japanese (P<10−8), but a different five-marker haplotype was associated (P=2.07 × 10−6) in the Dutch. Of the 22 SLC44A4 SNPs, rs2736428 (novel ulcerative colitis GWAS SNP) was found significantly associated in north Indians (P=4.94 × 10−10) and Japanese (P=3.37 × 10−9), but not among the Dutch. These results suggest (i) apparent allelic heterogeneity in CFB and genetic heterogeneity in SLC44A4 across different ethnic groups; (ii) shared ulcerative colitis genetic etiological factors among Asians; and finally (iii) re-exploration of GWAS findings together with high-density genotyping/sequencing and trans-ethnic fine mapping approaches may help identify shared and population-specific risk variants and enable to explain missing disease heritability. PMID:27759029

Real-Time Optical Fuel-to-Air Ratio Sensor for Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Nguyen, Quang-Viet; Mongia, Rajiv K.; Dibble, Robert W.

1999-01-01

The measurement of the temporal distribution of fuel in gas turbine combustors is important in considering pollution, combustion efficiency and combustor dynamics and acoustics. Much of the previous work in measuring fuel distributions in gas turbine combustors has focused on the spatial aspect of the distribution. The temporal aspect however, has often been overlooked, even though it is just as important. In part, this is due to the challenges of applying real-time diagnostic techniques in a high pressure and high temperature environment. A simple and low-cost instrument that non-intrusively measures the real-time fuel-to-air ratio (FAR) in a gas turbine combustor has been developed. The device uses a dual wavelength laser absorption technique to measure the concentration of most hydrocarbon fuels such as jet fuel, methane, propane, etc. The device can be configured to use fiber optics to measure the local FAR inside a high pressure test rig without the need for windows. Alternatively, the device can readily be used in test rigs that have existing windows without modifications. An initial application of this instrument was to obtain time-resolved measurements of the FAR in the premixer of a lean premixed prevaporized (LPP) combustor at inlet air pressures and temperatures as high as 17 atm at 800 K, with liquid JP-8 as the fuel. Results will be presented that quantitatively show the transient nature of the local FAR inside a LPP gas turbine combustor at actual operating conditions. The high speed (kHz) time resolution of this device, combined with a rugged fiber optic delivery system, should enable the realization of a flight capable active-feedback and control system for the abatement of noise and pollutant emissions in the future. Other applications that require an in-situ and time-resolved measurement of fuel vapor concentrations should also find this device to be of use.

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

NASA Technical Reports Server (NTRS)

Taylor, J. R.

1979-01-01

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

Wave combustors for trans-atmospheric vehicles

NASA Technical Reports Server (NTRS)

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

1989-01-01

The Wave Combustor is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture and thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter, lighter engine compared to the scramjet. This engine, which is called the Oblique Detonation Wave Engine (ODWE), can then be utilized to provide a smaller, lighter vehicle or to provide a higher payload capability for a given vehicle weight. An analysis of the performance of a conceptual trans-atmospheric vehicle powered by an ODWE is given here.

Numerical study of supersonic combustors by multi-block grids with mismatched interfaces

NASA Technical Reports Server (NTRS)

Moon, Young J.

1990-01-01

A three dimensional, finite rate chemistry, Navier-Stokes code was extended to a multi-block code with mismatched interface for practical calculations of supersonic combustors. To ensure global conservation, a conservative algorithm was used for the treatment of mismatched interfaces. The extended code was checked against one test case, i.e., a generic supersonic combustor with transverse fuel injection, examining solution accuracy, convergence, and local mass flux error. After testing, the code was used to simulate the chemically reacting flow fields in a scramjet combustor with parallel fuel injectors (unswept and swept ramps). Computational results were compared with experimental shadowgraph and pressure measurements. Fuel-air mixing characteristics of the unswept and swept ramps were compared and investigated.

Obseration of flow regime transition in CFB riser using an LDV

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

Yue, Paul C.; Mei, Joseph S.; Shadle, Lawrence J.

2011-01-01

The solids flow in a circulating fluidized bed (CFB) riser is often described to have a core-annular structure. For a given superficial gas velocity, at the initial introduction of solids into a riser a flow structure of dilute upflow regime exists. Continuing to increase the solids flow in the riser transitions the flow structure to the core-annular flow regime. However, with further increase of solids flow a condition is reached, depending on the superficial gas velocity, where all the solids across the riser cross section flow upwards, even those at the wall. When the solids flux, solids fraction and gasmore » velocity are relatively high, such a condition is described as the dense phase suspense upflow (DSU) regime. In this paper we report our observations of these flow regime transitions by using a laser Doppler velocimeter (LDV) to monitor the upward and downward particle flow velocities at and near the riser wall of the National Energy Technology Laboratory’s 30.4 centimeters diameter CFB cold flow model. The particles were high density polyethylene (PPE) spheres with a Sauter mean diameter of 861 micron and a density of 800 kg/m3. Three superficial gas velocities of 6.55 m/s, 10.67 m/s and 13.72 m/s were used in this study. For the case of superficial gas velocity 6.55 m/s, the experimental data show that the transition from dilute upflow to core-annular flow occurred when the solids flux was about 7 kg/m{sup 2}-s and the transition from core-annular flow to dense suspension upflow was about 147 kg/m{sup 2}-s. As the superficial gas velocity was increased to 10.67 m/s the corresponding flow regime transitions were at 34 kg/m{sup 2}-s and 205 kg/m{sup 2}-s, respectively. For the case of superficial gas velocity of 13.72 m/s the data showed no distinct transition of flow regimes. The particles were all upflow for the range of solids fluxes from 10 kg/m{sup 2}-s to 286 kg/m{sup 2}-s.« less

Effects of Cavity Configurations on Flameholding and Performances of Kerosene Fueled Scramjet Combustor

NASA Astrophysics Data System (ADS)

Shi, Deyong; Song, Wenyan; Wang, Yuhang; Wang, Yanhua

2017-08-01

In this work, the effects of cavity flameholder configurations on flameholding and performances of kerosene fueled scramjet combustor were studied experimentally and numerically. For experiments, a directly connected ground facility was used and clean high enthalpy air, with a total temperature of 800 K and a total pressure of 800 Kpa, was provided by an electricity resistance heater. To investigate the effects of cavity configurations on flameholding capacity and reacting-flow characteristics, three different flameholders, one single cavity flameholder and two tandem cavity flameholders, were used in experiments. For the two combustors with tandem cavity flameholders, the location and configurations of its up-stream cavity were same with the single cavity flameholder, and the length-to-depth ratios for down-stream cavities were 9 and 11 respectively. The experimental results showed that stabilize kerosene combustion were achieved for combustor with tandem cavity flameholders mounted, and none for that with single cavity flameholder. The none-reacting and reacting flows of combustor models with tandem cavity flameholders were compared and studied with numerical and experimental results. The results showed that higher combustion efficiencies and pressure recovery ratios were achieved for the combustor with down-stream cavity length-to-depth ratio of 9.

Preliminary investigation of the performance of a single tubular combustor at pressure up to 12 atmospheres

NASA Technical Reports Server (NTRS)

Wear, Jerrold D; Butze, Helmut F

1954-01-01

The effects of combustor operation at conditions representative of those encountered in high pressure-ratio turbojet engines or at high flight speeds on carbon deposition, exhaust smoke, and combustion efficiency were studied in a single tubular combustor. Carbon deposition and smoke formation tests were conducted over a range of combustor-inlet pressures from 33 to 173 pounds per square inch absolute and combustor reference velocities from 78 to 143 feet per second. Combustion efficiency tests were conducted over a range of pressures from 58 to 117 pounds per square inch absolute and velocities from 89 to 172 feet per second.

Nonintrusive transceiver and method for characterizing temperature and velocity fields in a gas turbine combustor

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

DeSilva, Upul P.; Claussen, Heiko

An acoustic transceiver is implemented for measuring acoustic properties of a gas in a turbine engine combustor. The transceiver housing defines a measurement chamber and has an opening adapted for attachment to a turbine engine combustor wall. The opening permits propagation of acoustic signals between the gas in the turbine engine combustor and gas in the measurement chamber. An acoustic sensor mounted to the housing receives acoustic signals propagating in the measurement chamber, and an acoustic transmitter mounted to the housing creates acoustic signals within the measurement chamber. An acoustic measurement system includes at least two such transceivers attached tomore » a turbine engine combustor wall and connected to a controller.« less

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.

Combustion-acoustic stability analysis for premixed gas turbine combustors

NASA Technical Reports Server (NTRS)

Darling, Douglas; Radhakrishnan, Krishnan; Oyediran, Ayo; Cowan, Lizabeth

1995-01-01

Lean, prevaporized, premixed combustors are susceptible to combustion-acoustic instabilities. A model was developed to predict eigenvalues of axial modes for combustion-acoustic interactions in a premixed combustor. This work extends previous work by including variable area and detailed chemical kinetics mechanisms, using the code LSENS. Thus the acoustic equations could be integrated through the flame zone. Linear perturbations were made of the continuity, momentum, energy, chemical species, and state equations. The qualitative accuracy of our approach was checked by examining its predictions for various unsteady heat release rate models. Perturbations in fuel flow rate are currently being added to the model.

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...

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

Propagation of a premixed flame in a divided-chamber combustor

NASA Technical Reports Server (NTRS)

Cattolica, R. J.; Barr, P. K.; Mansour, N. N.

1987-01-01

The propagation of premixed ethylene-air mixtures (of 0.5, 0.525, 0.55, and 0.65 equivalence ratios) in a divided-chamber combustor was investigated. The vessel, divided by a small cylindrical prechamber, had optical access (for laser-schlieren videography) and was instrumented by a pressure transducer. For the Reynolds numbers of 1870, 2300, and 2830, the observed spatial development of the laminar flames showed that the flame position and shape could be scaled by a characteristic time, based on the burned gas flame speed and the length of the prechamber. Above a Reynolds number of 4330, this scaling breaks down the appearance of Kelvin-Helmholtz instabilities. The observed flame propagation was compared with predictions obtained with a numerical model of flame propagation. The calculated spatial and temporal development of the flame in the main combustion chamber agreed with the experimental observations only for the lowest Reynolds number (1870).

Characterization of centrifugally-loaded flame migration for ultra-compact combustors

NASA Astrophysics Data System (ADS)

LeBay, Kenneth D.

The Air Force Research Laboratory (AFRL) has designed a centrifugally-loaded Ultra-Compact Combustor (UCC) showing viable merit for reducing gas turbine combustor length by as much as 66%. The overarching goal of this research was to characterize the migration of centrifugally-loaded flames in a sectional model of the UCC to enable scaling of the design from 15 cm to the 50--75 cm diameter of most engines. Two-line Planar Laser-Induced Fluorescence thermometry (PLIF) of OH, time-resolved Particle Image Velocimetry (PIV), and high-speed video data were collected. Using a sectional UCC model, the flame migration angle was determined to be a function of the UCC/core velocity ratio (VR) while both the VR and the centrifugal or "g-load" affected the migration quantity. Higher g-loads and lower VRs yielding higher migration but lower VRs had lower core flow temperatures due to higher core air mass flow. A comparison of the straight and curved UCC sections showed the centrifugal load increased the flame migration but increased unsteadiness. The flame migration into the core was estimated using pressure and temperature measurements upstream, and PIV measurements downstream of the core flow interface with constant density and velocity profile assumptions. The flame migration quantity was used to estimate the core flow temperature which was in relatively good agreement with the measured PLIF values. The migration quantity scaled relatively linearly with the UCC tangential velocity, which corresponds to the g-load value, with the slope determined by the VR. A simple analytical model resulted for the dependence of the migration quantity on the tangential velocity and VR. The quantitative relationships determined in this research provided a detailed description of the migration of centrifugally-loaded flames in a sectional UCC.

Stably operating pulse combustor and method

DOEpatents

Zinn, B.T.; Reiner, D.

1990-05-29

A pulse combustor apparatus is described which is adapted to burn either a liquid fuel or a pulverized solid fuel within a preselected volume of the combustion chamber. The combustion process is substantially restricted to an optimum combustion zone in order to attain effective pulse combustion operation. 4 figs.

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.

Transient analysis of a pulsed detonation combustor using the numerical propulsion system simulation

NASA Astrophysics Data System (ADS)

Hasler, Anthony Scott

The performance of a hybrid mixed flow turbofan (with detonation tubes installed in the bypass duct) is investigated in this study and compared with a baseline model of a mixed flow turbofan with a standard combustion chamber as a duct burner. Previous studies have shown that pulsed detonation combustors have the potential to be more efficient than standard combustors, but they also present new challenges that must be overcome before they can be utilized. The Numerical Propulsion System Simulation (NPSS) will be used to perform the analysis with a pulsed detonation combustor model based on a numerical simulation done by Endo, Fujiwara, et. al. Three different cases will be run using both models representing a take-off situation, a subsonic cruise and a supersonic cruise situation. Since this study investigates a transient analysis, the pulse detonation combustor is run in a rig setup first and then its pressure and temperature are averaged for the cycle to obtain quasi-steady results.

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

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Paxson, Daniel E.

2008-01-01

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

Optical Diagnosis of Gas Turbine Combustors Being Conducted

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

Fuel injection staged sectoral combustor for burning low-BTU fuel gas

DOEpatents

Vogt, Robert L.

1981-01-01

A high-temperature combustor for burning low-BTU coal gas in a gas turbine is described. The combustor comprises a plurality of individual combustor chambers. Each combustor chamber has a main burning zone and a pilot burning zone. A pipe for the low-BTU coal gas is connected to the upstream end of the pilot burning zone; this pipe surrounds a liquid fuel source and is in turn surrounded by an air supply pipe; swirling means are provided between the liquid fuel source and the coal gas pipe and between the gas pipe and the air pipe. Additional preheated air is provided by counter-current coolant air in passages formed by a double wall arrangement of the walls of the main burning zone communicating with passages of a double wall arrangement of the pilot burning zone; this preheated air is turned at the upstream end of the pilot burning zone through swirlers to mix with the original gas and air input (and the liquid fuel input when used) to provide more efficient combustion. One or more fuel injection stages (second stages) are provided for direct input of coal gas into the main burning zone. The countercurrent air coolant passages are connected to swirlers surrounding the input from each second stage to provide additional oxidant.

Fuel injection staged sectoral combustor for burning low-BTU fuel gas

DOEpatents

Vogt, Robert L.

1985-02-12

A high-temperature combustor for burning low-BTU coal gas in a gas turbine is described. The combustor comprises a plurality of individual combustor chambers. Each combustor chamber has a main burning zone and a pilot burning zone. A pipe for the low-BTU coal gas is connected to the upstream end of the pilot burning zone: this pipe surrounds a liquid fuel source and is in turn surrounded by an air supply pipe: swirling means are provided between the liquid fuel source and the coal gas pipe and between the gas pipe and the air pipe. Additional preheated air is provided by counter-current coolant air in passages formed by a double wall arrangement of the walls of the main burning zone communicating with passages of a double wall arrangement of the pilot burning zone: this preheated air is turned at the upstream end of the pilot burning zone through swirlers to mix with the original gas and air input (and the liquid fuel input when used) to provide more efficient combustion. One or more fuel injection stages (second stages) are provided for direct input of coal gas into the main burning zone. The countercurrent air coolant passages are connected to swirlers surrounding the input from each second stage to provide additional oxidant.

40 CFR 60.34b - Emission guidelines for municipal waste combustor operating practices.

Code of Federal Regulations, 2013 CFR

2013-07-01

... combustor operating practices. 60.34b Section 60.34b Protection of Environment ENVIRONMENTAL PROTECTION... September 20, 1994 § 60.34b Emission guidelines for municipal waste combustor operating practices. (a) For approval, a State plan shall include emission limits for carbon monoxide at least as protective as the...

40 CFR 60.34b - Emission guidelines for municipal waste combustor operating practices.

Code of Federal Regulations, 2011 CFR

2011-07-01

... combustor operating practices. 60.34b Section 60.34b Protection of Environment ENVIRONMENTAL PROTECTION... September 20, 1994 § 60.34b Emission guidelines for municipal waste combustor operating practices. (a) For approval, a State plan shall include emission limits for carbon monoxide at least as protective as the...

40 CFR 60.34b - Emission guidelines for municipal waste combustor operating practices.

Code of Federal Regulations, 2014 CFR

2014-07-01

... combustor operating practices. 60.34b Section 60.34b Protection of Environment ENVIRONMENTAL PROTECTION... September 20, 1994 § 60.34b Emission guidelines for municipal waste combustor operating practices. (a) For approval, a State plan shall include emission limits for carbon monoxide at least as protective as the...

40 CFR 60.34b - Emission guidelines for municipal waste combustor operating practices.

Code of Federal Regulations, 2012 CFR

2012-07-01

... combustor operating practices. 60.34b Section 60.34b Protection of Environment ENVIRONMENTAL PROTECTION... September 20, 1994 § 60.34b Emission guidelines for municipal waste combustor operating practices. (a) For approval, a State plan shall include emission limits for carbon monoxide at least as protective as the...

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.

Variable volume combustor with aerodynamic fuel flanges for nozzle mounting

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

McConnaughhay, Johnie Franklin; Keener, Christopher Paul; Johnson, Thomas Edward

2016-09-20

The present application provides a combustor for use with a gas turbine engine. The combustor may include a number of micro-mixer fuel nozzles and a fuel injection system for providing a flow of fuel to the micro-mixer fuel nozzles. The fuel injection system may include a number of support struts supporting the fuel nozzles and for providing the flow of fuel therethrough. The fuel injection system also may include a number of aerodynamic fuel flanges connecting the micro-mixer fuel nozzles and the support struts.

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.

System and method for reducing combustion dynamics and NO.sub.x in a combustor

DOEpatents

Uhm, Jong H.; Johnson, Thomas Edward

2015-11-20

A system for reducing combustion dynamics and NO.sub.x in a combustor includes a tube bundle that extends radially across at least a portion of the combustor, wherein the tube bundle comprises an upstream surface axially separated from a downstream surface. A shroud circumferentially surrounds the upstream and downstream surfaces. A plurality of tubes extends through the tube bundle from the upstream surface through the downstream surface, wherein the downstream surface is stepped to produce tubes having different lengths through the tube bundle. A method for reducing combustion dynamics and NO.sub.x in a combustor includes flowing a working fluid through a plurality of tubes radially arranged between an upstream surface and a downstream surface of an end cap that extends radially across at least a portion of the combustor, wherein the downstream surface is stepped.

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.

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.

An experimental study of the stable and unstable operation of an LPP gas turbine combustor

NASA Astrophysics Data System (ADS)

Dhanuka, Sulabh Kumar

A study was performed to better understand the stable operation of an LPP combustor and formulate a mechanism behind the unstable operation. A unique combustor facility was developed at the University of Michigan that incorporates the latest injector developed by GE Aircraft Engines and enables operation at elevated pressures with preheated air at flow-rates reflective of actual conditions. The large optical access has enabled the use of a multitude of state-of-the-art laser diagnostics such as PIV and PLIF, and has shed invaluable light not only into the GE injector specifically but also into gas turbine combustors in general. Results from Particle Imaging Velocimetry (PIV) have illustrated the role of velocity, instantaneous vortices, and key recirculation zones that are all critical to the combustor's operation. It was found that considerable differences exist between the iso-thermal and reacting flows, and between the instantaneous and mean flow fields. To image the flame, Planar Laser Induced Fluorescence (PLIF) of the formaldehyde radical was successfully utilized for the first time in a Jet-A flame. Parameters regarding the flame's location and structure have been obtained that assist in interpreting the velocity results. These results have also shown that some of the fuel injected from the main fuel injectors actually reacts in the diffusion flame of the pilot. The unstable operation of the combustor was studied in depth to obtain the stability limits of the combustor, behavior of the flame dynamics, and frequencies of the oscillations. Results from simultaneous pressure and high speed chemiluminescence images have shown that the low frequency dynamics can be characterized as flashback oscillations. The results have also shown that the stability of the combustor can be explained by simple and well established premixed flame stability mechanisms. This study has allowed the development of a model that describes the instability mechanism and accurately

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.

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.

Influence of the burner swirl on the azimuthal instabilities in an annular combustor

NASA Astrophysics Data System (ADS)

Mazur, Marek; Nygård, Håkon; Worth, Nicholas; Dawson, James

2017-11-01

Improving our fundamental understanding of thermoacoustic instabilities will aid the development of new low emission gas turbine combustors. In the present investigation the effects of swirl on the self-excited azimuthal combustion instabilities in a multi-burner annular annular combustor are investigated experimentally. Each of the burners features a bluff body and a swirler to stabilize the flame. The combustor is operated with an ethylene-air premixture at powers up to 100 kW. The swirl number of the burners is varied in these tests. For each case, dynamic pressure measurements at different azimuthal positions, as well as overhead imaging of OH* of the entire combustor are conducted simultaneously and at a high sampling frequency. The measurements are then used to determine the azimuthal acoustic and heat release rate modes in the chamber and to determine whether these modes are standing, spinning or mixed. Furthermore, the phase shift between the heat release rate and pressure and the shape of these two signals are analysed at different azimuthal positions. Based on the Rayleigh criterion, these investigations allow to obtain an insight about the effects of the swirl on the instability margins of the combustor. This project has received funding from the European Research Council (ERC) under the European Union's Horizon 2020 research and innovation programme (Grant agreement n° 677931 TAIAC).

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.

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.

CFD Modelling Applied to the Co-Combustion of Paper Sludge and Coal in a 130 t/h CFB Boiler

NASA Astrophysics Data System (ADS)

Yu, Z. S.; Ma, X. Q.; Lai, Z. Y.; Xiao, H. M.

Three-dimensional mathematical model has been developed as a tool for co-combustion of paper sludge and coal in a 130 tJh Circulating Fluidized Bed (CFB) boiler. Mathematical methods had been used based on a commercial software FLUENT for combustion. The predicted results of CFB furnace show that the co-combustion of paper sludge/coal is initially intensively at the bottom of bed; the temperature reaches its maximum in the dense-phase zone, around l400K. It indicates that paper sludge spout into furnace from the recycle inlet can increase the furnace maximum temperature (l396.3K), area-weighted average temperature (l109.6K) and the furnace gas outlet area-weighted average temperature(996.8K).The mathematical modeling also predicts that 15 mass% paper sludge co-combustion is the highest temperature at the flue gas outlet, it is 1000.8K. Moreover, it is proved that mathematical models can serve as a tool for detailed analysis of co-combustion of paper sludge and coal processes in a circulating fluidized bed furnace when in view of its convenience. The results gained from numerical simulation show that paper sludge enter into furnace from the recycle inlet excelled than mixing with coal and at the underside of phase interface.

Swept-Ramp Detonation Initiation Performance in a High-Pressure Pulse Detonation Combustor

DTIC Science & Technology

2010-12-01

conditions at sea level, but at elevated temperatures of 300–500°F in the combustor. The current work was motivated by a need to experimentally...The current work was motivated by a need to experimentally evaluate the detonation initiation performance of a PDC at elevated combustor pressures...High-Speed Propulsion Technologies (After [3]) .....................2 Figure 2. Stationary One-Dimensional Combustion Wave Model (From [7

CFD analyses of combustor and nozzle flowfields

NASA Astrophysics Data System (ADS)

Tsuei, Hsin-Hua; Merkle, Charles L.

1993-11-01

The objectives of the research are to improve design capabilities for low thrust rocket engines through understanding of the detailed mixing and combustion processes. A Computational Fluid Dynamic (CFD) technique is employed to model the flowfields within the combustor, nozzle, and near plume field. The computational modeling of the rocket engine flowfields requires the application of the complete Navier-Stokes equations, coupled with species diffusion equations. Of particular interest is a small gaseous hydrogen-oxygen thruster which is considered as a coordinated part of an ongoing experimental program at NASA LeRC. The numerical procedure is performed on both time-marching and time-accurate algorithms, using an LU approximate factorization in time, flux split upwinding differencing in space. The integrity of fuel film cooling along the wall, its effectiveness in the mixing with the core flow including unsteady large scale effects, the resultant impact on performance and the assessment of the near plume flow expansion to finite pressure altitude chamber are addressed.

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.

FEASIBILITY OF BURNING COAL IN CATALYTIC COMBUSTORS

EPA Science Inventory

The report gives results of a study, showing that pulverized coal can be burned in a catalytic combustor. Pulverized coal combustion in catalytic beds is markedly different from gaseous fuel combustion. Gas combustion gives uniform bed temperatures and reaction rates over the ent...

CFH haplotypes and ARMS2, C2, C3, and CFB alleles show association with susceptibility to age-related macular degeneration in Mexicans

PubMed Central

Zenteno, Juan Carlos; Fernández-López, Juan Carlos; Rodríguez-Corona, Ulises; Falfán-Valencia, Ramcés; Sebastian, Leticia; Morales, Fabiola; Ochoa-Contreras, Daniel; Carnevale, Alessandra; Silva-Zolezzi, Irma

2014-01-01

Purpose To evaluate the contribution of genetic variants of complement factor H (CFH), complement component 2 and 3 (C2 and C3), complement factor B (CFB), and age-related maculopathy susceptibility 2 (ARMS2) to age-related macular degeneration (AMD) risk in the Mexican Mestizo population. Methods Analysis included 282 unrelated Mexican patients with advanced AMD, 205 healthy controls, and 280 population controls. Stereoscopic fundus images were graded on the Clinical Age-Related Maculopathy System (CARMS). We designed a resequencing strategy using primers with M13 adaptor for the 23 exons of the CFH gene in a subgroup of 96 individuals clinically evaluated: 48 AMD cases and 48 age- and sex-matched healthy controls. Single nucleotide polymorphisms (SNPs) in C3 (Arg80Gly and Pro292Leu), C2 (rs547154), CFB (Leu9His), and ARMS2 (Ala69Ser) were genotyped in all patients, healthy and population controls using TaqMan assay. Results All evaluated individuals were Mexican Mestizos, and their genetic ancestry was validated using 224 ancestry informative markers and calculating Fst values. The CFH resequencing revealed 19 SNPs and a common variant in the intron 2 splice acceptor site; three CFH haplotypes inferred from individual genotypes, showed significant differences between cases and controls. The risk alleles in C3 (rs1047286, odds ratio [OR]=2.48, 95% confidence interval [CI]=1.64–3.75, p=1.59E-05; rs2230199, OR=2.15, 95% CI=1.48–3.13, p=6.28E-05) and in ARMS2 (rs10490924, OR=3.09, 95% CI=2.48–3.86, p=5.42E-23) were strongly associated with risk of AMD. The protective effect of alleles in C2 (rs547154) and CFB (rs4151667) showed a trend but was not significantly associated after correction for multiple testing. Conclusions Our results show that ARMS2 and C3 are major contributors to advanced AMD in Mexican patients, while the contributions of CFH, C2, and CFB are minor to those of other populations, reveling significant ethnic differences in minor allele

Discriminating between pyroxenite and peridotite sources for continental flood basalts (CFB) in southern Africa using olivine chemistry

NASA Astrophysics Data System (ADS)

Howarth, Geoffrey H.; Harris, Chris

2017-10-01

Continental Flood Basalts (CFB) result from voluminous outpourings of magma that often precede continental break-up. Notwithstanding the petrogenetic importance of CFBs, the nature of the mantle source for such magmas is contentious, particularly with regard to picrites with Ni-rich olivine phenocrysts. Previous studies have suggested that Ni-rich olivines associated with plume volcanism in regions of thickened (>90 km) lithosphere are related to either source mineralogy differences (peridotite versus pyroxenite) or change in olivine-melt partitioning due to pressure increase. In order to evaluate these two hypotheses, we present trace element data for olivines from the Karoo CFB Tuli and Mwenezi picrites and the Etendeka CFB Horingbaai/LTZ-L type picrites, all of which erupted in regions of thickened (>90 km) lithosphere in southern Africa. Karoo picrite olivines are Ni-rich, Ca- and Mn-poor, and have low (<1.4) 100*Mn/Fe. These compositions are consistent with a pyroxenitic source. Etendeka Horingbaai/LTZ-L picrite olivines do not show Ni-enrichment, but are characterized by high Al and Cr, and high (>1.4) 100*Mn/Fe, which is more consistent with high temperature melting of a dominantly peridotitic source. We also show that the Karoo and Etendeka olivines are characterized by distinct Mn/Zn ratios of <13 and >15, respectively. In addition, bulk rock geochemical data compilations and previously reported olivine δ18O for Karoo and Etendeka CFBs are discussed in order to further constrain source components based on previously described pyroxenite melt geochemical indices such as MgO-CaO systematics, FeO/MnO, Zn/Fe, and FC3MS (FeO/CaO-3*MgO/SiO2). These geochemical indices suggest a pyroxenite-dominated source for Karoo CFBs as well as for Etendeka ferropicrites whereas a peridotite-dominated source is indicated for Etendeka Horingbaai/LTZ-L type picrites analyzed in this study. Based on our data, Ni-enrichment of olivine in plume-related magmas in regions of

Steam Reformer With Fibrous Catalytic Combustor

NASA Technical Reports Server (NTRS)

Voecks, Gerald E.

1987-01-01

Proposed steam-reforming reactor derives heat from internal combustion on fibrous catalyst. Supplies of fuel and air to combustor controlled to meet demand for heat for steam-reforming reaction. Enables use of less expensive reactor-tube material by limiting temperature to value safe for material yet not so low as to reduce reactor efficiency.

Large eddy simulation of combustion characteristics in a kerosene fueled rocket-based combined-cycle engine combustor

NASA Astrophysics Data System (ADS)

Huang, Zhi-wei; He, Guo-qiang; Qin, Fei; Cao, Dong-gang; Wei, Xiang-geng; Shi, Lei

2016-10-01

This study reports combustion characteristics of a rocket-based combined-cycle engine combustor operating at ramjet mode numerically. Compressible large eddy simulation with liquid kerosene sprayed and vaporized is used to study the intrinsic unsteadiness of combustion in such a propulsion system. Results for the pressure oscillation amplitude and frequency in the combustor as well as the wall pressure distribution along the flow-path, are validated using experimental data, and they show acceptable agreement. Coupled with reduced chemical kinetics of kerosene, results are compared with the simultaneously obtained Reynolds-Averaged Navier-Stokes results, and show significant differences. A flow field analysis is also carried out for further study of the turbulent flame structures. Mixture fraction is used to determine the most probable flame location in the combustor at stoichiometric condition. Spatial distributions of the Takeno flame index, scalar dissipation rate, and heat release rate reveal that different combustion modes, such as premixed and non-premixed modes, coexisted at different sections of the combustor. The RBCC combustor is divided into different regions characterized by their non-uniform features. Flame stabilization mechanism, i.e., flame propagation or fuel auto-ignition, and their relative importance, is also determined at different regions in the combustor.

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.

Oil shale combustor model developed by Greek researchers

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

Not Available

1986-09-01

Work carried out in the Department of Chemical Engineering at the University of Thessaloniki, Thessaloniki, Greece has resulted in a model for the combustion of retorted oil shale in a fluidized bed combustor. The model is generally applicable to any hot-solids retorting process, whereby raw oil shale is retorted by mixing with a hot solids stream (usually combusted spent shale), and then the residual carbon is burned off the spent shale in a fluidized bed. Based on their modelling work, the following conclusions were drawn by the researchers. (1) For the retorted particle size distribution selected (average particle diameter 1600more » microns) complete carbon conversion is feasible at high pressures (2.7 atmosphere) and over the entire temperature range studied (894 to 978 K). (2) Bubble size was found to have an important effect, especially at conditions where reaction rates are high (high temperature and pressure). (3) Carbonate decomposition increases with combustor temperature and residence time. Complete carbon conversion is feasible at high pressures (2.7 atmosphere) with less than 20 percent carbonate decomposition. (4) At the preferred combustor operating conditions (high pressure, low temperature) the main reaction is dolomite decomposition while calcite decomposition is negligible. (5) Recombination of CO/sub 2/ with MgO occurs at low temperatures, high pressures, and long particle residence times.« less

Combustion stability with baffles, absorbers and velocity sensitive combustion. [liquid propellant rocket combustors

NASA Technical Reports Server (NTRS)

Mitchell, C. E.

1980-01-01

Analytical and computational techniques were developed to predict the stability behavior of liquid propellant rocket combustors using damping devices such as acoustic liners, slot absorbers, and injector face baffles. Models were developed to determine the frequency and decay rate of combustor oscillations, the spatial and temporal pressure waveforms, and the stability limits in terms of combustion response model parameters.

A three-dimensional algebraic grid generation scheme for gas turbine combustors with inclined slots

NASA Technical Reports Server (NTRS)

Yang, S. L.; Cline, M. C.; Chen, R.; Chang, Y. L.

1993-01-01

A 3D algebraic grid generation scheme is presented for generating the grid points inside gas turbine combustors with inclined slots. The scheme is based on the 2D transfinite interpolation method. Since the scheme is a 2D approach, it is very efficient and can easily be extended to gas turbine combustors with either dilution hole or slot configurations. To demonstrate the feasibility and the usefulness of the technique, a numerical study of the quick-quench/lean-combustion (QQ/LC) zones of a staged turbine combustor is given. Preliminary results illustrate some of the major features of the flow and temperature fields in the QQ/LC zones. Formation of co- and counter-rotating bulk flow and shape temperature fields can be observed clearly, and the resulting patterns are consistent with experimental observations typical of the confined slanted jet-in-cross flow. Numerical solutions show the method to be an efficient and reliable tool for generating computational grids for analyzing gas turbine combustors with slanted slots.

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.

Self-regulating fuel staging port for turbine combustor

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

Van Nieuwenhuizen, William F.; Fox, Timothy A.; Williams, Steven

2014-07-08

A port (60) for axially staging fuel and air into a combustion gas flow path 28 of a turbine combustor (10A). A port enclosure (63) forms an air path through a combustor wall (30). Fuel injectors (64) in the enclosure provide convergent fuel streams (72) that oppose each other, thus converting velocity pressure to static pressure. This forms a flow stagnation zone (74) that acts as a valve on airflow (40, 41) through the port, in which the air outflow (41) is inversely proportion to the fuel flow (25). The fuel flow rate is controlled (65) in proportion to enginemore » load. At high loads, more fuel and less air flow through the port, making more air available to the premixing assemblies (36).« less

Measurement of Turbulent Pressure and Temperature Fluctuations in a Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

Passaro, Andrea; LaGraff, John E.; Oldfield, Martin L. G.; Biagioni, Leonardo; Moss, Roger W.; Battelle, Ryan T.; Povinelli, Louis A. (Technical Monitor)

2003-01-01

The present research concerns the development of high-frequency pressure and temperature probes and related instrumentation capable of performing spectral characterization of unsteady pressure and temperature fluctuations over the 0.05 20 kHz range, at the exit of a gas turbine combustor operating at conditions close to nominal ones for large power generation turbomachinery. The probes used a transient technique pioneered at Oxford University; in order to withstand exposure to the harsh environment the probes were fitted on a rapid injection and cooling system jointly developed by Centrospazio CPR and Syracuse University. The experimental runs were performed on a large industrial test rig being operated by ENEL Produzione. The achieved results clearly show the satisfactory performance provided by this diagnostic tool, even though the poor location of the injection port prevented the tests from yielding more insight of the core flow turbulence characteristics. The pressure and temperature probes survived several dozen injections in the combustor hot jet, while consistently providing the intended high frequency performance. The apparatus was kept connected to the combustor during long duration firings, operating as an unobtrusive, self contained, piggy-back experiment: high frequency flow samplings were remotely recorded at selected moments corresponding to different combustor operating conditions.

Hydrogen injection scheme influence on flow structure in supersonic combustor of constant cross-section

NASA Astrophysics Data System (ADS)

Starov, A. V.; Goldfeld, M. A.

2017-10-01

The efficiency of using two variants of hydrogen injection (distributed and non-distributed injection from vertical pylons) is experimentally investigated. The tests are performed in the attached pipeline regime with the Mach number at the model combustor entrance M=2. The combustion chamber has a backward-facing step at the entrance and slotted channels for combustion stabilization. The tested variants of injection differ basically by the shapes of the fuel jets and, correspondingly, by the hydrogen distribution over the combustor. As a result, distributed injection is found to provide faster ignition, upstream displacement of the elevated pressure region, and more intense combustion over the entire combustor volume.

Performance prediction of a ducted rocket combustor

NASA Astrophysics Data System (ADS)

Stowe, Robert

2001-07-01

The ducted rocket is a supersonic flight propulsion system that takes the exhaust from a solid fuel gas generator, mixes it with air, and burns it to produce thrust. To develop such systems, the use of numerical models based on Computational Fluid Dynamics (CFD) is increasingly popular, but their application to reacting flow requires specific attention and validation. Through a careful examination of the governing equations and experimental measurements, a CFD-based method was developed to predict the performance of a ducted rocket combustor. It uses an equilibrium-chemistry Probability Density Function (PDF) combustion model, with a gaseous and a separate stream of 75 nm diameter carbon spheres to represent the fuel. After extensive validation with water tunnel and direct-connect combustion experiments over a wide range of geometries and test conditions, this CFD-based method was able to predict, within a good degree of accuracy, the combustion efficiency of a ducted rocket combustor.

Effect of broad properties fuel on injector performance in a reverse flow combustor

NASA Technical Reports Server (NTRS)

Raddlebaugh, S. M.; Norgren, C. T.

1983-01-01

The effect of fuel type on the performance of various fuel injectors was investigated in a reverse flow combustor. Combustor performance and emissions are documented for simplex pressure atomizing, spill flow, and airblast fuel injectors using a broad properties fuel and compared with performance using Jet A fuel. Test conditions simulated a range of flight conditions including sea level take off, low and high altitude cruise, as well as a parametric evaluation of the effect of increased combustor loading. The baseline simplex injector produced higher emission levels with corresponding lower combustion efficiency with the broad properties fuel. There was little or not loss in performance by the two advanced concept injectors with the broad properties fuel. The airblast injector proved to be especially insensitive to fuel type.

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.

Techniques for enhancing durability and equivalence ratio control in a rich-lean, three-stage ground power gas turbine combustor

NASA Technical Reports Server (NTRS)

Schultz, D. F.

1982-01-01

Rig tests of a can-type combustor were performed to demonstrate two advanced ground power engine combustor concepts: steam cooled rich-burn combustor primary zones for enhanced durability; and variable combustor geometry for three stage combustion equivalence ratio control. 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 offers the potential of both long life and reduced use of strategic materials for liner fabrication. Three degrees of variable geometry were successfully implemented to control airflow distribution within the combustor. One was a variable blade angle axial flow air swirler to control primary airflow while the other two consisted of rotating bands to control secondary and tertiary or dilution air flow.

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.

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

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.

Large eddy simulation of soot evolution in an aircraft combustor

NASA Astrophysics Data System (ADS)

Mueller, Michael E.; Pitsch, Heinz

2013-11-01

An integrated kinetics-based Large Eddy Simulation (LES) approach for soot evolution in turbulent reacting flows is applied to the simulation of a Pratt & Whitney aircraft gas turbine combustor, and the results are analyzed to provide insights into the complex interactions of the hydrodynamics, mixing, chemistry, and soot. The integrated approach includes detailed models for soot, combustion, and the unresolved interactions between soot, chemistry, and turbulence. The soot model is based on the Hybrid Method of Moments and detailed descriptions of soot aggregates and the various physical and chemical processes governing their evolution. The detailed kinetics of jet fuel oxidation and soot precursor formation is described with the Radiation Flamelet/Progress Variable model, which has been modified to account for the removal of soot precursors from the gas-phase. The unclosed filtered quantities in the soot and combustion models, such as source terms, are closed with a novel presumed subfilter PDF approach that accounts for the high subfilter spatial intermittency of soot. For the combustor simulation, the integrated approach is combined with a Lagrangian parcel method for the liquid spray and state-of-the-art unstructured LES technology for complex geometries. Two overall fuel-to-air ratios are simulated to evaluate the ability of the model to make not only absolute predictions but also quantitative predictions of trends. The Pratt & Whitney combustor is a Rich-Quench-Lean combustor in which combustion first occurs in a fuel-rich primary zone characterized by a large recirculation zone. Dilution air is then added downstream of the recirculation zone, and combustion continues in a fuel-lean secondary zone. The simulations show that large quantities of soot are formed in the fuel-rich recirculation zone, and, furthermore, the overall fuel-to-air ratio dictates both the dominant soot growth process and the location of maximum soot volume fraction. At the higher fuel

Fuel injection assembly for gas turbine engine combustor

NASA Technical Reports Server (NTRS)

Candy, Anthony J. (Inventor); Glynn, Christopher C. (Inventor); Barrett, John E. (Inventor)

2002-01-01

A fuel injection assembly for a gas turbine engine combustor, including at least one fuel stem, a plurality of concentrically disposed tubes positioned within each fuel stem, wherein a cooling supply flow passage, a cooling return flow passage, and a tip fuel flow passage are defined thereby, and at least one fuel tip assembly connected to each fuel stem so as to be in flow communication with the flow passages, wherein an active cooling circuit for each fuel stem and fuel tip assembly is maintained by providing all active fuel through the cooling supply flow passage and the cooling return flow passage during each stage of combustor operation. The fuel flowing through the active cooling circuit is then collected so that a predetermined portion thereof is provided to the tip fuel flow passage for injection by the fuel tip assembly.

Primary zone dynamics in a gas turbine combustor

NASA Technical Reports Server (NTRS)

Sullivan, J. P.; Barron, D.; Seal, M.; Morgan, D.; Murthy, S. N. B.

1989-01-01

Fluid mechanical investigations simulating the flow in the primary zone of a gas turbine combustor are presented using three generic test rigs: (1) rotating pipe yielding a swirling jet of air; (2) primary zone model with a single swirler and various primary jet configurations, operated with air; and (3) two rectangular models of a (stretched-out) annular combustor with five swirlers in the backwall and with various primary jet configurations, one operated with air and the other with water. Concentration measurements are obtained using laser sheet imaging techniques and velocity measurements using a laser Doppler velocimeter. The results show recirculation zones, intense mixing, instabilities of the interacting jets and the presence of large random vortical motions. The flowfields are shown to exhibit bimodal behavior, have asymmetries despite symmetrical geometry and inlet conditions and display strong jet/swirler and swirler/swirler interactions.

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.

Concentric catalytic combustor

DOEpatents

Bruck, Gerald J [Oviedo, FL; Laster, Walter R [Oviedo, FL

2009-03-24

A catalytic combustor (28) includes a tubular pressure boundary element (90) having a longitudinal flow axis (e.g., 56) separating a first portion (94) of a first fluid flow (e.g., 24) from a second portion (95) of the first fluid flow. The pressure boundary element includes a wall (96) having a plurality of separate longitudinally oriented flow paths (98) annularly disposed within the wall and conducting respective portions (100, 101) of a second fluid flow (e.g., 26) therethrough. A catalytic material (32) is disposed on a surface (e.g., 102, 103) of the pressure boundary element exposed to at least one of the first and second portions of the first fluid flow.

Combustion efficiency determined from wall pressure and temperature measurement in a Mach 2 combustor

NASA Technical Reports Server (NTRS)

Segal, Corin; Mcdaniel, James C.; Whitehurst, Robert B.; Krauss, Roland H.

1991-01-01

A study of transverse hydrogen injection behind a rearward facing step in a Mach 2 airflow was conducted to determine the combustion efficiency and the combustor/inlet interactions at the low temperature lean-mixture operational end of a scramjet combustor model. The fuel was injected at sonic conditions into the electrically heated airstream, which was maintained at 850 K or below. The static pressure delivered at the entrance of the combustor ranged between 0.25 to 0.5 atm. Injector configurations included single and staged injectors placed at 3 or 3-and-7 step-heights downstream of the step, respectively, with injector diameters of 1, 1.5, and 2 mm. Ignition was achieved by initially unstarting the test section. The constant area combustor and the low initial temperatures caused thermal choking and upstream interaction to occur at very low equivalence ratios. Typically, most of the fuel was burned in the recirculation region behind the step and around the jets. The effects of initial conditions (temperature and pressure), fuel-to-air dynamic pressure ratio, and boundaries (thermal vs adiabatic) are presented.

Assessment of swirl spray interaction in lab scale combustor using time-resolved measurements

NASA Astrophysics Data System (ADS)

Rajamanickam, Kuppuraj; Jain, Manish; Basu, Saptarshi

2017-11-01

Liquid fuel injection in highly turbulent swirling flows becomes common practice in gas turbine combustors to improve the flame stabilization. It is well known that the vortex bubble breakdown (VBB) phenomenon in strong swirling jets exhibits complicated flow structures in the spatial domain. In this study, the interaction of hollow cone liquid sheet with such coaxial swirling flow field has been studied experimentally using time-resolved measurements. In particular, much attention is focused towards the near field breakup mechanism (i.e. primary atomization) of liquid sheet. The detailed swirling gas flow field characterization is carried out using time-resolved PIV ( 3.5 kHz). Furthermore, the complicated breakup mechanisms and interaction of the liquid sheet are imaged with the help of high-speed shadow imaging system. Subsequently, proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) is implemented over the instantaneous data sets to retrieve the modal information associated with the interaction dynamics. This helps to delineate more quantitative nature of interaction process between the liquid sheet and swirling gas phase flow field.

Compact Laser-Based Sensors for Monitoring and Control of Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Hanson, Ronald K.; Jeffries, Jay B.

2003-01-01

Research is reported on the development of sensors for gas turbine combustor applications that measure real-time gas temperature using near-infrared water vapor absorption and concentration in the combustor exhaust of trace quantities of pollutant NO and CO using mid-infrared absorption. Gas temperature is extracted from the relative absorption strength of two near-infrared transitions of water vapor. From a survey of the water vapor absorption spectrum, two overtone transitions near 1800 nm were selected that can be rapidly scanned in wavelength by injection current tuning a single DFB diode laser. From the ratio of the absorbances on these selected transitions, a path-integrated gas temperature can be extracted in near-real time. Demonstration measurements with this new temperature sensor showed that combustor instabilities could be identified in the power spectrum of the temperature versus time record. These results suggest that this strategy is extremely promising for gas turbine combustor control applications. Measurements of the concentration of NO and CO in the combustor exhaust are demonstrated with mid-infrared transitions using thermo-electrically cooled, quantum cascade lasers operating near 5.26 and 4.62 microns respectively. Measurements of NO are performed in an insulated exhaust duct of a C2H4-air flame at temperatures of approximately 600 K. CO measurements are performed above a rich H2-air flame seeded with CO2 and cooled with excess N2 to 1150 K. Using a balanced ratiometric detection technique a sensitivity of 0.36 ppm-m was achieved for NO and 0.21 ppm-m for CO. Comparisons between measured and predicted water-vapor and CO2 interference are discussed. The mid-infrared laser quantum cascade laser technology is in its infancy; however, these measurements demonstrate the potential for pollutant monitoring in exhaust gases with mid-IR laser absorption.

Combustion Noise at Elevated Pressures in a Liquid-Fueled Premixed Combustor

NASA Technical Reports Server (NTRS)

Darling, Douglas; Radhakrishnan, Krishnan; Oyediran, Ayo

1997-01-01

Noise generated in gas turbine combustors can exist in several forms-broadband noise, sharp resonant peaks, and regular or intermittent nonlinear pulsing. In the present study, dynamic pressure measurements were made in several JP-5-fueled combustor configurations, at various mean pressures and temperatures. The fluctuating pressure was measured at mean pressures from 6 to 14 atm and inlet temperatures from 550 K to 850 K. The goal of the present work was to study the effect of changes in mean flow conditions on combustor noise: both broadband noise and sharp tones were considered. In general, the shape of the broadband noise spectrum was consistent from one configuration to another. The shape of the spectrum was influenced by the acoustic filtering of the combustion zone. This filtering ensured the basic consistency of the spectra. In general, the trends in broadband noise observed at low mean pressures were also seen at high mean pressures; that is, the total sound level decreased with both increasing equivalence ratio and increasing inlet temperature. The combustor configurations without a central pilot experienced higher broadband noise levels and were more susceptible to narrow peak resonances than configurations with a central pilot. The sharp peaks were more sensitive to the mean flow than was the broadband noise, and the effects were not always the same. In some situations, increasing the equivalence ratio made the sharp peaks grow, while at other conditions, increasing the equivalence ratio made the sharp peaks shrink. Thus, it was difficult to predict when resonances would occur; however, they were reproducible. Acoustic coupling between the upstream and downstream regions of the combustor may play a role in the sharp-peaked oscillations. Noise was also observed near lean blow out. As with other types of noise, lean blow out noise was affected by the combustion chamber acoustics, which apparently maintains the fluctuations at a uniform frequency. However

Reliable and Affordable Control Systems Active Combustor Pattern Factor Control

NASA Technical Reports Server (NTRS)

McCarty, Bob; Tomondi, Chris; McGinley, Ray

2004-01-01

Active, closed-loop control of combustor pattern factor is a cooperative effort between Honeywell (formerly AlliedSignal) Engines and Systems and the NASA Glenn Research Center to reduce emissions and turbine-stator vane temperature variations, thereby enhancing engine performance and life, and reducing direct operating costs. Total fuel flow supplied to the engine is established by the speed/power control, but the distribution to individual atomizers will be controlled by the Active Combustor Pattern Factor Control (ACPFC). This system consist of three major components: multiple, thin-film sensors located on the turbine-stator vanes; fuel-flow modulators for individual atomizers; and control logic and algorithms within the electronic control.

Techno-economic assessment of a hybrid solar receiver and combustor

NASA Astrophysics Data System (ADS)

Lim, Jin Han; Nathan, Graham; Dally, Bassam; Chinnici, Alfonso

2016-05-01

A techno-economic analysis is performed to compare two different configurations of hybrid solar thermal systems with fossil fuel backup to provide continuous electricity output. The assessment compares a Hybrid Solar Receiver Combustor (HSRC), in which the functions of a solar cavity receiver and a combustor are integrated into a single device with a reference conventional solar thermal system using a regular solar cavity receiver with a backup boiler, termed the Solar Gas Hybrid (SGH). The benefits of the integration is assessed by varying the size of the storage capacity and heliostat field while maintaining the same overall thermal input to the power block.

Wave combustors for trans-atmospheric vehicles

NASA Technical Reports Server (NTRS)

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

1989-01-01

A performance analysis is given of a conceptual transatmospheric vehicle (TAV). The TAV is powered by a an oblique detonation wave engine (ODWE). The ODWE is an airbreathing hypersonic propulsion system which utilizes shock and detonation waves to enhance fuel-air mixing and combustion in supersonic flow. In this wave combustor concept, an oblique shock wave in the combustor can act as a flameholder by increasing the pressure and temperature of the air-fuel mixture, thereby decreasing the ignition delay. If the oblique shock is sufficiently strong, then the combustion front and the shock wave can couple into a detonation wave. In this case, combustion occurs almost instantaneously in a thin zone behind the wave front. The result is a shorter lighter engine compared to the scramjet. The ODWE-powered hypersonic vehicle performance is compared to that of a scramjet-powered vehicle. Among the results outlined, it is found that the ODWE trades a better engine performance above Mach 15 for a lower performance below Mach 15. The overall higher performance of the ODWE results in a 51,000-lb weight savings and a higher payload weight fraction of approximately 12 percent.

Dilution jet configurations in a reverse flow combustor. M.S. Thesis Final Report

NASA Technical Reports Server (NTRS)

Zizelman, J.

1985-01-01

Results of measurements of both temperature and velocity fields within a reverse flow combustor are presented. Flow within the combustor is acted upon by perpendicularly injected cooling jets introduced at three different locations along the inner and outer walls of the combustor. Each experiment is typified by a group of parameters: density ratio, momentum ratio, spacing ratio, and confinement parameter. Measurements of both temperature and velocity are presented in terms of normalized profiles at azimuthal positions through the turn section of the combustion chamber. Jet trajectories defined by minimum temperature and maximum velocity give a qualitative indication of the location of the jet within the cross flow. Results of a model from a previous temperature study are presented in some of the plots of data from this work.

Numerical prediction of turbulent flame stability in premixed/prevaporized (HSCT) combustors

NASA Technical Reports Server (NTRS)

Winowich, Nicholas S.

1990-01-01

A numerical analysis of combustion instabilities that induce flashback in a lean, premixed, prevaporized dump combustor is performed. KIVA-II, a finite volume CFD code for the modeling of transient, multidimensional, chemically reactive flows, serves as the principal analytical tool. The experiment of Proctor and T'ien is used as a reference for developing the computational model. An experimentally derived combustion instability mechanism is presented on the basis of the observations of Proctor and T'ien and other investigators of instabilities in low speed (M less than 0.1) dump combustors. The analysis comprises two independent procedures that begin from a calculated stable flame: The first is a linear increase of the equivalence ratio and the second is the linear decrease of the inflow velocity. The objective is to observe changes in the aerothermochemical features of the flow field prior to flashback. It was found that only the linear increase of the equivalence ratio elicits a calculated flashback result. Though this result did not exhibit large scale coherent vortices in the turbulent shear layer coincident with a flame flickering mode as was observed experimentally, there were interesting acoustic effects which were resolved quite well in the calculation. A discussion of the k-e turbulence model used by KIVA-II is prompted by the absence of combustion instabilities in the model as the inflow velocity is linearly decreased. Finally, recommendations are made for further numerical analysis that may improve correlation with experimentally observed combustion instabilities.

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.

A Comparison of Combustor-Noise Models

NASA Technical Reports Server (NTRS)

Hultgren, Lennart, S.

2012-01-01

The current status of combustor-noise prediction in the NASA Aircraft Noise Prediction Program (ANOPP) for current-generation (N) turbofan engines is summarized. Best methods for near-term updates are reviewed. Long-term needs and challenges for the N+1 through N+3 timeframe are discussed. This work was carried out under the NASA Fundamental Aeronautics Program, Subsonic Fixed Wing Project, Quiet Aircraft Subproject.

The experimental clean combustor program: Description and status to November 1975

NASA Technical Reports Server (NTRS)

Niedzwiecki, R. W.

1975-01-01

The generation of technology was studied for the development of advanced commercial CTOL aircraft engines with lower exhaust emissions than current aircraft. The program is in three phases. Phase 1, already completed, consisted of screening tests of low pollution combustor concepts. Phase 2, currently in progress, consists of test rig refinement of the most promising combustor concepts. Phase 2 test results are reported. Phase 3, also currently in progress, consists of incorporating and evaluating the best combustors as part of a complete engine. Engine test plans and pollution sampling techniques are described in this report. Program pollution goals, specified at engine idle and take-off conditions, are idle emission index value of 20 and 4 for carbon monoxide (CO) and total unburned hydrocarbons (THC), respectively, and at take-off are an oxides of nitrogen (NOx) emission index level of 10 and a smoke number of 15. Pollution data were obtained at all engine operating conditions. Results are presented in terms of emission index and also in terms of the Environmental Protection Agency's 1979 Standards Parameter.

Small gas turbine combustor study - Fuel injector performance in a transpiration-cooled liner

NASA Technical Reports Server (NTRS)

Riddlebaugh, S. M.; Norgren, C. T.

1985-01-01

The effect of fuel injection technique on the performance of an advanced reverse flow combustor liner constructed of Lamilloy (a multilaminate transpiration type material) was determined. Performance and emission levels are documented over a range of simulated flight conditions using simplex pressure atomizing, spill return, and splash cone airblast injectors. A parametric evaluation of the effect of increased combustor loading with each of the fuel injector types is obtained.

Small gas turbine combustor study: Fuel injector performance in a transpiration-cooled liner

NASA Technical Reports Server (NTRS)

Riddlebaugh, S. M.; Norgren, C. T.

1985-01-01

The effect of fuel injection technique on the performance of an advanced reverse flow combustor liner constructed of Lamilloy (a multilaminate transpiration type material) was determined. Performance and emission levels are documented over a range of simulated flight conditions using simplex pressure atomizing, spill return, and splash cone airblast injectors. A parametric evaluation of the effect of increased combustor loading with each of the fuel injector types is obtained.

Advanced Low Emissions Subsonic Combustor Study

NASA Technical Reports Server (NTRS)

Smith, Reid

1998-01-01

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

CFD Investigation of Pollutant Emission in Can-Type Combustor Firing Natural Gas, LNG and Syngas

NASA Astrophysics Data System (ADS)

Hasini, H.; Fadhil, SSA; Mat Zian, N.; Om, NI

2016-03-01

CFD investigation of flow, combustion process and pollutant emission using natural gas, liquefied natural gas and syngas of different composition is carried out. The combustor is a can-type combustor commonly used in thermal power plant gas turbine. The investigation emphasis on the comparison of pollutant emission such in particular CO2, and NOx between different fuels. The numerical calculation for basic flow and combustion process is done using the framework of ANSYS Fluent with appropriate model assumptions. Prediction of pollutant species concentration at combustor exit shows significant reduction of CO2 and NOx for syngas combustion compared to conventional natural gas and LNG combustion.

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.

Investigations on the Influence of the In-Stream Pylon and Strut on the Performance of a Scramjet Combustor

PubMed Central

Liu, Weidong; Sun, Mingbo

2014-01-01

The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight. PMID:25254234

Numerical Modeling of Gas Turbine Combustor Utilizing One-Dimensional Acoustics

NASA Astrophysics Data System (ADS)

Caley, Thomas M.

This study focuses on the numerical modeling of a gas turbine combustor set-up with known regions of thermoacoustic instability. The proposed model takes the form of a hybrid thermoacoustic network, with lumped elements representing boundary conditions and the flame, and 3-dimensional geometry volumes representing the geometry. The model is analyzed using a commercial 3-D finite element method (FEM) software, COMSOL Multiphysics. A great deal of literature is available covering thermoacoustic modeling, but much of it utilizes more computationally expensive techniques such as Large-Eddy Simulations, or relies on analytical modeling that is limited to specific test cases or proprietary software. The present study models the 3-D geometry of a high-pressure combustion chamber accurately, and uses the lumped elements of a thermoacoustic network to represent parts of the combustor system that can be experimentally tested under stable conditions, ensuring that the recorded acoustic responses can be attributed to that element alone. The numerical model has been tested against the experimental model with and without an experimentally-determined impedance boundary condition. Eigenfrequency studies are used to compare the frequency and growth rates (and from that, the thermoacoustic stability) of resonant modes in the combustor. The flame in the combustor is modeled with a flame transfer function that was determined from experimental testing using frequency forcing. The effect of flow rate on the impedance boundary condition is also examined experimentally and numerically to qualify the practice of modeling an orifice plate as an acoustically-closed boundary. Using the experimental flame transfer function and boundary conditions in the numerical model produced results that closely matched previous experimental tests in frequency, but not in stability characteristics. The lightweight nature of the numerical model means additional lumped elements can be easily added when

Combustion Dynamic Characteristics Identification in a 9-point LDI Combustor Under Choked Outlet Boundary Conditions

NASA Technical Reports Server (NTRS)

He, Zhuohui J.; Chang, Clarence T.

2017-01-01

Combustion dynamics data were collected at the NASA Glenn Research Center's CE-5 flame tube test facility under combustor outlet choked conditions. Two 9-point Swirl-Venturi Lean Direct Injection (SV-LDI) configurations were tested in a rectangular cuboid combustor geometry. Combustion dynamic data were measured at different engine operational conditions up to inlet air pressure and temperature of 24.13 bar and 828 K, respectively. In this study, the effects of acoustic cavity resonance, precessing vortex core (PVC), and non-uniform thermal expansion on the dynamic noise spectrum are identified by comparing the dynamic data that collected at various combustor inlet conditions along with combustor geometric calculations. The results show that the acoustic cavity resonance noises were seen in the counter-rotating pilot configuration but not in the co-rotating pilot configuration. Dynamic pressure noise band at around 0.9 kHz was only detected at the P'41 location (9.8 cm after fuel injector face) but not at the P'42 location (29 cm after the fuel injector face); the amplitude of this noise band depended on the thermal expansion ratio (T4/T3). The noise band at around 1.8 kHz was found to depend on the inlet air pressure or the air density inside the combustor. The PVC frequency was not observed in these two configurations.

RQL Sector Rig Testing of SiC/SiC Combustor Liners

NASA Technical Reports Server (NTRS)

Verrilli, Michael J.; Martin, Lisa C.; Brewer, David N.

2002-01-01

Combustor liners, manufactured from silicon carbide fiber-reinforced silicon carbide (SiC/SiC) were tested for 260 hr using a simulated gas turbine engine cycle. This report documents the results of the last 56 hr of testing. Damage occurred in one of the six different components that make up the combustor liner set, the rich zone liner. Cracks in the rich zone liner initiated at the leading edge due to stresses resulting from the component attachment configuration. Thin film thermocouples and fiber optic pyrometers were used to measure the rich zone liner's temperature and these results are reported.

Apparatus and filtering systems relating to combustors in combustion turbine engines

DOEpatents

Johnson, Thomas Edward [Greer, SC; Zuo, Baifang [Simpsonville, SC; Stevenson, Christian Xavier [Inman, SC

2012-07-24

A combustor for a combustion turbine engine, the combustor that includes: a chamber defined by an outer wall and forming a channel between windows defined through the outer wall toward a forward end of the chamber and at least one fuel injector positioned toward an aft end of the chamber; a screen; and a standoff comprising a raised area on an outer surface of the outer wall near the periphery of the windows; wherein the screen extends over the windows and is supported by the standoff in a raised position in relation to the outer surface of the outer wall and the windows.

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.

Error Reduction Program. [combustor performance evaluation codes

NASA Technical Reports Server (NTRS)

Syed, S. A.; Chiappetta, L. M.; Gosman, A. D.

1985-01-01

The details of a study to select, incorporate and evaluate the best available finite difference scheme to reduce numerical error in combustor performance evaluation codes are described. The combustor performance computer programs chosen were the two dimensional and three dimensional versions of Pratt & Whitney's TEACH code. The criteria used to select schemes required that the difference equations mirror the properties of the governing differential equation, be more accurate than the current hybrid difference scheme, be stable and economical, be compatible with TEACH codes, use only modest amounts of additional storage, and be relatively simple. The methods of assessment used in the selection process consisted of examination of the difference equation, evaluation of the properties of the coefficient matrix, Taylor series analysis, and performance on model problems. Five schemes from the literature and three schemes developed during the course of the study were evaluated. This effort resulted in the incorporation of a scheme in 3D-TEACH which is usuallly more accurate than the hybrid differencing method and never less accurate.

Mercury emissions during cofiring of sub-bituminous coal and biomass (chicken waste, wood, coffee residue, and tobacco stalk) in a laboratory-scale fluidized bed combustor

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

Yan Cao; Hongcang Zhou; Junjie Fan

Four types of biomass (chicken waste, wood pellets, coffee residue, and tobacco stalks) were cofired at 30 wt % with a U.S. sub-bituminous coal (Powder River Basin Coal) in a laboratory-scale fluidized bed combustor. A cyclone, followed by a quartz filter, was used for fly ash removal during tests. The temperatures of the cyclone and filter were controlled at 250 and 150{sup o}C, respectively. Mercury speciation and emissions during cofiring were investigated using a semicontinuous mercury monitor, which was certified using ASTM standard Ontario Hydra Method. Test results indicated mercury emissions were strongly correlative to the gaseous chlorine concentrations, butmore » not necessarily correlative to the chlorine contents in cofiring fuels. Mercury emissions could be reduced by 35% during firing of sub-bituminous coal using only a quartz filter. Cofiring high-chlorine fuel, such as chicken waste (Cl = 22340 wppm), could largely reduce mercury emissions by over 80%. When low-chlorine biomass, such as wood pellets (Cl = 132 wppm) and coffee residue (Cl = 134 wppm), is cofired, mercury emissions could only be reduced by about 50%. Cofiring tobacco stalks with higher chlorine content (Cl = 4237 wppm) did not significantly reduce mercury emissions. Gaseous speciated mercury in flue gas after a quartz filter indicated the occurrence of about 50% of total gaseous mercury to be the elemental mercury for cofiring chicken waste, but occurrence of above 90% of the elemental mercury for all other cases. Both the higher content of alkali metal oxides or alkali earth metal oxides in tested biomass and the occurrence of temperatures lower than 650{sup o}C in the upper part of the fluidized bed combustor seemed to be responsible for the reduction of gaseous chlorine and, consequently, limited mercury emissions reduction during cofiring. 36 refs., 3 figs. 1 tab.« less

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.

N+2 Advanced Low NOx Combustor Technology Final Report

NASA Technical Reports Server (NTRS)

Herbon, John; Aicholtz, John; Hsieh, Shih-Yang; Viars, Philip; Birmaher, Shai; Brown, Dan; Patel, Nayan; Carper, Doug; Cooper, Clay; Fitzgerald, Russell

2017-01-01

In accordance with NASAs technology goals for future subsonic vehicles, this contract identified and developed new combustor concepts toward meeting N+2 generation (2020) LTO (landing and take-off) NOx emissions reduction goal of 75 from the standard adopted at Committee on Aviation Environmental Protection 6 (CAEP6). Based on flame tube emissions, operability, and autoignition testing, one concept was down selected for sector testing at NASA. The N+2 combustor sector successfully demonstrated 75 reduction for LTO NOx (vs. CAEP6) and cruise NOx (vs. 2005 B777-200 reference) while maintaining 99.9 cruise efficiency and no increase in CO and HC emissions.The program also developed enabling technologies for the combustion system including ceramic matrix composites (CMC) liner materials, active combustion control concepts, and laser ignition for improved altitude relight.

Development and testing of pulsed and rotating detonation combustors

NASA Astrophysics Data System (ADS)

St. George, Andrew C.

Detonation is a self-sustaining, supersonic, shock-driven, exothermic reaction. Detonation combustion can theoretically provide significant improvements in thermodynamic efficiency over constant pressure combustion when incorporated into existing cycles. To harness this potential performance benefit, countless studies have worked to develop detonation combustors and integrate these devices into existing systems. This dissertation consists of a series of investigations on two types of detonation combustors: the pulse detonation combustor (PDC) and the rotating detonation combustor (RDC). In the first two investigations, an array of air-breathing PDCs is integrated with an axial power turbine. The system is initially operated with steady and pulsed cold air flow to determine the effect of pulsed flow on turbine performance. Various averaging approaches are employed to calculate turbine efficiency, but only flow-weighted (e.g., mass or work averaging) definitions have physical significance. Pulsed flow turbine efficiency is comparable to steady flow efficiency at high corrected flow rates and low rotor speeds. At these conditions, the pulse duty cycle expands and the variation of the rotor incidence angle is constrained to a favorable range. The system is operated with pulsed detonating flow to determine the effect of frequency, fill fraction, and rotor speed on turbine performance. For some conditions, output power exceeds the maximum attainable value from steady constant pressure combustion due to a significant increase in available power from the detonation products. However, the turbine component efficiency estimated from classical thermodynamic analysis is four times lower than the steady design point efficiency. Analysis of blade angles shows a significant penalty due to the detonation, fill, and purge processes simultaneously imposed on the rotor. The latter six investigations focus on fundamental research of the RDC concept. A specially-tailored RDC data

Fuel property effects in stirred combustors

NASA Technical Reports Server (NTRS)

1980-01-01

Soot formation in strongly backmixed combustion was investigated using the jet-stirred combustor (JSC). This device provided a combustion volume in which temperature and combustion were uniform. It simulated the recirculating characteristics of the gas turbine primary zone; it was in this zone where mixture conditions were sufficiently rich to produce soot. Results indicate that the JSC allows study of soot formation in an aerodynamic situation revelant to gas turbines.

Analytical Fuel Property Effects-Small Combustors

DTIC Science & Technology

1984-10-01

Research Center Contract NAS 3-23165 ,_ _---, NASA CR-174738 AVSCOM TR-84-C-14 EDR 11683 Analytical Fuel Property Effects-- Small Combustors by R. D...National Aeronautics and Space Administration NASA Lewis Research Center Contract NAS 3-23165 TABLE OF CONTENTS Section Title ft Summary...would not sustain a flame at normal operating conditions. Frivate communication between Gary Seng of the NASA Lewis Research Center fuel labs and R. D

Pulsed atmospheric fluidized bed combustor apparatus

DOEpatents

Mansour, Momtaz N.

1993-10-26

A pulsed atmospheric fluidized bed reactor system is disclosed and claimed along with a process for utilization of same for the combustion of, e.g. high sulfur content coal. The system affords a economical, ecologically acceptable alternative to oil and gas fired combustors. The apparatus may also be employed for endothermic reaction, combustion of waste products, e.g., organic and medical waste, drying materials, heating air, calcining and the like.

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.

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.

Optimization of operation conditions and configurations for solid-propellant ducted rocket combustors

NASA Astrophysics Data System (ADS)

Onn, Shing-Chung; Chiang, Hau-Jei; Hwang, Hang-Che; Wei, Jen-Ko; Cherng, Dao-Lien

1993-06-01

The dynamic behavior of a 2D turbulent mixing and combustion process has been studied numerically in the main combustion chamber of a solid-propellant ducted rocket (SDR). The mathematical model is based on the Favre-averaged conservation equations developed by Cherng (1990). Combustion efficiency, rather than specific impulse from earlier studies, is applied successfully to optimize the effects of two parameters by a multiple linear regression model. Specifically, the fuel-air equivalence ratio of the operating conditions and the air inlet location of configurations for the SDR combustor have been studied. For a equivalence ratio near the stoichiometric condition, the use of specific impulse or combustion efficiency will show similar trend in characterizing the reacting flow field in the combustor. For the overall fuel lean operating conditions, the change of combustion efficiency is much more sensitive to that of air inlet location than specific impulse does, suggesting combustion efficiency a better property than specific impulse in representing the condition toward flammability limits. In addition, the air inlet for maximum efficiency, in general, appears to be located at downstream of that for highest specific impulse. The optimal case for the effects of two parameters occurs at fuel lean condition, which shows a larger recirculation zone in front, deeper penetration of ram air into the combustor and much larger high temperature zone near the centerline of the combustor exit than those shown in the optimal case for overall equivalence ratio close to stoichiometric.

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.

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.

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.

Flame Stability in a Trapped-Vortex Spray-Combustor

NASA Astrophysics Data System (ADS)

Chakka, P.; Mancilla, P. C.; Acharya, S.

1999-11-01

Flame stabilization mechanisms in a Trapped-Vortex (TV) cavity is investigated experimentally and computationally in the current research. The TV-cavity is placed coaxially in the combustor and the flame is maintained through injection of liquid fuel spray and air from the inside face of the afterbody. This concept was introduced by Roquemore and company of Wright-Patterson AFB for gaseous fuel injection into the cavity and is extended for liquid fuel sprays in the current research. The flame holding capability of the TV-cavity is studied for different equivalence ratios of the secondary injection and overall Lean Blow-Out (LBO) limits are presented for different primary and secondary flow rates. The interaction and mixing of the main flow with the secondary vortex flow is investigated through the Laser Doppler Velocimetry measurements taken through a quartz window near the cavity. Also, temperature distribution through IR measurements and pressure fluctuations inside the chamber are presented for complete performance analysis of the TV cavity combustor.

Simulations of NOx Emissions from Low Emissions Discrete Jet Injector Combustor Tests

NASA Technical Reports Server (NTRS)

Ajmani, Kumud; Breisacher, Kevin

2014-01-01

An experimental and computational study was conducted to evaluate the performance and emissions characteristics of a candidate Lean Direct Injection (LDI) combustor configuration with a mix of simplex and airblast injectors. The National Combustion Code (NCC) was used to predict the experimentally measured EINOx emissions for test conditions representing low power, medium power, and high-power engine cycle conditions. Of the six cases modeled with the NCC using a reduced-kinetics finite-rate mechanism and lagrangian spray modeling, reasonable predictions of combustor exit temperature and EINOx were obtained at two high-power cycle conditions.

Active Suppression of Instabilities in Engine Combustors

NASA Technical Reports Server (NTRS)

Kopasakis, George

2004-01-01

A method of feedback control has been proposed as a means of suppressing thermo-acoustic instabilities in a liquid- fueled combustor of a type used in an aircraft engine. The basic principle of the method is one of (1) sensing combustor pressure oscillations associated with instabilities and (2) modulating the rate of flow of fuel to the combustor with a control phase that is chosen adaptively so that the pressure oscillations caused by the modulation oppose the sensed pressure oscillations. The need for this method arises because of the planned introduction of advanced, lean-burning aircraft gas turbine engines, which promise to operate with higher efficiencies and to emit smaller quantities of nitrogen oxides, relative to those of present aircraft engines. Unfortunately, the advanced engines are more susceptible to thermoacoustic instabilities. These instabilities are hard to control because they include large dead-time phase shifts, wide-band noise characterized by amplitudes that are large relative to those of the instabilities, exponential growth of the instabilities, random net phase walks, and amplitude fluctuations. In this method (see figure), the output of a combustion-pressure sensor would be wide-band-pass filtered and then further processed to generate a control signal that would be applied to a fast-actuation valve to modulate the flow of fuel. Initially, the controller would rapidly take large phase steps in order to home in, within a fraction of a second, to a favorable phase region within which the instability would be reduced. Then the controller would restrict itself to operate within this phase region and would further restrict itself to operate within a region of stability, as long as the power in the instability signal was decreasing. In the phase-shifting scheme of this method, the phase of the control vector would be made to continuously bounce back and forth from one boundary of an effective stability region to the other. Computationally

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.

Experimental clean combustor program, phase 3: Noise measurement addendum. [CF6-50 high bypass turbofan engine noise

NASA Technical Reports Server (NTRS)

Doyle, V. L.

1978-01-01

The acoustic characteristics of the double annular combustor in a CF6-50 high bypass turbofan engine were investigated. Internal fluctuating pressure measurements were made in the combustor region and in the core exhaust. The transmission loss across the turbine and nozzle was determined from the measurements and compared to previous component results and present theory. The primary noise source location in the combustor was investigated. Spectral comparisons of test rig results were made with the engine results. The measured overall power level was compared with component and engine correlating parameters.

Flow conditioner for fuel injector for combustor and method for low-NO.sub.x combustor

DOEpatents

Dutta, Partha; Smith, Kenneth O.; Ritz, Frank J.

2013-09-10

An injector for a gas turbine combustor including a catalyst coated surface forming a passage for feed gas flow and a channel for oxidant gas flow establishing an axial gas flow through a flow conditioner disposed at least partially within an inner wall of the injector. The flow conditioner includes a length with an interior passage opening into upstream and downstream ends for passage of the axial gas flow. An interior diameter of the interior passage smoothly reduces and then increases from upstream to downstream ends.

Ignition of Hydrogen-Oxygen Rocket Combustor with Chlorine Trifluoride and Triethylaluminum

NASA Technical Reports Server (NTRS)

Gregory, John W.; Straight, David M.

1961-01-01

Ignition of a nominal-125-pound-thrust cold (2000 R) gaseous-hydrogen - liquid-oxygen rocket combustor with chlorine trifluoride (hypergolic with hydrogen) and triethylaluminum (hypergolic with oxygen) resulted in consistently smooth starting transients for a wide range of combustor operating conditions. The combustor exhaust nozzle discharged into air at ambient conditions. Each starting transient consisted of the following sequence of events: injection of the lead main propellant, injection of the igniter chemical, ignition of these two chemicals, injection of the second main propellant, ignition of the two main propellants, increase in chamber pressure to its terminal value, and cutoff of igniter-chemical flow. Smooth ignition was obtained with an ignition delay of less than 100 milliseconds for the reaction of the lead propellant with the igniter chemical using approximately 0.5 cubic inch (0-038 lb) of chlorine trifluoride or 1.0 cubic inch (0-031 lb) of triethylaluminum. These quantities of igniter chemical were sufficient to ignite a 20-percent-fuel hydrogen-oxygen mixture with a delay time of less than 15 milliseconds. Test results indicated that a simple, light weight chemical ignition system for hydrogen-oxygen rocket engines may be possible.

On the modelling of non-reactive and reactive turbulent combustor flows

NASA Technical Reports Server (NTRS)

Nikjooy, Mohammad; So, Ronald M. C.

1987-01-01

A study of non-reactive and reactive axisymmetric combustor flows with and without swirl is presented. Closure of the Reynolds equations is achieved by three models: kappa-epsilon, algebraic stress and Reynolds stress closure. Performance of two locally nonequilibrium and one equilibrium algebraic stress models is analyzed assuming four pressure strain models. A comparison is also made of the performance of a high and a low Reynolds number model for combustor flow calculations using Reynolds stress closures. Effects of diffusion and pressure-strain models on these closures are also investigated. Two models for the scalar transport are presented. One employs the second-moment closure which solves the transport equations for the scalar fluxes, while the other solves the algebraic equations for the scalar fluxes. In addition, two cases of non-premixed and one case of premixed combustion are considered. Fast- and finite-rate chemistry models are applied to non-premixed combustion. Both show promise for application in gas turbine combustors. However, finite rate chemistry models need to be examined to establish a suitable coupling of the heat release effects on turbulence field and rate constants.

A Very Large Eddy Simulation of the Nonreacting Flow in a Single-Element Lean Direct Injection Combustor Using PRNS with a Nonlinear Subscale Model

NASA Technical Reports Server (NTRS)

Shih, Tsan-Hsing; Liu, Nan-Suey

2009-01-01

Very large eddy simulation (VLES) of the nonreacting turbulent flow in a single-element lean direct injection (LDI) combustor has been successfully performed via the approach known as the partially resolved numerical simulation (PRNS/VLES) using a nonlinear subscale model. The grid is the same as the one used in a previous RANS simulation, which was considered as too coarse for a traditional LES simulation. In this study, we first carry out a steady RANS simulation to provide the initial flow field for the subsequent PRNS/VLES simulation. We have also carried out an unsteady RANS (URANS) simulation for the purpose of comparing its results with that of the PRNS/VLES simulation. In addition, these calculated results are compared with the experimental data. The present effort has demonstrated that the PRNS/VLES approach, while using a RANS type of grid, is able to reveal the dynamically important, unsteady large-scale turbulent structures occurring in the flow field of a single-element LDI combustor. The interactions of these coherent structures play a critical role in the dispersion of the fuel, hence, the mixing between the fuel and the oxidizer in a combustor.

A mathematical model for jet engine combustor pollutant emissions

NASA Technical Reports Server (NTRS)

Boccio, J. L.; Weilerstein, G.; Edelman, R. B.

1973-01-01

Mathematical modeling for the description of the origin and disposition of combustion-generated pollutants in gas turbines is presented. A unified model in modular form is proposed which includes kinetics, recirculation, turbulent mixing, multiphase flow effects, swirl and secondary air injection. Subelements of the overall model were applied to data relevant to laboratory reactors and practical combustor configurations. Comparisons between the theory and available data show excellent agreement for basic CO/H2/Air chemical systems. For hydrocarbons the trends are predicted well including higher-than-equilibrium NO levels within the fuel rich regime. Although the need for improved accuracy in fuel rich combustion is indicated, comparisons with actual jet engine data in terms of the effect of combustor-inlet temperature is excellent. In addition, excellent agreement with data is obtained regarding reduced NO emissions with water droplet and steam injection.

Features of Ignition and Stable Combustion in Supersonic Combustor

NASA Astrophysics Data System (ADS)

Goldfeld, M.; Starov, A.; Timofeev, K.

2009-01-01

Present paper describes the results of experimental investigations of the supersonic combustor with entrance Mach numbers from 2 to 4 at static pressure from 0.8 to 2.5 bars, total temperature from 2000K to 3000K. Hydrogen and kerosene were used as fuel. The conditions, under which the self-ignition and intensive combustion of the fuel realized were found. Position of ignition area in the channel was determined and features of flame propagation in the channel presented. A possibility to ensure an efficient combustion of hydrogen and kerosene at a high supersonic flow velocity at the combustor entrance without special throttling and/or pseudo-shock introduction was shown. Analysis of applicability of existing methods of criterion descriptions of conditions of self-ignition and extinction of combustion is executed for generalization of experimental results on the basis of results obtained.

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.

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

NASA Technical Reports Server (NTRS)

Hulka, James

2008-01-01

The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

Combustor for fine particulate coal

DOEpatents

Carlson, L.W.

1988-01-26

A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover. 4 figs.

Combustor for fine particulate coal

DOEpatents

Carlson, Larry W.

1988-01-01

A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover.

Combustor for fine particulate coal

DOEpatents

Carlson, L.W.

1988-11-08

A particulate coal combustor with two combustion chambers is provided. The first combustion chamber is toroidal; air and fuel are injected, mixed, circulated and partially combusted. The air to fuel ratio is controlled to avoid production of soot or nitrogen oxides. The mixture is then moved to a second combustion chamber by injection of additional air where combustion is completed and ash removed. Temperature in the second chamber is controlled by cooling and gas mixing. The clean stream of hot gas is then delivered to a prime mover. 4 figs.

Application of mixing-controlled combustion models to gas turbine combustors

NASA Technical Reports Server (NTRS)

Nguyen, Hung Lee

1990-01-01

Gas emissions were studied from a staged Rich Burn/Quick-Quench Mix/Lean Burn combustor were studied under test conditions encountered in High Speed Research engines. The combustor was modeled at conditions corresponding to different engine power settings, and the effect of primary dilution airflow split on emissions, flow field, flame size and shape, and combustion intensity, as well as mixing, was investigated. A mathematical model was developed from a two-equation model of turbulence, a quasi-global kinetics mechanism for the oxidation of propane, and the Zeldovich mechanism for nitric oxide formation. A mixing-controlled combustion model was used to account for turbulent mixing effects on the chemical reaction rate. This model assumes that the chemical reaction rate is much faster than the turbulent mixing rate.

Experimental study on combustion modes and thrust performance of a staged-combustor of the scramjet with dual-strut

NASA Astrophysics Data System (ADS)

Yang, Qingchun; Chetehouna, Khaled; Gascoin, Nicolas; Bao, Wen

2016-05-01

To enable the scramjet operate in a wider flight Mach number, a staged-combustor with dual-strut is introduced to hold more heat release at low flight Mach conditions. The behavior of mode transition was examined using a direct-connect model scramjet experiment along with pressure measurements. The typical operating modes of the staged-combustor are analyzed. Fuel injection scheme has a significant effect on the combustor operating modes, particularly for the supersonic combustion mode. Thrust performances of the combustor with different combustion modes and fuel distributions are reported in this paper. The first-staged strut injection has a better engine performance in the operation of subsonic combustion mode. On the contrast, the second-staged strut injection has a better engine performance in the operation of supersonic combustion mode.

Design of a laboratory scale fluidized bed reactor

NASA Astrophysics Data System (ADS)

Wikström, E.; Andersson, P.; Marklund, S.

1998-04-01

The aim of this project was to construct a laboratory scale fluidized bed reactor that simulates the behavior of full scale municipal solid waste combustors. The design of this reactor is thoroughly described. The size of the laboratory scale fluidized bed reactor is 5 kW, which corresponds to a fuel-feeding rate of approximately 1 kg/h. The reactor system consists of four parts: a bed section, a freeboard section, a convector (postcombustion zone), and an air pollution control (APC) device system. The inside diameter of the reactor is 100 mm at the bed section and it widens to 200 mm in diameter in the freeboard section; the total height of the reactor is 1760 mm. The convector part consists of five identical sections; each section is 2700 mm long and has an inside diameter of 44.3 mm. The reactor is flexible regarding the placement and number of sampling ports. At the beginning of the first convector unit and at the end of each unit there are sampling ports for organic micropollutants (OMP). This makes it possible to study the composition of the flue gases at various residence times. Sampling ports for inorganic compounds and particulate matter are also placed in the convector section. All operating parameters, reactor temperatures, concentrations of CO, CO2, O2, SO2, NO, and NO2 are continuously measured and stored at selected intervals for further evaluation. These unique features enable full control over the fuel feed, air flows, and air distribution as well as over the temperature profile. Elaborate details are provided regarding the configuration of the fuel-feeding systems, the fluidized bed, the convector section, and the APC device. This laboratory reactor enables detailed studies of the formation mechanisms of OMP, such as polychlorinated dibenzo-p-dioxins (PCDDs), polychlorinated dibenzofurans (PCDFs), poly-chlorinated biphenyls (PCBs), and polychlorinated benzenes (PCBzs). With this system formation mechanisms of OMP occurring in both the combustion

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.

System for reducing combustion dynamics and NO.sub.x in a combustor

DOEpatents

Uhm, Jong Ho; Ziminsky, Willy Steve; Johnson, Thomas Edward; Hughes, Michael John; York, William David

2016-05-31

A combustor includes an end cap that extends radially across at least a portion of the combustor. The end cap includes an upstream surface axially separated from a downstream surface. A plurality of tubes extend from the upstream surface through the downstream surface of the end cap to provide fluid communication through the end cap. Each tube in a first set of the plurality of tubes has an inlet proximate to the upstream surface and an outlet downstream from the downstream surface. Each outlet has a first portion that extends a different axial distance from the inlet than a second portion.

Odor intensity and characterization studies of exhaust from a turbojet engine combustor

NASA Technical Reports Server (NTRS)

Butze, H. F.; Kendall, D. A.

1973-01-01

Sensory odor tests of the exhaust from a turbojet combustor operating at simulated idle conditions were made by a human panel sniffing diluted exhaust gas. Simultaneously, samples of undiluted exhaust gas were collected on adsorbent substrates, subsequently removed by solvent flushing, and analyzed chemically by liquid chromatographic methods. The concentrations of the principal malodorous species, the aromatic (unburned fuel-related) and the oxygenated (partially burned fuel) fractions, as determined chromatographically, correlated well with the intensity of the odor as determined by sniffing. Odor intensity increased as combustion efficiency decreased. Combustor modifications which increased combustion efficiency decreased odor intensity.

Performance gains by using heated natural-gas fuel in an annular turbojet combustor

NASA Technical Reports Server (NTRS)

Marchionna, N. R.

1973-01-01

A full-scale annular turbojet combustor was tested with natural gas fuel heated from ambient temperature to 800 K (980 F). In all tests, heating the fuel improved combustion efficiency. Two sets of gaseous fuel nozzles were tested. Combustion instabilities occurred with one set of nozzles at two conditions: one where the efficiency approached 100 percent with the heated fuel; the other where the efficiency was very poor with the unheated fuel. The second set of nozzles exhibited no combustion instability. Altitude relight tests with the second set showed that relight was improved and was achievable at essentially the same condition as blowout when the fuel temperature was 800 K (980 F).

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...

DEMONSTRATION BULLETIN: CIRCULATING BED COMBUSTOR - OGDEN ENVIRONMENTAL SERVICES, INC.

EPA Science Inventory

An evaluation of the Ogden Environmental Services (OES) circulating bed combustor (CBC) technology was carried out under the superfund Innovative Technology Evaluation (SITE) Program to determine its applicabilitY as an on-site treatment method for waste site cleanups, and more s...

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.

Imaging of Combustion Species in a Radially-Staged Gas Turbine Combustor

NASA Technical Reports Server (NTRS)

Locke, Randy J.; Hicks, Yolanda R.; Anderson, Robert C.; Ockunzzi, Kelly A.; Schock, Harold J.

1996-01-01

Plannar laser-induced fluorescence (PLIF) is used to characterize the complex flowfield of a unique fuel-lean, radially-staged high pressure gas turbine combustor. PLIF images of OH are presented for two fuel injector configurations. PLIF images of NO, the first acquired at these conditions, are presented and compared with gas sample extraction probe measurements. Flow field imaging of nascent C2 chemiluminescence is also investigated. An examination is made of the interaction between adjoining lean premixed prevaporized (LPP) injectors. Fluorescence interferences at conditions approaching 2000 K and 15 atm are observed and attributed to polycyclic aromatic hydrocarbon (PAH) emissions. All images are acquired at a position immediately downstream of the fuel injectors with combustor burning JP-5 fuel.

Systems and methods for preventing flashback in a combustor assembly

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

Johnson, Thomas Edward; Ziminsky, Willy Steve; Stevenson, Christian Xavier

2016-04-05

Embodiments of the present application include a combustor assembly. The combustor assembly may include a combustion chamber, a first plenum, a second plenum, and one or more elongate air/fuel premixing injection tubes. Each of the elongate air/fuel premixing injection tubes may include a first length at least partially disposed within the first plenum and configured to receive a first fluid from the first plenum. Moreover, each of the elongate air/fuel premixing injection tubes may include a second length disposed downstream of the first length and at least partially disposed within the second plenum. The second length may be formed ofmore » a porous wall configured to allow a second fluid from the second plenum to enter the second length and create a boundary layer about the porous wall.« less

Sensitivity Analysis to Turbulent Combustion Models for Combustor-Turbine Interactions

NASA Astrophysics Data System (ADS)

Miki, Kenji; Moder, Jeff; Liou, Meng-Sing

2017-11-01

The recently-updated Open National CombustionCode (Open NCC) equipped with alarge-eddy simulation (LES) is applied to model the flow field inside the Energy Efficient Engine (EEE) in conjunction with sensitivity analysis to turbulent combustion models. In this study, we consider three different turbulence-combustion interaction models, the Eddy-Breakup model (EBU), the Linear-Eddy Model (LEM) and the Probability Density Function (PDF)model as well as the laminar chemistry model. Acomprehensive comparison of the flow field and the flame structure will be provided. One of our main interests isto understand how a different model predicts thermal variation on the surface of the first stage vane. Considering that these models are often used in combustor/turbine communities, this study should provide some guidelines on numerical modeling of combustor-turbine interactions.

Effect of fuel vapor concentrations on combustor emissions and performance

NASA Technical Reports Server (NTRS)

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

1973-01-01

Effects of fuel vaporization on the exhaust emission levels of oxides of nitrogen, carbon monoxide, total hydrocarbons, and smoke number were obtained in an experimental turbojet combustor segment. Two different fuel injectors were used in which liquid ASTM A-1 jet fuel and vapor propane fuel were independently controlled to simulate varying degrees of vaporization. Tests were conducted over a range of inlet-air temperatures from 478 to 700 K, pressures from 4 to 20 atm, and combustor reference velocities from 15.3 to 27.4 m/sec. Converting from liquid to complete vapor fuel resulted in oxides of nitrogen reductions of as much as 22 percent and smoke number reductions up to 51 percent. Supplement data are also presented on flame emissivity, flame temperature, and primary-zone liner wall temperatures.

Heat transfer to horizontal tubes in a pilot-scale fluidized-bed combustor burning low-rank coals

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

Grewal, N.S.; Goblirsch, G.

Experimental data are obtained for the heat transfer coefficient between immersed horizontal tube bundles and an atmospheric-fluidized-bed combustor burning low-rank coals. Silica sand and limestone are used as bed material. The tests are conducted, with and without limestone addition and ash recycle, at average bed temperatures ranging from 1047 to 1125 K, superficial fluidizing velocity of 1.66 to 2.04 m/s, and excess air levels of 15 to 40 percent. The experimental data are examined in the light of the existing correlations for the heat transfer coefficient. The predicted values of heat transfer coefficient from the correlations proposed by Grewal andmore » Bansal et al. are found to be within + or - 25 percent of the experimental values of heat transfer coefficient, when the contribution due to radiation is also included.« less

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

Low NO.sub.x multistage combustor

DOEpatents

Becker, Frederick E.; Breault, Ronald W.; Litka, Anthony F.; McClaine, Andrew W.; Shukla, Kailash

2000-01-01

A high efficiency, Vortex Inertial Staged Air (VIStA) combustor provides ultra-low NO.sub.X production of about 20 ppmvd or less with CO emissions of less than 50 ppmvd, both at 3% O.sub.2. Prompt NO.sub.X production is reduced by partially reforming the fuel in a first combustion stage to CO and H.sub.2. This is achieved in the first stage by operating with a fuel rich mixture, and by recirculating partially oxidized combustion products, with control over stoichiometry, recirculation rate and residence time. Thermal NO.sub.X production is reduced in the first stage by reducing the occurrence of high temperature combustion gas regions. This is achieved by providing the first stage burner with a thoroughly pre-mixed fuel/oxidant composition, and by recirculating part of the combustion products to further mix the gases and provide a more uniform temperature in the first stage. In a second stage combustor thermal NO.sub.X production is controlled by inducing a large flow of flue gas recirculation in the second stage combustion zone to minimize the ultimate temperature of the flame. One or both of the first and second stage burners can be cooled to further reduce the combustion temperature and to improve the recirculation efficiency. Both of these factors tend to reduce production of NO.sub.X.

COMBUSTION CONTROL OF ORGANIC EMISSIONS FROM MUNICIPAL WASTE COMBUSTORS

EPA Science Inventory