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Effect of fuel distribution and flame stabilization techniques on combustion stability limits for swirl-can combustor modules
Chemical reaction knock rating of fuels for internal combustion engines
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Recent theoretical investigations on graphite particle combustion have employed several levels of heterogeneous reaction models, ranging from global to elementary models, to describe the oxidation of carbon to gaseous products. Unlike the counterpart homogeneous reaction models, these heterogeneous reaction models are not well developed because of the difficulties associated with decoupling the physical characteristics of the solid (e.g. surface area taking part in combustion) from the chemical kinetic data. This is certainly true for porous graphite particle combustion, where heterogeneous and homogeneous reactions occur within the pores and play an important role in the overall oxidation process. As a result, there are considerable uncertainties of physical phenomena predicted using different heterogeneous kinetic models available in the literature. A good example, discussed later in this paper, is the predicted critical particle size below which the mass burning rate becomes exponentially small. The main goal of this study is to understand the basic mechanism controlling such rapid changes in burning rates, by developing a model where physical contributions are decoupled from chemical rate constants in a consistent manner. Another important goal of the proposed study is to develop a truly intrinsic, detailed heterogeneous reaction model for porous graphite combustion at high-temperatures, and to derive a systematically reduced heterogeneous reaction model in terms of the elementary reaction rate constants of the detailed model. The validation of chemical kinetic models describing the heterogeneous and homogeneous combustion in and around a spherically symmetric porous graphite particle can be considerably simplified by experimental measurements obtained under microgravity conditions. A vital component of this study is to conduct such supporting experiments on particle burning rate and surface temperature using NASA microgravity facilities, in close coordination with the theoretical effort. The basic understanding obtained and models developed as part of this project will be useful for optimal design of coal combustion devices. These models can also be extended to investigate the role of heterogeneous chemistry on pollutant formation pathways in combustion devices. The theoretical approach developed here, with pore diffusion effects decoupled from the chemical effects, can also be extended to understand the heterogeneous combustion of other porous fuels, for example, combustion of magnesium in a CO2 environment for propulsion in the Martian atmosphere.
Droplet combustion experiments carried out onboard the International Space Station (ISS), using pure fuels and fuel mixtures, have shown that quasi-steady burning can be sustained by a non-traditional flame configuration, namely a "cool flame" burning in the "partial-burning" regime where both fuel and oxygen leak through the low-temperature controlled flame-sheet. Recent experiments involving large, bi-component fuel (n-decane and hexanol, 50/50 by volume) droplets at elevated pressures show that the visible, hot flame becomes extremely weak while the burning rate remains relatively high, suggesting the possibility of simultaneous presence of "cool" and "hot" flames of roughly equal importance. The radiant output from these bi-component droplets is relatively high and cannot be accounted for only by the presence of a visible hot-flame. In this analysis we explore the theoretical possibility of a dual-flame structure, where one flame lies close to the droplet surface called the "cool-flame," and other farther away from the droplet surface, termed the "hot-flame." A Burke-Schumann analysis of this dual-structure seems to indicate such flame structures are possible over a narrow range of initial conditions. Theoretical results can be compared against available experimental data for pure and bi-component fuel droplet combustion to test how realistic the model may be.
An overview is provided of a CFD assessment with advanced fuel blends to assess fuel impacts on NOx emissions and flame structure in an axially-staged combustor for NASA’s Commercial Supersonic Transport (CST) Program. The National Combustor Code (OpenNCC) was used to perform two-phase reacting flow computations with various blending ratios of two different aviation fuels, ‘average’ Jet-A (A2) and GEVO-ATJ (C1), for UTRC’s Axially Controlled Stoichiometry (ACS) combustor at CST cruise conditions. The predicted flame structures in the ACS combustor with three different blending ratios of A2 and C1 fuel were very similar to each other. The predicted NOx emissions for all fuel blends were within 10% of the experimentally measured NOx emissions for 100% A2 fuel.
An overview is provided of a CFD assessment with Jet-A replacement fuels and advanced fuel blends to assess fuel impacts on NOx emissions and flame structure in an axially-staged combustor for NASA’s Commercial Supersonic Transport (CST) program. The National Combustor Code (OpenNCC) was used to perform two-phase reacting flow computations with two replacement fuels for Jet-A, and various blending ratios of ‘average’ Jet-A (A2) and GEVO-ATJ (C1), for UTRC’s Axially Controlled Stoichiometry (ACS) combustor at CST cruise conditions. The predicted flame structures in the ACS combustor with for the replacement fuels, and for three different blending ratios of A2 and C1 fuel were very similar to each other. The predicted NOx emissions for the two replacement fuels, and three blending ratios of A2 and C1 were within 15% of the experimentally measured NOx emissions for 100% A2 fuel.
Rotating detonation engines (RDEs) have gained attention as a promising technology for future aviation engines. However, the numerical studies of these systems pose severe challenges due to the broad range of spatial and temporal scales. In this study, we use an adaptive mesh refinement based compressible, reactive solver PeleC to resolve the broad range of scales and accurately capture shock and detonation waves using high-resolution numerical schemes. Multi- species transport along with compressible Navier-Stokes equations are solved in the model along with a finite-rate based chemistry model. Embedded boundary method is used to model the complex geometry consisting of discrete fuel nozzles and the combustion chamber. The fuel consisting predominantly of hydrogen is doped with varying levels of methane while air is used as the oxidizer. For a specified total pressure and temperature, the number of stable detonation waves is found to decrease with increasing methane concentration in the fuel mixture. Additionally, no stable detonation solutions are observed for methane composition higher than 20% by volume for the range of operating conditions studied (300-900K, 10 Atm). The increased presence of high temperature zones is also indicative of higher thermal NOx emissions at low methane concentrations. The effect of fuel-air mixture composition and temperature on the detonability, detonation wave structure, mode transitions and their stability are analyzed in this study in addition to their implication on NOx emission.
The effects of select fuel property variations on two major engine classifications are summarized. Thirteen refined and blended fuels were used which exhibited significant variations in hydrogen content, aromatic type, initial boiling point, final boiling point, and viscosity. Trends were very similar but the degree of fuel sensitivity was not constant. For both systems the dominant fuel property during high pressure operation was found to be fuel hydrogen content. For operation at low pressure test points the fuel volatility and viscosity became the dominant fuel properties for both systems. Aromatic type and final boiling point did not significantly affect combustion data. Correlations of other fuel properties with these and other performance parameters are presented.
Combustion efficiency of hydrogen fuel for varying afterburner configurations
Results of an evaluation of high pressure combustion of fuel rich and oxidizer rich LOX/RP-1 propellants using 4.0 inch diameter prototype preburner injectors and chambers are presented. Testing covered a pressure range from 8.9 to 17.5 MN/square meters (1292 to 2540 psia). Fuel rich mixture ratios ranged from 0.238 to 0.367; oxidizer rich mixture ratios ranged from 27.2 to 47.5. Performance, gas temperature uniformity, and stability data for two fuel rich and two ozidizer rich preburner injectors are presented for a conventional like-on-like (LOL) design and a platelet design injector. Kinetically limited combustion is shown by the excellent agreement of measured fuel rich gas composition and C performance data with kinetic model predictions. The oxidizer rich test results support previous equilibrium combustion predictions.
A series of distinct phenomena have recently been observed in single-fuel-droplet combustion tests performed on the International Space Station (ISS). This study attempts to simulate the observed flame behavior numerically using a gaseous n-heptane fuel source in zero gravity and a time-dependent axisymmetric (2D) code, which includes a detailed reaction mechanism (127 species and 1130 reactions), diffusive transport, and a radiation model (for CH4, CO, CO2, H2O, and soot). The calculated combustion characteristics depend strongly on the air velocity around the fuel source. In a near-quiescent air environment (< or = 2 mm/s), with a sufficiently large fuel injection velocity (1 cm/s), a growing spherical diffusion flame extinguishes at ≈1200 K due to radiative heat losses. This is typically followed by a transition to the low-temperature (cool-flame) regime with a reaction zone (at ≈700 K) in close proximity to the fuel source. The 'cool flame' regime is formed due to the negative temperature coefficient in the low-temperature chemistry. After a relatively long period (≈18 s) of the cool flame regime, a flash re-ignition occurs, associated with flame-edge propagation and subsequent extinction of the re-ignited flame. In a low-speed (≈3 mm/s) airstream (which simulates the slight droplet movement), the diffusion flame is enhanced upstream and experiences a local extinction downstream at ≈1200 K, followed by steady flame pulsations (≈0.4 Hz). At higher air velocities (4-10 mm/s), the locally extinguished flame becomes steady state. The present axisymmetric computational approach helps in revealing the non-premixed 'cool flame' structure and 2D flame-flow interactions observed in recent microgravity droplet combustion experiments.
The results of an investigation of the effect of operating conditions and fuel properties on emission for the two-stage combustion of fuels with significant organic nitrogen content are presented. The way in which the emissions of nitrogen oxides and carbon monoxide are affected by the decreased hydrogen content and the increased organic nitrogen content of coal-derived fuels is discussed. Limited measurements of smoke from the rich-lean combustion of simulated syncrude fuels indicate relatively high smoke emissions in spite of the very lean second-stage burning. This fact, together with the high observed carbon monoxide emissions, suggests that trade-offs will be necessary between the conditions that minimize NOx and those that control CO and smoke emissions.
Experimental and analytical investigations focusing on aluminum/hydrocarbon gel droplet secondary atomization and its effects on gel-fueled rocket engine performance are being conducted. A single laser sheet sizing/velocimetry diagnostic technique, which should eliminate sizing bias in the data collection process, has been designed and constructed to overcome limitations of the two-color forward-scatter technique used in previous work. Calibration of this system is in progress and the data acquisition/validation code is being written. Narrow-band measurements of radiant emission, discussed in previous reports, will be used to determine if aluminum ignition has occurred in a gel droplet. A one-dimensional model of a gel-fueled rocket combustion chamber, described in earlier reports, has been exercised in conjunction with a two-dimensional, two-phase nozzle code to predict the performance of an aluminum/hydrocarbon fueled engine. Estimated secondary atomization effects on propellant burnout distance, condensed particle radiation losses to the chamber walls, and nozzle two phase flow losses are also investigated. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size, and radiation heat losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated, depending on secondary atomization intensity. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two phase flow effects on overall engine performance. Radiation losses yielded a one percent decrease in engine Isp. Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine Isp was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine Isp efficiencies, accounting for radiation and two phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of five.
Test results of a full annulus swirl-can combustor operated at simulated engine idle conditions indicated that significant improvements in combustion efficiency and accompanying reductions in pollutants could be realized with radial scheduling of fuel. Test conditions were an inlet air temperature of 478 K, a pressure of 4 atmospheres, and a reference velocity of 26 meters per second. At a fuel-air ratio of 0.008, radial scheduling of fuel increased combustion efficiency from less than 50 percent with no scheduling to nearly 100 percent and produced emission index values of unburned hydrocarbons and carbon monoxide of 10 and 40, respectively.
Solid oxide fuel cell – internal combustion engine (SOFC-ICE) hybrid systems are an attractive solution for electricity generation. The system can achieve up to 70% theoretical electric power conversion efficiency through energy cascading enabled by utilizing the anode off-gas from the SOFC as the fuel source for the ICE. Experimental investigations were conducted with a single cylinder Cooperative Fuel Research (CFR) engine by altering fuel-air equivalence ratio (φ), and compression ratio (CR) to study the engine load, combustion characteristics, and emissions levels of dry SOFC anode off-gas consisting of 33.9% H 2 , 15.6% CO, and 50.5% CO 2 . The combustion efficiency of the anode off-gas was directly evaluated by measuring the engine-out CO emissions. The highest net-indicated fuel conversion efficiency of 31.3% occurred at φ = 0.90 and CR = 13:1. These results demonstrate that the anode off-gas can be successfully oxidized using a spark ignition combustion mode. The fuel conversion efficiency of the anode tail gas is expected to further increase in a more modern engine architecture that can achieve increased burn rates in comparison to the CFR engine. NO x emissions from the combustion of anode off-gas were minimal as the cylinder peak temperatures never exceeded 1800 K. This experimental study ultimately demonstrates the viability of an ICE to operate using an anode off-gas, thus creating a complementary role for an ICE to be paired with a SOFC in a hybrid power generation plant.
Combustion of H2/hydrocarbon (HC) fuel mixtures may be considered in certain volume-limited supersonic airbreathing propulsion applications. Effects of HC addition to H2 were evaluated, using a recent argon-bathed, coaxial, tubular opposed jet burner (OJB) technique to measure the extinction limits of counterflow diffusion flames. The OJB flames were formed by a laminar jet of (N2 and/or HC)-diluted H2 mixture opposed by a similar jet of air at ambient conditions. The OJB data, derived from respective binary mixtures of H2 and methane, ethylene, or propane HCs, were used to characterize BLOWOFF and RESTORE. BLOWOFF is a sudden breaking of the dish-shaped OJB flame to a stable torus or ring shape, and RESTORE marks sudden restoration of the central flame by radial inward flame propagation. BLOWOFF is a measure of kinetically-limited flame reactivity/speed under highly stretched, but relatively ideal impingement flow conditions. RESTORE measures inward radial flame propagation rate, which is sensitive to ignition processes in the cool central core. It is concluded that relatively small molar amounts of added HC greatly reduce the reactivity characteristics of counterflow hydrogen-air diffusion flames, for ambient initial conditions.
Overview presentation with examples of how fuel-air mixer changes affect fluid and flame structure, given to an undergraduate students in a combustion course.