A Method for Large Eddy Simulation of Acoustic Combustion Instabilities

NASA Astrophysics Data System (ADS)

Wall, Clifton; Moin, Parviz

2003-11-01

A method for performing Large Eddy Simulation of acoustic combustion instabilities is presented. By extending the low Mach number pressure correction method to the case of compressible flow, a numerical method is developed in which the Poisson equation for pressure is replaced by a Helmholtz equation. The method avoids the acoustic CFL condition by using implicit time advancement, leading to large efficiency gains at low Mach number. The method also avoids artificial damping of acoustic waves. The numerical method is attractive for the simulation of acoustics combustion instabilities, since these flows are typically at low Mach number, and the acoustic frequencies of interest are usually low. Additionally, new boundary conditions based on the work of Poinsot and Lele have been developed to model the acoustic effect of a long channel upstream of the computational inlet, thus avoiding the need to include such a channel in the computational domain. The turbulent combustion model used is the Level Set model of Duchamp de Lageneste and Pitsch for premixed combustion. Comparison of LES results to the reacting experiments of Besson et al. will be presented.

A Method for Large Eddy Simulation of Acoustic Combustion Instabilities

NASA Astrophysics Data System (ADS)

Wall, Clifton; Pierce, Charles; Moin, Parviz

2002-11-01

A method for performing Large Eddy Simulation of acoustic combustion instabilities is presented. By extending the low Mach number pressure correction method to the case of compressible flow, a numerical method is developed in which the Poisson equation for pressure is replaced by a Helmholtz equation. The method avoids the acoustic CFL condition by using implicit time advancement, leading to large efficiency gains at low Mach number. The method also avoids artificial damping of acoustic waves. The numerical method is attractive for the simulation of acoustic combustion instabilities, since these flows are typically at low Mach number, and the acoustic frequencies of interest are usually low. Both of these characteristics suggest the use of larger time steps than those allowed by an acoustic CFL condition. The turbulent combustion model used is the Combined Conserved Scalar/Level Set Flamelet model of Duchamp de Lageneste and Pitsch for partially premixed combustion. Comparison of LES results to the experiments of Besson et al will be presented.

Numerical parametric studies of spray combustion instability

NASA Technical Reports Server (NTRS)

Pindera, M. Z.

1993-01-01

A coupled numerical algorithm has been developed for studies of combustion instabilities in spray-driven liquid rocket engines. The model couples gas and liquid phase physics using the method of fractional steps. Also introduced is a novel, efficient methodology for accounting for spray formation through direct solution of liquid phase equations. Preliminary parametric studies show marked sensitivity of spray penetration and geometry to droplet diameter, considerations of liquid core, and acoustic interactions. Less sensitivity was shown to the combustion model type although more rigorous (multi-step) formulations may be needed for the differences to become apparent.

Catalyst Bed Instability Within the USFE H2O2/JP-8 Rocket Engine

NASA Technical Reports Server (NTRS)

Johnson, Curtis W.; Anderson, William; Ross, Robert; Lyles, G. (Technical Monitor)

2000-01-01

Orbital Sciences Corporation has been awarded a contract by NASA's Marshall Space Flight Center, in cooperation with the U.S. Air Force Research Laboratory's Military Space Plane Technology Program Office, for the Upper Stage Flight Experiment (USFE) program. Orbital is designing, developing, and will flight test a new low-cost, 10,000 lbf hydrogen peroxide/ JP-8 pressure fed liquid rocket. During combustion chamber tests at NASA Stennis Space Center (SSC) of the USFE engine, the catalyst bed showed a low frequency instability occurring as the H202 flow reached about 1/3 its design rate. This paper reviews the USFE catalyst bed and combustion chamber and its operation, then discusses the dynamics of the instability. Next the paper describes the dynamic computer model used to recreate the instability. The model was correlated to the SSC test data, and used to investigate possible solutions to the problem. The combustion chamber configuration which solved the instability is shown, and the subsequent stable operation presented.

Application of Detailed Chemical Kinetics to Combustion Instability Modeling

DTIC Science & Technology

2016-01-04

Modeling 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Harvazinski, Matt; Talley, Doug; Sankaran, Venke 5d. PROJECT...Chemical Kinetics to Combustion Instability Modeling Matt Harvazinski, Doug Talley, Venke Sankaran Air Force Research Laboratory Edwards AFB, CA...distribution unlimited. 3 Prior Work – Kinetics Used • Simulations : 1) 3D real geometry 2) Unsteady 3) Long run-times 4) Coupled physics • 1- 4

Active Control of High Frequency Combustion Instability in Aircraft Gas-Turbine Engines

NASA Technical Reports Server (NTRS)

Corrigan, Bob (Technical Monitor); DeLaat, John C.; Chang, Clarence T.

2003-01-01

Active control of high-frequency (greater than 500 Hz) combustion instability has been demonstrated in the NASA single-nozzle combustor rig at United Technologies Research Center. The combustor rig emulates an actual engine instability and has many of the complexities of a real engine combustor (i.e. actual fuel nozzle and swirler, dilution cooling, etc.) In order to demonstrate control, a high-frequency fuel valve capable of modulating the fuel flow at up to 1kHz was developed. Characterization of the fuel delivery system was accomplished in a custom dynamic flow rig developed for that purpose. Two instability control methods, one model-based and one based on adaptive phase-shifting, were developed and evaluated against reduced order models and a Sectored-1-dimensional model of the combustor rig. Open-loop fuel modulation testing in the rig demonstrated sufficient fuel modulation authority to proceed with closed-loop testing. During closed-loop testing, both control methods were able to identify the instability from the background noise and were shown to reduce the pressure oscillations at the instability frequency by 30%. This is the first known successful demonstration of high-frequency combustion instability suppression in a realistic aero-engine environment. Future plans are to carry these technologies forward to demonstration on an advanced low-emission combustor.

Sensitivity of Combustion-Acoustic Instabilities to Boundary Conditions for Premixed Gas Turbine Combustors

NASA Technical Reports Server (NTRS)

Darling, Douglas; Radhakrishnan, Krishnan; Oyediran, Ayo

1995-01-01

Premixed combustors, which are being considered for low NOx engines, are susceptible to instabilities due to feedback between pressure perturbations and combustion. This feedback can cause damaging mechanical vibrations of the system as well as degrade the emissions characteristics and combustion efficiency. In a lean combustor instabilities can also lead to blowout. A model was developed to perform linear combustion-acoustic stability analysis using detailed chemical kinetic mechanisms. The Lewis Kinetics and Sensitivity Analysis Code, LSENS, was used to calculate the sensitivities of the heat release rate to perturbations in density and temperature. In the present work, an assumption was made that the mean flow velocity was small relative to the speed of sound. Results of this model showed the regions of growth of perturbations to be most sensitive to the reflectivity of the boundary when reflectivities were close to unity.

Combustion Stability Analyses for J-2X Gas Generator Development

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Protz, C. S.; Casiano, M. J.; Kenny, R. J.

2010-01-01

The National Aeronautics and Space Administration (NASA) is developing a liquid oxygen/liquid hydrogen rocket engine for upper stage and trans-lunar applications of the Ares vehicles for the Constellation program. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. Development was contracted to Pratt & Whitney Rocketdyne in 2006. Over the past several years, development of the gas generator for the J-2X engine has progressed through a variety of workhorse injector, chamber, and feed system configurations. Several of these configurations have resulted in injection-coupled combustion instability of the gas generator assembly at the first longitudinal mode of the combustion chamber. In this paper, the longitudinal mode combustion instabilities observed on the workhorse test stand are discussed in detail. Aspects of this combustion instability have been modeled at the NASA Marshall Space Flight Center with several codes, including the Rocket Combustor Interaction Design and Analysis (ROCCID) code and a new lumped-parameter MatLab model. To accurately predict the instability characteristics of all the chamber and injector geometries and test conditions, several features of the submodels in the ROCCID suite of calculations required modification. Finite-element analyses were conducted of several complicated combustion chamber geometries to determine how to model and anchor the chamber response in ROCCID. A large suite of sensitivity calculations were conducted to determine how to model and anchor the injector response in ROCCID. These modifications and their ramification for future stability analyses of this type are discussed in detail. The lumped-parameter MatLab model of the gas generator assembly was created as an alternative calculation to the ROCCID methodology. This paper also describes this model and the stability calculations.

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

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Breisacher, Kevin J.

2000-01-01

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

The 17th JANNAF Combustion Meeting, Volume 1

NASA Technical Reports Server (NTRS)

Eggleston, D. S. (Editor)

1980-01-01

The combustion of solid rocket propellants and combustion in ramjets is addressed. Subjects discussed include metal burning, steady-state combustion of composite propellants, velocity coupling and nonlinear instability, vortex shedding and flow effects on combustion instability, combustion instability in solid rocket motors, combustion diagnostics, subsonic and supersonic ramjet combustion, characterization of ramburner flowfields, and injection and combustion of ramjet fuels.

Multi-injector modeling of transverse combustion instability experiments

NASA Astrophysics Data System (ADS)

Shipley, Kevin J.

Concurrent simulations and experiments are used to study combustion instabilities in a multiple injector element combustion chamber. The experiments employ a linear array of seven coaxial injector elements positioned atop a rectangular chamber. Different levels of instability are driven in the combustor by varying the operating and geometry parameters of the outer driving injector elements located near the chamber end-walls. The objectives of the study are to apply a reduced three-injector model to generate a computational test bed for the evaluation of injector response to transverse instability, to apply a full seven-injector model to investigate the inter-element coupling between injectors in response to transverse instability, and to further develop this integrated approach as a key element in a predictive methodology that relies heavily on subscale test and simulation. To measure the effects of the transverse wave on a central study injector element two opposing windows are placed in the chamber to allow optical access. The chamber is extensively instrumented with high-frequency pressure transducers. High-fidelity computational fluid dynamics simulations are used to model the experiment. Specifically three-dimensional, detached eddy simulations (DES) are used. Two computational approaches are investigated. The first approach models the combustor with three center injectors and forces transverse waves in the chamber with a wall velocity function at the chamber side walls. Different levels of pressure oscillation amplitudes are possible by varying the amplitude of the forcing function. The purpose of this method is to focus on the combustion response of the study element. In the second approach, all seven injectors are modeled and self-excited combustion instability is achieved. This realistic model of the chamber allows the study of inter-element flow dynamics, e.g., how the resonant motions in the injector tubes are coupled through the transverse pressure waves in the chamber. The computational results are analyzed and compared with experiment results in the time, frequency and modal domains. Results from the three injector model show how applying different velocity forcing amplitudes change the amplitude and spatial location of heat release from the center injector. The instability amplitudes in the simulation are able to be tuned to experiments and produce similar modal combustion responses of the center injector. The reaction model applied was found to play an important role in the spatial and temporal heat release response. Only when the model was calibrated to ignition delay measurements did the heat release response reflect measurements in the experiment. While insightful the simulations are not truly predictive because the driving frequency and forcing function amplitude are input into the simulation. However, the use of this approach as a tool to investigate combustion response is demonstrated. Results from the seven injector simulations provide an insightful look at the mechanisms driving the instability in the combustor. The instability was studied over a range of pressure fluctuations, up to 70% of mean chamber pressure produced in the self-exited simulation. At low amplitudes the transverse instability was found to be supported by both flame impingement with the side wall as well as vortex shedding at the primary acoustic frequency. As instability level grew the primary supporting mechanism shifted to just vortex impingement on the side walls and the greatest growth was seen as additional vortices began impinging between injector elements at the primary acoustic frequency. This research reveals the advantages and limitations of applying these two modeling techniques to simulate multiple injector experiments. The advantage of the three injector model is a simplified geometry which results in faster model development and the ability to more rapidly study the injector response under varying velocity amplitudes. The possibly faster run time is offset though by the need to run multiple cases to calibrate the model to the experiment. The model is also limited to studying the central injector effect and lacks heat release sources from the outer injectors and additional vortex interactions as shown in the seven injector simulation. The advantage of the seven injector model is that the whole domain can be explored to provide a better understanding about influential processes but does require longer development and run time due to the extensive gridding requirement. Both simulations have proven useful in exploring transverse combustion instability and show the need to further develop subscale experiments and companions simulations in developing a full-scale combustion instability prediction capability.

Analyses of Injection-Coupled Combustion Instability from J-2X Gas Generator Development

NASA Technical Reports Server (NTRS)

Hulka, James R.; Kenny, R. Jeremy; Protz, Chris; Casiano, Matthew

2011-01-01

During development of the gas generator for the liquid oxygen/liquid hydrogen propellant J-2X rocket engine, combustion instabilities were observed near the frequency of the first longitudinal acoustic mode of the hot gas combustion chamber duct. These instabilities were similar to intermediate-frequency or buzz-type instabilities as described in historical programs, except for several aspects: 1) the frequencies were low, in the realm of chug; 2) at times the instability oscillation amplitudes were quite large, with peak-to-peak amplitudes exceeding 50% of the mean chamber pressure along with the appearance of harmonics; 3) the chamber excitation was related to but not exactly at the first longitudinal combustion chamber acoustic mode; and 4) the injector provided mass flow rate oscillations induced by capacitance and inertance effects in the injector rather than by organ pipe resonances of the coaxial oxidizer posts. This type of combustion instability is referred to as "injection coupling" because one critical driving source of the instability is mass flow rate oscillations from the injector. However, the type of injection coupling observed here is different than observed in previous instances of buzz instability with coaxial injectors, because of the lower frequencies and lack of influence from the oxidizer post organ pipe resonances. Test data and preliminary analyses of the initial combustion instabilities were presented in several papers at the 5th Liquid Propulsion Subcommittee meeting. Since that time, additional hot-fire tests with several new hardware configurations have been conducted, and additional analyses have been completed. The analytical models described in previous papers have been updated to include the influences of new geometrical configurations, including a different oxidizer injector manifold configuration and a branch pipe in the hot gas duct that supplies gaseous helium during the start transient to pre-spin the turbine. In addition, the analysis methodology has been revisited to evaluate the potential influence of a combustion response as well as an injection response.

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.

Combustion Instability in an Acid-Heptane Rocket with a Pressurized-Gas Propellant Pumping System

NASA Technical Reports Server (NTRS)

Tischler, Adelbert O.; Bellman, Donald R.

1951-01-01

Results of experimental measurements of low-frequency combustion instability of a 300-pound thrust acid-heptane rocket engine were compared to the trends predicted by an analysis of combustion instability in a rocket engine with a pressurized-gas propellant pumping system. The simplified analysis, which assumes a monopropellant model, was based on the concept of a combustion the delay occurring from the moment of propellant injection to the moment of propellant combustion. This combustion time delay was experimentally measured; the experimental values were of approximately half the magnitude predicted by the analysis. The pressure-fluctuation frequency for a rocket engine with a characteristic length of 100 inches and operated at a combustion-chamber pressure of 280 pounds per square inch absolute was 38 cycles per second; the analysis indicated. a frequency of 37 cycles per second. Increasing combustion-chamber characteristic length decreased the pressure-fluctuation frequency, in conformity to the analysis. Increasing the chamber operating pressure or increasing the injector pressure drop increased the frequency. These latter two effects are contrary to the analysis; the discrepancies are attributed to the conflict between the assumptions made to simplify the analysis and the experimental conditions. Oxidant-fuel ratio had no apparent effect on the experimentally measured pressure-fluctuation frequency for acid-heptane ratios from 3.0 to 7.0. The frequencies decreased with increased amplitude of the combustion-chamber pressure variations. The analysis indicated that if the combustion time delay were sufficiently short, low-frequency combustion instability would be eliminated.

A Review of LOX/Kerosene Combustion Instability in American and Russian Combustion Devices in Application to Next-Generation Launch Technology

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Nesman, Tomas E.; Hulka, James R.; Dougherty, N. Sam

2003-01-01

The Next-Generation Launch Technology (NGLT) project was introduced with its objectives. To meet the objectives, NASA has directed aerospace industry to perform advances and risk reduction of relevant technologies, including propulsion. Originally, the propulsion industry focused on producing both LOWLH2 and LOWkerosene flight engine technology demonstrators. These flight engine technology demonstrators were briefly reviewed. NASA recently redirected this focus to Lowkerosene only. Discussion of LOWkerosene combustion devices was and is prefaced by grave concerns about combustion instability. These concerns have prompted a review of LOWkerosene combustion instability in American and Russian combustion devices. In the review of the Russian propulsion industry's experience in eliminating LOWkerosene combustion instabilities, the history of principal Russian rocket scientists and their role in the development of LOXkerosene combustion devices is presented. The innovative methods implemented by the Russians of eliminations combustion instabilities in LOXkerosene combustion devices were reviewed. The successful elimination of these combustion instabilities has resulted in two generations of Russian-produced, high-performance LOWkerosene combustion devices.

Analysis of combustion instability in liquid fuel rocket motors. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Wong, K. W.

1979-01-01

The development of an analytical technique used in the solution of nonlinear velocity-sensitive combustion instability problems is presented. The Galerkin method was used and proved successful. The pressure wave forms exhibit a strong second harmonic distortion and a variety of behaviors are possible depending on the nature of the combustion process and the parametric values involved. A one dimensional model provides insight into the problem by allowing a comparison of Galerkin solutions with more exact finite difference computations.

Control of Thermo-Acoustics Instabilities: The Multi-Scale Extended Kalman Approach

NASA Technical Reports Server (NTRS)

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

2003-01-01

"Multi-Scale Extended Kalman" (MSEK) is a novel model-based control approach recently found to be effective for suppressing combustion instabilities in gas turbines. A control law formulated in this approach for fuel modulation demonstrated steady suppression of a high-frequency combustion instability (less than 500Hz) in a liquid-fuel combustion test rig under engine-realistic conditions. To make-up for severe transport-delays on control effect, the MSEK controller combines a wavelet -like Multi-Scale analysis and an Extended Kalman Observer to predict the thermo-acoustic states of combustion pressure perturbations. The commanded fuel modulation is composed of a damper action based on the predicted states, and a tones suppression action based on the Multi-Scale estimation of thermal excitations and other transient disturbances. The controller performs automatic adjustments of the gain and phase of these actions to minimize the Time-Scale Averaged Variances of the pressures inside the combustion zone and upstream of the injector. The successful demonstration of Active Combustion Control with this MSEK controller completed an important NASA milestone for the current research in advanced combustion technologies.

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 2: a Study of Low Frequency Combustion Instability in Rocket Engine Preburners Using a Heterogeneous Stirred Tank Reactor Model. Final Report M.S. Thesis - Aug. 1987

NASA Technical Reports Server (NTRS)

Bartrand, Timothy A.

1988-01-01

During the shutdown of the space shuttle main engine, oxygen flow is shut off from the fuel preburner and helium is used to push the residual oxygen into the combustion chamber. During this process a low frequency combustion instability, or chug, occurs. This chug has resulted in damage to the engine's augmented spark igniter due to backflow of the contents of the preburner combustion chamber into the oxidizer feed system. To determine possible causes and fixes for the chug, the fuel preburner was modeled as a heterogeneous stirred tank combustion chamber, a variable mass flow rate oxidizer feed system, a constant mass flow rate fuel feed system and an exit turbine. Within the combustion chamber gases were assumed perfectly mixed. To account for liquid in the combustion chamber, a uniform droplet distribution was assumed to exist in the chamber, with mean droplet diameter determined from an empirical relation. A computer program was written to integrate the resulting differential equations. Because chamber contents were assumed perfectly mixed, the fuel preburner model erroneously predicted that combustion would not take place during shutdown. The combustion rate model was modified to assume that all liquid oxygen that vaporized instantaneously combusted with fuel. Using this combustion model, the effect of engine parameters on chamber pressure oscillations during the SSME shutdown was calculated.

Elimination of Intermediate-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Nesman, Tomas E.; Turner, Jim E. (Technical Monitor)

2001-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. The thrust chamber exhibited benign, yet marginally unstable combustion. The marginally unstable combustion was characterized by chamber pressure oscillations with large amplitudes and a frequency that was too low to be identified as acoustic or high-frequency combustion instability and too high to be identified as chug or low-frequency combustion instability. The source of the buzz or intermediate-frequency combustion instability was traced to the fuel venturi whose violently noisy cavitation caused resonance in the feedline downstream. Combustion was stabilized by increasing the throat diameter of the fuel venturi such that the cavitation would occur more quietly.

Hydrodynamic Instability in an Extended Landau/Levich Model of Liquid-Propellant Combustion at Normal and Reduced Gravity

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.

1998-01-01

The classical Landau/Levich models of liquid-propellant combustion, despite their relative simplicity, serve as seminal examples that correctly describe the onset of hydrodynamic instability in reactive systems. Recently, these two separate models have been combined and extended to account for a dynamic dependence, absent in the original formulations, of the local burning rate on the local pressure and temperature fields. The resulting model admits an extremely rich variety of both hydrodynamic and reactive/diffusive instabilities that can be analyzed either numerically or analytically in various limiting parameter regimes. In the present work, a formal asymptotic analysis, based on the realistic smallness of the gas-to-liquid density ratio, is developed to investigate the combined effects of gravity and other parameters on the hydrodynamic instability of the propagating liquid/gas interface. In particular, an analytical expression is derived for the neutral stability boundary A(sub p)(k), where A(sub p) is the pressure sensitivity of the burning rate and k is the wavenumber of the disturbance. The results demonstrate explicitly the stabilizing effect of gravity on long-wave disturbances, the stabilizing effect of viscosity (both liquid and gas) and surface tension on short-wave perturbations, and the instability associated with intermediate wavenumbers for critical negative values of A(sub p). In the limiting case of weak gravity, it is shown that hydrodynamic instability in liquid-propellant combustion is a long-wave instability phenomenon, whereas at normal gravity, this instability is first manifested through O(1) wavenumber disturbances. It is also demonstrated that, in general, surface tension and the viscosity of both the liquid and gas phases each produce comparable stabilizing effects in the large-wavenumber regime, thereby providing important modifications to previous analyses in which one or more of these effects were neglected.

Hydrodynamic Instability in an Extended Landau/Levich Model of Liquid-Propellant Combustion at Normal and Reduced Gravity

NASA Technical Reports Server (NTRS)

Margolis, S. B.

1997-01-01

The classical Landau/Levich models of liquid-propellant combustion, despite their relative simplicity, serve as seminal examples that correctly describe the onset of hydrodynamic instability in reactive systems. Recently, these two separate models have been combined and extended to account for a dynamic dependence, absent in the original formulations, of the local burning rate on the local pressure and temperature fields. The resulting model admits an extremely rich variety of both hydrodynamic and reactive/diffusive instabilities that can be analyzed either numerically or analytically in various limiting parameter regimes. In the present work, a formal asymptotic analysis, based on the realistic smallness of the gas-to-liquid density ratio, is developed to investigate the combined effects of gravity and other parameters on the hydrodynamic instability of the propagating liquid/gas interface. In particular, an analytical expression is derived for the neutral stability boundary A(p)(k), where A(p) is the pressure sensitivity of the burning rate and k is the wavenumber of the disturbance. The results demonstrate explicitly the stabilizing effect of gravity on long-wave disturbances, the stabilizing effect of viscosity (both liquid and gas) and surface tension on short-wave perturbations, and the instability associated with intermediate wavenumbers for negative values of A(p). In the limiting case of weak gravity, it is shown that hydrodynamic instability in liquid-propellant combustion is a long-wave instability phenomenon, whereas at normal gravity, this instability is first manifested through O(1) wavenumber disturbances. it is also demonstrated that, in general, surface tension and the viscosity of both the liquid and gas phases each produce comparable stabilizing effects in the long-wavenumber regime, thereby providing important modifications to previous analyses in which one or more of these effects were neglected.

LOX/Hydrocarbon Combustion Instability Investigation

NASA Technical Reports Server (NTRS)

Jensen, R. J.; Dodson, H. C.; Claflin, S. E.

1989-01-01

The LOX/Hydrocarbon Combustion Instability Investigation Program was structured to determine if the use of light hydrocarbon combustion fuels with liquid oxygen (LOX) produces combustion performance and stability behavior similar to the LOX/hydrogen propellant combination. In particular methane was investigated to determine if that fuel can be rated for combustion instability using the same techniques as previously used for LOX/hydrogen. These techniques included fuel temperature ramping and stability bomb tests. The hot fire program probed the combustion behavior of methane from ambient to subambient temperatures. Very interesting results were obtained from this program that have potential importance to future LOX/methane development programs. A very thorough and carefully reasoned documentation of the experimental data obtained is contained. The hot fire test logic and the associated tests are discussed. Subscale performance and stability rating testing was accomplished using 40,000 lb. thrust class hardware. Stability rating tests used both bombs and fuel temperature ramping techniques. The test program was successful in generating data for the evaluation of the methane stability characteristics relative to hydrogen and to anchor stability models. Data correlations, performance analysis, stability analyses, and key stability margin enhancement parameters are discussed.

A novel approach to predict the stability limits of combustion chambers with large eddy simulation

NASA Astrophysics Data System (ADS)

Pritz, B.; Magagnato, F.; Gabi, M.

2010-06-01

Lean premixed combustion, which allows for reducing the production of thermal NOx, is prone to combustion instabilities. There is an extensive research to develop a reduced physical model, which allows — without time-consuming measurements — to calculate the resonance characteristics of a combustion system consisting of Helmholtz resonator type components (burner plenum, combustion chamber). For the formulation of this model numerical investigations by means of compressible Large Eddy Simulation (LES) were carried out. In these investigations the flow in the combustion chamber is isotherm, non-reacting and excited with a sinusoidal mass flow rate. Firstly a combustion chamber as a single resonator subsequently a coupled system of a burner plenum and a combustion chamber were investigated. In this paper the results of additional investigations of the single resonator are presented. The flow in the combustion chamber was investigated without excitation at the inlet. It was detected, that the mass flow rate at the outlet cross section is pulsating once the flow in the chamber is turbulent. The fast Fourier transform of the signal showed that the dominant mode is at the resonance frequency of the combustion chamber. This result sheds light on a very important source of self-excited combustion instabilities. Furthermore the LES can provide not only the damping ratio for the analytical model but the eigenfrequency of the resonator also.

Development of a computational testbed for numerical simulation of combustion instability

NASA Technical Reports Server (NTRS)

Grenda, Jeffrey; Venkateswaran, Sankaran; Merkle, Charles L.

1993-01-01

A synergistic hierarchy of analytical and computational fluid dynamic techniques is used to analyze three-dimensional combustion instabilities in liquid rocket engines. A mixed finite difference/spectral procedure is employed to study the effects of a distributed vaporization zone on standing and spinning instability modes within the chamber. Droplet atomization and vaporization are treated by a variety of classical models found in the literature. A multi-zone, linearized analytical solution is used to validate the accuracy of the numerical simulations at small amplitudes for a distributed vaporization region. This comparison indicates excellent amplitude and phase agreement under both stable and unstable operating conditions when amplitudes are small and proper grid resolution is used. As amplitudes get larger, expected nonlinearities are observed. The effect of liquid droplet temperature fluctuations was found to be of critical importance in driving the instabilities of the combustion chamber.

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

High-frequency combustion instability control through acoustic modulation at the inlet boundary for liquid rocket engine applications

NASA Astrophysics Data System (ADS)

Bennewitz, John William

This research investigation encompasses experimental tests demonstrating the control of a high-frequency combustion instability by acoustically modulating the propellant flow. A model rocket combustor burned gaseous oxygen and methane using a single-element, pentad-style injector. Flow conditions were established that spontaneously excited a 2430 Hz first longitudinal combustion oscillation at an amplitude up to p'/pc ≈ 6%. An acoustic speaker was placed at the base of the oxidizer supply to modulate the flow and alter the oscillatory behavior of the combustor. Two speaker modulation approaches were investigated: (1) Bands of white noise and (2) Pure sinusoidal tones. The first approach adjusted 500 Hz bands of white noise ranging from 0-500 Hz to 2000-2500 Hz, while the second implemented single-frequency signals with arbitrary phase swept from 500-2500 Hz. The results showed that above a modulation signal amplitude threshold, both approaches suppressed 95+% of the spontaneous combustion oscillation. By increasing the applied signal amplitude, a wider frequency range of instability suppression became present for these two acoustic modulation approaches. Complimentary to these experiments, a linear modal analysis was undertaken to investigate the effects of acoustic modulation at the inlet boundary on the longitudinal instability modes of a dump combustor. The modal analysis employed acoustically consistent matching conditions with a specific impedance boundary condition at the inlet to represent the acoustic modulation. From the modal analysis, a naturally unstable first longitudinal mode was predicted in the absence of acoustic modulation, consistent with the spontaneously excited 2430 Hz instability observed experimentally. Subsequently, a detailed investigation involving variation of the modulation signal from 0-2500 Hz and mean combustor temperature from 1248-1685 K demonstrated the unstable to stable transition of a 2300-2500 Hz first longitudinal mode. The model-predicted mode stability transition was consistent with experimental observations, supporting the premise that inlet acoustic modulation is a means to control high-frequency combustion instabilities. From the modal analysis, it may be deduced that the inlet impedance provides a damping mechanism for instability suppression. Combined, this work demonstrates the strategic application of acoustic modulation within an injector as a potential method to control high-frequency combustion instabilities for liquid rocket engine applications.

Causes of Combustion Instabilities with Passive and Active Methods of Control for practical application to Gas Turbine Engines

NASA Astrophysics Data System (ADS)

Cornwell, Michael D.

Combustion at high pressure in applications such as rocket engines and gas turbine engines commonly experience destructive combustion instabilities. These instabilities results from interactions between combustion heat release, fluid mechanics and acoustics. This research explores the significant affect of unstable fluid mechanics processes in augmenting unstable periodic combustion heat release. The frequency of the unstable heat release may shift to match one of the combustors natural acoustic frequencies which then can result in significant energy exchange from chemical to acoustic energy resulting in thermoacoustic instability. The mechanisms of the fluid mechanics in coupling combustion to acoustics are very broad with many varying mechanisms explained in detail in the first chapter. Significant effort is made in understanding these mechanisms in this research in order to find commonalities, useful for mitigating multiple instability mechanisms. The complexity of combustion instabilities makes mitigation of combustion instabilities very difficult as few mitigation methods have historically proven to be very effective for broad ranges of combustion instabilities. This research identifies turbulence intensity near the forward stagnation point and movement of the forward stagnation point as a common link in what would otherwise appear to be very different instabilities. The most common method of stabilization of both premixed and diffusion flame combustion is through the introduction of swirl. Reverse flow along the centerline is introduced to transport heat and chemically active combustion products back upstream to sustain combustion. This research develops methods to suppress the movement of the forward stagnation point without suppressing the development of the vortex breakdown process which is critical to the transport of heat and reactive species necessary for flame stabilization. These methods are useful in suppressing the local turbulence at the forward stagnation point, limiting dissipation of heat and reactive species significantly improving stability. Combustion hardware is developed and tested to demonstrate the stability principles developed as part of this research. In order to more completely understand combustion instability a very unique method of combustion was researched where there are no discrete points of combustion initiation such as the forward stagnation point typical in many combustion systems including swirl and jet wake stabilized combustion. This class of combustion which has empirical evidence of great stability and efficient combustion with low CO, NOx and UHC emissions is described as high oxidization temperature distributed combustion. This mechanism of combustion is shown to be stable largely because there are no stagnations points susceptible to fluid mechanic perturbations. The final topic of research is active combustion control by fuel modulation. This may be the only practical method of controlling most instabilities with a single technique. As there are many papers reporting active combustion control algorithms this research focused on the complexities of the physics of fuel modulation at frequencies up to 1000 Hz with proportionally controlled flow amplitude. This research into the physics of high speed fluid movement, oscillation mechanical mechanisms and electromagnetics are demonstrated by development and testing of a High Speed Latching Oscillator Valve.

Active control: an investigation method for combustion instabilities

NASA Astrophysics Data System (ADS)

Poinsot, T.; Yip, B.; Veynante, D.; Trouvé, A.; Samaniego, J. M.; Candel, S.

1992-07-01

Closed-loop active control methods and their application to combustion instabilities are discussed. In these methods the instability development is impeded with a feedback control loop: the signal provided by a sensor monitoring the flame or pressure oscillations is processed and sent back to actuators mounted on the combustor or on the feeding system. Different active control systems tested on a non-premixed multiple-flame turbulent combustor are described. These systems can suppress all unstable plane modes of oscillation (i.e. low frequency modes). The active instability control (AIC) also constitutes an original and powerful technique for studies of mechanisms leading to instability or resulting from the instability. Two basic applications of this kind are described. In the first case the flame is initially controlled with AIC, the feedback loop is then switched off and the growth of the instability is analysed through high speed Schlieren cinematography and simultaneous sound pressure and reaction rate measurements. Three phases are identified during th growth of the oscillations: (1) a linear phase where acoustic waves induce a flapping motion of the flame sheets without interaction between sheets, (2) a modulation phase, where flame sheets interact randomly and (3) a nonlinear phase where the flame sheets are broken and a limit cycle is reached. In the second case we investigate different types of flame extinctions associated with combustion instability. It is shown that pressure oscillations may lead to partial or total extinctions. Extinctions occur in various forms but usually follow a rapid growth of pressure oscillations. The flame is extinguished during the modulation phase observed in the initiation experiments. In these studies devoted to transient instability phenomena, the control system constitutes a unique investigation tool because it is difficult to obtain the same information by other means. Implications for modelling and prediction of combustion instabilities are discussed.

Mechanism of instabilities in turbulent combustion leading to flashback

NASA Astrophysics Data System (ADS)

Keller, J. O.; Vaneveld, L.; Ghoniem, A. F.; Daily, J. W.; Oppenheim, A. K.; Korschelt, D.; Hubbard, G. L.

1981-01-01

High-speed schlieren cinematography, combined with synchronized pressure transducer records, was used to investigate the mechanism of combustion instabilities leading to flashback. The combustion chamber had an oblong rectangular cross-section to model the essential features of planar flow, and was provided with a rearward facing step acting as a flameholder. As the rich limit was approached, three instability modes were observed: (1) humming - a significant increase in the amplitude of the vortex pattern; (2) buzzing - a large-scale oscillation of the flame; and (3) chucking - a cyclic reformation of the flame, which results in flashback. The mechanism of these phenomena is ascribed to the action of vortices in the recirculation zone and their interactions with the trailing vortex pattern of the turbulent mixing layer behind the step.

Hydrodynamic Instability and Thermal Coupling in a Dynamic Model of Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, S. B.

1999-01-01

For liquid-propellant combustion, the Landau/Levich hydrodynamic models have been combined and extended to account for a dynamic dependence of the burning rate on the local pressure and temperature fields. Analysis of these extended models is greatly facilitated by exploiting the realistic smallness of the gas-to-liquid density ratio rho. Neglecting thermal coupling effects, an asymptotic expression was then derived for the cellular stability boundary A(sub p)(k) where A(sub p) is the pressure sensitivity of the burning rate and k is the disturbance wavenumber. The results explicitly indicate the stabilizing effects of gravity on long-wave disturbances, and those of viscosity and surface tension on short-wave perturbations, and the instability associated with intermediate wavenumbers for critical negative values of A(sub p). In the limit of weak gravity, hydrodynamic instability in liquid-propellant combustion becomes a long-wave, instability phenomenon, whereas at normal gravity, this instability is first manifested through O(1) wavenumbers. In addition, surface tension and viscosity (both liquid and gas) each produce comparable effects in the large-wavenumber regime, thereby providing important modifications to the previous analyses in which one or more of these effects was neglected. For A(sub p)= O, the Landau/Levich results are recovered in appropriate limiting cases, although this typically corresponds to a hydrodynamically unstable parameter regime for p << 1. In addition to the classical cellular form of hydrodynamic stability, there exists a pulsating form corresponding to the loss of stability of steady, planar burning to time-dependent perturbations. This occurs for negative values of the parameter A(sub p), and is thus absent from the original Landau/Levich models. In the extended model, however, there exists a stable band of negative pressure sensitivities bounded above by the Landau type of instability, and below by this pulsating form of hydrodynamic instability. Indeed, nonsteady modes of combustion have been observed at low pressures in hydroxylammonium nitrate (HAN)-based liquid propellants, which often exhibit negative pressure sensitivities. While nonsteady combustion may correspond to secondary and higher-order bifurcations above the cellular boundary, it may also be a manifestation of this pulsating type of hydrodynamic instability. In the present work, a nonzero temperature sensitivity is incorporated into our previous asymptotic analyses. This entails a coupling of the energy equation to the previous purely hydrodynamic problem, and leads to a significant modification of the pulsating boundary such that, for sufficiently large values of the temperature-sensitivity parameter, liquid-propellant combustion can become intrinsically unstable to this alternative form of hydrodynamic instability. For simplicity, further attention is confined here to the inviscid version of the problem since, despite the fact that viscous and surface-tension effects are comparable, the qualitative nature of the cellular boundary remains preserved in the zero-viscosity limit, as does the existence of the pulsating boundary. The mathematical model adopts the classical assumption that there is no distributed reaction in either the liquid or gas phases, but now the reaction sheet, representing either a pyrolysis reaction or an exothermic decomposition at the liquid/gas interface, is assumed to depend on local conditions there.

A methodology to study the possible occurrence of chugging in liquid rocket engines during transient start-up

NASA Astrophysics Data System (ADS)

Leonardi, Marco; Nasuti, Francesco; Di Matteo, Francesco; Steelant, Johan

2017-10-01

An investigation on the low frequency combustion instabilities due to the interaction of combustion chamber and feed line dynamics in a liquid rocket engine is carried out implementing a specific module in the system analysis software EcosimPro. The properties of the selected double time lag model are identified according to the two classical assumptions of constant and variable time lag. Module capabilities are evaluated on a literature experimental set up consisting of a combustion chamber decoupled from the upstream feed lines. The computed stability map results to be in good agreement with both experimental data and analytical models. Moreover, the first characteristic frequency of the engine is correctly predicted, giving confidence on the use of the module for the analysis of chugging instabilities. As an example of application, a study is carried out on the influence of the feed lines on the system stability, correctly capturing that the lines extend the stable regime of the combustion chamber and that the propellant domes play a key role in coupling the dynamics of combustion chamber and feed lines. A further example is presented to discuss on the role of pressure growth rate and of the combustion chamber properties on the possible occurrence of chug instability during engine start-up and on the conditions that lead to its damping or growth.

Dynamic properties of combustion instability in a lean premixed gas-turbine combustor.

PubMed

Gotoda, Hiroshi; Nikimoto, Hiroyuki; Miyano, Takaya; Tachibana, Shigeru

2011-03-01

We experimentally investigate the dynamic behavior of the combustion instability in a lean premixed gas-turbine combustor from the viewpoint of nonlinear dynamics. A nonlinear time series analysis in combination with a surrogate data method clearly reveals that as the equivalence ratio increases, the dynamic behavior of the combustion instability undergoes a significant transition from stochastic fluctuation to periodic oscillation through low-dimensional chaotic oscillation. We also show that a nonlinear forecasting method is useful for predicting the short-term dynamic behavior of the combustion instability in a lean premixed gas-turbine combustor, which has not been addressed in the fields of combustion science and physics.

Effects of Fuel Spray Modeling on Combustion Instability Predictions in a Single-Element Lean Direct Injection (LDI) Gas Turbine Combustor

DTIC Science & Technology

2014-09-01

evaporation in the vicinity of the injector . Recently, Kim and Menon 9 applied the same approach to study the characteristics of longitudinal...phenomena that govern the occurrence of combustion instabilities. The experiments involve a single injector element in a longitudinal mode combustor with...well characterized inflow conditions and a choked nozzle exit condition. Varying parameters such as the length of the air plenum, the combustor length

Modeling of microgravity combustion experiments

NASA Technical Reports Server (NTRS)

Buckmaster, John

1995-01-01

This program started in February 1991, and is designed to improve our understanding of basic combustion phenomena by the modeling of various configurations undergoing experimental study by others. Results through 1992 were reported in the second workshop. Work since that time has examined the following topics: Flame-balls; Intrinsic and acoustic instabilities in multiphase mixtures; Radiation effects in premixed combustion; Smouldering, both forward and reverse, as well as two dimensional smoulder.

Navier-Stokes Entropy Controlled Combustion Instability Analysis for Liquid Propellants

NASA Technical Reports Server (NTRS)

Chung, T. J.; Yoon, W. S.

1990-01-01

Navier-Stokes solutions are used to calculate oscillatory components of pressure, velocity, and density, which in turn provide necessary data to compute energy growth factors to determine combustion instability. It is shown that wave instabilities are associated with changes in entropy and the space and time averages of oscillatory components of pressure, velocity and density, together with the mean flow field in the energy equation. Compressible laminar and turbulent flows and reacting flows with hydrogen/oxygen combustion are considered. The SSME combustion/thrust chamber is used for illustration of the theory. The analysis shows that the increase of mean pressure and disturbances consistently results in the increase of instability. It is shown that adequate combustion instability analysis requires at least third order nonlinearity in energy growth or decay.

Method and apparatus for detecting combustion instability in continuous combustion systems

DOEpatents

Benson, Kelly J.; Thornton, Jimmy D.; Richards, George A.; Straub, Douglas L.

2006-08-29

An apparatus and method to sense the onset of combustion stability is presented. An electrode is positioned in a turbine combustion chamber such that the electrode is exposed to gases in the combustion chamber. A control module applies a voltage potential to the electrode and detects a combustion ionization signal and determines if there is an oscillation in the combustion ionization signal indicative of the occurrence of combustion stability or the onset of combustion instability. A second electrode held in a coplanar but spaced apart manner by an insulating member from the electrode provides a combustion ionization signal to the control module when the first electrode fails. The control module broadcasts a notice if the parameters indicate the combustion process is at the onset of combustion instability or broadcasts an alarm signal if the parameters indicate the combustion process is unstable.

Pulsating Hydrodynamic Instability in a Dynamic Model of Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sacksteder, Kurt (Technical Monitor)

1999-01-01

Hydrodynamic (Landau) instability in combustion is typically associated with the onset of wrinkling of a flame surface, corresponding to the formation of steady cellular structures as the stability threshold is crossed. In the context of liquid-propellant combustion, such instability has recently been shown to occur for critical values of the pressure sensitivity of the burning rate and the disturbance wavenumber, significantly generalizing previous classical results for this problem that assumed a constant normal burning rate. Additionally, however, a pulsating form of hydrodynamic instability has been shown to occur as well, corresponding to the onset of temporal oscillations in the location of the liquid/gas interface. In the present work, we consider the realistic influence of a nonzero temperature sensitivity in the local burning rate on both types of stability thresholds. It is found that for sufficiently small values of this parameter, there exists a stable range of pressure sensitivities for steady, planar burning such that the classical cellular form of hydrodynamic instability and the more recent pulsating form of hydrodynamic instability can each occur as the corresponding stability threshold is crossed. For larger thermal sensitivities, however, the pulsating stability boundary evolves into a C-shaped curve in the disturbance-wavenumber/ pressure-sensitivity plane, indicating loss of stability to pulsating perturbations for all sufficiently large disturbance wavelengths. It is thus concluded, based on characteristic parameter values, that an equally likely form of hydrodynamic instability in liquid-propellant combustion is of a nonsteady, long-wave nature, distinct from the steady, cellular form originally predicted by Landau.

Pulsating Hydrodynamic Instability and Thermal Coupling in an Extended Landau/Levich Model of Liquid-Propellant Combustion. 1; Inviscid Analysis

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sacksteder, Kurt (Technical Monitor)

1999-01-01

Hydrodynamic (Landau) instability in combustion is typically associated with the onset of wrinkling of a flame surface, corresponding to the formation of steady cellular structures as the stability threshold is crossed. In the context of liquid-propellant combustion, such instability has recently been shown to occur for critical values of the pressure sensitivity of the burning rate and the disturbance wavenumber, significantly generalizing previous classical results for this problem that assumed a constant normal burning rate. Additionally, however, a pulsating form of hydrodynamic instability has been shown to occur as well, corresponding to the onset of temporal oscillations in the location of the liquid/gas interface. In the present work, we consider the realistic influence of a non-zero temperature sensitivity in the local burning rate on both types of stability thresholds. It is found that for sufficiently small values of this parameter, there exists a stable range of pressure sensitivities for steady, planar burning such that the classical cellular form of hydrodynamic instability and the more recent pulsating form of hydrodynamic instability can each occur as the corresponding stability threshold is crossed. For larger thermal sensitivities, however, the pulsating stability boundary evolves into a C-shaped curve in the (disturbance-wavenumber, pressure-sensitivity) plane, indicating loss of stability to pulsating perturbations for all sufficiently large disturbance wavelengths. It is thus concluded, based on characteristic parameter values, that an equally likely form of hydrodynamic instability in liquid-propellant combustion is of a non-steady, long-wave nature, distinct from the steady, cellular form originally predicted by Landau.

Liquid propellant rocket combustion instability

NASA Technical Reports Server (NTRS)

Harrje, D. T.

1972-01-01

The solution of problems of combustion instability for more effective communication between the various workers in this field is considered. The extent of combustion instability problems in liquid propellant rocket engines and recommendations for their solution are discussed. The most significant developments, both theoretical and experimental, are presented, with emphasis on fundamental principles and relationships between alternative approaches.

Triggering of longitudinal combustion instabilities in solid rocket motors: Nonlinear combustion response

NASA Technical Reports Server (NTRS)

Wicker, J. M.; Greene, W. D.; Kim, S. I.; Yang, V.

1995-01-01

Pulsed oscillations in solid rocket motors are investigated with emphasis on nonlinear combustion response. The study employs a wave equation governing the unsteady motions in a two-phase flow, and a solution technique based on spatial- and time-averaging. A wide class of combustion response functions is studied to second-order in fluctuation amplitude to determine if, when, and how triggered instabilities arise. Conditions for triggering are derived from analysis of limit cycles, and regions of triggering are found in parametric space. Based on the behavior of model dynamical systems, introduction of linear cross-coupling and quadratic self-coupling among the acoustic modes appears to be the manner in which the nonlinear combustion response produces triggering to a stable limit cycle. Regions of initial conditions corresponding to stable pulses were found, suggesting that stability depends on initial phase angle and harmonic content, as well as the composite amplitude, of the pulse.

Extensions to the time lag models for practical application to rocket engine stability design

NASA Astrophysics Data System (ADS)

Casiano, Matthew J.

The combustion instability problem in liquid-propellant rocket engines (LREs) has remained a tremendous challenge since their discovery in the 1930s. Improvements are usually made in solving the combustion instability problem primarily using computational fluid dynamics (CFD) and also by testing demonstrator engines. Another approach is to use analytical models. Analytical models can be used such that design, redesign, or improvement of an engine system is feasible in a relatively short period of time. Improvements to the analytical models can greatly aid in design efforts. A thorough literature review is first conducted on liquid-propellant rocket engine (LRE) throttling. Throttling is usually studied in terms of vehicle descent or ballistic missile control however there are many other cases where throttling is important. It was found that combustion instabilities are one of a few major issues that occur during deep throttling (other major issues are heat transfer concerns, performance loss, and pump dynamics). In the past and again recently, gas injected into liquid propellants has shown to be a viable solution to throttle engines and to eliminate some forms of combustion instability. This review uncovered a clever solution that was used to eliminate a chug instability in the Common Extensible Cryogenic Engine (CECE), a modified RL10 engine. A separate review was also conducted on classic time lag combustion instability models. Several new stability models are developed by incorporating important features to the classic and contemporary models, which are commonly used in the aerospace rocket industry. The first two models are extensions of the original Crocco and Cheng concentrated combustion model with feed system contributions. A third new model is an extension to the Wenzel and Szuch double-time lag model also with feed system contributions. The first new model incorporates the appropriate injector acoustic boundary condition which is neglected in contemporary models. This new feature shows that the injector boundary can play a significant role for combustion stability, especially for gaseous injection systems or a system with an injector orifice on the order of the size of the chamber. The second new model additionally accounts for resistive effects. Advanced signal analysis techniques are used to extract frequency-dependent damping from a gas generator component data set. The damping values are then used in the new stability model to more accurately represent the chamber response of the component. The results show a more realistic representation of stability margin by incorporating the appropriate damping effects into the chamber response from data. The original Crocco model, a contemporary model, and the two new models are all compared and contrasted to a marginally stable test case showing their applicability. The model that incorporates resistive aspects shows the best comparison to the test data. Parametrics are also examined to show the influence of the new features and their applicability. The new features allow a more accurate representation of stability margin to be obtained. The third new model is an extension to the Wenzel and Szuch double-time lag chug model. The feed system chug model is extended to account for generic propellant flow rates. This model is also extended to incorporate aspects due to oxygen boiling and helium injection in the feed system. The solutions to the classic models, for the single-time lag and the double-time lag models, are often plotted on a practical engine operating map, however the models have presented some difficulties for numerical algorithms for several reasons. Closed-form solutions for use on these practical operating maps are formulated and developed. These models are incorporated in a graphical user interface tool and the new model is compared to an extensive data set. It correctly predicts the stability behavior at various operating conditions incorporating the influence of injected helium and boiling oxygen in the feed system.

Combustion Stability Assessments of the Black Brant Solid Rocket Motor

NASA Technical Reports Server (NTRS)

Fischbach, Sean

2014-01-01

The Black Brant variation of the Standard Brant developed in the 1960's has been a workhorse motor of the NASA Sounding Rocket Project Office (SRPO) since the 1970's. In March 2012, the Black Brant Mk1 used on mission 36.277 experienced combustion instability during a flight at White Sands Missile Range, the third event in the last four years, the first occurring in November, 2009, the second in April 2010. After the 2010 event the program has been increasing the motor's throat diameter post-delivery with the goal of lowering the chamber pressure and increasing the margin against combustion instability. During the most recent combustion instability event, the vibrations exceeded the qualification levels for the Flight Termination System. The present study utilizes data generated from T-burner testing of multiple Black Brant propellants at the Naval Air Warfare Center at China Lake, to improve the combustion stability predictions for the Black Brant Mk1 and to generate new predictions for the Mk2. Three unique one dimensional (1-D) stability models were generated, representing distinct Black Brant flights, two of which experienced instabilities. The individual models allowed for comparison of stability characteristics between various nozzle configurations. A long standing "rule of thumb" states that increased stability margin is gained by increasing the throat diameter. In contradiction to this experience based rule, the analysis shows that little or no margin is gained from a larger throat diameter. The present analysis demonstrates competing effects resulting from an increased throat diameter accompanying a large response function. As is expected, more acoustic energy was expelled through the nozzle, but conversely more acoustic energy was generated due to larger gas velocities near the propellant surfaces.

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

NASA Technical Reports Server (NTRS)

Hulka, James R.; Jones, Gregg W.

2010-01-01

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

Active suppression of vortex-driven combustion instability using controlled liquid-fuel injection

NASA Astrophysics Data System (ADS)

Pang, Bin

Combustion instabilities remain one of the most challenging problems encountered in developing propulsion and power systems. Large amplitude pressure oscillations, driven by unsteady heat release, can produce numerous detrimental effects. Most previous active control studies utilized gaseous fuels to suppress combustion instabilities. However, using liquid fuel to suppress combustion instabilities is more realistic for propulsion applications. Active instability suppression in vortex-driven combustors using a direct liquid fuel injection strategy was theoretically established and experimentally demonstrated in this dissertation work. Droplet size measurements revealed that with pulsed fuel injection management, fuel droplet size could be modulated periodically. Consequently, desired heat release fluctuation could be created. If this oscillatory heat release is coupled with the natural pressure oscillation in an out of phase manner, combustion instabilities can be suppressed. To identify proper locations of supplying additional liquid fuel for the purpose of achieving control, the natural heat release pattern in a vortex-driven combustor was characterized in this study. It was found that at high Damkohler number oscillatory heat release pattern closely followed the evolving vortex front. However, when Damkohler number became close to unity, heat release fluctuation wave no longer coincided with the coherent structures. A heat release deficit area was found near the dump plane when combustor was operated in lean premixed conditions. Active combustion instability suppression experiments were performed in a dump combustor using a controlled liquid fuel injection strategy. High-speed Schlieren results illustrated that vortex shedding plays an important role in maintaining self-sustained combustion instabilities. Complete combustion instability control requires total suppression of these large-scale coherent structures. The sound pressure level at the excited dominant frequency was reduced by more than 20 dB with controlled liquid fuel injection method. Scaling issues were also investigated in this dump combustor to test the effectiveness of using pulsed liquid fuel injection strategies to suppress instabilities at higher power output conditions. With the liquid fuel injection control method, it was possible to suppress strong instabilities with initial amplitude of +/-5 psi down to the background noise level. The stable combustor operating range was also expanded from equivalence ratio of 0.75 to beyond 0.9.

Hydrodynamic Instability in an Extended Landau/Levich Model of Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sackesteder, Kurt (Technical Monitor)

1998-01-01

The classical Landau/Levich models of liquid propellant combustion, which serve as seminal examples of hydrodynamic instability in reactive systems, have been combined and extended to account for a dynamic dependence, absent in the original formulations, of the local burning rate on the local pressure and/or temperature fields. The resulting model admits an extremely rich variety of both hydrodynamic and reactive/diffusive instabilities that can be analyzed in various limiting parameter regimes. In the present work, a formal asymptotic analysis, based on the realistic smallness of the gas-to-liquid density ratio, is developed to investigate the combined effects of gravity, surface tension and viscosity on the hydrodynamic instability of the propagating liquid/gas interface. In particular, a composite asymptotic expression, spanning three distinguished wavenumber regimes, is derived for both cellular and pulsating hydrodynamic neutral stability boundaries A(sub p)(k), where A(sub p) is the pressure sensitivity of the burning rate and k is the disturbance wavenumber. For the case of cellular (Landau) instability, the results demonstrate explicitly the stabilizing effect of gravity on long-wave disturbances, the stabilizing effect of viscosity and surface tension on short-wave perturbations, and the instability associated with intermediate wavenumbers for critical negative values of A(sub p). In the limiting case of weak gravity, it is shown that cellular hydrodynamic instability in this context is a long-wave instability phenomenon, whereas at normal gravity, this instability is first manifested through O(l) wavenumber disturbances. It is also demonstrated that, in the large wavenumber regime, surface tension and both liquid and gas viscosity all produce comparable stabilizing effects in the large-wavenumber regime, thereby providing significant modifications to previous analyses of Landau instability in which one or more of these effects were neglected. In contrast, the pulsating hydrodynamic stability boundary is found to be insensitive to gravitational and surface-tension effects, but is more sensitive to the effects of liquid viscosity, which is a significant stabilizing effect for O(l) and higher wavenumbers. Liquid-propellant combustion is predicted to be stable (i.e., steady and planar) only for a range of negative pressure sensitivities that lie between the two types of hydrodynamic stability boundaries.

Transient processes in the combustion of nitramine propellants

NASA Technical Reports Server (NTRS)

Cohen, N. S.; Strand, L. D.

1978-01-01

A transient combustion model of nitramine propellants is combined with an isentropic compression shock formation model to determine the role of nitramine propellant combustion in DDT, excluding effects associated with propellant structural properties or mechanical behavior. The model is derived to represent the closed pipe experiment that is widely used to characterize explosives, except that the combustible material is a monolithic charge rather than compressed powder. Computations reveal that the transient combustion process cannot by itself produce DDT by this model. Compressibility of the solid at high pressure is the key factor limiting pressure buildups created by the combustion. On the other hand, combustion mechanisms which promote pressure buildups are identified and related to propellant formulation variables. Additional combustion instability data for nitramine propellants are presented. Although measured combustion response continues to be low, more data are required to distinguish HMX and active binder component contributions. A design for a closed vessel apparatus for experimental studies of high pressure combustion is discussed.

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

PubMed Central

Morgans, Aimee S.

2016-01-01

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

Advances in Turbulent Combustion Dynamics Simulations in Bluff-Body Stabilized Flames-Body Stabilized Flames

DTIC Science & Technology

2015-11-30

Master’s Thesis 3. DATES COVERED (From - To) 01 Nov 2015 – 30 Nov 2015 4. TITLE AND SUBTITLE Advances in Turbulent Combustion Dynamics Simulations...the three main aspects of bluff-body stabilized flames: stationary combustion , lean blow-out, and thermo-acoustic instabilities. For the cases of...stationary combustion and lean blow-out, an improved version of the Linear Eddy Model approach is used, while in the case of thermo-acoustic

Control Strategies for HCCI Mixed-Mode Combustion

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

Wagner, Robert M; Edwards, Kevin Dean

2010-03-01

Delphi Automotive Systems and ORNL established this CRADA to expand the operational range of Homogenous Charge Compression Ignition (HCCI) mixed-mode combustion for gasoline en-gines. ORNL has extensive experience in the analysis, interpretation, and control of dynamic engine phenomena, and Delphi has extensive knowledge and experience in powertrain compo-nents and subsystems. The partnership of these knowledge bases was important to address criti-cal barriers associated with the realistic implementation of HCCI and enabling clean, efficient operation for the next generation of transportation engines. The foundation of this CRADA was established through the analysis of spark-assisted HCCI data from a single-cylinder research engine.more » This data was used to (1) establish a conceptual kinetic model to better understand and predict the development of combustion instabilities, (2) develop a low-order model framework suitable for real-time controls, and (3) provide guidance in the initial definition of engine valve strategies for achieving HCCI operation. The next phase focused on the development of a new combustion metric for real-time characterization of the combustion process. Rapid feedback on the state of the combustion process is critical to high-speed decision making for predictive control. Simultaneous to the modeling/analysis studies, Delphi was focused on the development of engine hardware and the engine management system. This included custom Delphi hardware and control systems allowing for flexible control of the valvetrain sys-tem to enable HCCI operation. The final phase of this CRADA included the demonstration of conventional and spark assisted HCCI on the multi-cylinder engine as well as the characterization of combustion instabilities, which govern the operational boundaries of this mode of combustion. ORNL and Delphi maintained strong collaboration throughout this project. Meetings were held on a bi-weekly basis with additional reports, presentation, and meetings as necessary to maintain progress. Delphi provided substantial support through modeling, hardware, data exchange, and technical consultation. This CRADA was also successful at establishing important next steps to further expanding the use of an HCCI engine for improved fuel efficiency and emissions. These topics will be address in a follow-on CRADA. The objectives are: (1) Improve fundamental understanding of the development of combustion instabilities with HCCI operation through modeling and experiments; (2) Develop low-order model and feedback combustion metrics which are well suited to real-time predictive controls; and (3) Construct multi-cylinder engine system with advanced Delphi technologies and charac-terize HCCI behavior to better understand limitations and opportunities for expanded high-efficiency operation.« less

Operating condition and geometry effects on low-frequency afterburner combustion instability in a turbofan at altitude

NASA Technical Reports Server (NTRS)

Cullom, R. R.; Johnsen, R. L.

1979-01-01

Three afterburner configurations were tested in a low-bypass-ratio turbofan engine to determine the effect of various fuel distributions, inlet conditions, flameholder geometry, and fuel injection location on combustion instability. Tests were conducted at simulated flight conditions of Mach 0.75 and 1.3 at altitudes from 11,580 to 14,020 m (38,000 to 46,000 ft). In these tests combustion instability with frequency from 28 to 90 Hz and peak-to-peak pressure amplitude up to 46.5 percent of the afterburner inlet total pressure level was encountered. Combustion instability was suppressed in these tests by varying the fuel distribution in the afterburner.

Large Eddy Simulations of Transverse Combustion Instability in a Multi-Element Injector

DTIC Science & Technology

2016-07-27

plagued the development of liquid rocket engines and remains a large riskin the development and acquisition of new liquid rocket engines. Combustion...simulations to better understand the physics that can lead combustion instability in liquid rocket engines. Simulations of this type are able to...instabilities found in liquid rocket engines are transverse. The motivating of the experiment behind the current work is to subject the CVRC injector

Space Propulsion and Power

DTIC Science & Technology

2013-03-08

crystals with tunable band gaps possible Refractive index N is imaginary - Bulk Electromagnetic waves cannot propogate But surface plasmons...Directional wave radiation through plasmon resonances Directional wave guiding through mid-band defect wave localization Distribution A: Approved for... acoustic damping, shear- layer instability (PERTURBATION EXPANSION EXAMPLE) classical wave equation for combustion instability: model

A Study of Flame Physics and Solid Propellant Rocket Physics

DTIC Science & Technology

2007-10-01

and ellipsoids, and the packing of pellets relevant to igniter modeling. Other topics are the instabilities of smolder waves, premixed flame...instabilities in narrow tubes, and flames supported by a spinning porous plug burner . Much of this work has been reported in the high-quality archival...perchlorate in fuel binder, the combustion of model propellant packs of ellipses and ellipsoids, and the packing of pellets relevant to igniter modeling

Review of Combustion-acoustic Instabilities

NASA Technical Reports Server (NTRS)

Oyediran, Ayo; Darling, Douglas; Radhakrishnan, Krishnan

1995-01-01

Combustion-acoustic instabilities occur when the acoustic energy increase due to the unsteady heat release of the flame is greater than the losses of acoustic energy from the system. The problem of combustion-acoustic instability is a concern in many devices for various reasons, as each device may have a unique mechanism causing unsteady heat release rates and many have unique boundary conditions. To accurately predict and quantify combustion-acoustic stabilities, the unsteady heat release rate and boundary conditions need to be accurately determined. The present review brings together work performed on a variety of practical combustion devices. Many theoretical and experimental investigations of the unsteady heat release rate have been performed, some based on perturbations in the fuel delivery system particularly for rocket instabilities, while others are based on hydrodynamic processes as in ramjet dump combustors. The boundary conditions for rocket engines have been analyzed and measured extensively. However, less work has been done to measure acoustic boundary conditions in many other combustion systems.

Experimental and analytical investigations of longitudinal combustion instability in a continuously variable resonance combustor (CVRC)

NASA Astrophysics Data System (ADS)

Yu, Yen Ching

An analytical model based on linearized Euler equations (LEE) is developed and used in conjunction with a validating experiment to study combustion instability. The LEE model features mean flow effects, entropy waves, adaptability for more physically-realistic boundary conditions, and is generalized for multiple-domain conditions. The model calculates spatial modes, resonant frequencies and linear growth rates of the overall system. The predicted resonant frequencies and spatially-resolved mode shapes agree with the experimental data from a longitudinally-unstable model rocket combustor to within 7%. Different gaseous fuels (methane, ethylene, and hydrogen) were tested under fixed geometry. Tests with hydrogen were stable, whereas ethylene, methane, and JP-8 were increasingly unstable. A novel method for obtaining large amounts of stability data under variable resonance conditions in a single test was demonstrated. The continuously variable resonance combustor (CVRC) incorporates a traversing choked axial oxidizer inlet to vary the overall combustion system resonance. The CVRC experiment successfully demonstrates different level of instability, transitions between stability levels, and identifies the most stable and unstable geometric combination. Pressure oscillation amplitudes ranged from less than 10% of mean pressure to greater than 60%. At low amplitudes, measured resonant frequency changed with inlet location but at high amplitude the measured resonance frequency matched the frequency of the combustion chamber. As the system transitions from linear to non-linear instability, the higher harmonics of the fundamental resonant mode appear nearly simultaneously. Transient, high-amplitude, broadband noise, at lower frequencies (on the order of 200 Hz) are also observed. Conversely, as the system transitions back to a more linear stability regime, the higher harmonics disappear sequentially, led by the highest order. Good agreements between analytical and experimental results are attained by treating the experiment as quasi-stationary. The stability characteristics from the high frequency measurements are further analyzed using filtered pressure traces, spectrograms, power spectral density plots, and oscillation decrements. Future works recommended include: direct measurements, such as chemiluminescence or high-speed imaging to examine the unsteady combustion processes; three-way comparisons between the acoustic-based, linear Euler-based, and non-linear Euler/RANS model; use the high fidelity computation to investigate the forcing terms modeled in the acoustic-based model.

Transient combustion in hybrid rockets

NASA Astrophysics Data System (ADS)

Karabeyoglu, Mustafa Arif

1998-09-01

Hybrid rockets regained interest recently as an alternative chemical propulsion system due to their advantages over the solid and liquid systems that are currently in use. Development efforts on hybrids revealed two important problem areas: (1) low frequency instabilities and (2) slow transient response. Both of these are closely related to the transient behavior which is a poorly understood aspect of hybrid operation. This thesis is mainly involved with a theoretical study of transient combustion in hybrid rockets. We follow the methodology of identifying and modeling the subsystems of the motor such as the thermal lags in the solid, boundary layer combustion and chamber gasdynamics from a dynamic point of view. We begin with the thermal lag in the solid which yield the regression rate for any given wall heat flux variation. Interesting phenomena such as overshooting during throttling and the amplification and phase lead regions in the frequency domain are discovered. Later we develop a quasi-steady transient hybrid combustion model supported with time delays for the boundary layer processes. This is integrated with the thermal lag system to obtain the thermal combustion (TC) coupled response. The TC coupled system with positive delays generated low frequency instabilities. The scaling of the instabilities are in good agreement with actual motor test data. Finally, we formulate a gasdynamic model for the hybrid chamber which successfully resolves the filling/emptying and longitudinal acoustic behavior of the motor. The TC coupled system is later integrated to the gasdynamic model to obtain the overall response (TCG coupled system) of gaseous oxidizer motors with stiff feed systems. Low frequency instabilities were also encountered for the TCG coupled system. Apart from the transient investigations, the regression rate behavior of liquefying hybrid propellants such as solid cryogenic materials are also studied. The theory is based on the possibility of enhancement of regression rate by the entrainment mass transfer from a liquid layer formed on the fuel surface. The predicted regression rates are in good agreement with the cryogenic experimental findings obtained recently at Edwards Airforce Base with a frozen pentane and gaseous oxygen system.

Motor Flow Instabilities - Part 1

DTIC Science & Technology

2004-01-01

by the flow, the structure motions (as possibly affecting the mean and unsteady flows). Finally, the model should be able: a) to propagate the...combustion responses function determinations, Dedicated models for combustion mechanisms and fluid- structure couplings, Dedicated and documented test...associated with these large motors (recall that f1L ≈ a/2L) rendered such oscillations undesirable since they were able to couple to the structural modes

Experimental Replication of an Aeroengine Combustion Instability

NASA Technical Reports Server (NTRS)

Cohen, J. M.; Hibshman, J. R.; Proscia, W.; Rosfjord, T. J.; Wake, B. E.; McVey, J. B.; Lovett, J.; Ondas, M.; DeLaat, J.; Breisacher, K.

2000-01-01

Combustion instabilities in gas turbine engines are most frequently encountered during the late phases of engine development, at which point they are difficult and expensive to fix. The ability to replicate an engine-traceable combustion instability in a laboratory-scale experiment offers the opportunity to economically diagnose the problem (to determine the root cause), and to investigate solutions to the problem, such as active control. The development and validation of active combustion instability control requires that the causal dynamic processes be reproduced in experimental test facilities which can be used as a test bed for control system evaluation. This paper discusses the process through which a laboratory-scale experiment was designed to replicate an instability observed in a developmental engine. The scaling process used physically-based analyses to preserve the relevant geometric, acoustic and thermo-fluid features. The process increases the probability that results achieved in the single-nozzle experiment will be scalable to the engine.

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.

Parametric study of shock-induced combustion in a hydrogen air system

NASA Technical Reports Server (NTRS)

Ahuja, J. K.; Tiwari, Surendra N.

1994-01-01

A numerical parametric study is conducted to simulate shock-induced combustion under various free-stream conditions and varying blunt body diameter. A steady combustion front is established if the free-stream Mach number is above the Chapman-Jouguet speed of the mixture, whereas an unsteady reaction front is established if the free-stream Mach number is below or at the Chapman-Jouguet speed of the mixture. The above two cases have been simulated for Mach 5.11 and Mach 6.46 with a projectile diameter of 15 mm. Mach 5.11, which is an underdriven case, shows an unsteady reaction front, whereas Mach 6.46, which is an overdriven case, shows a steady reaction front. Next for Mach 5. 11 reducing the diameter to 2.5 mm causes the instabilities to disappear, whereas, for Mach 6.46 increasing the diameter of the projectile to 225 mm causes the instabilities to reappear, indicating that Chapman-Jouguet speed is not the only deciding factor for these instabilities to trigger. The other key parameters are the projectile diameter, induction time, activation energy and the heat release. The appearance and disappearance of the instabilities have been explained by the one-dimensional wave interaction model.

Assessing Spontaneous Combustion Instability with Nonlinear Time Series Analysis

NASA Technical Reports Server (NTRS)

Eberhart, C. J.; Casiano, M. J.

2015-01-01

Considerable interest lies in the ability to characterize the onset of spontaneous instabilities within liquid propellant rocket engine (LPRE) combustion devices. Linear techniques, such as fast Fourier transforms, various correlation parameters, and critical damping parameters, have been used at great length for over fifty years. Recently, nonlinear time series methods have been applied to deduce information pertaining to instability incipiency hidden in seemingly stochastic combustion noise. A technique commonly used in biological sciences known as the Multifractal Detrended Fluctuation Analysis has been extended to the combustion dynamics field, and is introduced here as a data analysis approach complementary to linear ones. Advancing, a modified technique is leveraged to extract artifacts of impending combustion instability that present themselves a priori growth to limit cycle amplitudes. Analysis is demonstrated on data from J-2X gas generator testing during which a distinct spontaneous instability was observed. Comparisons are made to previous work wherein the data were characterized using linear approaches. Verification of the technique is performed by examining idealized signals and comparing two separate, independently developed tools.

Reduced Order Modeling of Combustion Instability in a Gas Turbine Model Combustor

NASA Astrophysics Data System (ADS)

Arnold-Medabalimi, Nicholas; Huang, Cheng; Duraisamy, Karthik

2017-11-01

Hydrocarbon fuel based propulsion systems are expected to remain relevant in aerospace vehicles for the foreseeable future. Design of these devices is complicated by combustion instabilities. The capability to model and predict these effects at reduced computational cost is a requirement for both design and control of these devices. This work focuses on computational studies on a dual swirl model gas turbine combustor in the context of reduced order model development. Full fidelity simulations are performed utilizing URANS and Hybrid RANS-LES with finite rate chemistry. Following this, data decomposition techniques are used to extract a reduced basis representation of the unsteady flow field. These bases are first used to identify sensor locations to guide experimental interrogations and controller feedback. Following this, initial results on developing a control-oriented reduced order model (ROM) will be presented. The capability of the ROM will be further assessed based on different operating conditions and geometric configurations.

Assessing Spontaneous Combustion Instability with Recurrence Quantification Analysis

NASA Technical Reports Server (NTRS)

Eberhart, Chad J.; Casiano, Matthew J.

2016-01-01

Spontaneous instabilities can pose a significant challenge to verification of combustion stability, and characterizing its onset is an important avenue of improvement for stability assessments of liquid propellant rocket engines. Recurrence Quantification Analysis (RQA) is used here to explore nonlinear combustion dynamics that might give insight into instability. Multiple types of patterns representative of different dynamical states are identified within fluctuating chamber pressure data, and markers for impending instability are found. A class of metrics which describe these patterns is also calculated. RQA metrics are compared with and interpreted against another metric from nonlinear time series analysis, the Hurst exponent, to help better distinguish between stable and unstable operation.

Fast and slow active control of combustion instabilities in liquid-fueled combustors

NASA Astrophysics Data System (ADS)

Lee, Jae-Yeon

This thesis describes an experimental investigation of two different novel active control approaches that are employed to suppress combustion instabilities in liquid-fueled combustors. A "fast" active controller requires continuous modulation of the fuel injection rate at the frequency of the instability with proper phase and gain. Use of developed optical tools reveals that the "fast" active control system suppresses the instability by changing the nearly flat distribution of the phase between pressure and heat release oscillations to a gradually varying phase distribution, thus dividing the combustion zone into regions that alternately damp and drive combustor oscillations. The effects of these driving/damping regions tend to counter one another, which result in significant damping of the unstable oscillations. In contrast, a "slow" active controller operates at a rate commensurate with that at which operating conditions change during combustor operation. Consequently, "slow" controllers need infrequent activation in response to changes in engine operating conditions to assure stable operation at all times. Using two types of fuel injectors that can produce large controllable variation of fuel spray properties, it is shown that by changing the spray characteristics it is possible to significantly damp combustion instabilities. Similar to the aforementioned result of the "fast" active control study, "slow" change of the fuel spray properties also modifies the nearly flat phase distribution during unstable operation to a gradually varying phase distribution, resulting in combustor "stabilization". Furthermore, deconvolutions of CH*-chemiluminescence images reveal the presence of vortex-flame interaction during unstable operation. Strong driving of instabilities occurs where the mean axial velocity of the flow is approximately zero, a short distance downstream of the flame holder where a significant fraction of the fuel burns in phase with the pressure oscillations. It is shown that the "fast" and "slow" active control approaches suppress combustion instabilities in a different manner. Nevertheless, the both control approaches successfully suppress combustion instabilities by modifying the temporal and spatial behavior of the combustion process heat release that is responsible for driving the instability.

Mean Flow Augmented Acoustics in Rocket Systems

NASA Technical Reports Server (NTRS)

Fischbach, Sean

2014-01-01

Combustion instability in solid rocket motors and liquid engines has long been a subject of concern. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process and gas dynamics. Recent advances in energy based modeling of combustion instabilities require accurate determination of acoustic frequencies and mode shapes. Of particular interest is the acoustic mean flow interactions within the converging section of a rocket nozzle, where gradients of pressure, density, and velocity become large. The expulsion of unsteady energy through the nozzle of a rocket is identified as the predominate source of acoustic damping for most rocket systems. Recently, an approach to address nozzle damping with mean flow effects was implemented by French [1]. This new approach extends the work originated by Sigman and Zinn [2] by solving the acoustic velocity potential equation (AVPE) formulated by perturbing the Euler equations [3]. The present study aims to implement the French model within the COMSOL Multiphysiscs framework and analyzes one of the author's presented test cases.

Three-step approach for prediction of limit cycle pressure oscillations in combustion chambers of gas turbines

NASA Astrophysics Data System (ADS)

Iurashev, Dmytro; Campa, Giovanni; Anisimov, Vyacheslav V.; Cosatto, Ezio

2017-11-01

Currently, gas turbine manufacturers frequently face the problem of strong acoustic combustion driven oscillations inside combustion chambers. These combustion instabilities can cause extensive wear and sometimes even catastrophic damages to combustion hardware. This requires prevention of combustion instabilities, which, in turn, requires reliable and fast predictive tools. This work presents a three-step method to find stability margins within which gas turbines can be operated without going into self-excited pressure oscillations. As a first step, a set of unsteady Reynolds-averaged Navier-Stokes simulations with the Flame Speed Closure (FSC) model implemented in the OpenFOAM® environment are performed to obtain the flame describing function of the combustor set-up. The standard FSC model is extended in this work to take into account the combined effect of strain and heat losses on the flame. As a second step, a linear three-time-lag-distributed model for a perfectly premixed swirl-stabilized flame is extended to the nonlinear regime. The factors causing changes in the model parameters when applying high-amplitude velocity perturbations are analysed. As a third step, time-domain simulations employing a low-order network model implemented in Simulink® are performed. In this work, the proposed method is applied to a laboratory test rig. The proposed method permits not only the unsteady frequencies of acoustic oscillations to be computed, but the amplitudes of such oscillations as well. Knowing the amplitudes of unstable pressure oscillations, it is possible to determine how these oscillations are harmful to the combustor equipment. The proposed method has a low cost because it does not require any license for computational fluid dynamics software.

On the Behavior of a Shear-Coaxial Jet, Spanning Sub- to Supercritical Pressures, with and without an Externally Imposed Transverse Acoustic Field

DTIC Science & Technology

2006-05-01

rocket engines (LRE) have experienced high-frequency combustion instability, which impose an acoustic field in the combustion chamber. The acoustic...Graduate School iii ABSTRACT In the past, liquid rocket engines (LRE) have experienced high-frequency combustion instability, which impose an...49 3.5 Instrumentation

Characterization of Low-Frequency Combustion Stability of the Fastrac Engine

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Jones, Preston (Technical Monitor)

2002-01-01

A series of tests were conducted to measure the combustion performance of the Fastrac engine thrust chamber. During mainstage, the thrust chamber exhibited no large-amplitude chamber pressure oscillations that could be identified as low-frequency combustion instability or 'chug'. However, during start-up and shutdown, the thrust chamber very briefly exhibited large-amplitude chamber pressure oscillations that were identified as chug. These instabilities during start-up and shutdown were regarded as benign due to their brevity. Linear models of the thrust chamber and the propellant feed systems were formulated for both the thrust chamber component tests and the flight engine tests. These linear models determined the frequency and decay rate of chamber pressure oscillations given the design and operating conditions of the thrust chamber and feed system. The frequency of chamber pressure oscillations determined from the model closely matched the frequency of low-amplitude, low-frequency chamber pressure oscillations exhibited in some of the later thrust chamber mainstage tests. The decay rate of the chamber pressure oscillations determined from the models indicated that these low-frequency oscillations were stable. Likewise, the decay rate, determined from the model of the flight engine tests indicated that the low-frequency chamber pressure oscillations would be stable.

Nonlinear rocket motor stability prediction: Limit amplitude, triggering, and mean pressure shifta)

NASA Astrophysics Data System (ADS)

Flandro, Gary A.; Fischbach, Sean R.; Majdalani, Joseph

2007-09-01

High-amplitude pressure oscillations in solid propellant rocket motor combustion chambers display nonlinear effects including: (1) limit cycle behavior in which the fluctuations may dwell for a considerable period of time near their peak amplitude, (2) elevated mean chamber pressure (DC shift), and (3) a triggering amplitude above which pulsing will cause an apparently stable system to transition to violent oscillations. Along with the obvious undesirable vibrations, these features constitute the most damaging impact of combustion instability on system reliability and structural integrity. The physical mechanisms behind these phenomena and their relationship to motor geometry and physical parameters must, therefore, be fully understood if instability is to be avoided in the design process, or if effective corrective measures must be devised during system development. Predictive algorithms now in use have limited ability to characterize the actual time evolution of the oscillations, and they do not supply the motor designer with information regarding peak amplitudes or the associated critical triggering amplitudes. A pivotal missing element is the ability to predict the mean pressure shift; clearly, the designer requires information regarding the maximum chamber pressure that might be experienced during motor operation. In this paper, a comprehensive nonlinear combustion instability model is described that supplies vital information. The central role played by steep-fronted waves is emphasized. The resulting algorithm provides both detailed physical models of nonlinear instability phenomena and the critically needed predictive capability. In particular, the origin of the DC shift is revealed.

Engine Hydraulic Stability. [injector model for analyzing combustion instability

NASA Technical Reports Server (NTRS)

Kesselring, R. C.; Sprouse, K. M.

1977-01-01

An analytical injector model was developed specifically to analyze combustion instability coupling between the injector hydraulics and the combustion process. This digital computer dynamic injector model will, for any imposed chamber of inlet pressure profile with a frequency ranging from 100 to 3000 Hz (minimum) accurately predict/calculate the instantaneous injector flowrates. The injector system is described in terms of which flow segments enter and leave each pressure node. For each flow segment, a resistance, line lengths, and areas are required as inputs (the line lengths and areas are used in determining inertance). For each pressure node, volume and acoustic velocity are required as inputs (volume and acoustic velocity determine capacitance). The geometric criteria for determining inertances of flow segments and capacitance of pressure nodes was set. Also, a technique was developed for analytically determining time averaged steady-state pressure drops and flowrates for every flow segment in an injector when such data is not known. These pressure drops and flowrates are then used in determining the linearized flow resistance for each line segment of flow.

Large Eddy Simulations of Transverse Combustion Instability in a Multi-Element Injector

DTIC Science & Technology

2016-07-27

Instability in a Multi- Element Injector 5a. CONTRACT NUMBER 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Matthew Harvazinski, Yogin...Simulations of Transverse  Combustion Instability in a Multi‐ Element  Injector 2 History Damaged engine injector  faceplate caused by combustion...Clearance #16346 3 Single  Element  Studies Short Post Marginally Stable Intermediate Post Unstable Long Post Stable Long Post Unstable CVRC Experiment

Research on combustion instability and application to solid propellant rocket motors. II.

NASA Technical Reports Server (NTRS)

Culick, F. E. C.

1972-01-01

Review of the current state of analyses of combustion instability in solid-propellant rocket motors, citing appropriate measurements and observations. The work discussed has become increasingly important, both for the interpretation of laboratory data and for predicting the transient behavior of disturbances in full-scale motors. Two central questions are considered - namely, linear stability and nonlinear behavior. Several classes of problems are discussed as special cases of a general approach to the analysis of combustion instability. Application to motors, and particularly the limitations presently understood, are stressed.

Modeling of vortex generated sound in solid propellant rocket motors

NASA Technical Reports Server (NTRS)

Flandro, G. A.

1980-01-01

There is considerable evidence based on both full scale firings and cold flow simulations that hydrodynamically unstable shear flows in solid propellant rocket motors can lead to acoustic pressure fluctuations of significant amplitude. Although a comprehensive theoretical understanding of this problem does not yet exist, procedures were explored for generating useful analytical models describing the vortex shedding phenomenon and the mechanisms of coupling to the acoustic field in a rocket combustion chamber. Since combustion stability prediction procedures cannot be successful without incorporation of all acoustic gains and losses, it is clear that a vortex driving model comparable in quality to the analytical models currently employed to represent linear combustion instability must be formulated.

Mixing in Shear Coaxial Jets with and without Acoustics

DTIC Science & Technology

2012-03-29

Distribution Unlimited Combustion Instability Lab - Background • Combustion instability is an unsustainable growth of pressure and heat transfer ...beyond liquid, gas states. Shear coaxial injectors are a common choice for cryogenic liquid rocket engines. Interactions of transverse acoustics with...and combustion beyond liquid, gas states • Shear coaxial injectors are a common choice for cryogenic liquid rocket engines • Interactions of

Demonstration of Active Combustion Control

NASA Technical Reports Server (NTRS)

Lovett, Jeffrey A.; Teerlinck, Karen A.; Cohen, Jeffrey M.

2008-01-01

The primary objective of this effort was to demonstrate active control of combustion instabilities in a direct-injection gas turbine combustor that accurately simulates engine operating conditions and reproduces an engine-type instability. This report documents the second phase of a two-phase effort. The first phase involved the analysis of an instability observed in a developmental aeroengine and the design of a single-nozzle test rig to replicate that phenomenon. This was successfully completed in 2001 and is documented in the Phase I report. This second phase was directed toward demonstration of active control strategies to mitigate this instability and thereby demonstrate the viability of active control for aircraft engine combustors. This involved development of high-speed actuator technology, testing and analysis of how the actuation system was integrated with the combustion system, control algorithm development, and demonstration testing in the single-nozzle test rig. A 30 percent reduction in the amplitude of the high-frequency (570 Hz) instability was achieved using actuation systems and control algorithms developed within this effort. Even larger reductions were shown with a low-frequency (270 Hz) instability. This represents a unique achievement in the development and practical demonstration of active combustion control systems for gas turbine applications.

A Nonlinear Model for Fuel Atomization in Spray Combustion

NASA Technical Reports Server (NTRS)

Liu, Nan-Suey (Technical Monitor); Ibrahim, Essam A.; Sree, Dave

2003-01-01

Most gas turbine combustion codes rely on ad-hoc statistical assumptions regarding the outcome of fuel atomization processes. The modeling effort proposed in this project is aimed at developing a realistic model to produce accurate predictions of fuel atomization parameters. The model involves application of the nonlinear stability theory to analyze the instability and subsequent disintegration of the liquid fuel sheet that is produced by fuel injection nozzles in gas turbine combustors. The fuel sheet is atomized into a multiplicity of small drops of large surface area to volume ratio to enhance the evaporation rate and combustion performance. The proposed model will effect predictions of fuel sheet atomization parameters such as drop size, velocity, and orientation as well as sheet penetration depth, breakup time and thickness. These parameters are essential for combustion simulation codes to perform a controlled and optimized design of gas turbine fuel injectors. Optimizing fuel injection processes is crucial to improving combustion efficiency and hence reducing fuel consumption and pollutants emissions.

Pressure oscillations and instability of working processes in the combustion chambers of solid rocket motors

NASA Astrophysics Data System (ADS)

Emelyanov, V. N.; Teterina, I. V.; Volkov, K. N.; Garkushev, A. U.

2017-06-01

Metal particles are widely used in space engineering to increase specific impulse and to supress acoustic instability of intra-champber processes. A numerical analysis of the internal injection-driven turbulent gas-particle flows is performed to improve the current understanding and modeling capabilities of the complex flow characteristics in the combustion chambers of solid rocket motors (SRMs) in presence of forced pressure oscillations. The two-phase flow is simulated with a combined Eulerian-Lagrangian approach. The Reynolds-averaged Navier-Stokes equations and transport equations of k - ε model are solved numerically for the gas. The particulate phase is simulated through a Lagrangian deterministic and stochastic tracking models to provide particle trajectories and particle concentration. The results obtained highlight the crucial significance of the particle dispersion in turbulent flowfield and high potential of statistical methods. Strong coupling between acoustic oscillations, vortical motion, turbulent fluctuations and particle dynamics is observed.

The effect of convection and shear on the damping and propagation of pressure waves

NASA Astrophysics Data System (ADS)

Kiel, Barry Vincent

Combustion instability is the positive feedback between heat release and pressure in a combustion system. Combustion instability occurs in the both air breathing and rocket propulsion devices, frequently resulting in high amplitude spinning waves. If unchecked, the resultant pressure fluctuations can cause significant damage. Models for the prediction of combustion instability typically include models for the heat release, the wave propagation and damping. Many wave propagation models for propulsion systems assume negligible flow, resulting in the wave equation. In this research the effect of flow on wave propagation was studied both numerically and experimentally. Two experiential rigs were constructed, one with axial flow to study the longitudinal waves, the other with swirling flow to study circumferential waves. The rigs were excited with speakers and the resultant pressure was measured simultaneously at many locations. Models of the rig were also developed. Equations for wave propagation were derived from the Euler Equations. The resultant resembled the wave equation with three additional terms, two for the effect of the convection and a one for the effect of shear of the mean flow on wave propagation. From the experimental and numerical data several conclusions were made. First, convection and shear both act as damping on the wave propagation, reducing the magnitude of the Frequency Response Function and the resonant frequency of the modes. Second, the energy extracted from the mean flow as a result of turbulent shear for a given condition is frequency dependent, decreasing with increasing frequency. The damping of the modes, measured for the same shear flow, also decreased with frequency. Finally, the two convective terms cause the anti-nodes of the modes to no longer be stationary. For both the longitudinal and circumferential waves, the anti-nodes move through the domain even for mean flow Mach numbers less than 0.10. It was concluded that convection causes the spinning waves documented in inlets and exhausts of gas turbine engines, rocket combustion chambers, and afterburner chambers. As a result, the effects of shear must be included when modeling wave propagation, even for mean flows less than < Mach 0.10.

Nitramine smokeless propellant research

NASA Technical Reports Server (NTRS)

1977-01-01

A transient ballistics and combustion model was derived to represent the closed vessel experiment that is widely used to characterize propellants. The model incorporates the nitramine combustion mechanisms. A computer program was developed to solve the time dependent equations, and was applied to explain aspects of closed vessel behavior. It is found that the rate of pressurization in the closed vessel is insufficient at pressures of interest to augment the burning rate by time dependent processes. Series of T-burner experiments were performed to compare the combustion instability characteristics of nitramine (HMX) containing propellants and ammonium perchlorate (AP) propellants. It is found that the inclusion of HMX consistently renders the propellant more stable.

Baseline Computational Fluid Dynamics Methodology for Longitudinal-Mode Liquid-Propellant Rocket Combustion Instability

NASA Technical Reports Server (NTRS)

Litchford, R. J.

2005-01-01

A computational method for the analysis of longitudinal-mode liquid rocket combustion instability has been developed based on the unsteady, quasi-one-dimensional Euler equations where the combustion process source terms were introduced through the incorporation of a two-zone, linearized representation: (1) A two-parameter collapsed combustion zone at the injector face, and (2) a two-parameter distributed combustion zone based on a Lagrangian treatment of the propellant spray. The unsteady Euler equations in inhomogeneous form retain full hyperbolicity and are integrated implicitly in time using second-order, high-resolution, characteristic-based, flux-differencing spatial discretization with Roe-averaging of the Jacobian matrix. This method was initially validated against an analytical solution for nonreacting, isentropic duct acoustics with specified admittances at the inflow and outflow boundaries. For small amplitude perturbations, numerical predictions for the amplification coefficient and oscillation period were found to compare favorably with predictions from linearized small-disturbance theory as long as the grid exceeded a critical density (100 nodes/wavelength). The numerical methodology was then exercised on a generic combustor configuration using both collapsed and distributed combustion zone models with a short nozzle admittance approximation for the outflow boundary. In these cases, the response parameters were varied to determine stability limits defining resonant coupling onset.

Hydrodynamic and thermal mechanisms of filtration combustion inclinational instability based on non-uniform distribution of initial preheating temperature

NASA Astrophysics Data System (ADS)

Xia, Yongfang; Shi, Junrui; Xu, Youning; Ma, Rui

2018-03-01

Filtration combustion (FC) is one style of porous media combustion with inert matrix, in which the combustion wave front propagates, only downstream or reciprocally. In this paper, we investigate the FC flame front inclinational instability of lean methane/air mixtures flowing through a packed bed as a combustion wave front perturbation of the initial preheating temperature non-uniformity is assumed. The predicted results show that the growth rate of the flame front inclinational angle is proportional to the magnitude of the initial preheating temperature difference. Additionally, depending on gas inlet gas velocity and equivalence ratio, it is demonstrated that increase of gas inlet gas velocity accelerates the FC wave front deformation, and the inclinational instability evolves faster at lower equivalence ratio. The development of the flame front inclinational angle may be regarded as a two-staged evolution, which includes rapid increase, and approaching maximum value of inclinational angle due to the quasi-steady condition of the combustion system. The hydrodynamic and thermal mechanisms of the FC inclinational instability are analyzed. Consequently, the local propagation velocity of the FC wave front is non-uniform to result in the development of inclinational angle at the first stage of rapid increase.

Analysis of Self-Excited Combustion Instabilities Using Decomposition Techniques

DTIC Science & Technology

2016-07-05

are evaluated for the study of self-excited longitudinal combustion instabilities in laboratory-scaled single-element gas turbine and rocket...Air Force Base, California 93524 DOI: 10.2514/1.J054557 Proper orthogonal decomposition and dynamic mode decomposition are evaluated for the study of...instabilities. In addition, we also evaluate the capabilities of the methods to deal with data sets of different spatial extents and temporal resolution

Robust control of combustion instabilities

NASA Astrophysics Data System (ADS)

Hong, Boe-Shong

Several interactive dynamical subsystems, each of which has its own time-scale and physical significance, are decomposed to build a feedback-controlled combustion- fluid robust dynamics. On the fast-time scale, the phenomenon of combustion instability is corresponding to the internal feedback of two subsystems: acoustic dynamics and flame dynamics, which are parametrically dependent on the slow-time-scale mean-flow dynamics controlled for global performance by a mean-flow controller. This dissertation constructs such a control system, through modeling, analysis and synthesis, to deal with model uncertainties, environmental noises and time- varying mean-flow operation. Conservation law is decomposed as fast-time acoustic dynamics and slow-time mean-flow dynamics, served for synthesizing LPV (linear parameter varying)- L2-gain robust control law, in which a robust observer is embedded for estimating and controlling the internal status, while achieving trade- offs among robustness, performances and operation. The robust controller is formulated as two LPV-type Linear Matrix Inequalities (LMIs), whose numerical solver is developed by finite-element method. Some important issues related to physical understanding and engineering application are discussed in simulated results of the control system.

The Stability and Structure of Lean Hydrogen-Air Flames: Effects of Gravity

DTIC Science & Technology

1990-05-17

INTRODUCTION ................................................................................................. 1 MULTIDIMENSIONAL FLAME MODEL ...combustion, molecular diffusion between the reactants, intermediates, and products, thermal conduction, convection, and gravity. Such a detailed model allows...instabil- ity, generally called the Rayleigh-Taylor instability5 . A numerical model of the premixed hydrogen flame that includes all the physical

Active Control of Combustor Instability Shown to Help Lower Emissions

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Chang, Clarence T.

2002-01-01

In a quest to reduce the environmental impact of aerospace propulsion systems, extensive research is being done in the development of lean-burning (low fuel-to-air ratio) combustors that can reduce emissions throughout the mission cycle. However, these lean-burning combustors have an increased susceptibility to thermoacoustic instabilities, or high-pressure oscillations much like sound waves, that can cause severe high-frequency vibrations in the combustor. These pressure waves can fatigue the combustor components and even the downstream turbine blades. This can significantly decrease the safe operating life of the combustor and turbine. Thus, suppression of the thermoacoustic combustor instabilities is an enabling technology for lean, low-emissions combustors. Under the Aerospace Propulsion and Power Base Research and Technology Program, the NASA Glenn Research Center, in partnership with Pratt & Whitney and United Technologies Research Center, is developing technologies for the active control of combustion instabilities. With active combustion control, the fuel is pulsed to put pressure oscillations into the system. This cancels out the pressure oscillations being produced by the instabilities. Thus, the engine can have lower pollutant emissions and long life.The use of active combustion instability control to reduce thermo-acoustic-driven combustor pressure oscillations was demonstrated on a single-nozzle combustor rig at United Technologies. This rig has many of the complexities of a real engine combustor (i.e., an actual fuel nozzle and swirler, dilution cooling, etc.). Control was demonstrated through modeling, developing, and testing a fuel-delivery system able to the 280-Hz instability frequency. The preceding figure shows the capability of this system to provide high-frequency fuel modulations. Because of the high-shear contrarotating airflow in the fuel injector, there was some concern that the fuel pulses would be attenuated to the point where they would not be effective for control. Testing in the combustor rig showed that open-loop pulsing of the fuel was, in fact, able to effectively modulate the combustor pressure. To suppress the combustor pressure oscillations due to thermoacoustic instabilities, it is desirable to time the injection of the fuel so that it interferes with the instability. A closed-loop control scheme was developed that uses combustion pressure feedback and a phase-shifting controller to time the fuel-injection pulses. Some suppression of the pressure oscillations at the 280-Hz instability frequency was demonstrated (see the next figure). However, the overall peak-to- peak pressure oscillations in the combustor were only mildly reduced. Improvements to control hardware and control methods are being continued to gain improved closed-loop reduction of the pressure oscillations.pulse the fuel at

Investigation of combustion control in a dump combustor using the feedback free fluidic oscillator

NASA Astrophysics Data System (ADS)

Meier, Eric J.

The feedback free fluidic oscillator uses the unsteady nature of two colliding jets to create a single oscillating outlet jet with a wide sweep angle. These devices have the potential to provide additional combustion control, boundary layer control, thrust vectoring, and industrial flow deflection. Two-dimensional computational fluid dynamics, CFD, was used to analyze the jet oscillation frequency over a range of operating conditions and to determine the effect that geometric changes in the oscillator design have on the frequency. Results presented illustrate the changes in jet oscillation frequency with gas type, gas temperature, operating pressure, pressure ratio across the oscillator, aspect ratio of the oscillator, and the frequency trends with various changes to the oscillator geometry. A fluidic oscillator was designed and integrated into single element rocket combustor with the goal of suppressing longitudinal combustion instabilities. An array of nine fluidic oscillators was tested to mimic modulated secondary oxidizer injection into the dump plane using 15% of the oxidizer flow. The combustor has a coaxial injector that uses gaseous methane and decomposed hydrogen peroxide at an O/F of 11.66. A sonic choke plate on an actuator arm allows for continuous adjustment of the oxidizer post acoustics for studying a variety of instability magnitudes. The fluidic oscillator unsteady outlet jet performance is compared with 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 60% reduction in the instability pressure oscillation magnitude when compared to the steady jet and baseline data. The results indicate open loop propellant modulation for combustion control can be achieved through fluidic devices that require no moving parts or electrical power to operate. Three-dimensional computational fluid dynamics, 3-D CFD, was conducted to determine the mechanism by which the fluidic oscillators were able to suppress the combustion instability. Results for steady jet secondary injection, showed a strong coupling between the jet injection and the combustion instability pressure pulse. The computational results were able to closely match the experimental results and previous CFD data. The model with the oscillating fluidic oscillator injection was unable to match the stable combustion seen in the experimental data. Further investigation is needed to determine the role higher order chemistry kinetics play in the process and the role of manifolds on the un-choked fuel and fluidic oscillator inlets. This research demonstrates the ability to modulate propellant injection and suppress combustion instabilities using fluidic devices that require no electrical power or moving parts. The advent of advanced manufacturing technologies such as direct metal laser sintering will allow for integration of fluidic devices into combustors to provide open loop active control with a high degree of reliability. Additionally, 2-D CFD analysis is demonstrated to be a valid tool for predicting the feedback free fluidic oscillator oscillation mechanism.

Study on Combustion Characteristics and Propelling Projectile Motion Process of Bulk-Loaded Liquid Propellant

NASA Astrophysics Data System (ADS)

Xue, Xiaochun; Yu, Yonggang; Mang, Shanshan

2017-07-01

Data are presented showing that the problem of gas-liquid interaction instability is an important subject in the combustion and the propellant projectile motion process of a bulk-loaded liquid propellant gun (BLPG). The instabilities themselves arise from the sources, including fluid motion, to form a combustion gas cavity called Taylor cavity, fluid turbulence and breakup caused by liquid motion relative to the combustion chamber walls, and liquid surface breakup arising from a velocity mismatch on the gas-liquid interface. Typically, small disturbances that arise early in the BLPG combustion interior ballistic cycle can become amplified in the absence of burn rate limiting characteristics. Herein, significant attention has been given to developing and emphasizing the need for better combustion repeatability in the BLPG. Based on this goal, the concept of using different geometries of the combustion chamber is introduced and the concept of using a stepped-wall structure on the combustion chamber itself as a useful means of exerting boundary control on the combustion evolution to thus restrain the combustion instability has been verified experimentally in this work. Moreover, based on this background, the numerical simulation is devoted to a special combustion issue under transient high-pressure and high-temperature conditions, namely, studying the combustion mechanism in a stepped-wall combustion chamber with full monopropellant on one end that is stationary and the other end can move at high speed. The numerical results also show that the burning surface of the liquid propellant can be defined geometrically and combustion is well behaved as ignition and combustion progressivity are in a suitable range during each stage in this combustion chamber with a stepped-wall structure.

A Computational Study of Transverse Combustion Instability Mechanisms

DTIC Science & Technology

2014-07-01

April 2001. 7. Selle, L ., Benoit , L ., Poinsot, T., Nicoud, F., Krebs, W., “Joint use of compressible large-eddy simulation and Helmholtz solvers for...Mechanisms Kevin J. Shipley1, William E. Anderson2 Purdue University, West Lafayette, IN, 47906 Matthew E. Harvazinski3, and Venkateswaran Sankaran4...Lafayette, IN, August 2010. 9. Xia, G., Harvazinski, M., Anderson, W., Merkle, C. L ., “Investigation of Modeling and Physical Parameters on Instability

Integrated Physics-based Modeling and Experiments for Improved Prediction of Combustion Dynamics in Low-Emission Systems

NASA Technical Reports Server (NTRS)

Anderson, William E.; Lucht, Robert P.; Mongia, Hukam

2015-01-01

Concurrent simulation and experiment was undertaken to assess the ability of a hybrid RANS-LES model to predict combustion dynamics in a single-element lean direct-inject (LDI) combustor showing self-excited instabilities. High frequency pressure modes produced by Fourier and modal decomposition analysis were compared quantitatively, and trends with equivalence ratio and inlet temperature were compared qualitatively. High frequency OH PLIF and PIV measurements were also taken. Submodels for chemical kinetics and primary and secondary atomization were also tested against the measured behavior. For a point-wise comparison, the amplitudes matched within a factor of two. The dependence on equivalence ratio was matched. Preliminary results from simulation using an 18-reaction kinetics model indicated instability amplitudes closer to measurement. Analysis of the simulations suggested a band of modes around 1400 Hz were due to a vortex bubble breakdown and a band of modes around 6 kHz were due to a precessing vortex core hydrodynamic instability. The primary needs are directly coupled and validated ab initio models of the atomizer free surface flow and the primary atomization processes, and more detailed study of the coupling between the 3D swirling flow and the local thermoacoustics in the diverging venturi section.

Tripropellant combustion process

NASA Technical Reports Server (NTRS)

Kmiec, T. D.; Carroll, R. G.

1988-01-01

The addition of small amounts of hydrogen to the combustion of LOX/hydrocarbon propellants in large rocket booster engines has the potential to enhance the system stability. Programs being conducted to evaluate the effects of hydrogen on the combustion of LOX/hydrocarbon propellants at supercritical pressures are described. Combustion instability has been a problem during the development of large hydrocarbon fueled rocket engines. At the higher combustion chamber pressures expected for the next generation of booster engines, the effect of unstable combustion could be even more destructive. The tripropellant engine cycle takes advantage of the superior cooling characteristics of hydrogen to cool the combustion chamber and a small amount of the hydrogen coolant can be used in the combustion process to enhance the system stability. Three aspects of work that will be accomplished to evaluate tripropellant combustion are described. The first is laboratory demonstration of the benefits through the evaluation of drop size, ignition delay and burning rate. The second is analytical modeling of the combustion process using the empirical relationship determined in the laboratory. The third is a subscale demonstration in which the system stability will be evaluated. The approach for each aspect is described and the analytical models that will be used are presented.

Fundamental Insights into Combustion Instability Predictions in Aerospace Propulsion

NASA Astrophysics Data System (ADS)

Huang, Cheng

Integrated multi-fidelity modeling has been performed for combustion instability in aerospace propulsion, which includes two levels of analysis: first, computational fluid dynamics (CFD) using hybrid RANS/LES simulations for underlying physics investigations (high-fidelity modeling); second, modal decomposition techniques for diagnostics (analysis & validation); third, development of flame response model using model reduction techniques for practical design applications (low-order model). For the high-fidelity modeling, the relevant CFD code development work is moving towards combustion instability prediction for liquid propulsion system. A laboratory-scale single-element lean direct injection (LDI) gas turbine combustor is used for configuration that produces self-excited combustion instability. The model gas turbine combustor is featured with an air inlet section, air plenum, swirler-venturi-injector assembly, combustion chamber, and exit nozzle. The combustor uses liquid fuel (Jet-A/FT-SPK) and heated air up to 800K. Combustion dynamics investigations are performed with the same geometry and operating conditions concurrently between the experiment and computation at both high (φ=0.6) and low (φ=0.36) equivalence ratios. The simulation is able to reach reasonable agreement with experiment measurements in terms of the pressure signal. Computational analyses are also performed using an acoustically-open geometry to investigate the characteristic hydrodynamics in the combustor with both constant and perturbed inlet mass flow rates. Two hydrodynamic modes are identified by using Dynamic Mode Decomposition (DMD) analysis: Vortex Breakdown Bubble (VBB) and swirling modes. Following that, the closed geometry simulation results are analyzed in three steps. In step one, a detailed cycle analysis shows two physically important couplings in the combustor: first, the acoustic compression enhances the spray drop breakup and vaporization, and generates more gaseous fuel for reaction; second, the acoustic compression couples with the unsteady hydrodynamics found in the open-geometry simulation, enhances the fuel/air mixing, and triggers a large amount of heat addition. In step two, a modal analysis using DMD extracts the dynamic features of important modes in the combustor, and identifies the presence of Precessing Vortex Core (PVC) mode and its nonlinear interactions with acoustic modes. Moreover, the DMD analysis helps to establish the couplings between the hydrodynamics and acoustics in terms of frequencies. In step 3, Rayleigh index analysis provides a quantitative assessment of acoustics/combustion couplings and identifies local regions for instability driving/damping. Two modal decomposition techniques, Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD), are assessed in terms of their capabilities in extracting important information from the original simulation dataset and in validating the computational results using the experiment measurement. A POD analysis provides a series of modes with decreasing energy content and it offers an efficient and optimized way to represent a large dataset. The frequency-based DMD technique provides modes that correspond to all single frequencies. For the low-order modeling, fundamental aspects are examined to study necessary conditions, criteria and approaches to develop a reduced-order model (ROM) that is able to represent generic combustion/flame responses, which then can be used in an engineering level tool to provide efficient predictions of combustion instability for practical design applications. Explorations are focused on model reduction techniques by using the so-called POD/Galerkin method. The method uses the numerical solutions of the model equations as the database for building a set of POD eigen-bases. Specifically, the numerical solutions are calculated by perturbing quantities of interest such as the inlet conditions. The POD-derived eigen-bases are, in turn, used in conjunction with a Galerkin procedure to reduce the governing partial differential equation to an ordinary differential equation, which constitutes the ROM. Once the ROM is established, it can then be used as a lower-order test-bed to predict detailed results within certain parametric ranges at a fraction of the cost of solving the full governing equations. A detailed assessment is performed on the method in two parts. In part one, a one-dimensional scalar reaction-advection model equation is used for fundamental investigations, which include verification of the POD eigen-basis calculation and of the ROM development procedure. Moreover, certain criteria during ROM development are established: 1. a necessary number of POD modes that should be included to guarantee a stable ROM; 2. the need for the numerical discretization scheme to be consistent between the original CFD and the developed ROM. Furthermore, the predictive capabilities of the resulting ROM are evaluated to test its limits and to validate the values of applying broadband forcing in improving the ROM performance. In part two, the exploration is extended to a vector system of equations. Using the one-dimensional Euler equation is used as a model equation. A numerical stability issue is identified during the ROM development, the cause of which is further studied and attributed to the normalization methods implemented to generate coupled POD eigen-bases for vector variables. (Abstract shortened by UMI.).

Apparatus and method for combusting low quality fuel

DOEpatents

Brushwood, John Samuel; Pillsbury, Paul; Foote, John; Heilos, Andreas

2003-11-04

A gas turbine (12) capable of combusting a low quality gaseous fuel having a ratio of flammability limits less than 2, or a heat value below 100 BTU/SCF. A high quality fuel is burned simultaneously with the low quality fuel to eliminate instability in the combustion flame. A sensor (46) is used to monitor at least one parameter of the flame indicative of instability. A controller (50) having the sensor signal (48) as input is programmed to control the relative flow rates of the low quality and high quality fuels. When instability is detected, the flow rate of high quality fuel is automatically increased in relation to the flow rate of low quality fuel to restore stability.

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.

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.

Application of Detailed Chemical Kinetics to Combustion Instability Modeling

DTIC Science & Technology

2016-01-04

the stability characteristics. 15. SUBJECT TERMS N /A 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT 18. NUMBER OF PAGES 19a...include area code) N /A Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std. 239.18 Application of Detailed Chemical Kinetics to Combustion...Frenklach, M., Moriarty, N ., Eiteneer, B., Goldenberg, M., Bowman, C., Hanson, R., Song, S., W. Gardiner, J., Lissianski, V., and Qin, Z., “GRI-Mech 3.0

Evaluation and Improvement of Liquid Propellant Rocket Chugging Analysis Techniques. Part 1: A One-Dimensional Analysis of Low Frequency Combustion Instability in the Fuel Preburner of the Space Shuttle Main Engine. Final Report M.S. Thesis - Aug. 1986

NASA Technical Reports Server (NTRS)

Lim, Kair Chuan

1986-01-01

Low frequency combustion instability, known as chugging, is consistently experienced during shutdown in the fuel and oxidizer preburners of the Space Shuttle Main Engines. Such problems always occur during the helium purge of the residual oxidizer from the preburner manifolds during the shutdown sequence. Possible causes and triggering mechanisms are analyzed and details in modeling the fuel preburner chug are presented. A linearized chugging model, based on the foundation of previous models, capable of predicting the chug occurrence is discussed and the predicted results are presented and compared to experimental work performed by NASA. Sensitivity parameters such as chamber pressure, fuel and oxidizer temperatures, and the effective bulk modulus of the liquid oxidizer are considered in analyzing the fuel preburner chug. The computer program CHUGTEST is utilized to generate the stability boundary for each sensitivity study and the region for stable operation is identified.

Combustion Instability Analysis and the Effects of Drop Size on Acoustic Driving Rocket Flow

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Ellison, L. Renea; Moser, Marlow D.

2004-01-01

High frequency combustion instability, the most destructive kind, is generally solved on a per engine basis. The instability often is the result of compounding acoustic oscillations, usually from the propellant combustion itself. To counteract the instability the chamber geometry can be changed and/or the method of propellant injection can be altered. This experiment will alter the chamber dimensions slightly; using a cylindrical shape of constant diameter and the length will be varied from six to twelve inches in three-inch increments. The main flowfield will be the products of a high OF hydrogen/oxygen flow. The liquid fuel will be injected into this flowfield using a modulated injector. It will allow for varied droplet size, feed rate, spray pattern, and location for the mixture within the chamber. The response will be deduced from the chamber pressure oscillations.

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

NASA Technical Reports Server (NTRS)

DeLaat, John C.

2011-01-01

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

Analysis of Hydrodynamic (Landau) Instability in Liquid-Propellant Combustion at Normal and Reduced Gravity

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.

1997-01-01

The burning of liquid propellants is a fundamental combustion problem that is applicable to various types of propulsion and energetic systems. The deflagration process is often rather complex, with vaporization and pyrolysis occurring at the liquid/gas interface and distributed combustion occurring either in the gas phase or in a spray. Nonetheless, there are realistic limiting cases in which combustion may be approximated by an overall reaction at the liquid/gas interface. In one such limit, the gas flame occurs under near-breakaway conditions, exerting little thermal or hydrodynamic influence on the burning propellant. In another such limit, distributed combustion occurs in an intrusive regime, the reaction zone lying closer to the liquid/gas interface than the length scale of any disturbance of interest. Finally, the liquid propellant may simply undergo exothermic decomposition at the surface without any significant distributed combustion, such as appears to occur in some types of HydroxylAmmonium Nitrate (HAN)-based liquid propellants at low pressures. Such limiting models have recently been formulated,thereby significantly generalizing earlier classical models that were originally introduced to study the hydrodynamic stability of a reactive liquid/gas interface. In all of these investigations, gravity appears explicitly and plays a significant role, along with surface tension, viscosity, and, in the more recent models, certain reaction-rate parameters associated with the pressure and temperature sensitivities of the reaction itself. In particular, these parameters determine the stability of the deflagration with respect to not only classical hydrodynamic disturbances, but also with respect to reactive/diffusive influences as well. Indeed, the inverse Froude number, representing the ratio of buoyant to inertial forces, appears explicitly in all of these models, and consequently, in the dispersion relation that determines the neutral stability boundaries beyond which steady, planar burning is unstable to nonsteady, and/or nonplanar (cellular) modes of burning. These instabilities thus lead to a number of interesting phenomena, such as the sloshing type of waves that have been observed in mixtures of HAN and TriEthanolAmmonium Nitrate (TEAN) with water. Although the Froude number was treated as an O(1) quantity in these studies, the limit of small inverse Froude number corresponding to the microgravity regime is increasingly of interest and can be treated explicitly, leading to various limiting forms of the models, the neutral stability boundaries, and, ultimately, the evolution equations that govern the nonlinear dynamics of the propagating reaction front. In the present work, we formally exploit this limiting parameter regime to compare some of the features of hydrodynamic instability of liquid-propellant combustion at reduced gravity with the same phenomenon at normal gravity.

Active Combustion Control for Aircraft Gas Turbine Engines

NASA Technical Reports Server (NTRS)

DeLaat, John C.; Breisacher, Kevin J.; Saus, Joseph R.; Paxson, Daniel E.

2000-01-01

Lean-burning combustors are susceptible to combustion instabilities. Additionally, due to non-uniformities in the fuel-air mixing and in the combustion process, there typically exist hot areas in the combustor exit plane. These hot areas limit the operating temperature at the turbine inlet and thus constrain performance and efficiency. Finally, it is necessary to optimize the fuel-air ratio and flame temperature throughout the combustor to minimize the production of pollutants. In recent years, there has been considerable activity addressing Active Combustion Control. NASA Glenn Research Center's Active Combustion Control Technology effort aims to demonstrate active control in a realistic environment relevant to aircraft engines. Analysis and experiments are tied to aircraft gas turbine combustors. Considerable progress has been shown in demonstrating technologies for Combustion Instability Control, Pattern Factor Control, and Emissions Minimizing Control. Future plans are to advance the maturity of active combustion control technology to eventual demonstration in an engine environment.

Characterizing dilute combustion instabilities in a multi-cylinder spark-ignited engine using symbolic analysis

DOE PAGES

Daw, C. Stuart; Finney, Charles E. A.; Kaul, Brian C.; ...

2014-12-29

Spark-ignited internal combustion engines have evolved considerably in recent years in response to increasingly stringent regulations for emissions and fuel-economy. One new advanced engine strategy utilizes high levels of exhaust gas recirculation (EGR) to reduce combustion temperatures, thereby increasing thermodynamic efficiency and reducing nitrogen oxide emissions. While this strategy can be highly effective, it also poses major control and design challenges due to the large combustion oscillations that develop at sufficiently high EGR levels. Previous research has documented that combustion instabilities can propagate between successive engine cycles in individual cylinders via self-generated feedback of reactive species and thermal energy inmore » the retained residual exhaust gases. In this work, we use symbolic analysis to characterize multi-cylinder combustion oscillations in an experimental engine operating with external EGR. At low levels of EGR, intra-cylinder oscillations are clearly visible and appear to be associated with brief, intermittent coupling among cylinders. As EGR is increased further, a point is reached where all four cylinders lock almost completely in phase and alternate simultaneously between two distinct bi-stable combustion states. From a practical perspective, it is important to understand the causes of this phenomenon and develop diagnostics that might be applied to ameliorate its effects. We demonstrate here that two approaches for symbolizing the engine combustion measurements can provide useful probes for characterizing these instabilities.« less

Linear analysis of the Richtmyer-Meshkov instability in shock-flame interactions

NASA Astrophysics Data System (ADS)

Massa, L.; Jha, P.

2012-05-01

Shock-flame interactions enhance supersonic mixing and detonation formation. Therefore, their analysis is important to explosion safety, internal combustion engine performance, and supersonic combustor design. The fundamental process at the basis of the interaction is the Richtmyer-Meshkov instability supported by the density difference between burnt and fresh mixtures. In the present study we analyze the effect of reactivity on the Richtmyer-Meshkov instability with particular emphasis on combustion lengths that typify the scaling between perturbation growth and induction. The results of the present linear analysis study show that reactivity changes the perturbation growth rate by developing a pressure gradient at the flame surface. The baroclinic torque based on the density gradient across the flame acts to slow down the instability growth of high wave-number perturbations. A gasdynamic flame representation leads to the definition of a Peclet number representing the scaling between perturbation and thermal diffusion lengths within the flame. Peclet number effects on perturbation growth are observed to be marginal. The gasdynamic model also considers a finite flame Mach number that supports a separation between flame and contact discontinuity. Such a separation destabilizes the interface growth by augmenting the tangential shear.

Numerical methods for large eddy simulation of acoustic combustion instabilities

NASA Astrophysics Data System (ADS)

Wall, Clifton T.

Acoustic combustion instabilities occur when interaction between the combustion process and acoustic modes in a combustor results in periodic oscillations in pressure, velocity, and heat release. If sufficiently large in amplitude, these instabilities can cause operational difficulties or the failure of combustor hardware. In many situations, the dominant instability is the result of the interaction between a low frequency acoustic mode of the combustor and the large scale hydrodynamics. Large eddy simulation (LES), therefore, is a promising tool for the prediction of these instabilities, since both the low frequency acoustic modes and the large scale hydrodynamics are well resolved in LES. Problems with the tractability of such simulations arise, however, due to the difficulty of solving the compressible Navier-Stokes equations efficiently at low Mach number and due to the large number of acoustic periods that are often required for such instabilities to reach limit cycles. An implicit numerical method for the solution of the compressible Navier-Stokes equations has been developed which avoids the acoustic CFL restriction, allowing for significant efficiency gains at low Mach number, while still resolving the low frequency acoustic modes of interest. In the limit of a uniform grid the numerical method causes no artificial damping of acoustic waves. New, non-reflecting boundary conditions have also been developed for use with the characteristic-based approach of Poinsot and Lele (1992). The new boundary conditions are implemented in a manner which allows for significant reduction of the computational domain of an LES by eliminating the need to perform LES in regions where one-dimensional acoustics significantly affect the instability but details of the hydrodynamics do not. These new numerical techniques have been demonstrated in an LES of an experimental combustor. The new techniques are shown to be an efficient means of performing LES of acoustic combustion instabilities and are shown to accurately predict the occurrence and frequency of the dominant mode of the instability observed in the experiment.

The nature of combustion noise: Stochastic or chaotic?

NASA Astrophysics Data System (ADS)

Gupta, Vikrant; Lee, Min Chul; Li, Larry K. B.

2016-11-01

Combustion noise, which refers to irregular low-amplitude pressure oscillations, is conventionally thought to be stochastic. It has therefore been modeled using a stochastic term in the analysis of thermoacoustic systems. Recently, however, there has been a renewed interest in the validity of that stochastic assumption, with tests based on nonlinear dynamical theory giving seemingly contradictory results: some show combustion noise to be stochastic while others show it to be chaotic. In this study, we show that this contradiction arises because those tests cannot distinguish between noise amplification and chaos. We further show that although there are many similarities between noise amplification and chaos, there are also some subtle differences. It is these subtle differences, not the results of those tests, that should be the focus of analyses aimed at determining the true nature of combustion noise. Recognizing this is an important step towards improved understanding and modeling of combustion noise for the study of thermoacoustic instabilities. This work was supported by the Research Grants Council of Hong Kong (Project No. 16235716 and 26202815).

A review of active control approaches in stabilizing combustion systems in aerospace industry

NASA Astrophysics Data System (ADS)

Zhao, Dan; Lu, Zhengli; Zhao, He; Li, X. Y.; Wang, Bing; Liu, Peijin

2018-02-01

Self-sustained combustion instabilities are one of the most plaguing challenges and problems in lean-conditioned propulsion and land-based engine systems, such as rocket motors, gas turbines, industrial furnace and boilers, and turbo-jet thrust augmenters. Either passive or active control in open- or closed-loop configurations can be implemented to mitigate such instabilities. One of the classical disadvantages of passive control is that it is only implementable to a designed combustor over a limited frequency range and can not respond to the changes in operating conditions. Compared with passive control approaches, active control, especially in closed-loop configuration is more adaptive and has inherent capacity to be implemented in practice. The key components in closed-loop active control are 1) sensor, 2) controller (optimization algorithm) and 3) dynamic actuator. The present work is to outline the current status, technical challenges and development progress of the active control approaches (in open- or closed-loop configurations). A brief description of feedback control, adaptive control, model-based control and sliding mode control are provided first by introducing a simplified Rijke-type combustion system. The modelled combustion system provides an invaluable platform to evaluate the performance of these feedback controllers and a transient growth controller. The performance of these controllers are compared and discussed. An outline of theoretical, numerical and experimental investigations are then provided to overview the research and development progress made during the last 4 decades. Finally, potential, challenges and issues involved with the design, application and implementation of active combustion control strategies on a practical engine system are highlighted.

Active control of reheat buzz

NASA Astrophysics Data System (ADS)

Dowling, A. P.; Hooper, N.; Langhorne, P. J.; Bloxsidge, G. J.

1987-01-01

Reheat buzz is a low-frequency combustion instability involving the propagation of longitudinal pressure waves inside a duct in which a flame is anchored. Active control has been successfully applied to this instability. The controller alters the upstream acoustic boundary condition and thereby changes the energy balance in duct. Control is found to reduce the peak in the pressure spectrum due to the combustion instability by 20 dB. The acoustic energy in the whole 0-800-Hz bandwidth is reduced to about 10 percent of its uncontrolled value. A comparison with numerical calculations is presented.

Engine-Level Simulation of Liquid Rocket Combustion Instabilities: Transcritical Combustion Simulations in Single Injector Configurations

DTIC Science & Technology

2012-03-01

simple 1-step mechanism taking into account 4 species: CH4, O2, CO2 and H2O. Figure 2. Multiblock grid for the CVRC experiment. Left: Overall view, Right... Supercritical (and subcritical) fluid behavior and modeling: drops, streams, shear and mixing layers, jets and sprays. Progress in Energy and...hydrogen shear-coaxial jet flames at supercritical pressure. Com- bustion science and technology, 178(1-3):229–252, 2006. 12 B. E. Poling, J. M. Prausnitz

Method and device for diagnosing and controlling combustion instabilities in internal combustion engines operating in or transitioning to homogeneous charge combustion ignition mode

DOEpatents

Wagner, Robert M [Knoxville, TN; Daw, Charles S [Knoxville, TN; Green, Johney B [Knoxville, TN; Edwards, Kevin D [Knoxville, TN

2008-10-07

This invention is a method of achieving stable, optimal mixtures of HCCI and SI in practical gasoline internal combustion engines comprising the steps of: characterizing the combustion process based on combustion process measurements, determining the ratio of conventional and HCCI combustion, determining the trajectory (sequence) of states for consecutive combustion processes, and determining subsequent combustion process modifications using said information to steer the engine combustion toward desired behavior.

Active Control of Mixing and Combustion, from Mechanisms to Implementation

NASA Astrophysics Data System (ADS)

Ghoniem, Ahmed F.

2001-11-01

Implementation of active control in complex processes, of the type encountered in high Reynolds number mixing and combustion, is predicated upon the identification of the underlying mechanisms and the construction of reduced order models that capture their essential characteristics. The mechanisms of interest must be shown to be amenable to external actuations, allowing optimal control strategies to exploit the delicate interactions that lead to the desired outcome. Reduced order models are utilized in defining the form and requisite attributes of actuation, its relationship to the monitoring system and the relevant control algorithms embedded in a feedforward or a feedback loop. The talk will review recent work on active control of mixing in combustion devices in which strong shear zones concur with mixing, combustion stabilization and flame anchoring. The underlying mechanisms, e.g., stability of shear flows, formation/evolution of large vortical structures in separating and swirling flows, their mutual interactions with acoustic fields, flame fronts and chemical kinetics, etc., are discussed in light of their key roles in mixing, burning enhancement/suppression, and combustion instability. Subtle attributes of combustion mechanisms are used to suggest the requisite control strategies.

Overview of Current Activities in Combustion Instability

DTIC Science & Technology

2015-10-02

and avoid liquid rocket engine combustion stability problems Approach:  1) Develop a  SOA  combustion stability software package  called Stable...phase II will invest in Multifidelity Tools and Methodologies – CSTD will develop a SOA combustion stability software package called Stable Combustion

Characterizing dilute combustion instabilities in a multi-cylinder spark-ignited engine using symbolic analysis.

PubMed

Daw, C S; Finney, C E A; Kaul, B C; Edwards, K D; Wagner, R M

2015-02-13

Spark-ignited internal combustion engines have evolved considerably in recent years in response to increasingly stringent regulations for emissions and fuel economy. One new advanced engine strategy ustilizes high levels of exhaust gas recirculation (EGR) to reduce combustion temperatures, thereby increasing thermodynamic efficiency and reducing nitrogen oxide emissions. While this strategy can be highly effective, it also poses major control and design challenges due to the large combustion oscillations that develop at sufficiently high EGR levels. Previous research has documented that combustion instabilities can propagate between successive engine cycles in individual cylinders via self-generated feedback of reactive species and thermal energy in the retained residual exhaust gases. In this work, we use symbolic analysis to characterize multi-cylinder combustion oscillations in an experimental engine operating with external EGR. At low levels of EGR, intra-cylinder oscillations are clearly visible and appear to be associated with brief, intermittent coupling among cylinders. As EGR is increased further, a point is reached where all four cylinders lock almost completely in phase and alternate simultaneously between two distinct bi-stable combustion states. From a practical perspective, it is important to understand the causes of this phenomenon and develop diagnostics that might be applied to ameliorate its effects. We demonstrate here that two approaches for symbolizing the engine combustion measurements can provide useful probes for characterizing these instabilities. © 2014 The Author(s) Published by the Royal Society. All rights reserved.

Computational Fluid Dynamics Modeling of the Operation of a Flame Ionization Sensor

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

Huckaby, E.D.; Chorpening, B.T.; Thornton, J.D.

The sensors and controls research group at the United States Department of Energy (DOE) National Energy Technology Laboratory (NETL) is continuing to develop the Combustion Control and Diagnostics Sensor (CCADS) for gas turbine applications. CCADS uses the electrical conduction of the charged species generated during the combustion process to detect combustion instabilities and monitor equivalence ratio. As part of this effort, combustion models are being developed which include the interaction between the electric field and the transport of charged species. The primary combustion process is computed using a flame wrinkling model (Weller et. al. 1998) which is a component ofmore » the OpenFOAM toolkit (Jasak et. al. 2004). A sub-model for the transport of charged species is attached to this model. The formulation of the charged-species model similar that applied by Penderson and Brown (1993) for the simulation of laminar flames. The sub-model consists of an additional flux due to the electric field (drift flux) added to the equations for the charged species concentrations and the solution the electric potential from the resolved charge density. The subgrid interactions between the electric field and charged species transport have been neglected. Using the above procedure, numerical simulations are performed and the results compared with several recent CCADS experiments.« less

Investigation of hypersonic shock-induced combustion in a hydrogen-air system

NASA Technical Reports Server (NTRS)

Ahuja, J. K.; Tiwari, S. N.; Singh, D. J.

1992-01-01

A numerical study is conducted to simulate the ballistic range experiments at Mach 5.11 and 6.46. The flow field is found to be unsteady with periodic instabilities originating in the stagnation zone. The unsteadiness of the flow field decreased with increase in the Mach number, thus indicating that it is possible to stabilize such flow fields with a high degree of overdrive. The frequency of periodic instability is determined using Fourier power spectrum and is found to be in good agreement with the experimental data. The physics of the instability is explained by the wave interaction models available in the literature.

Role of buoyancy and heat release in fire modeling, propagation, and instability

Treesearch

Shahid M. Mughal; Yousuff M. Hussaini; Scott L. Goodrick; Philip Cunningham

2007-01-01

In an investigation of the dynamics of coupled fluid-combustion-buoyancy driven problems, an idealised model formulation is used to investigate the role of buoyancy and heat release in an evolving boundary layer, with particular emphasis on examining underlying fluid dynamics to explain observed phenomena arising in forest fire propagation. The role played by the...

An experimental study of the effect of a pilot flame on technically pre-mixed, self-excited combustion instabilities

NASA Astrophysics Data System (ADS)

O'Meara, Bridget C.

Combustion instabilities are a problem facing the gas turbine industry in the operation of lean, pre-mixed combustors. Secondary flames known as "pilot flames" are a common passive control strategy for eliminating combustion instabilities in industrial gas turbines, but the underlying mechanisms responsible for the pilot flame's stabilizing effect are not well understood. This dissertation presents an experimental study of a pilot flame in a single-nozzle, swirl-stabilized, variable length atmospheric combustion test facility and the effect of the pilot on combustion instabilities. A variable length combustor tuned the acoustics of the system to excite instabilities over a range of operating conditions without a pilot flame. The inlet velocity was varied from 25 -- 50 m/s and the equivalence ratio was varied from 0.525 -- 0.65. This range of operating conditions was determined by the operating range of the combustion test facility. Stability at each operating condition and combustor length was characterized by measurements of pressure oscillations in the combustor. The effect of the pilot flame on the magnitude and frequency of combustor stability was then investigated. The mechanisms responsible for the pilot flame effect were studied using chemiluminescence flame images of both stable and unstable flames. Stable flame structure was investigated using stable flame images of CH* chemiluminescence emission. The effect of the pilot on stable flame metrics such as flame length, flame angle, and flame width was investigated. In addition, a new flame metric, flame base distance, was defined to characterize the effect of the pilot flame on stable flame anchoring of the flame base to the centerbody. The effect of the pilot flame on flame base anchoring was investigated because the improved stability with a pilot flame is usually attributed to improved flame anchoring through the recirculation of hot products from the pilot to the main flame base. Chemiluminescence images of unstable flames were used to identify several instability mechanisms and infer how these mechanisms are affected by the pilot flame. Flame images of cases in which the pilot flame did not eliminate the instability were investigated to understand why the pilot flame is not effective in certain cases. The phase of unstable pilot flame oscillations was investigated to determine how the phase of pilot flame oscillations may affect its ability to interfere with instability mechanisms in the main flame. A forced flame response study was conducted to determine the effect of inlet velocity oscillation amplitude on the pilot flame. The flame response was characterized by measurements of velocity oscillations in the injector and chemiluminescence intensity oscillations determined from flame images. As the forcing amplitude increases, the pilot flame's effect on the flame transfer function magnitude becomes weaker. Flame images show that as the forcing amplitude increases, the pilot flame oscillations increase, leading to an ineffective pilot. The results of the flame response portion of this study highlight the effect of instability amplitude on the ability of a pilot flame to eliminate a combustion instability.

Lectures on combustion theory

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

Burstein, S.Z.; Lax, P.D.; Sod, G.A.

1978-09-01

Eleven lectures are presented on mathematical aspects of combustion: fluid dynamics, deflagrations and detonations, chemical kinetics, gas flows, combustion instability, flame spread above solids, spark ignition engines, burning rate of coal particles and hydrocarbon oxidation. Separate abstracts were prepared for three of the lectures. (DLC)

A high accuracy sequential solver for simulation and active control of a longitudinal combustion instability

NASA Technical Reports Server (NTRS)

Shyy, W.; Thakur, S.; Udaykumar, H. S.

1993-01-01

A high accuracy convection scheme using a sequential solution technique has been developed and applied to simulate the longitudinal combustion instability and its active control. The scheme has been devised in the spirit of the Total Variation Diminishing (TVD) concept with special source term treatment. Due to the substantial heat release effect, a clear delineation of the key elements employed by the scheme, i.e., the adjustable damping factor and the source term treatment has been made. By comparing with the first-order upwind scheme previously utilized, the present results exhibit less damping and are free from spurious oscillations, offering improved quantitative accuracy while confirming the spectral analysis reported earlier. A simple feedback type of active control has been found to be capable of enhancing or attenuating the magnitude of the combustion instability.

Abstracts. 1978 AFOSR Contractors Meeting on Air-Breathing Combustion Dynamics and Kinetics, Ramada Inn-Downtown Dayton, Ohio, 10 - 13 October 1978

DTIC Science & Technology

1978-10-13

Combustion in G.D. Smith, C.E. Peters High Speed Flows AEDC/ARO (PO-78-0012) 5:00 ADJOURN 6:30 Social Hour (Cash Bar) Ramada Inn Banquet 12 Oct. 78...which would sustain the instability structures observed in a number of problemA . During the initial phase of the development of the instabilities, the

On Nonlinear Combustion Instability in Liquid Propellant Rocket Motors

NASA Technical Reports Server (NTRS)

Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

2004-01-01

All liquid propellant rocket instability calculations in current use have limited value in the predictive sense and serve mainly as a correlating framework for the available data sets. The well-known n-t model first introduced by Crocco and Cheng in 1956 is still used as the primary analytical tool of this type. A multitude of attempts to establish practical analytical methods have achieved only limited success. These methods usually produce only stability boundary maps that are of little use in making critical design decisions in new motor development programs. Recent progress in understanding the mechanisms of combustion instability in solid propellant rockets"' provides a firm foundation for a new approach to prediction, diagnosis, and correction of the closely related problems in liquid motor instability. For predictive tools to be useful in the motor design process, they must have the capability to accurately determine: 1) time evolution of the pressure oscillations and limit amplitude, 2) critical triggering pulse amplitude, and 3) unsteady heat transfer rates at injector surfaces and chamber walls. The method described in this paper relates these critical motor characteristics directly to system design parameters. Inclusion of mechanisms such as wave steepening, vorticity production and transport, and unsteady detonation wave phenomena greatly enhance the representation of key features of motor chamber oscillatory behavior. The basic theoretical model is described and preliminary computations are compared to experimental data. A plan to develop the new predictive method into a comprehensive analysis tool is also described.

An Experimental Investigation of Self-Excited Combustion Dynamics in a Single Element Lean Direct Injection (LDI) Combustor

NASA Astrophysics Data System (ADS)

Gejji, Rohan M.

The management of combustion dynamics in gas turbine combustors has become more challenging as strict NOx/CO emission standards have led to engine operation in a narrow, lean regime. While premixed or partially premixed combustor configurations such as the Lean Premixed Pre-vaporized (LPP), Rich Quench Lean burn (RQL), and Lean Direct Injection (LDI) have shown a potential for reduced NOx emissions, they promote a coupling between acoustics, hydrodynamics and combustion that can lead to combustion instabilities. These couplings can be quite complex, and their detailed understanding is a pre-requisite to any engine development program and for the development of predictive capability for combustion instabilities through high-fidelity models. The overarching goal of this project is to assess the capability of high-fidelity simulation to predict combustion dynamics in low-emissions gas turbine combustors. A prototypical lean-direct-inject combustor was designed in a modular configuration so that a suitable geometry could be found by test. The combustor comprised a variable length air plenum and combustion chamber, air swirler, and fuel nozzle located inside a subsonic venturi. The venturi cross section and the fuel nozzle were consistent with previous studies. Test pressure was 1 MPa and variables included geometry and acoustic resonance, inlet temperatures, equivalence ratio, and type of liquid fuel. High-frequency pressure measurements in a well-instrumented metal chamber yielded frequencies and mode shapes as a function of inlet air temperature, equivalence ratio, fuel nozzle placement, and combustor acoustic resonances. The parametric survey was a significant effort, with over 105 tests on eight geometric configurations. A good dataset was obtained that could be used for both operating-point-dependent quantitative comparisons, and testing the ability of the simulation to predict more global trends. Results showed a very strong dependence of instability amplitude on the geometric configuration of the combustor, i.e., its acoustic resonance characteristics, with measured pressure fluctuation amplitudes ranged from 5 kPa (0.5% of mean pressure) to 200 kPa ( 20% of mean pressure) depending on combustor geometry. The stability behavior also showed a consistent and pronounced dependence on equivalence ratio and inlet air temperature. Instability amplitude increased with higher equivalence ratio and with lower inlet air temperature. A pronounced effect of fuel nozzle location on the combustion dynamics was also observed. Combustion instabilities with the fuel nozzle at the throat of the venturi throat were stronger than in the configuration with fuel nozzle 2.6 mm upstream of the nozzle. A second set of dynamics data was based on high-response-rate laser-based combustion diagnostics using an optically accessible combustor section. High-frequency measurements of OH*-chemiluminescence and OH-PLIF and velocity fields using PIV were obtained at a relatively stable, low equivalence ratio case and a less stable case at higher equivalence ratio. PIV measurements were performed at 5 kHz for non-reacting flow but glare from the cylindrical quartz chamber limited the field of view to a small region in the combustor. Quantitative and qualitative comparisons were made for five different combinations of geometry and operating condition that yielded discriminating stability behavior in the experiment with simulations that were carried out concurrently. Comparisons were made on the basis of trends and pressure mode data as well as with OH-PLIF measurements for the baseline geometry at equivalence ratios of 0.44 and 0.6. Overall, the ability of the simulation to match experimental data and trends was encouraging. Dynamic Mode Decomposition (DMD) analysis was performed on two sets of computations - a global 2-step chemistry mechanism and an 18-step chemistry mechanism - and the OH-PLIF images to allow comparison of dynamic patterns of heat release and OH distribution in the combustion zone. The DMD analysis was able to identify similar dominant unstable modes in the combustor. Recommendations for future work are based on the continued requirement for quantitative and spatio-temporally resolved data for direct comparison with computational efforts to develop predictive capabilities for combustion instabilities at relevant operating conditions. Discriminating instability behavior for the prototypical combustor demonstrated in this study is critical for any robust validation effort Unit physics based scaling of the current effort to multi-element combustors along with improvement in diagnostic techniques and analysis efforts are recommended for advancement in understanding of the complex physics in the multi-phase, three dimensional and turbulent combustion processes in the LDI combustor.

Multi-Fidelity Framework for Modeling Combustion Instability

DTIC Science & Technology

2016-07-27

generated from the reduced-domain dataset. Evaluations of the framework are performed based on simplified test problems for a model rocket combustor showing...generated from the reduced-domain dataset. Evaluations of the framework are performed based on simplified test problems for a model rocket combustor...of Aeronautics and Astronautics and Associate Fellow AIAA. ‡ Professor Emeritus. § Senior Scientist, Rocket Propulsion Division and Senior Member

Measurement and analysis of combustion response to transverse combustion instability

NASA Astrophysics Data System (ADS)

Pomeroy, Brian R.

This research aimed to gain a better understanding of the response of a gas-centered swirl coaxial injector to transverse combustion instability. The goals of the research were to develop a combustion chamber that would be able to spontaneously produce transverse combustion instability at elevated pressures and temperatures. Methods were also developed to analyze high-speed video images to understand the response of the injector. A combustion chamber was designed that produced high levels of instabilities. The chamber was capable of pressures as high as 1034 kPa (150 psi) and operated using decomposed 90% hydrogen peroxide and JP-8. The chamber used an array of seven gas-centered swirl coaxial injectors that exhibited linear instability to drive the transverse oscillations. The injector elements would operate in a monopropellant configuration flowing only decomposed hydrogen peroxide or in a bipropellant configuration. The location of the bipropellant injectors could be varied to change the level of the instability in the chamber from 10% of the chamber pressure up to 70% of the chamber pressure. A study element was placed in the center of the chamber where it was observed simultaneously by two high-speed video cameras which recorded a backlit video to show the location of the fuel spray and the location of the emitted CH* chemiluminescence. The videos were synchronized with high frequency pressure measurements to gain a full understanding of the physics in the combustion chamber. Results showed that the study element was coupled with the first mode velocity wave. This was expected due to the first mode velocity anti-node being located in the center of the chamber. The velocity is an absolute maximum twice during each cycle so the coupling with the second mode pressure was also investigated showing a possible coupling with both the velocity and pressure. The results of the first mode velocity showed that, as the velocity wave traveled through the chamber, the fuel spray was first displaced into an oxidizer rich region and secondly followed by a reaction in the direction of travel of the velocity wave as the peak velocity traveled through the region. The deflection into the oxidizer rich region was especially apparent in high-level instabilities. In low-level instabilities, the velocity wave was not strong enough to fully displace the fuel, and instead the oxidizer core was deflected into the fuel annulus causing a reaction in the direction of travel of the velocity wave. Neighboring oxidizer only injectors caused a lower reaction upstream as the neighboring oxidizer was deflected into the fuel annulus. The region of the fluctuating emitted light agreed well in size, shape and location with a correlation between the first mode velocity and combustion leading to the conclusion that the first mode is highly coupled with velocity. The second mode variance did not agree well with either the velocity or pressure correlation leading to a conclusion that it is coupled with both velocity and pressure. When comparing the variance to the pressure or velocity correlation, parts of the variance compared in shape and location to the pressure or velocity correlation, however, this was not true for all regions of response. This leads to a conclusion that both the pressure and velocity can be affecting the second mode. The second mode chemiluminescence emission occurs when the velocity is nearly zero in the chamber leading to the reaction to not be deflected and occurring downstream of the injector. At the same time, the second mode pressure is a minimum so an increase in mass flow could be responsible for the increased reaction. The methods and combustion chamber used to study the response of an injector can be used in the future to study any injector or combination of injectors placed at various locations in the chamber to study pressure or velocity coupling. The chemiluminescence data can be used to develop transfer functions for use in low fidelity computational models and can be used to validate high fidelity CFD.

A preliminary analysis of low frequency pressure oscillations in hybrid rocket motors

NASA Technical Reports Server (NTRS)

Jenkins, Rhonald M.

1994-01-01

Past research with hybrid rockets has suggested that certain motor operating conditions are conducive to the formation of pressure oscillations, or flow instabilities, within the motor combustion chamber. These combustion-related vibrations or pressure oscillations may be encountered in virtually any type of rocket motor and typically fall into three frequency ranges: low frequency oscillations (0-300 Hz); intermediate frequency oscillations (400-1000 Hz); and high frequency oscillations (greater than 1000 Hz). In general, combustion instability is characterized by organized pressure oscillations occurring at well-defined intervals with pressure peaks that may maintain themselves, grow, or die out. Usually, such peaks exceed +/- 5% of the mean chamber pressure. For hybrid motors, these oscillations have been observed to grow to a limiting amplitude which may be dependent on factors such as fuel characteristics, oxidizer injector characteristics, average chamber pressure, oxidizer mass flux, combustion chamber length, and grain geometry. The approach taken in the present analysis is to develop a modified chamber length, L, instability theory which accounts for the relationship between pressure and oxidizer to fuel concentration ratio in the motor.

Axisymmetric single shear element combustion instability experiment

NASA Technical Reports Server (NTRS)

Breisacher, Kevin J.

1993-01-01

The combustion stability characteristics of a combustor consisting of a single shear element and a cylindrical chamber utilizing LOX and gaseous hydrogen as propellants are presented. The combustor geometry and the resulting longitudinal mode instability are axisymmetric. Hydrogen injection temperature and pyrotechnic pulsing were used to determine stability boundaries. Mixture ratio, fuel annulus gap, and LOX post configuration were varied. Performance and stability data were obtained for chamber pressures of 300 and 1000 psia.

Axisymmetric single shear element combustion instability experiment

NASA Technical Reports Server (NTRS)

Breisacher, Kevin J.

1993-01-01

The combustion stability characteristics of a combustor consisting of a single shear element and a cylindrical chamber utilizing LOX and gaseous hydrogen as propellants are presented. The combustor geometry and the resulting longitudinal mode instability are axisymmetric. Hydrogen injection temperature and pyrotechnic pulsing were used to determine stability boundaries. Mixture ratio, fuel annulus gap, and LOX post configuration were varied. Performance and stability data are presented for chamber pressures of 300 and 1000 psia.

On Pulsating and Cellular Forms of Hydrodynamic Instability in Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sacksteder, Kurt (Technical Monitor)

1998-01-01

An extended Landau-Levich model of liquid-propellant combustion, one that allows for a local dependence of the burning rate on the (gas) pressure at the liquid-gas interface, exhibits not only the classical hydrodynamic cellular instability attributed to Landau but also a pulsating hydrodynamic instability associated with sufficiently negative pressure sensitivities. Exploiting the realistic limit of small values of the gas-to-liquid density ratio p, analytical formulas for both neutral stability boundaries may be obtained by expanding all quantities in appropriate powers of p in each of three distinguished wave-number regimes. In particular, composite analytical expressions are derived for the neutral stability boundaries A(sub p)(k), where A, is the pressure sensitivity of the burning rate and k is the wave number of the disturbance. For the cellular boundary, the results demonstrate explicitly the stabilizing effect of gravity on long-wave disturbances, the stabilizing effect of viscosity (both liquid and gas) and surface tension on short-wave perturbations, and the instability associated with intermediate wave numbers for negative values of A(sub p), which is characteristic of many hydroxylammonium nitrate-based liquid propellants over certain pressure ranges. In contrast, the pulsating hydrodynamic stability boundary is insensitive to gravitational and surface-tension effects but is more sensitive to the effects of liquid viscosity because, for typical nonzero values of the latter, the pulsating boundary decreases to larger negative values of A(sub p) as k increases through O(l) values. Thus, liquid-propellant combustion is predicted to be stable (that is, steady and planar) only for a range of negative pressure sensitivities that lie below the cellular boundary that exists for sufficiently small negative values of A(sub p) and above the pulsating boundary that exists for larger negative values of this parameter.

Control of Combustion-Instabilities Through Various Passive Devices

NASA Technical Reports Server (NTRS)

Frendi, Abdelkader; Nesman, Tom; Canabal, Francisco

2005-01-01

Results of a computational study on the effectiveness of various passive devices for the control of combustion instabilities are presented. An axi-symmetric combustion chamber is considered. The passive control devices investigated are, baffles, Helmholtz resonators and quarter-waves. The results show that a Helmholtz resonator with a smooth orifice achieves the best control results, while a baffle is the least effective for the frequency tested. At high sound pressure levels, the Helmholtz resonator is less effective. It is also found that for a quarter wave, the smoothness of the orifice has the opposite effect than the Helmholtz resonator, i.e. results in less control.

Perturbation solutions of combustion instability problems

NASA Technical Reports Server (NTRS)

Googerdy, A.; Peddieson, J., Jr.; Ventrice, M.

1979-01-01

A method involving approximate modal analysis using the Galerkin method followed by an approximate solution of the resulting modal-amplitude equations by the two-variable perturbation method (method of multiple scales) is applied to two problems of pressure-sensitive nonlinear combustion instability in liquid-fuel rocket motors. One problem exhibits self-coupled instability while the other exhibits mode-coupled instability. In both cases it is possible to carry out the entire linear stability analysis and significant portions of the nonlinear stability analysis in closed form. In the problem of self-coupled instability the nonlinear stability boundary and approximate forms of the limit-cycle amplitudes and growth and decay rates are determined in closed form while the exact limit-cycle amplitudes and growth and decay rates are found numerically. In the problem of mode-coupled instability the limit-cycle amplitudes are found in closed form while the growth and decay rates are found numerically. The behavior of the solutions found by the perturbation method are in agreement with solutions obtained using complex numerical methods.

Nonlinear Longitudinal Mode Instability in Liquid Propellant Rocket Engine Preburners

NASA Technical Reports Server (NTRS)

Sims, J. D. (Technical Monitor); Flandro, Gary A.; Majdalani, Joseph; Sims, Joseph D.

2004-01-01

Nonlinear pressure oscillations have been observed in liquid propellant rocket instability preburner devices. Unlike the familiar transverse mode instabilities that characterize primary combustion chambers, these oscillations appear as longitudinal gas motions with frequencies that are typical of the chamber axial acoustic modes. In several respects, the phenomenon is similar to longitudinal mode combustion instability appearing in low-smoke solid propellant motors. An important feature is evidence of steep-fronted wave motions with very high amplitude. Clearly, gas motions of this type threaten the mechanical integrity of associated engine components and create unacceptably high vibration levels. This paper focuses on development of the analytical tools needed to predict, diagnose, and correct instabilities of this type. For this purpose, mechanisms that lead to steep-fronted, high-amplitude pressure waves are described in detail. It is shown that such gas motions are the outcome of the natural steepening process in which initially low amplitude standing acoustic waves grow into shock-like disturbances. The energy source that promotes this behavior is a combination of unsteady combustion energy release and interactions with the quasi-steady mean chamber flow. Since shock waves characterize the gas motions, detonation-like mechanisms may well control the unsteady combustion processes. When the energy gains exceed the losses (represented mainly by nozzle and viscous damping), the waves can rapidly grow to a finite amplitude limit cycle. Analytical tools are described that allow the prediction of the limit cycle amplitude and show the dependence of this wave amplitude on the system geometry and other design parameters. This information can be used to guide corrective procedures that mitigate or eliminate the oscillations.

Combustion Instabilities In Solid Propellant Rocket Motors

DTIC Science & Technology

2004-01-01

instability in a self-excited system, sketched in Figure 1.5(a). In contrast, the initial transient in a linear system forced by an invariant external agent ...because the driving agent supplies only ¯nite power. (a) (b) Figure 1.5. Transient behavior of (a) Self Excited Linearly Unstable Motions; (b) Forced...the vibration of a Helmholtz resonator obtained, for example, by blowing over the open end of a bottle. The cause in a combustion chamber may be the

Modeling of Nonlinear Combustion Instability in Solid Propellant Rocket Motors

DTIC Science & Technology

1984-02-01

34. .. .°. .., . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . .... . . . . ..°.... . .°-""... ’o.’ . . °o: :--, - .:" . "" . °° - - 54. Flandro , 0. A., "Solid Propellant Acoustic Admittance...such as those due to Gary , 2 1) Gourlay and Morris ( 2 2 ) and Mas- (23)son are more involved, both from a program development, and computational

Spray formation processes of impinging jet injectors

NASA Technical Reports Server (NTRS)

Anderson, W. E.; Ryan, H. M.; Pal, S.; Santoro, R. J.

1993-01-01

A study examining impinging liquid jets has been underway to determine physical mechanisms responsible for combustion instabilities in liquid bi-propellant rocket engines. Primary atomization has been identified as an important process. Measurements of atomization length, wave structure, and drop size and velocity distribution were made under various ambient conditions. Test parameters included geometric effects and flow effects. It was observed that pre-impingement jet conditions, specifically whether they were laminar or turbulent, had the major effect on primary atomization. Comparison of the measurements with results from a two dimensional linear aerodynamic stability model of a thinning, viscous sheet were made. Measured turbulent impinging jet characteristics were contrary to model predictions; the structure of waves generated near the point of jet impingement were dependent primarily on jet diameter and independent of jet velocity. It has been postulated that these impact waves are related to pressure and momentum fluctuations near the impingement region and control the eventual disintegration of the liquid sheet into ligaments. Examination of the temporal characteristics of primary atomization (ligament shedding frequency) strongly suggests that the periodic nature of primary atomization is a key process in combustion instability.

Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors

NASA Technical Reports Server (NTRS)

Padmanabhan, M. S.; Powell, E. A.; Zinn, B. T.

1975-01-01

A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.

Liquid rocket engine combustion stabilization devices

NASA Technical Reports Server (NTRS)

1974-01-01

Combustion instability, which results from a coupling of the combustion process and the fluid dynamics of the engine system, was investigated. The design of devices which reduce coupling (combustion chamber baffles) and devices which increase damping (acoustic absorbers) are described. Included in the discussion are design criteria and recommended practices, structural and mechanical design, thermal control, baffle geometry, baffle/engine interactions, acoustic damping analysis, and absorber configurations.

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).

Computational Study of Combustion Dynamics in a Single-Element Lean Direct Injection Gas Turbine Combustor

DTIC Science & Technology

2013-12-01

instabilities for different equivalence ratios and fuel injector locations. Comparisons of the computational and experimental results are carried out using...the fuel injector and swirler as the full geometry. The full geometry in Fig. 1 (b) is the same as the one that was used in the experiments. In Fig. 1...combustion instabilities in both the simulations and the experiments. Fuel injector (Detail B) sits in the converging-diverging section connecting the air

Combustion Instability Phenomena of Importance to Liquid Propellant Engines

DTIC Science & Technology

1993-07-31

July 31, 1993 !Annual 01 July 92-30 June 93 4 TITLE AND SUBTITLE s . FUNDING NUMBERS (U) Combustion Instability Phenomena of Importance to Liquid...Propellant Engines PE - 61102F IPR- 2308 6. AUTHOR( S ) SA - Al G - AFOSR-91-0336 R. J. Santoro and W. E. Anderson 7. PERFORMING ORGANIZATION NAME( S ) AND...technology are not being used; instead, current engines are essentially being built with the same injector designs that were developed in the 1960’ s . I e The

Combustion Instability Phenomena of Importance to Liquid Propellant Engines

DTIC Science & Technology

1994-08-31

TITLE AND SUBTITLE S . FUNDING NUMBERS PE - 61102F (U) Combustion Instability Phenomena of Importance to PR - 2308 Liquid Propellant Engines SA - Al...6. AUTHOR( S ) G - AFOSR-91-0336 R.J. Santoro and W.E. Anderson 7. PERFORMING ORGANIZATION NAME( S ) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT...IMONITORING AGENCY NAME( S ) AND ADDRESS(ES) . -. .r • ,.,. AGE ¶77~~UME AFOSR/NA A &--LMCT- 110 Duncan Avenue, Suite B115 m ELEC IL B911ing AFB, DC 20332-0001

Identification and quantification analysis of nonlinear dynamics properties of combustion instability in a diesel engine

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

Yang, Li-Ping, E-mail: yangliping302@hrbeu.edu.cn; Ding, Shun-Liang; Song, En-Zhe

The cycling combustion instabilities in a diesel engine have been analyzed based on chaos theory. The objective was to investigate the dynamical characteristics of combustion in diesel engine. In this study, experiments were performed under the entire operating range of a diesel engine (the engine speed was changed from 600 to 1400 rpm and the engine load rate was from 0% to 100%), and acquired real-time series of in-cylinder combustion pressure using a piezoelectric transducer installed on the cylinder head. Several methods were applied to identify and quantitatively analyze the combustion process complexity in the diesel engine including delay-coordinate embedding, recurrencemore » plot (RP), Recurrence Quantification Analysis, correlation dimension (CD), and the largest Lyapunov exponent (LLE) estimation. The results show that the combustion process exhibits some determinism. If LLE is positive, then the combustion system has a fractal dimension and CD is no more than 1.6 and within the diesel engine operating range. We have concluded that the combustion system of diesel engine is a low-dimensional chaotic system and the maximum values of CD and LLE occur at the lowest engine speed and load. This means that combustion system is more complex and sensitive to initial conditions and that poor combustion quality leads to the decrease of fuel economy and the increase of exhaust emissions.« less

Identification and quantification analysis of nonlinear dynamics properties of combustion instability in a diesel engine.

PubMed

Yang, Li-Ping; Ding, Shun-Liang; Litak, Grzegorz; Song, En-Zhe; Ma, Xiu-Zhen

2015-01-01

The cycling combustion instabilities in a diesel engine have been analyzed based on chaos theory. The objective was to investigate the dynamical characteristics of combustion in diesel engine. In this study, experiments were performed under the entire operating range of a diesel engine (the engine speed was changed from 600 to 1400 rpm and the engine load rate was from 0% to 100%), and acquired real-time series of in-cylinder combustion pressure using a piezoelectric transducer installed on the cylinder head. Several methods were applied to identify and quantitatively analyze the combustion process complexity in the diesel engine including delay-coordinate embedding, recurrence plot (RP), Recurrence Quantification Analysis, correlation dimension (CD), and the largest Lyapunov exponent (LLE) estimation. The results show that the combustion process exhibits some determinism. If LLE is positive, then the combustion system has a fractal dimension and CD is no more than 1.6 and within the diesel engine operating range. We have concluded that the combustion system of diesel engine is a low-dimensional chaotic system and the maximum values of CD and LLE occur at the lowest engine speed and load. This means that combustion system is more complex and sensitive to initial conditions and that poor combustion quality leads to the decrease of fuel economy and the increase of exhaust emissions.

Model-Based Self-Tuning Multiscale Method for Combustion Control

NASA Technical Reports Server (NTRS)

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

2006-01-01

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

Daniel Sokolowski in the Rocket Operations Building

NASA Image and Video Library

1966-06-21

Dan Sokolowski worked as an engineering coop student at the National Aeronautics and Space Administration (NASA) Lewis Research Center from 1962 to 1966 while earning his Mechanical Engineering degree from Purdue. At the time of this photograph Sokolowski had just been hired as a permanent NASA employee in the Chemical Rocket Evaluation Branch of the Chemical Rocket Division. He had also just won a regional American Institute of Aeronautics and Astronautics competition for his paper on high and low-frequency combustion instability. The resolution of the low-frequency combustion instability, or chugging, in liquid hydrogen rocket systems was one of Lewis’ more significant feats of the early 1960s. In most rocket engine combustion chambers, the pressure, temperature, and flows are in constant flux. The engine is considered to be operating normally if the fluctuations remain random and within certain limits. Lewis researchers used high-speed photography to study and define Pratt and Whitney’s RL-10’s combustion instability by throttling the engine under the simulated flight conditions. They found that the injection of a small stream of helium gas into the liquid-oxygen tank immediately stabilized the system. Sokolowski’s later work focused on combustion in airbreathing engines. In 1983 was named Manager of a multidisciplinary program aimed at improving durability of combustor and turbine components. After 39 years Sokolowski retired from NASA in September 2002.

Viscous Analysis of Pulsating Hydrodynamic Instability and Thermal Coupling Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sacksteder, Kurt (Technical Monitor)

2000-01-01

A pulsating form of hydrodynamic instability has recently been shown to arise during liquid-propellant deflagration in those parameter regimes where the pressure-dependent burning rate is characterized by a negative pressure sensitivity. This type of instability can coexist with the classical cellular, or Landau form of hydrodynamic instability, with the occurrence of either dependent on whether the pressure sensitivity is sufficiently large or small in magnitude. For the inviscid problem, it has been shown that, when the burning rate is realistically allowed to depend on temperature as well as pressure, sufficiently large values of the temperature sensitivity relative to the pressure sensitivity causes like pulsating form of hydrodynamic instability to become dominant. In that regime, steady, planar burning becomes intrinsically unstable to pulsating disturbances whose wave numbers are sufficiently small. This analysis is extended to the fully viscous case, where it is shown that although viscosity is stabilizing for intermediate and larger wave number perturbations, the intrinsic pulsating instability for small wave numbers remains. Under these conditions, liquid-propellant combustion is predicted to be characterized by large unsteady cells along the liquid/gas interface.

Nonlinear Combustion Instability Prediction

NASA Technical Reports Server (NTRS)

Flandro, Gary

2010-01-01

The liquid rocket engine stability prediction software (LCI) predicts combustion stability of systems using LOX-LH2 propellants. Both longitudinal and transverse mode stability characteristics are calculated. This software has the unique feature of being able to predict system limit amplitude.

High-Speed Linear Raman Spectroscopy for Instability Analysis of a Bluff Body Flame

NASA Technical Reports Server (NTRS)

Kojima, Jun; Fischer, David

2013-01-01

We report a high-speed laser diagnostics technique based on point-wise linear Raman spectroscopy for measuring the frequency content of a CH4-air premixed flame stabilized behind a circular bluff body. The technique, which primarily employs a Nd:YLF pulsed laser and a fast image-intensified CCD camera, successfully measures the time evolution of scalar parameters (N2, O2, CH4, and H2O) in the vortex-induced flame instability at a data rate of 1 kHz. Oscillation of the V-shaped flame front is quantified through frequency analysis of the combustion species data and their correlations. This technique promises to be a useful diagnostics tool for combustion instability studies.

Numerical Modelling of Staged Combustion Aft-Injected Hybrid Rocket Motors

NASA Astrophysics Data System (ADS)

Nijsse, Jeff

The staged combustion aft-injected hybrid (SCAIH) rocket motor is a promising design for the future of hybrid rocket propulsion. Advances in computational fluid dynamics and scientific computing have made computational modelling an effective tool in hybrid rocket motor design and development. The focus of this thesis is the numerical modelling of the SCAIH rocket motor in a turbulent combustion, high-speed, reactive flow framework accounting for solid soot transport and radiative heat transfer. The SCAIH motor is modelled with a shear coaxial injector with liquid oxygen injected in the center at sub-critical conditions: 150 K and 150 m/s (Mach ≈ 0.9), and a gas-generator gas-solid mixture of one-third carbon soot by mass injected in the annual opening at 1175 K and 460 m/s (Mach ≈ 0.6). Flow conditions in the near injector region and the flame anchoring mechanism are of particular interest. Overall, the flow is shown to exhibit instabilities and the flame is shown to anchor directly on the injector faceplate with temperatures in excess of 2700 K.

Comparison of Laminar and Linear Eddy Model Closures for Combustion Instability Simulations

DTIC Science & Technology

2015-07-01

14. ABSTRACT Unstable liquid rocket engines can produce highly complex dynamic flowfields with features such as rapid changes in temperature and...applicability. In the present study, the linear eddy model (LEM) is applied to an unstable single element liquid rocket engine to assess its performance and to...Sankaran‡ Air Force Research Laboratory, Edwards AFB, CA, 93524 Unstable liquid rocket engines can produce highly complex dynamic flowfields with features

Active control of combustion instabilities

NASA Astrophysics Data System (ADS)

Al-Masoud, Nidal A.

A theoretical analysis of active control of combustion thermo-acoustic instabilities is developed in this dissertation. The theoretical combustion model is based on the dynamics of a two-phase flow in a liquid-fueled propulsion system. The formulation is based on a generalized wave equation with pressure as the dependent variable, and accommodates all influences of combustion, mean flow, unsteady motions and control inputs. The governing partial differential equations are converted to an equivalent set of ordinary differential equations using Galerkin's method by expressing the unsteady pressure and velocity fields as functions of normal mode shapes of the chamber. This procedure yields a representation of the unsteady flow field as a system of coupled nonlinear oscillators that is used as a basis for controllers design. Major research attention is focused on the control of longitudinal oscillations with both linear and nonlinear processes being considered. Starting with a linear model using point actuators, the optimal locations of actuators and sensors are developed. The approach relies on the quantitative measures of the degree of controllability and component cost. These criterion are arrived at by considering the energies of the system's inputs and outputs. The optimality criteria for sensor and actuator locations provide a balance between the importance of the lower order (controlled) and the higher (residual) order modes. To address the issue of uncertainties in system's parameter, the minimax principles based controller is used. The minimax corresponds to finding the best controller for the worst parameter deviation. In other words, choosing controller parameters to minimize, and parameter deviation to maximize some quadratic performance metric. Using the minimax-based controller, a remarkable improvement in the control system's ability to handle parameter uncertainties is achieved when compared to the robustness of the regular control schemes such as LQR and LQG. Since the observed instabilities are harmonic, the concept of "harmonic input" is successfully implemented using a parametric controller to eliminate the thermo-acoustic instability. This control scheme relies on the determination of a phase-shift to maximize the energy dissipation and a controller gain to assure stability and minimize a pre-specified performance index. The closed loop control law design is based on finding an optimal phase angle such that the heat release produced by secondary oscillatory fuel injection is out of phase with the mode's pressure oscillations, thus maximizing energy dissipation, and on finding the limits on the controller gain that ensures system stability. The optimal gains are determined using ITA, ISE, ITAE performance indices. Simulations show successful implementation of the proposed technique.

Triggered instabilities in rocket motors and active combustion control for an incinerator afterburner

NASA Astrophysics Data System (ADS)

Wicker, Josef M.

1999-11-01

Two branches of research are conducted in this thesis. The first deals with nonlinear combustion response as a mechanism for triggering combustion instabilities in solid rocket motors. A nonlinear wave equation is developed to study a wide class of combustion response functions to second-order in fluctuation amplitude. Conditions for triggering are derived from analysis of limit cycles, and regions of triggering are found in parametric space. Introduction of linear cross-coupling and quadratic self-coupling among the acoustic modes appears to be how the nonlinear combustion response produces triggering to a stable limit cycle. Regions of initial conditions corresponding to stable pulses were found, suggesting that stability depends on initial phase angle and harmonic content, as well as the composite amplitude, of the pulse. Also, dependence of nonlinear stability upon system parameters is considered. The second part of this thesis presents research for a controller to improve the emissions of an incinerator afterburner. The developed controller was experimentally tested at the Naval Air Warfare Center (NAWC), on a 50kW-scale model of an afterburner for Naval shipboard incinerator applications. Acoustic forcing of the combustor's reacting shear layer is used to control the formation of coherent vortical structures, within which favorable fuel-air mixing and efficient combustion can occur. Laser-based measurements of CO emissions are used as the performance indicator for the combustor. The controller algorithm is based on the downhill simplex method and adjusts the shear layer forcing parameters in order to minimize the CO emissions. The downhill simplex method was analyzed with respect to its behavior in the face of time-variation of the plant and noise in the sensor signal, and was modified to account for these difficulties. The control system has experimentally demonstrated the ability (1) to find optimal control action for single- and multi-variable control, (2) to maintain optimal control for time-varying operating states, and (3) to automatically adjust auxiliary fuel in response to changing stoichiometry of the incoming waste pyrolysis gas. Also presented but not tested in the experiments are an expert-type model-guidance feature to aid convergence of the controller to optimum control, and methodology for maintaining flammability.

Elimination of High-Frequency Combustion Instability in the Fastrac Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Rocker, Marvin; Nesman, Thomas E.

1998-01-01

NASA's Marshall Space Flight Center(MSFC) has been tasked with developing a 60,000 pound thrust, pump-fed, LOX/RP-1 engine under the Advanced Space Transportation Program(ASTP). This government-led design has been designated the Fastrac engine. The X-34 vehicle will use the Fastrac engine as the main propulsion system. The X-34 will be a suborbital vehicle developed by the Orbital Sciences Corporation. The X-34 vehicle will be launched from an L-1011 airliner. After launch, the X-34 vehicle will be able to climb to altitudes up to 250,000 feet and reach speeds up to Mach 8, over a mission range of 500 miles. The overall length, wingspan, and gross takeoff weight of the X-34 vehicle are 58.3 feet, 27.7 feet and 45,000 pounds, respectively. This report summarizes the plan of achieving a Fastrac thrust chamber assembly(TCA) stable bomb test that meets the JANNAF standards, the Fastrac TCA design, and the combustion instabilities exhibited by the Fastrac TCA during testing at MSFC's test stand 116 as determined from high-frequency fluctuating pressure measurements. This report also summarizes the characterization of the combustion instabilities from the pressure measurements and the steps taken to eliminate the instabilities.

Propellant injection strategy for suppressing acoustic combustion instability

NASA Astrophysics Data System (ADS)

Diao, Qina

Shear-coaxial injector elements are often used in liquid-propellant-rocket thrust chambers, where combustion instabilities remain a significant problem. A conventional solution to the combustion instability problem relies on passive control techniques that use empirically-developed hardware such as acoustic baffles and tuned cavities. In addition to adding weight and decreasing engine performance, these devices are designed using trial-and-error methods, which do not provide the capability to predict the overall system stability characteristics in advance. In this thesis, two novel control strategies that are based on propellant fluid dynamics were investigated for mitigating acoustic instability involving shear-coaxial injector elements. The new control strategies would use a set of controlled injectors allowing local adjustment of propellant flow patterns for each operating condition, particularly when instability could become a problem. One strategy relies on reducing the oxidizer-fuel density gradient by blending heavier methane with the main fuel, hydrogen. Another strategy utilizes modifying the equivalence ratio to affect the acoustic impedance through mixing and reaction rate changes. The potential effectiveness of these strategies was assessed by conducting unit-physics experiments. Two different model combustors, one simulating a single-element injector test and the other a double-element injector test, were designed and tested for flame-acoustic interaction. For these experiments, the Reynolds number of the central oxygen jet was kept between 4700 and 5500 making the injector flames sufficiently turbulent. A compression driver, mounted on one side of the combustor wall, provided controlled acoustic excitation to the injector flames, simulating the initial phase of flame-acoustic interaction. Acoustic excitation was applied either as band-limited white noise forcing between 100 Hz and 5000 Hz or as single-frequency, fixed-amplitude forcing at 1150 Hz which represented a frequency least amplified by any resonance. Effects of each control strategy on flame-acoustic interaction were assessed in terms of modifying the acoustic resonance characteristics subject to white-noise excitation and changes in flame brush thickness under single-frequency excitation. In the methane blending experiments, the methane mole fraction was varied between 0% and 63%. Under white noise excitation, up to 16% shift in a resonant frequency was observed but the acoustic pressure spectra remained qualitatively similar. For the fixed frequency forcing, the spatial extent of flame-acoustic interaction was substantially reduced. In the other experiments, the equivalence ratio of the control injector was varied between zero and infinity, causing up to 40% shift in a resonant frequency as well as changes in the acoustic pressure spectrum. These results open up the possibility of employing flow-based control to prevent combustion instabilities in liquid-fueled rockets.

Historical problem areas lessons learned

NASA Technical Reports Server (NTRS)

Sackheim, Bob; Fester, Dale A.

1991-01-01

Historical problem areas in space transportation propulsion technology are identified in viewgraph form. Problem areas discussed include materials compatibility, contamination, pneumatic/feed system flow instabilities, instabilities in rocket engine combustion and fuel sloshing, exhaust plume interference, composite rocket nozzle failure, and freeze/thaw damage.

Flame propagation in two-dimensional solids: Particle-resolved studies with complex plasmas

NASA Astrophysics Data System (ADS)

Yurchenko, S. O.; Yakovlev, E. V.; Couëdel, L.; Kryuchkov, N. P.; Lipaev, A. M.; Naumkin, V. N.; Kislov, A. Yu.; Ovcharov, P. V.; Zaytsev, K. I.; Vorob'ev, E. V.; Morfill, G. E.; Ivlev, A. V.

2017-10-01

Using two-dimensional (2D) complex plasmas as an experimental model system, particle-resolved studies of flame propagation in classical 2D solids are carried out. Combining experiments, theory, and molecular dynamics simulations, we demonstrate that the mode-coupling instability operating in 2D complex plasmas reveals all essential features of combustion, such as an activated heat release, two-zone structure of the self-similar temperature profile ("flame front"), as well as thermal expansion of the medium and temperature saturation behind the front. The presented results are of relevance for various fields ranging from combustion and thermochemistry, to chemical physics and synthesis of materials.

Spray combustion under oscillatory pressure conditions

NASA Technical Reports Server (NTRS)

Jacobs, H. R.; Santoro, R. J.

1991-01-01

The performance and stability of liquid rocket engines is often argued to be significantly impacted by atomization and droplet vaporization processes. In particular, combustion instability phenomena may result from the interactions between the oscillating pressure field present in the rocket combustor and the fuel and oxidizer injection process. Few studies have been conducted to examine the effects of oscillating pressure fields on spray formation and its evolution under rocket engine conditions. The pressure study is intended to address the need for such studies. In particular, two potentially important phenomena are addressed in the present effort. The first involves the enhancement of the atomization process for a liquid jet subjected to an oscillating pressure field of known frequency and amplitude. The objective of this part of the study is to examine the coupling between the pressure field and or the resulting periodically perturbed velocity field on the breakup of the liquid jet. In particular, transverse mode oscillations are of interest since such modes are considered of primary importance in combustion instability phenomena. The second aspect of the project involves the effects of an oscillating pressure on droplet coagulation and secondary atomization. The objective of this study is to examine the conditions under which phenomena following the atomization process are affected by perturbations to the pressure or velocity fields. Both coagulation and represent a coupling mechanism between the pressure field and the energy release process in rocket combustors. It is precisely this coupling which drives combustion instability phenomena. Consequently, the present effort is intended to provide the fundamental insights needed to evaluate these processes as important mechanisms in liquid rocket instability phenomena.

Viscous and Thermal Effects on Hydrodynamic Instability in Liquid-Propellant Combustion

NASA Technical Reports Server (NTRS)

Margolis, Stephen B.; Sacksteder, Kurt (Technical Monitor)

2000-01-01

A pulsating form of hydrodynamic instability has recently been shown to arise during the deflagration of liquid propellants in those parameter regimes where the pressure-dependent burning rate is characterized by a negative pressure sensitivity. This type of instability can coexist with the classical cellular, or Landau, form of hydrodynamic instability, with the occurrence of either dependent on whether the pressure sensitivity is sufficiently large or small in magnitude. For the inviscid problem, it has been shown that when the burning rate is realistically allowed to depend on temperature as well as pressure, that sufficiently large values of the temperature sensitivity relative to the pressure sensitivity causes the pulsating form of hydrodynamic instability to become dominant. In that regime, steady, planar burning becomes intrinsically unstable to pulsating disturbances whose wavenumbers are sufficiently small. In the present work, this analysis is extended to the fully viscous case, where it is shown that although viscosity is stabilizing for intermediate and larger wavenumber perturbations, the intrinsic pulsating instability for small wavenumbers remains. Under these conditions, liquid-propellant combustion is predicted to be characterized by large unsteady cells along the liquid/gas interface.

The causes of unstable engine idle speed and their solutions

NASA Astrophysics Data System (ADS)

Yang, Fan

2018-06-01

There are many types of engines. The most commonly used engine for automobiles is the internal combustion engine. Internal combustion engines use a four-stroke combustion cycle to convert gasoline into motion. The four-stroke approach, also known as the "Ototo cycle," commemorates Nicklaus Otto, who invented it in 1867. The working cycle of a four-stroke engine consists of four piston strokes, ie, intake stroke, compression stroke, power stroke, and exhaust stroke. This article focuses on the cause of the instability of the four-stroke engine and its solution. There are many reasons for the instability of the engine, so this article will be divided into four areas: intake system, fuel system, ignition system and mechanical structure. Based on the above reasons, the corresponding solution is proposed.

Effects of porous insert on flame dynamics in a lean premixed swirl-stabilized combustor

NASA Astrophysics Data System (ADS)

Brown, Marcus; Agrawal, Ajay; Allen, James; Kornegay, John

2016-11-01

In this study, we investigated different methods of determining the effect a porous insert has on flame dynamics during lean premixed combustion. A metallic porous insert is used to mitigate instabilities in a swirl-stabilized combustor. Thermoacoustic instabilities are seen as negative consequences of lean premixed combustion and eliminating them is the motivation for our research. Three different diagnostics techniques with high-speed Photron SA5 cameras were used to monitor flame characteristics. Particle image velocimetry (PIV) was used to observe vortical structures and recirculation zones within the combustor. Using planar laser induced fluorescence (PLIF), we were able to observe changes in the reaction zones during instabilities. Finally, utilizing a color high-speed camera, visual images depicting a flame's oscillations during the instability were captured. Using these monitoring techniques, we are able to support the claims made in previous studies stating that the porous insert in the combustor significantly reduces the thermoacoustic instability. Funding for this research was provided by the NSF REU site Grant EEC 1358991 and NASA Grant NNX13AN14A.

Nitramine smokeless propellant research

NASA Technical Reports Server (NTRS)

Cohen, N. S.; Strand, L. P.

1977-01-01

A transient ballistics and combustion model is derived to represent the closed vessel experiment that is widely used to characterize propellants. A computer program is developed to solve the time-dependent equations, and is applied to explain aspects of closed vessel behavior. In the case of nitramine propellants the cratering of the burning surface associated with combustion above break-point pressures augments the effective burning rate as deduced from the closed vessel experiment. Low pressure combustion is significantly affected by the ignition process and, in the case of nitramine propellants, by the developing and changing surface structure. Thus, burning rates deduced from the closed vessel experiment may or may not agree with those measured in the equilibrium strand burner. Series of T burner experiments are performed to compare the combustion instability characteristics of nitramine (HMX) containing propellants and ammonium perchlorate (AP)propellants. Although ash produced by more fuel rich propellants could have provided mechanical suppression, results from clean-burning propellants permit the conclusion that HMX reduces the acoustic driving.

Richtmyer-Meshkov instability in shock-flame interactions

NASA Astrophysics Data System (ADS)

Massa, Luca; Pallav Jha Collaboration

2011-11-01

Shock-flame interactions occur in supersonic mixing and detonation formation. Therefore, their analysis is important to explosion safety, internal combustion engine performance, and supersonic combustor design. The fundamental process at the basis of the interaction is the Richtmyer-Meshkov instability supported by the density difference between burnt and fresh mixtures. In the present study we analyze the effect of reactivity on the Richtmyer- Meshkov instability with particular emphasis on combustion lengths that typify the scaling between perturbation growth and induction. The results of the present linear analysis study show that reactivity changes the perturbation growth rate by developing a non-zero pressure gradient at the flame surface. The baroclinic torque based on the density gradient across the flame acts to slow down the instability growth for high wave numbers. A non-hydrodynamic flame representation leads to the definition of an additional scaling Peclet number, the effects of which are investigated. It is found that an increased flame-contact separation destabilizes the contact discontinuity by augmenting the tangential shear.

Evaluation of Start Transient Oscillations with the J-2X Engine Gas Generator Assembly

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Morgan, C. J.; Casiano, M. J.

2015-01-01

During development of the gas generator for the liquid oxygen/liquid hydrogen propellant J-2X rocket engine, distinctive and oftentimes high-amplitude pressure oscillations and hardware vibrations occurred during the start transient of nearly every workhorse gas generator assembly test, as well as during many tests of engine system hardware. These oscillations appeared whether the steady-state conditions exhibited stable behavior or not. They occurred similarly with three different injector types, and with every combustion chamber configuration tested, including chamber lengths ranging over a 5:1 range, several different nozzle types, and with or without a side branch line simulating a turbine spin start gas supply line. Generally, two sets of oscillations occurred, one earlier in the start transient and at higher frequencies, and the other almost immediately following and at lower frequencies. Multiple dynamic pressure measurements in the workhorse combustion chambers indicated that the oscillations were associated with longitudinal acoustic modes of the combustion chambers, with the earlier and higher frequency oscillation usually related to the second longitudinal acoustic mode and the later and lower frequency oscillation usually related to the first longitudinal acoustic mode. Given that several early development gas generator assemblies exhibited unstable behavior at frequencies near the first longitudinal acoustic modes of longer combustion chambers, the start transient oscillations are presumed to provide additional insight into the nature of the combustion instability mechanisms. Aspects of the steadystate oscillations and combustion instabilities from development and engine system test programs have been reported extensively in the three previous JANNAF Liquid Propulsion Subcommittee meetings (see references below). This paper describes the hardware configurations, start transient sequence operations, and transient and dynamic test data during the start transient. The implications of these results on previous analyses and understanding of the combustion instability observed during steady-state conditions, especially the effects of injector influences, is discussed.

Combustion Stability Verification for the Thrust Chamber Assembly of J-2X Developmental Engines 10001, 10002, and 10003

NASA Technical Reports Server (NTRS)

Morgan, C. J.; Hulka, J. R.; Casiano, M. J.; Kenny, R. J.; Hinerman, T. D.; Scholten, N.

2015-01-01

The J-2X engine, a liquid oxygen/liquid hydrogen propellant rocket engine available for future use on the upper stage of the Space Launch System vehicle, has completed testing of three developmental engines at NASA Stennis Space Center. Twenty-one tests of engine E10001 were conducted from June 2011 through September 2012, thirteen tests of the engine E10002 were conducted from February 2013 through September 2013, and twelve tests of engine E10003 were conducted from November 2013 to April 2014. Verification of combustion stability of the thrust chamber assembly was conducted by perturbing each of the three developmental engines. The primary mechanism for combustion stability verification was examining the response caused by an artificial perturbation (bomb) in the main combustion chamber, i.e., dynamic combustion stability rating. No dynamic instabilities were observed in the TCA, although a few conditions were not bombed. Additional requirements, included to guard against spontaneous instability or rough combustion, were also investigated. Under certain conditions, discrete responses were observed in the dynamic pressure data. The discrete responses were of low amplitude and posed minimal risk to safe engine operability. Rough combustion analyses showed that all three engines met requirements for broad-banded frequency oscillations. Start and shutdown transient chug oscillations were also examined to assess the overall stability characteristics, with no major issues observed.

Transient Simulation of Pressure Oscillations in the Fuel Feedline of the Fastrac Engine Thrust Chamber

NASA Technical Reports Server (NTRS)

Bullard, Brad

1998-01-01

During mainstage testing of the 60,000 lbf thrust Fastrac thrust chamber at MSFC's Test Stand 116 (TS 116), sustained, large amplitude oscillations near 530 Hz were observed in the pressure data. These oscillations were detected both in the RP-1 feedline, downstream of the cavitating venturi, and in the combustion chamber. The driver of the instability is believed to be feedline excitation driven by either periodic cavity collapse at the exit of the cavitating venturi or combustion instability. In covitating venturi, static pressure drops as the flow passes through a constriction resembling a converging-diverging nozzle until the vapor pressure is reached. At the venturi throat, the flow is essentially choked, which is why these devices are typically used for mass flow rate control and disturbance isolation. Typically, a total pressure drop of 15% or more across the venturi is required for cavitation. For much larger pressure differentials, unstable cavities can form and subsequently collapse downstream of the throat. Although the disturbances generated by cavitating venturis is generally considered to be broad-band, this type of phenomena could generate periodic behavior capable of exciting the feedline. An excitation brought about by combustion instability would result from the coupling of a combustion chamber acoustic mode and a feedline resonance frequency. This type of coupling is referred to as "buzz" and is not uncommon for engines in this thrust range.

Modeling and simulation of combustion chamber and propellant dynamics and issues in active control of combustion instabilities

NASA Astrophysics Data System (ADS)

Isella, Giorgio Carlo

A method for a comprehensive approach to analysis of the dynamics of an actively controlled combustion chamber, with detailed analysis of the combustion models for the case of a solid rocket propellant, is presented here. The objective is to model the system as interconnected blocks describing the dynamics of the chamber, combustion and control. The analytical framework for the analysis of the dynamics of a combustion chamber is based on spatial averaging, as introduced by Culick. Combustion dynamics are analyzed for the case of a solid propellant. Quasi-steady theory is extended to include the dynamics of the gas-phase and also of a surface layer. The models are constructed so that they produce a combustion response function for the solid propellant that can be immediately introduced in the our analytical framework. The principal objective mechanisms responsible for the large sensitivity, observed experimentally, of propellant response to small variations. We show that velocity coupling, and not pressure coupling, has the potential to be the mechanism responsible for that high sensitivity. We also discuss the effect of particulate modeling on the global dynamics of the chamber and revisit the interpretation of the intrinsic stability limit for burning of solid propellants. Active control is also considered. Particular attention is devoted to the effect of time delay (between sensing and actuation); several methods to compensate for it are discussed, with numerical examples based on the approximate analysis produced by our framework. Experimental results are presented for the case of a Dump Combustor. The combustor exhibits an unstable burning mode, defined through the measurement of the pressure trace and shadowgraph imaging. The transition between stable and unstable modes of operation is characterized by the presence of hysteresis, also observed in other experimental works, and hence not a special characteristic of this combustor. Control is introduced in the form of pulsed secondary fuel. We show the capability of forcing the transition from unstable to stable burning, hence extending the stable operating regime of the combustor. The transition, characterized by the use of a shadowgraph movie sequence, is attributed to a combined fluid-mechanic and combustion mechanism.

CFD code evaluation for internal flow modeling

NASA Technical Reports Server (NTRS)

Chung, T. J.

1990-01-01

Research on the computational fluid dynamics (CFD) code evaluation with emphasis on supercomputing in reacting flows is discussed. Advantages of unstructured grids, multigrids, adaptive methods, improved flow solvers, vector processing, parallel processing, and reduction of memory requirements are discussed. As examples, researchers include applications of supercomputing to reacting flow Navier-Stokes equations including shock waves and turbulence and combustion instability problems associated with solid and liquid propellants. Evaluation of codes developed by other organizations are not included. Instead, the basic criteria for accuracy and efficiency have been established, and some applications on rocket combustion have been made. Research toward an ultimate goal, the most accurate and efficient CFD code, is in progress and will continue for years to come.

Combustion stability analysis of preburners in liquid propellant rocket engines during shutdown

NASA Technical Reports Server (NTRS)

Lim, Kair-Chuan; George, Paul E., II

1987-01-01

A linearized one-dimensional lumped-parameter model capable of predicting the occurrence of the low frequency combustion instability (chugging) experienced during preburner shutdown in the Space Shuttle Main Engines is discussed, and predictions are compared with NASA experimental results. Results from a parametric study of parameters including chamber pressure, fuel and oxygen temperatures, and the effective bulk modulus of the liquid oxidizer suggest that chugging is probably affected by conditions at shutdown through the fuel and oxidizer temperatures. It is suggested that chugging is initiated when the fuel, oxidizer, and helium temperature and flow rates pass into an unstable region, and that chugging may be terminated by decaying pressures.

Combustion instability investigations on the BR710 jet engine

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

Konrad, W.; Brehm, N.; Kameier, F.

1998-01-01

During the development of the BR710 jet engine, audible combustor instabilities (termed rumble) occurred. Amplitudes measured with test cell microphones were up to 130 dB at around 100 Hz. Disturbances of this amplitude are clearly undesirable, even if only present during start-up, and a research program was initiated to eliminate the problem. Presented here is the methodical and structured approach used to identify, understand, and remove the instability. Some reference is made to theory, which was used for guidance, but the focus of the work is on the research done to find the cause of the problem and to correctmore » it. The investigation followed two separate, but parallel, paths--one looking in detail at individual components of the engine to identify possible involvement in the instability and the other looking at the pressure signals from various parts of a complete engine to help pinpoint the source of the disturbance. The main cause of the BR710 combustor rumble was found to be a self-excited aerodynamic instability arising from the design of the fuel injector head. In the end, minor modifications lead to spray pattern changes, which greatly reduced the combustor noise. As a result of this work, new recommendation are made for reducing the risk of combustion instabilities in jet engines.« less

Characterisation of acoustic energy content in an experimental combustion chamber with and without external forcing

NASA Astrophysics Data System (ADS)

Webster, S.; Hardi, J.; Oschwald, M.

2015-03-01

The influence of injection conditions on rocket engine combustion stability is investigated for a sub-scale combustion chamber with shear coaxial injection elements and the propellant combination hydrogen-oxygen. The experimental results presented are from a series of tests conducted at subcritical and supercritical pressures for oxygen and for both ambient and cryogenic temperature hydrogen. The stability of the system is characterised by the root mean squared amplitude of dynamic combustion chamber pressure in the upper part of the acoustic spectrum relevant for high frequency combustion instabilities. Results are presented for both unforced and externally forced combustion chamber configurations. It was found that, for both the unforced and externally forced configurations, the injection velocity had the strongest influence on combustion chamber stability. Through the use of multivariate linear regression the influence of hydrogen injection temperature and hydrogen injection mass flow rate were best able to explain the variance in stability for dependence on injection velocity ratio. For unforced tests turbulent jet noise from injection was found to dominate the energy content of the signal. For the externally forced configuration a non-linear regression model was better able to predict the variance, suggesting the influence of non-linear behaviour. The response of the system to variation of injection conditions was found to be small; suggesting that the combustion chamber investigated in the experiment is highly stable.

Optical diagnostics in gas turbine combustors

NASA Astrophysics Data System (ADS)

Woodruff, Steven D.

1999-01-01

Deregulation of the power industry and increasingly tight emission controls are pushing gas turbine manufacturers to develop engines operating at high pressure for efficiency and lean fuel mixtures to control NOx. This combination also gives rise to combustion instabilities which threaten engine integrity through acoustic pressure oscillations and flashback. High speed imaging and OH emission sensors have been demonstrated to be invaluable tools in characterizing and monitoring unstable combustion processes. Asynchronous imaging technique permit detailed viewing of cyclic flame structure in an acoustic environment which may be modeled or utilized in burner design . The response of the flame front to the acoustic pressure cycle may be tracked with an OH emission monitor using a sapphire light pipe for optical access. The OH optical emission can be correlated to pressure sensor data for better understanding of the acoustical coupling of the flame. Active control f the combustion cycle can be implemented using an OH emission sensor for feedback.

A theoretical evaluation of rigid baffles in suppression of combustion instability

NASA Technical Reports Server (NTRS)

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

1976-01-01

An analytical technique for the prediction of the effects of rigid baffles on the stability of liquid propellant combustors is presented. A three dimensional combustor model characterized by a concentrated combustion source at the chamber injector and a constant Mach number nozzle is used. The linearized partial differential equations describing the unsteady flow field are solved by an eigenfunction matching method. Boundary layer corrections to this unsteady flow are used to evaluate viscous and turbulence effects within the flow. An integral stability relationship is then employed to predict the decay rate of the oscillations. Results show that sufficient dissipation exists to indicate that the proper mechanism of baffle damping is a fluid dynamic loss. The response of the dissipation model to varying baffle blade length, mean flow Mach number and oscillation amplitude is examined.

Microgravity Combustion Science and Fluid Physics Experiments and Facilities for the ISS

NASA Technical Reports Server (NTRS)

Lauver, Richard W.; Kohl, Fred J.; Weiland, Karen J.; Zurawski, Robert L.; Hill, Myron E.; Corban, Robert R.

2001-01-01

At the NASA Glenn Research Center, the Microgravity Science Program supports both ground-based and flight experiment research in the disciplines of Combustion Science and Fluid Physics. Combustion Science research includes the areas of gas jet diffusion flames, laminar flames, burning of droplets and misting fuels, solids and materials flammability, fire and fire suppressants, turbulent combustion, reaction kinetics, materials synthesis, and other combustion systems. The Fluid Physics discipline includes the areas of complex fluids (colloids, gels, foams, magneto-rheological fluids, non-Newtonian fluids, suspensions, granular materials), dynamics and instabilities (bubble and drop dynamics, magneto/electrohydrodynamics, electrochemical transport, geophysical flows), interfacial phenomena (wetting, capillarity, contact line hydrodynamics), and multiphase flows and phase changes (boiling and condensation, heat transfer, flow instabilities). A specialized International Space Station (ISS) facility that provides sophisticated research capabilities for these disciplines is the Fluids and Combustion Facility (FCF). The FCF consists of the Combustion Integrated Rack (CIR), the Fluids Integrated Rack (FIR) and the Shared Accommodations Rack and is designed to accomplish a large number of science investigations over the life of the ISS. The modular, multiuser facility is designed to optimize the science return within the available resources of on-orbit power, uplink/downlink capacity, crew time, upmass/downmass, volume, etc. A suite of diagnostics capabilities, with emphasis on optical techniques, will be provided to complement the capabilities of the subsystem multiuser or principal investigator-specific experiment modules. The paper will discuss the systems concept, technical capabilities, functionality, and the initial science investigations in each discipline.

Invited Review. Combustion instability in spray-guided stratified-charge engines. A review

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

Fansler, Todd D.; Reuss, D. L.; Sick, V.

2015-02-02

Our article reviews systematic research on combustion instabilities (principally rare, random misfires and partial burns) in spray-guided stratified-charge (SGSC) engines operated at part load with highly stratified fuel -air -residual mixtures. Results from high-speed optical imaging diagnostics and numerical simulation provide a conceptual framework and quantify the sensitivity of ignition and flame propagation to strong, cyclically varying temporal and spatial gradients in the flow field and in the fuel -air -residual distribution. For SGSC engines using multi-hole injectors, spark stretching and locally rich ignition are beneficial. Moreover, combustion instability is dominated by convective flow fluctuations that impede motion of themore » spark or flame kernel toward the bulk of the fuel, coupled with low flame speeds due to locally lean mixtures surrounding the kernel. In SGSC engines using outwardly opening piezo-electric injectors, ignition and early flame growth are strongly influenced by the spray's characteristic recirculation vortex. For both injection systems, the spray and the intake/compression-generated flow field influence each other. Factors underlying the benefits of multi-pulse injection are identified. Finally, some unresolved questions include (1) the extent to which piezo-SGSC misfires are caused by failure to form a flame kernel rather than by flame-kernel extinction (as in multi-hole SGSC engines); (2) the relative contributions of partially premixed flame propagation and mixing-controlled combustion under the exceptionally late-injection conditions that permit SGSC operation on E85-like fuels with very low NO x and soot emissions; and (3) the effects of flow-field variability on later combustion, where fuel-air-residual mixing within the piston bowl becomes important.« less

Combustion-transition interaction in a jet flame

NASA Astrophysics Data System (ADS)

Yule, A. J.; Chigier, N. A.; Ralph, S.; Boulderstone, R.; Ventura, J.

1980-01-01

The transition between laminar and turbulent flow in a round jet flame is studied experimentally. Comparison is made between transition in non-burning and burning jets and between jet flames with systematic variation in initial Reynolds number and equivalence ratio. Measurements are made using laser anemometry, miniature thermocouples, ionization probes, laser-schlieren and high speed cine films. Compared with the cold jet, the jet flame has a longer potential core, undergoes a slower transition to turbulence, has lower values of fluctuating velocity near the burner but higher values further downstream, contains higher velocity gradients in the mixing layer region although the total jet width does not alter greatly in the first twenty diameters. As in the cold jet, transitional flow in the flame contains waves and vortices and these convolute and stretch the initially laminar interface burning region. Unlike the cold jet, which has Kelvin-Helmholtz instabilities, the jet flame can contain at least two initial instabilities; an inner high frequency combustion driven instability and an outer low frequency instability which may be influenced by buoyancy forces.

Combustion Stability Analyses of Coaxial Element Injectors with Liquid Oxygen/Liquid Methane Propellants

NASA Technical Reports Server (NTRS)

Hulka, J. R.

2010-01-01

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for in-space vehicles. This propellant combination has not been previously used in a flight-qualified engine system, so limited test data and analysis results are available at this stage of early development. NASA has funded several hardware-oriented activities with oxygen and methane propellants over the past several years with the Propulsion and Cryogenic Advanced Development (PCAD) project, under the Exploration Technology Development Program. As part of this effort, the NASA Marshall Space Flight Center has conducted combustion stability analyses of several of the configurations. This paper presents test data and analyses of combustion stability from the recent PCAD-funded test programs at the NASA MSFC. These test programs used swirl coaxial element injectors with liquid oxygen and liquid methane propellants. Oxygen was injected conventionally in the center of the coaxial element, and swirl was provided by tangential entry slots. Injectors with 28-element and 40-element patterns were tested with several configurations of combustion chambers, including ablative and calorimeter spool sections, and several configurations of fuel injection design. Low frequency combustion instability (chug) occurred with both injectors, and high-frequency combustion instability occurred at the first tangential (1T) transverse mode with the 40-element injector. In most tests, a transition between high-amplitude chug with gaseous methane flow and low-amplitude chug with liquid methane flow was readily observed. Chug analyses of both conditions were conducted using techniques from Wenzel and Szuch and from the Rocket Combustor Interactive Design and Analysis (ROCCID) code. The 1T mode instability occurred in several tests and was apparent by high-frequency pressure measurements as well as dramatic increases in calorimeter-measured heat flux throughout the chamber. Analyses of the transverse mode were conducted with ROCCID and empirical methods such as Hewitt d/V. This paper describes the test hardware configurations, test data, analysis methods, and presents results of the various analyses.

Droplet-turbulence interactions in subcritical and supercritical evaporating sprays

NASA Technical Reports Server (NTRS)

Santavicca, Domenic A.; Coy, Edward; Greenfield, Stuart; Song, Young-Hoon

1991-01-01

The objective of this research is to obtain an improved understanding of droplet turbulence interactions in vaporizing liquid sprays under conditions typical of those encountered in liquid fueled rocket engines. The interaction between liquid droplets and the surrounding turbulent gas flow affects droplet dispersion, droplet collisions, droplet vaporization and gas-phase, fuel-oxidant mixing, and therefore has a significant effect on the engine's combustion characteristics. An example of this is the role which droplet-turbulence interactions are believed to play in combustion instabilities. Despite their importance, droplet-turbulence interactions and their effect on liquid fueled rocket engine performance are not well understood. This is particularly true under supercritical conditions, where many conventional concepts, such as surface tension, no longer apply. Our limited understanding of droplet-turbulence interactions, under both subcritical conditions, represents a major limitation in our ability to design improved liquid previously unavailable information and valuable new insights which will directly impact the design of future liquid fueled rocket engines, as well as, allow for the development of significantly improved spray combustion models, making such models useful design tools.

Annual Research Briefs

NASA Technical Reports Server (NTRS)

Spinks, Debra (Compiler)

1997-01-01

This report contains the 1997 annual progress reports of the research fellows and students supported by the Center for Turbulence Research (CTR). Titles include: Invariant modeling in large-eddy simulation of turbulence; Validation of large-eddy simulation in a plain asymmetric diffuser; Progress in large-eddy simulation of trailing-edge turbulence and aeronautics; Resolution requirements in large-eddy simulations of shear flows; A general theory of discrete filtering for LES in complex geometry; On the use of discrete filters for large eddy simulation; Wall models in large eddy simulation of separated flow; Perspectives for ensemble average LES; Anisotropic grid-based formulas for subgrid-scale models; Some modeling requirements for wall models in large eddy simulation; Numerical simulation of 3D turbulent boundary layers using the V2F model; Accurate modeling of impinging jet heat transfer; Application of turbulence models to high-lift airfoils; Advances in structure-based turbulence modeling; Incorporating realistic chemistry into direct numerical simulations of turbulent non-premixed combustion; Effects of small-scale structure on turbulent mixing; Turbulent premixed combustion in the laminar flamelet and the thin reaction zone regime; Large eddy simulation of combustion instabilities in turbulent premixed burners; On the generation of vorticity at a free-surface; Active control of turbulent channel flow; A generalized framework for robust control in fluid mechanics; Combined immersed-boundary/B-spline methods for simulations of flow in complex geometries; and DNS of shock boundary-layer interaction - preliminary results for compression ramp flow.

Experimental, theoretical, and numerical studies of small scale combustion

NASA Astrophysics Data System (ADS)

Xu, Bo

Recently, the demand increased for the development of microdevices such as microsatellites, microaerial vehicles, micro reactors, and micro power generators. To meet those demands the biggest challenge is obtaining stable and complete combustion at relatively small scale. To gain a fundamental understanding of small scale combustion in this thesis, thermal and kinetic coupling between the gas phase and the structure at meso and micro scales were theoretically, experimentally, and numerically studied; new stabilization and instability phenomena were identified; and new theories for the dynamic mechanisms of small scale combustion were developed. The reduction of thermal inertia at small scale significantly reduces the response time of the wall and leads to a strong flame-wall coupling and extension of burning limits. Mesoscale flame propagation and extinction in small quartz tubes were theoretically, experimentally and numerically studied. It was found that wall-flame interaction in mesoscale combustion led to two different flame regimes, a heat-loss dominant fast flame regime and a wall-flame coupling slow flame regime. The nonlinear transition between the two flame regimes was strongly dependent on the channel width and flow velocity. It is concluded that the existence of multiple flame regimes is an inherent phenomenon in mesoscale combustion. In addition, all practical combustors have variable channel width in the direction of flame propagation. Quasi-steady and unsteady propagations of methane and propane-air premixed flames in a mesoscale divergent channel were investigated experimentally and theoretically. The emphasis was the impact of variable cross-section area and the flame-wall coupling on the flame transition between different regimes and the onset of flame instability. For the first time, spinning flames were experimentally observed for both lean and rich methane and propane-air mixtures in a broad range of equivalence ratios. An effective Lewis number to describe the competition between the mass transport in gas phase and the heat conduction in gas and solid phases was defined. Experimental observation and theoretical analysis suggested that the flame-wall coupling significantly increased the effective Lewis number and led to a new mechanism to promote the thermal diffusion instability. Due to the short flow residence time in small scale combustion, reactants, and oxidizers may not be able to be fully premixed before combustion. As such, non-premixed combustion plays an important role. Non-premixed mixing layer combustion within a constrained mesoscale channel was studied. Depending on the flow rate, it was found that there were two different flame regimes, an unsteady bimodal flame regime and a flame street regime with multiple stable triple flamelets. This multiple triple flame structure was identified experimentally for the first time. A scaling analytical model was developed to qualitatively explain the mechanism of flame streets. The effects of flow velocity, wall temperature, and Lewis number on the distance between flamelets and the diffusion flame length were also investigated. The results showed that the occurrence of flame street regimes was a combined effect of heat loss, curvature, diffusion, and dilution. To complete this thesis, experiments were conducted to measure the OH concentration using Planar Laser Induced Fluorescence (PLIF) in a confined mesoscale combustor. Some preliminary results have been obtained for the OH concentration of flamelets in a flame street. When the scale of the micro reactor is further reduced, the rarefied gas effect may become significant. In this thesis, a new concentration slip model to describe the rarefied gas effect on the species transport in microscale chemical reactors was obtained. The present model is general and recovers the existing models in the limiting cases. The analytical results showed the concentration slip was dominated by two different mechanisms, the surface reaction induced concentration slip (RIC) and the temperature slip induced concentration slip (TIC). It is found that the magnitude of RIC slip was proportional to the product of the Damkohler number and Knudsen number. The results showed the impact of reaction induced concentration slip (RIC slip) effects on catalytic reactions strongly depended on the Damkohler number, the Knudsen number, and the surface accommodation coefficient.

Solid Rocket Motor Combustion Instability Modeling in COMSOL Multiphysics

NASA Technical Reports Server (NTRS)

Fischbach, Sean R.

2015-01-01

Combustion instability modeling of Solid Rocket Motors (SRM) remains a topic of active research. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process, acoustics, and steady-state gas dynamics. Recent advances in defining the energy transport of disturbances within steady flow-fields have been applied by combustion stability modelers to improve the analysis framework [1, 2, 3]. Employing this more accurate global energy balance requires a higher fidelity model of the SRM flow-field and acoustic mode shapes. The current industry standard analysis tool utilizes a one dimensional analysis of the time dependent fluid dynamics along with a quasi-three dimensional propellant grain regression model to determine the SRM ballistics. The code then couples with another application that calculates the eigenvalues of the one dimensional homogenous wave equation. The mean flow parameters and acoustic normal modes are coupled to evaluate the stability theory developed and popularized by Culick [4, 5]. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The current study employs the COMSOL multiphysics finite element framework to model the steady flow-field parameters and acoustic normal modes of a generic SRM. The study requires one way coupling of the CFD High Mach Number Flow (HMNF) and mathematics module. The HMNF module evaluates the gas flow inside of a SRM using St. Robert's law to model the solid propellant burn rate, no slip boundary conditions, and the hybrid outflow condition. Results from the HMNF model are verified by comparing the pertinent ballistics parameters with the industry standard code outputs (i.e. pressure drop, thrust, ect.). These results are then used by the coefficient form of the mathematics module to determine the complex eigenvalues of the Acoustic Velocity Potential Equation (AVPE). The mathematics model is truncated at the nozzle sonic line, where a zero flux boundary condition is self-satisfying. The remaining boundaries are modeled with a zero flux boundary condition, assuming zero acoustic absorption on all surfaces. The results of the steady-state CFD and AVPE analyses are used to calculate the linear acoustic growth rate as is defined by Flandro and Jacob [2, 3]. In order to verify the process implemented within COMSOL we first employ the Culick theory and compare the results with the industry standard. After the process is verified, the Flandro/Jacob energy balance theory is employed and results displayed.

Impact of Variations on 1-D Flow in Gas Turbine Engines via Monte Carlo Simulations

NASA Technical Reports Server (NTRS)

Ngo, Khiem Viet; Tumer, Irem

2004-01-01

The unsteady compressible inviscid flow is characterized by the conservations of mass, momentum, and energy; or simply the Euler equations. In this paper, a study of the subsonic one-dimensional Euler equations with local preconditioning is presented using a modal analysis approach. Specifically, this study investigates the behavior of airflow in a gas turbine engine using the specified conditions at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine, to determine the impact of variations in pressure, velocity, temperature, and density at low Mach numbers. Two main questions motivate this research: 1) Is there any aerodynamic problem with the existing gas turbine engines that could impact aircraft performance? 2) If yes, what aspect of a gas turbine engine could be improved via design to alleviate that impact and to optimize aircraft performance? This paper presents an initial attempt to model the flow behavior in terms of their eigenfrequencies subject to the assumption of the uncertainty or variation (perturbation). The flow behavior is explored using simulation outputs from a customer-deck model obtained from Pratt & Whitney. Variations of the main variables (i.e., pressure, temperature, velocity, density) about their mean states at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine are modeled. Flow behavior is analyzed for the high-pressure compressor and combustion chamber utilizing the conditions on their left and right boundaries. In the same fashion, similar analyses are carried out for the high-pressure and low-pressure turbines. In each case, the eigenfrequencies that are obtained for different boundary conditions are examined closely based on their probabilistic distributions, a result of a Monte Carlo 10,000 sample simulation. Furthermore, the characteristic waves and wave response are analyzed and contrasted among different cases, with and without preconditioners. The results reveal the existence of flow instabilities due to the combined effect of variations and excessive pressures in the case of the combustion chamber and high-pressure turbine. Finally, a discussion is presented on potential impacts of the instabilities and what can be improved via design to alleviate them for a better aircraft performance.

Particle-Image Velocimetry in Microgravity Laminar Jet Diffusion Flames

NASA Technical Reports Server (NTRS)

Sunderland, P. B.; Greenberg, P. S.; Urban, D. L.; Wernet, M. P.; Yanis, W.

1999-01-01

This paper discusses planned velocity measurements in microgravity laminar jet diffusion flames. These measurements will be conducted using Particle-Image Velocimetry (PIV) in the NASA Glenn 2.2-second drop tower. The observations are of fundamental interest and may ultimately lead to improved efficiency and decreased emissions from practical combustors. The velocity measurements will support the evaluation of analytical and numerical combustion models. There is strong motivation for the proposed microgravity flame configuration. Laminar jet flames are fundamental to combustion and their study has contributed to myriad advances in combustion science, including the development of theoretical, computational and diagnostic combustion tools. Nonbuoyant laminar jet flames are pertinent to the turbulent flames of more practical interest via the laminar flamelet concept. The influence of gravity on these flames is deleterious: it complicates theoretical and numerical modeling, introduces hydrodynamic instabilities, decreases length scales and spatial resolution, and limits the variability of residence time. Whereas many normal-gravity laminar jet diffusion flames have been thoroughly examined (including measurements of velocities, temperatures, compositions, sooting behavior and emissive and absorptive properties), measurements in microgravity gas-jet flames have been less complete and, notably, have included only cursory velocity measurements. It is envisioned that our velocity measurements will fill an important gap in the understanding of nonbuoyant laminar jet flames.

Combustion dynamics in cryogenic rocket engines: Research programme at DLR Lampoldshausen

NASA Astrophysics Data System (ADS)

Hardi, Justin S.; Traudt, Tobias; Bombardieri, Cristiano; Börner, Michael; Beinke, Scott K.; Armbruster, Wolfgang; Nicolas Blanco, P.; Tonti, Federica; Suslov, Dmitry; Dally, Bassam; Oschwald, Michael

2018-06-01

The Combustion Dynamics group in the Rocket Propulsion Department at the German Aerospace Center (DLR), Lampoldshausen, strives to advance the understanding of dynamic processes in cryogenic rocket engines. Leveraging the test facilities and experimental expertise at DLR Lampoldshausen, the group has taken a primarily experimental approach to investigating transient flows, ignition, and combustion instabilities for over one and a half decades. This article provides a summary of recent achievements, and an overview of current and planned research activities.

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.

Study of Unsteady, Sphere-Driven, Shock-Induced Combustion for Application to Hypervelocity Airbreathing Propulsion

NASA Technical Reports Server (NTRS)

Axdahl, Erik; Kumar, Ajay; Wilhite, Alan

2011-01-01

A premixed, shock-induced combustion engine has been proposed in the past as a viable option for operating in the Mach 10 to 15 range in a single stage to orbit vehicle. In this approach, a shock is used to initiate combustion in a premixed fuel/air mixture. Apparent advantages over a conventional scramjet engine include a shorter combustor that, in turn, results in reduced weight and heating loads. There are a number of technical challenges that must be understood and resolved for a practical system: premixing of fuel and air upstream of the combustor without premature combustion, understanding and control of instabilities of the shock-induced combustion front, ability to produce sufficient thrust, and the ability to operate over a range of Mach numbers. This study evaluated the stability of the shock-induced combustion front in a model problem of a sphere traveling in a fuel/air mixture at high Mach numbers. A new, rapid analysis method was developed and applied to study such flows. In this method the axisymmetric, body-centric Navier-Stokes equations were expanded about the stagnation streamline of a sphere using the local similarity hypothesis in order to reduce the axisymmetric equations to a quasi-1D set of equations. These reduced sets of equations were solved in the stagnation region for a number of flow conditions in a premixed, hydrogen/air mixture. Predictions from the quasi-1D analysis showed very similar stable or unstable behavior of the shock-induced combustion front as compared to experimental studies and higher-fidelity computational results. This rapid analysis tool could be used in parametric studies to investigate effects of fuel rich/lean mixtures, non-uniformity in mixing, contaminants in the mixture, and different chemistry models.

Hybrid propulsion for launch vehicle boosters: A program status update

NASA Technical Reports Server (NTRS)

Carpenter, R. L.; Boardman, T. A.; Claflin, S. E.; Harwell, R. J.

1995-01-01

Results obtained in studying the origin and suppression of large-amplitude pressure oscillations in a 24 in. diameter hybrid motor using a liquid oxygen/hydroxylterminated polybutadiene/polycyclopentadiene propellant system are discussed. Tests conducted with liquid oxygen flow rates varying from 10 to 40 lbm/sec were designed to gauge the effectiveness of various vaporization chamber flow fields, injector designs, and levels of heat addition in suppressing high-frequency longitudinal mode oscillations. Longitudinal acoustic modes did not arise in any tests. However, initial testing revealed the presence of high-amplitude, sinusoidal, nonacoustic oscillations persisting throughout the burn durations. Analysis showed this to be analogous to chug mode instability in liquid rocket engines brought about by a coupling of motor combustion processes and the liquid oxygen feed system. Analytical models were developed and verified by test data to predict the amplitude and frequency of feed-system-coupled combustion pressure oscillations. Subsequent testing showed that increasing the feed system impedance eliminated the bulk mode instability. This paper documents the work completed to date in performance of the Hybrid Propulsion Technology for Launch Vehicle Boosters Program (NAS8-39942) sponsored by NASA's George C. Marshall Space Flight Center.

Theoretical Acoustic Absorber Design Approach for LOX/LCH4 Pintle Injector Rocket Engines

NASA Astrophysics Data System (ADS)

Candelaria, Jonathan

Liquid rocket engines, or LREs, have served a key role in space exploration efforts. One current effort involves the utilization of liquid oxygen (LOX) and liquid methane (LCH4) LREs to explore Mars with in-situ resource utilization for propellant production. This on-site production of propellant will allow for greater payload allocation instead of fuel to travel to the Mars surface, and refueling of propellants to travel back to Earth. More useable mass yields a greater benefit to cost ratio. The University of Texas at El Paso's (UTEP) Center for Space Exploration and Technology Research Center (cSETR) aims to further advance these methane propulsion systems with the development of two liquid methane - liquid oxygen propellant combination rocket engines. The design of rocket engines, specifically liquid rocket engines, is complex in that many variables are present that must be taken into consideration in the design. A problem that occurs in almost every rocket engine development program is combustion instability, or oscillatory combustion. It can result in the destruction of the rocket, subsequent destruction of the vehicle and compromise the mission. These combustion oscillations can vary in frequency from 100 to 20,000 Hz or more, with varying effects, and occur from different coupling phenomena. It is important to understand the effects of combustion instability, its physical manifestations, how to identify the instabilities, and how to mitigate or dampen them. Linear theory methods have been developed to provide a mathematical understanding of the low- to mid-range instabilities. Nonlinear theory is more complex and difficult to analyze mathematically, therefore no general analytical method that yields a solution exists. With limited resources, time, and the advice of our NASA mentors, a data driven experimental approach utilizing quarter wave acoustic dampener cavities was designed. This thesis outlines the methodology behind the design of an acoustic dampening system for a 500 lbf and a 2000 lbf throttleable liquid oxygen liquid methane pintle injector rocket engine.

Development and validation of spray models for investigating diesel engine combustion and emissions

NASA Astrophysics Data System (ADS)

Som, Sibendu

Diesel engines intrinsically generate NOx and particulate matter which need to be reduced significantly in order to comply with the increasingly stringent regulations worldwide. This motivates the diesel engine manufacturers to gain fundamental understanding of the spray and combustion processes so as to optimize these processes and reduce engine emissions. Strategies being investigated to reduce engine's raw emissions include advancements in fuel injection systems, efficient nozzle orifice design, injection and combustion control strategies, exhaust gas recirculation, use of alternative fuels such as biodiesel etc. This thesis explores several of these approaches (such as nozzle orifice design, injection control strategy, and biodiesel use) by performing computer modeling of diesel engine processes. Fuel atomization characteristics are known to have a significant effect on the combustion and emission processes in diesel engines. Primary fuel atomization is induced by aerodynamics in the near nozzle region as well as cavitation and turbulence from the injector nozzle. The breakup models that are currently used in diesel engine simulations generally consider aerodynamically induced breakup using the Kelvin-Helmholtz (KH) instability model, but do not account for inner nozzle flow effects. An improved primary breakup (KH-ACT) model incorporating cavitation and turbulence effects along with aerodynamically induced breakup is developed and incorporated in the computational fluid dynamics code CONVERGE. The spray simulations using KH-ACT model are "quasi-dynamically" coupled with inner nozzle flow (using FLUENT) computations. This presents a novel tool to capture the influence of inner nozzle flow effects such as cavitation and turbulence on spray, combustion, and emission processes. Extensive validation is performed against the non-evaporating spray data from Argonne National Laboratory. Performance of the KH and KH-ACT models is compared against the evaporating and combusting data from Sandia National Laboratory. The KH-ACT model is observed to provide better predictions for spray dispersion, axial velocity decay, sauter mean diameter, and liquid and lift-off length interplay which is attributed to the enhanced primary breakup predicted by this model. In addition, experimentally observed trends with changing nozzle conicity could only be captured by the KH-ACT model. Results further indicate that the combustion under diesel engine conditions is characterized by a double-flame structure with a rich premixed reaction zone near the flame stabilization region and a non-premixed reaction zone further downstream. Finally, the differences in inner nozzle flow and spray characteristics of petrodiesel and biodiesel are quantified. The improved modeling capability developed in this work can be used for extensive diesel engine simulations to further optimize injection, spray, combustion, and emission processes.

Naval Research Reviews. Volume 35, Number 3,

DTIC Science & Technology

1983-01-01

are under way that deal wave structure, to determine the shock oscillation properties . with turbulent mixing and combustion in airbreathing systems...article are perimental and theoretical means were used to determine the concerned with combustion instability in liquid fuel ramjet relative importance...together, and it imparts mechanical properties to the mixture. Additives are used to adjust the chemical, physical, and explosive properties of the

Thermo-kinetic instabilities in model reactors. Examples in experimental tests

NASA Astrophysics Data System (ADS)

Lavadera, Marco Lubrano; Sorrentino, Giancarlo; Sabia, Pino; de Joannon, Mara; Cavaliere, Antonio; Ragucci, Raffaele

2017-11-01

The use of advanced combustion technologies (such as MILD, LTC, etc.) is among the most promising methods to reduce emission of pollutants. For such technologies, working temperatures are enough low to boost the formation of several classes of pollutants, such as NOx and soot. To access this temperature range, a significant dilution as well as preheating of reactants is required. Such conditions are usually achieved by a strong recirculation of exhaust gases that simultaneously dilute and pre-heat the fresh reactants. These peculiar operative conditions also imply strong fuel flexibility, thus allowing the use of low calorific value (LCV) energy carriers with high efficiency. However, the intersection of low combustion temperatures and highly diluted mixtures with intense pre-heating alters the evolution of the combustion process with respect to traditional flames, leading to features such as the susceptibility to oscillations, which are undesirable during combustion. Therefore, an effective use of advanced combustion technologies requires a thorough analysis of the combustion kinetic characteristics in order to identify optimal operating conditions and control strategies with high efficiency and low pollutant emissions. The present work experimentally and numerically characterized the ignition and oxidation processes of methane and propane, highly diluted in nitrogen, at atmospheric pressure, in a Plug Flow Reactor and a Perfectly Stirred Reactor under a wide range of operating conditions involving temperatures, mixture compositions and dilution levels. The attention was focused particularly on the chemistry of oscillatory phenomena and multistage ignitions. The global behavior of these systems can be qualitatively and partially quantitatively modeled using the detailed kinetic models available in the literature. Results suggested that, for diluted conditions and lower adiabatic flame temperatures, the competition among several pathways, i.e. intermediate- and high-temperature branching, branching and recombination channels, oxidation and recombination/pyrolysis pathways, is enhanced, thus permitting the onset of phenomena that are generally hidden during conventional combustion processes.

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen / Liquid Methane Main Engine

NASA Technical Reports Server (NTRS)

Melcher, John C.; Morehead, Robert L.

2014-01-01

The project Morpheus liquid oxygen (LOX) / liquid methane (LCH4) main engine is a Johnson Space Center (JSC) designed 5,000 lbf-thrust, 4:1 throttling, pressure-fed cryogenic engine using an impinging element injector design. The engine met or exceeded all performance requirements without experiencing any in- ight failures, but the engine exhibited acoustic-coupled combustion instabilities during sea-level ground-based testing. First tangential (1T), rst radial (1R), 1T1R, and higher order modes were triggered by conditions during the Morpheus vehicle derived low chamber pressure startup sequence. The instability was never observed to initiate during mainstage, even at low power levels. Ground-interaction acoustics aggravated the instability in vehicle tests. Analysis of more than 200 hot re tests on the Morpheus vehicle and Stennis Space Center (SSC) test stand showed a relationship between ignition stability and injector/chamber pressure. The instability had the distinct characteristic of initiating at high relative injection pressure drop at low chamber pressure during the start sequence. Data analysis suggests that the two-phase density during engine start results in a high injection velocity, possibly triggering the instabilities predicted by the Hewitt stability curves. Engine ignition instability was successfully mitigated via a higher-chamber pressure start sequence (e.g., 50% power level vs 30%) and operational propellant start temperature limits that maintained \\cold LOX" and \\warm methane" at the engine inlet. The main engine successfully demonstrated 4:1 throttling without chugging during mainstage, but chug instabilities were observed during some engine shutdown sequences at low injector pressure drop, especially during vehicle landing.

Orbital maneuvering engine feed system coupled stability investigation

NASA Technical Reports Server (NTRS)

Kahn, D. R.; Schuman, M. D.; Hunting, J. K.; Fertig, K. W.

1975-01-01

A digital computer model used to analyze and predict engine feed system coupled instabilities over a frequency range of 10 to 1000 Hz was developed and verified. The analytical approach to modeling the feed system hydrodynamics, combustion dynamics, chamber dynamics, and overall engineering model structure is described and the governing equations in each of the technical areas are presented. This is followed by a description of the generalized computer model, including formulation of the discrete subprograms and their integration into an overall engineering model structure. The operation and capabilities of the engineering model were verified by comparing the model's theoretical predictions with experimental data from an OMS-type engine with a known feed system/engine chugging history.

Control of Combustion-Instabilities Through Various Passive Devices

NASA Technical Reports Server (NTRS)

Frendi, Kader

2005-01-01

It is well known that under some operating conditions, rocket engines (using solid or liquid fuels) exhibit unstable modes of operation that can lead to engine malfunction and shutdown. The sources of these instabilities are diverse and are dependent on fuel, chamber geometry and various upstream sources such as pumps, valves and injection mechanism. It is believed that combustion-acoustic instabilities occur when the acoustic energy increase due to the unsteady heat release of the flame is greater than the losses of acoustic energy from the system [1, 2]. Giammar and Putnam [3] performed a comprehensive study of noise generated by gasfired industrial burners and made several key observations; flow noise was sometimes more intense than combustion roar, which tended to have a characteristic frequency spectrum. Turbulence was amplified by the flame. The noise power varied directly with combustion intensity and also with the product of pressure drop and heat release rate. Karchmer [4] correlated the noise emitted from a turbofan jet engine with that in the combustion chamber. This is important, since it quantified how much of the noise from an engine originates in the combustor. A physical interpretation of the interchange of energy between sound waves and unsteady heat release rates was given by Rayleigh [5] for inviscid, linear perturbations. Bloxidge et al [6] extended Rayleigh s criterion to describe the interaction of unsteady combustion with one-dimensional acoustic waves in a duct. Solutions to the mass, momentum and energy conservation equations in the pre- and post-flame zones were matched by making several assumptions about the combustion process. They concluded that changes in boundary conditions affect the energy balance of acoustic waves in the combustor. Abouseif et al [7] also solved the one-dimensional flow equations, but they used a onestep reaction to evaluate the unsteady heat release rate by relating it to temperature and velocity perturbations. Their analysis showed that oscillations arise from coupling between entropy waves produced at the flame and pressure waves originating from the nozzle. Yang and Culick [8] assumed a thin flame sheet, which is distorted by velocity and pressure oscillations. Conservation equations were expressed in integral form and solutions for the acoustic wave equations and complex frequencies were obtained. The imaginary part of the frequency indicated stability regions of the flame. Activation energy asymptotics together with a one-step reaction were used by McIntosh [9] to study the effects of acoustic forcing and feedback on unsteady, one-dimensional flames. He found that the flame stability was altered by the upstream acoustic feedback. Shyy et al [10] used a high-accuracy TVD scheme to simulate unsteady, one-dimensional longitudinal, combustion instabilities. However, numerical diffusion was not completely eliminated. Recently, Prasad [11] investigated numerically the interactions of pressure perturbations with premixed flames. He used complex chemistry to study responses of pressure perturbations in one-dimensional combustors. His results indicated that reflected and transmitted waves differed significantly from incident waves.

Acoustic effects of sprays

NASA Technical Reports Server (NTRS)

Pindera, Maciej Z.; Przekwas, Andrzej J.

1994-01-01

Since the early 1960's, it has been known that realistic combustion models for liquid fuel rocket engines should contain at least a rudimentary treatment of atomization and spray physics. This is of particular importance in transient operations. It has long been recognized that spray characteristics and droplet vaporization physics play a fundamental role in determining the stability behavior of liquid fuel rocket motors. This paper gives an overview of work in progress on design of a numerical algorithm for practical studies of combustion instabilities in liquid rocket motors. For flexibility, the algorithm is composed of semi-independent solution modules, accounting for different physical processes. Current findings are report and future work is indicated. The main emphasis of this research is the development of an efficient treatment to interactions between acoustic fields and liquid fuel/oxidizer sprays.

Vortex/Flame Interactions in Microgravity Pulsed Jet Diffusion Flames

NASA Technical Reports Server (NTRS)

Bahadori, M. Y.; Hegde, U.; Stocker, D. P.

1999-01-01

The problem of vortex/flame interaction is of fundamental importance to turbulent combustion. These interactions have been studied in normal gravity. It was found that due to the interactions between the imposed disturbances and buoyancy induced instabilities, several overall length scales dominated the flame. The problem of multiple scales does not exist in microgravity for a pulsed laminar flame, since there are no buoyancy induced instabilities. The absence of buoyant convection therefore provides an environment to study the role of vortices interacting with flames in a controlled manner. There are strong similarities between imposed and naturally occurring perturbations, since both can be described by the same spatial instability theory. Hence, imposing a harmonic disturbance on a microgravity laminar flame creates effects similar to those occurring naturally in transitional/turbulent diffusion flames observed in microgravity. In this study, controlled, large-scale, axisymmetric vortices are imposed on a microgravity laminar diffusion flame. The experimental results and predictions from a numerical model of transient jet diffusion flames are presented and the characteristics of pulsed flame are described.

Measurements of admittances and characteristic combustion times of reactive gaseous propellant coaxial injectors

NASA Technical Reports Server (NTRS)

Janardan, B. A.; Daniel, B. R.; Zinn, B. T.

1979-01-01

The results of an experimental investigation that was concerned with the quantitative determination of the capabilities of combustion processes associated with coaxial injectors to amplify and sustain combustor oscillations was described. The driving provided by the combustion process was determined by employing the modified standing-wave method utilizing coaxial injectors and air-acetylene mixtures. Analyses of the measured data indicate that the investigated injectors are capable of initiating and amplifying combustion instabilities under favorable conditions of injector-combustion coupling and over certain frequency ranges. These frequency ranges and the frequency at which an injector's driving capacity is maximum are observed to depend upon the equivalence ratio, the pressure drop across the injector orifices and the number of injector elements. The characteristic combustion times of coaxial injectors were determined from steady state temperature measurements.

Characterization of flame stabilization technologies

NASA Astrophysics Data System (ADS)

Bush, Scott Matthew

To experimentally explore and characterize a V-gutter stabilized flame, this research study developed a Combustion Wind Tunnel Test Facility capable of effectively simulating the freestream Mach #'s and temperatures achieved within the back end of a gas turbine jet engine. After validating this facility, it was then used to gain a better understanding of the flow dynamics and combustion dynamics associated with the V-gutter configuration. The motivation for studying the V-gutter stabilized flame is due to the concern in industry today with combustion instabilities that are encountered in military aircraft. To gain a better understanding of the complex flow field associated with the V-gutter stabilized flame, this research study utilized Particle Image Velocimetry to capture both non-reacting and reacting instantaneous and mean flow structures formed in the wake region of the three dimensional V-gutter bluff body. The results of this study showed significant differences between the non-reacting and reacting flow fields. The non-reacting case resulted in asymmetric shedding of large scale vortices from the V-gutter edges while the reacting case resulted in a combination of both symmetric and asymmetric shedding of smaller scale vortical structures. A comparison of the mean velocity components shows that the reacting case results in a larger region of reversed flow, experiences an acceleration of the freestream flow due to combustion, and results in a slower dissipation of the wake region. Simultaneous dynamic pressure and CH* chemiluminescence measurements were also recorded to determine the coupling between the flow dynamics and combustion dynamics. The results of this study showed that only low frequency combustion instabilities were encountered at various conditions within the envelope of stable operation because of the interaction between longitudinal acoustic waves and unsteady heat release. When approaching rich blow out, rms pressure amplitudes were as high as 2 psi, and approaching lean blow out lead to rms pressure amplitudes around 0.2 psi. These studies also showed the instability frequency increasing with increases in either inlet temperature or inlet Mach #. Additionally, increasing the inlet velocity or the DeZubay parameter reduced the stability limits of operation for the V-gutter stabilized flame.

Effects of thermoacoustic oscillations on spray combustion dynamics with implications for lean direct injection systems

NASA Astrophysics Data System (ADS)

Chishty, Wajid Ali

Thermoacoustic instabilities in modern high-performance, low-emission gas turbine engines are often observable as large amplitude pressure oscillations and can result in serious performance and structural degradations. These acoustic oscillations can cause oscillations in combustor through-flows and given the right phase conditions, can also drive unsteady heat release. To curb the potential harms caused by the existence of thermoacoustic instabilities, recent efforts have focused on the active suppression of these instabilities. Intuitively, development of effective active combustion control methodologies is strongly dependent on the knowledge of the onset and sustenance of thermoacoustic instabilities. Specially, non-premixed spray combustion environment pose additional challenges due to the inherent unstable dynamics of sprays. The understanding of the manner in which the combustor acoustics affect the spray characteristics, which in turn result in heat release oscillation, is therefore, of paramount importance. The experimental investigations and the modeling studies conducted towards achieving this knowledge have been presented in this dissertation. Experimental efforts comprise both reacting and non-reacting flow studies. Reacting flow experiments were conducted on a overall lean direct injection, swirl-stabilized combustor rig. The investigations spanned combustor characterization and stability mapping over the operating regime. The onset of thermoacoustic instability and the transition of the combustor to two unstable regimes were investigated via phase-locked chemiluminescence imaging and measurement and phase-locked acoustic characterization. It was found that the onset of the thermoacoustic instability is a function of the energy gain of the system, while the sustenance of instability is due to the in-phase relationship between combustor acoustics and unsteady heat release driven by acoustic oscillations. The presence of non-linearities in the system between combustor acoustic and heat release and also between combustor acoustics and air through-flow were found to exist. The impact of high amplitude limit-cycle pressure on droplet breakdown under very low mean airflow and the localized effects of forced primary fuel modulations on heat release were also investigated. The non-reacting flow experiments were conducted to study the spray behavior under the presence of an acoustic field. An isothermal acoustic rig was specially fabricated, where the pressure oscillations were generated using an acoustic driver. Phase Doppler Anemometry was used to measure the droplet velocities and sizes under varying acoustic forcing conditions and spray feed pressures. Measurements made at different locations in the spray were related to these variations in mean and unsteady inputs. The droplet velocities were found to show a second order response to acoustic forcing with the cut-off frequency equal to the relaxation time corresponding to mean droplet size. It was also found that under acoustic forcing the droplets migrate radially away from the spray centerline and show oscillatory excursions in their movement. Modeling efforts were undertaken to gain physical insights of spray dynamics under the influence of acoustic forcing and to explain the experimental findings. The radial migration of droplets and their oscillatory movement were validated. The flame characteristics in the two unstable regimes and the transition between them were explained. It was found that under certain acoustic and mean air-flow condition, bands of high droplet densities were formed which resulted in diffusion type group burning of droplets. It was also shown that very high acoustic amplitudes cause secondary breakup of droplets.

Active Control of Combustion Instability in a Ramjet Using Large-Eddy Simulations

DTIC Science & Technology

1992-09-01

model is also used to determine the turbulent subgrid fluxes appearing in the momentum equations. Thus, the subgrid stresses in the momentum transport...flows and in flows with complex geometries. To include the effect of walls, an additional correction has been used to ensure that the subgrid stress ...subgrid stress Ty varies as y+3 near the wall. A major issue for LES of complex flows is whether the primary assumption that the subgrid scales are

Acoustic cavity technology for high performance injectors

NASA Technical Reports Server (NTRS)

1976-01-01

The feasibility of damping more than one mode of rocket engine combustion instability by means of differently tuned acoustic cavities sharing a common entrance was shown. Analytical procedures and acoustic modeling techniques for predicting the stability behavior of acoustic cavity designs in hot firings were developed. Full scale testing of various common entrance, dual cavity configurations, and subscale testing for the purpose of obtaining motion pictures of the cavity entrance region, to aid in determining the mechanism of cavity damping were the two major aspects of the program.

Some effects of swirl on turbulent mixing and combustion

NASA Technical Reports Server (NTRS)

Rubel, A.

1972-01-01

A general formulation of some effects of swirl on turbulent mixing is given. The basis for the analysis is that momentum transport is enhanced by turbulence resulting from rotational instability of the fluid field. An appropriate form for the turbulent eddy viscosity is obtained by mixing length type arguments. The result takes the form of a corrective factor that is a function of the swirl and acts to increase the eddy viscosity. The factor is based upon the initial mixing conditions implying that the rotational turbulence decays in a manner similar to that of free shear turbulence. Existing experimental data for free jet combustion are adequately matched by using the modifying factor to relate the effects of swirl on eddy viscosity. The model is extended and applied to the supersonic combustion of a ring jet of hydrogen injected into a constant area annular air stream. The computations demonstrate that swirling the flow could: (1) reduce the burning length by one half, (2) result in more uniform burning across the annulus width, and (3) open the possibility of optimization of the combustion characteristics by locating the fuel jet between the inner wall and center of the annulus width.

Initiation of Gaseous Detonation by Conical Projectiles

NASA Astrophysics Data System (ADS)

Verreault, Jimmy

Initiation and stabilization of detonation by hypersonic conical projectiles launched into combustible gas mixtures is investigated. This phenomenon must be understood for the design and optimization of specific hypersonic propulsion devices, such as the oblique detonation wave engine and the ram accelerator. The criteria for detonation initiation by a projectile is also related to fundamental aspects of detonation research, such as the requirement for direct initiation of a detonation by a blast wave. Experimental results of this problem also offer useful references for validation of numerical and theoretical modeling. Projectiles with cone half angles varying from 15° to 60° were launched into stoichiometric mixtures of hydrogen/oxygen with 70% argon dilution at initial pressures between 10 and 200 kPa. The projectiles were launched from a combustion-driven gas gun at velocities up to 2.2 km/s (corresponding to 133% of the Chapman Jouguet velocity). Pictures of the flowfields generated by the projectiles were taken via Schlieren photography. Five combustion regimes were observed about the projectile ranging from prompt and delayed oblique detonation wave formation, combustion instabilities, a wave splitting, and an inert shock wave. Two types of transition from the prompt oblique detonation wave regime to the inert shock regime were observed. The first (the delayed oblique detonation wave regime) showed an inert shock attached to the tip of the projectile followed by a sharp kink at the onset of an oblique detonation wave; this regime occurred by decreasing the cone angle at high mixture pressures. The second (the combustion instabilities regime) exhibited large density gradients due to combustion ignition and quenching phenomena; this regime occurred by decreasing the mixture pressure at large cone angles. A number of theoretical models were considered to predict critical conditions for the initiation of oblique detonations. The Lee-Vasiljev model agreed qualitatively well with the experimental results for relatively blunt projectiles (cone half-angle larger than 35°) and low mixture pressures (lower than 100 kPa). The trend of the critical Damköhler number calculated along the projectile cone surface was similar to that of the experimental results for slender cones (cone half-angles lower 35°) and high mixture pressures (higher than 100 kPa). Steady 2D simulations of reacting flows over finite wedges using the method of characteristics with a one-step Arrhenius chemical reaction model reproduced the three regimes observed for direct initiation of a detonation: the subcritical, critical and supercritical regimes. It is shown that in order for a 2D wedge to be equivalent to the problem of blast initiation of a detonation (which is the essence of the Lee-Vasiljev model), the Mach number normal to the oblique shock needs to be greater than 50 and the wedge angle has to be smaller than 30°. Simulations of reacting flows over semi-infinite wedges and cones were validated with CFD results. Excellent agreement was reached between the angle of overdriven oblique detonations obtained from the simulations and those from a polar analysis. For wedge or cone angles equal or lower than the minimum angle for which an oblique detonation is attached (according to the polar analysis), a Chapman-Jouguet oblique detonation was initiated. In the conical configuration, the curvature around the cone axis allowed an oblique detonation to be self-sustained at an angle less than without the curvature effect. At larger activation energies, the initiation process of an oblique detonation wave at the tip of a semi-infinite wedge or cone was identified. Unsteady 2D computational simulations were also conducted and showed the cellular structure of an oblique detonation wave. Instabilities in the form of transverse shock waves along the oblique detonation front arise for large activation energies.

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.

An ignition-temperature model with two free interfaces in premixed flames

NASA Astrophysics Data System (ADS)

Brauner, Claude-Michel; Gordon, Peter V.; Zhang, Wen

2016-11-01

In this paper we consider an ignition-temperature zero-order reaction model of thermo-diffusive combustion. This model describes the dynamics of thick flames, which have recently received considerable attention in the physical and engineering literature. The model admits a unique (up to translations) planar travelling wave solution. This travelling wave solution is quite different from those usually studied in combustion theory. The main qualitative feature of this travelling wave is that it has two interfaces: the ignition interface where the ignition temperature is attained and the trailing interface where the concentration of deficient reactants reaches zero. We give a new mathematical framework for studying the cellular instability of such travelling front solutions. Our approach allows the analysis of a free boundary problem to be converted into the analysis of a boundary value problem having a fully nonlinear system of parabolic equations. The latter is very suitable for both mathematical and numerical analysis. We prove the existence of a critical Lewis number such that the travelling wave solution is stable for values of Lewis number below the critical one and is unstable for Lewis numbers that exceed this critical value. Finally, we discuss the results of numerical simulations of a fully nonlinear system that describes the perturbation dynamics of planar fronts. These simulations reveal, in particular, some very interesting 'two-cell' steady patterns of curved combustion fronts.

The Response of Cryogenic H2/O2 Coaxial Jet Flames to Acoustic Disturbances (POST PRINT)

DTIC Science & Technology

2015-01-05

in the presence of tuned or induced acoustic resonance . Anderson and coworkers have been investigating longitudinal combustion ... resonance that involved periodic blocking of a secondary exhaust nozzle. The periodic flow through the second- ary nozzle results in chamber pressure...Angeles, 2009. 10. Yu, Y., Koeglmeier, S., Sisco, J., and Anderson, W., " Combustion instability of gaseous fuels in a continuously variable resonance

Numerical Modeling of Fuel Injection into an Accelerating, Turning Flow with a Cavity

NASA Astrophysics Data System (ADS)

Colcord, Ben James

Deliberate continuation of the combustion in the turbine passages of a gas turbine engine has the potential to increase the efficiency and the specific thrust or power of current gas-turbine engines. This concept, known as a turbine-burner, must overcome many challenges before becoming a viable product. One major challenge is the injection, mixing, ignition, and burning of fuel within a short residence time in a turbine passage characterized by large three-dimensional accelerations. One method of increasing the residence time is to inject the fuel into a cavity adjacent to the turbine passage, creating a low-speed zone for mixing and combustion. This situation is simulated numerically, with the turbine passage modeled as a turning, converging channel flow of high-temperature, vitiated air adjacent to a cavity. Both two- and three-dimensional, reacting and non-reacting calculations are performed, examining the effects of channel curvature and convergence, fuel and additional air injection configurations, and inlet conditions. Two-dimensional, non-reacting calculations show that higher aspect ratio cavities improve the fluid interaction between the channel flow and the cavity, and that the cavity dimensions are important for enhancing the mixing. Two-dimensional, reacting calculations show that converging channels improve the combustion efficiency. Channel curvature can be either beneficial or detrimental to combustion efficiency, depending on the location of the cavity and the fuel and air injection configuration. Three-dimensional, reacting calculations show that injecting fuel and air so as to disrupt the natural motion of the cavity stimulates three-dimensional instability and improves the combustion efficiency.

Effect of Electric Field in the Stabilized Premixed Flame on Combustion Process Emissions

NASA Astrophysics Data System (ADS)

Otto, Krickis

2017-10-01

The effect of the AC and DC electrical field on combustion processes has been investigated by various researchers. The results of these experiments do not always correlate, due to different experiment conditions and experiment equipment variations. The observed effects of the electrical field impact on the combustion process depends on the applied voltage polarity, flame speed and combustion physics. During the experiment was defined that starting from 1000 V the ionic wind takes the effect on emissions in flue gases, flame shape and combustion instabilities. Simulation combustion process in hermetically sealed chamber with excess oxygen amount 3 % in flue gases showed that the positive effect of electrical field on emissions lies in region from 30 to 400 V. In aforementioned voltage range carbon monoxide emissions were reduced by 6 % and at the same time the nitrogen oxide emissions were increased by 3.5 %.

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.

Effects of non-thermal plasmas and electric field on hydrocarbon/air flames

NASA Astrophysics Data System (ADS)

Ganguly, Biswa

2009-10-01

Need to improve fuel efficiency, and reduce emission from hydrocarbon combustor in automotive and gas turbine engines have reinvigorated interest in reducing combustion instability of a lean flame. The heat generation rate in a binary reaction is HQ =N^2 c1c2 Q exp(-E/RT), where N is the density, c1 and c2 are mol fractions of the reactants, Q is the reaction heat release, E is the activation energy, R is the gas constant and T is the average temperature. For hydrocarbon-air reactions, the typical value of E/R ˜20, so most heat release reactions are confined to a thin reaction sheet at T >=1400 K. The lean flame burning condition is susceptible to combustion instability due to a critical balance between heat generation and heat loss rates, especially at high gas flow rate. Radical injection can increase flame speed by reducing the hydrocarbon oxidation reaction activation barrier and it can improve flame stability. Advances in nonequilibrium plasma generation at high pressure have prompted its application for energy efficient radical production to enhance hydrocarbon-air combustion. Dielectric barrier discharges and short pulse excited corona discharges have been used to enhance combustion stability. Direct electron impact dissociation of hydrocarbon and O2 produces radicals with lower fuel oxidation reaction activation barriers, initiating heat release reaction CnHm+O <-> CnHm-1+ OH (and other similar sets of reactions with partially dissociated fuel) below the typical cross-over temperature. Also, N2 (A) produced in air discharge at a moderate E/n can dissociate O2 leading to oxidation of fuel at lower gas temperature. Low activation energy reactions are also possible by dissociation of hydrocarbon CnHm+e -> CnHm-2+H2+e, where a chain propagation reaction H2+ O<-> OH+H can be initiated at lower gas temperature than possible under thermal equilibrium kinetics. Most of heat release comes from the reaction CO+OH-> CO2 +H, nonthermal OH production seem to improve combustion stability The effect of applied voltage in a flame below self-sustained plasma generation is known to enhance flame holding through induced turbulence. Review of recent results will be presented to show future research opportunities in quantitative measurements and modeling of hydrocarbon/air plasma enhanced combustion.

Studies in premixed combustion. [Benjamin Levich Inst. for Physico-Chemical Hydrodynamics, City College of CUNY, New York, New York

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

Sivashinsky, G.I.

1993-01-01

During the period under review, significant progress was been made in studying the intrinsic dynamics of premixed flames and the problems of flame-flow interaction. (1) A weakly nonlinear model for Bunsen burner stabilized flames was proposed and employed for the simulation of three-dimensional polyhedral flames -- one of the most graphic manifestations of thermal-diffusive instability in premixed combustion. (2) A high-precision large-scale numerical simulation of Bunsen burner tip structure was conducted. The results obtained supported the earlier conjecture that the tip opening observed in low Lewis number systems is a purely optical effect not involving either flame extinction or leakagemore » of unburned fuel. (3) A one-dimensional model describing a reaction wave moving through a unidirectional periodic flow field is proposed and studied numerically. For long-wavelength fields the system exhibits a peculiar non-uniqueness of possible propagation regimes. The transition from one regime to another occurs in a manner of hysteresis.« less

Numerical results on noise-induced dynamics in the subthreshold regime for thermoacoustic systems

NASA Astrophysics Data System (ADS)

Gupta, Vikrant; Saurabh, Aditya; Paschereit, Christian Oliver; Kabiraj, Lipika

2017-03-01

Thermoacoustic instability is a serious issue in practical combustion systems. Such systems are inherently noisy, and hence the influence of noise on the dynamics of thermoacoustic instability is an aspect of practical importance. The present work is motivated by a recent report on the experimental observation of coherence resonance, or noise-induced coherence with a resonance-like dependence on the noise intensity as the system approaches the stability margin, for a prototypical premixed laminar flame combustor (Kabiraj et al., Phys. Rev. E, 4 (2015)). We numerically investigate representative thermoacoustic models for such noise-induced dynamics. Similar to the experiments, we study variation in system dynamics in response to variations in the noise intensity and in a critical control parameter as the systems approach their stability margins. The qualitative match identified between experimental results and observations in the representative models investigated here confirms that coherence resonance is a feature of thermoacoustic systems. We also extend the experimental results, which were limited to the case of subcritical Hopf bifurcation, to the case of supercritical Hopf bifurcation. We identify that the phenomenon has qualitative differences for the systems undergoing transition via subcritical and supercritical Hopf bifurcations. Two important practical implications are associated with the findings. Firstly, the increase in noise-induced coherence as the system approaches the onset of thermoacoustic instability can be considered as a precursor to the instability. Secondly, the dependence of noise-induced dynamics on the bifurcation type can be utilised to distinguish between subcritical and supercritical bifurcation prior to the onset of the instability.

Shock Driven Multiphase Instabilities in Scramjet Applications

NASA Astrophysics Data System (ADS)

McFarland, Jacob

2016-11-01

Shock driven multiphase instabilities (SDMI) arise in many applications from dust production in supernovae to ejecta distribution in explosions. At the limit of small, fast reacting particles the instability evolves similar to the Richtmyer-Meshkov (RM) instability. However, as additional particle effects such as lag, phase change, and collisions become significant the required parameter space becomes much larger and the instability deviates significantly from the RM instability. In scramjet engines the SDMI arises during a cold start where liquid fuel droplets are injected and processed by shock and expansion waves. In this case the particle evaporation and mixing is important to starting and sustaining combustion, but the particles are large and slow to react, creating significant multiphase effects. This talk will examine multiphase mixing in scramjet relevant conditions in 3D multiphase hydrodynamic simulations using the FLASH code from the University of Chicago FLASH center.

Natural oscillations of a gas in an elongated combustion chamber

NASA Astrophysics Data System (ADS)

Nesterov, S. V.; Akulenko, L. D.; Baydulov, V. G.

2017-02-01

For the analysis of the frequencies and shapes of the natural oscillations of a gas in an elongated rectilinear combustion chamber, this chamber can be treated as a kind of an organ pipe that has the following specific features: 1. the chamber has an inlet and outlet nozzles; 2. a gas mixture burns in the combustion chamber; 3. the combustion materials flow out from the outlet nozzle; 4. the gas flows in such a way that its velocity in the larger part (closer to the outlet nozzle) of the chamber exceeds the speed of sound (Mach number M > 1). There are only separate domains (one or several), where M < 1. The excitation of the natural oscillations of the gas and an increase in the amplitude of such oscillations can lead to instability of the combustion process [1].

Investigation of critical burning of fuel droplets. [monopropellants

NASA Technical Reports Server (NTRS)

Faeth, G. M.; Chanin, S.

1974-01-01

The steady combustion characteristics of droplets were considered in combustion chamber environments at various pressures, flow conditions, and ambient oxidizer concentrations for a number of hydrocarbon fuels. Using data obtained earlier, predicted gasification rates were within + or - 30% of measurements when the correction for convection was based upon average properties between the liquid surface and the flame around the droplet. Analysis was also completed for the open loop response of monopropellant droplets, based upon earlier strand combustion results. At the limit of large droplets, where the effect of flame curvature is small, the results suggest sufficient response to provide a viable mechanism for combustion instability in the frequency and droplet size range appropriate to practical combustors. Calculations are still in progress for a broader range of droplet sizes, including conditions where active combustion effects are small.

Aeroelasticity Analysis of AN Industrial Gas Turbine Combustor Using a Simplified Combustion Model

NASA Astrophysics Data System (ADS)

Bréard, C.; Sayma, A. I.; Vahdati, M.; Imregun, M.

2002-12-01

Lean premixed industrial gas turbine combustors are susceptible to flame instabilities, resulting in large unsteady pressure waves that may cause the discharge nozzle to experience excessive vibration levels. A detailed aeroelasticity analysis, aimed at investigating possible structural failure mechanisms, was undertaken using a time-accurate unsteady flow representation, a simplified combustion disturbance and a structural model of the discharge nozzle. The computational domain included the lower part of the combustor geometry as well as the nozzle guide vanes (NGVs) at the HP turbine inlet. A pressure perturbation, representing the unsteadiness due to the combustion process, was applied below the tertiary fuel inlet and its frequency was set to each structural natural frequency in turn. The propagation of the pressure perturbation through the combustor nozzle, its reflection from the NGVs and further reflections were monitored using two different models. The first one, the so-called ``open'' system, ignored the reflections from the upper part of the combustion chamber while the second one, the ``closed'' system, assumed full reflection with an appropriate time shift. The calculations have shown that the imposed excitation could generate unsteady pressure shapes that were correlated with the ``flap'' modes of the discharge nozzle. In addition, an acoustic resonance condition was observed when the forcing pressure wave had a frequency close to 550 Hz, the experimentally observed failure frequency of the nozzle. The co-existence of these two factors, i.e., excitation/structural-mode match and the possibility of acoustic resonance, was thought to have the potential of producing very high vibration response.

Computational Investigation of Combustion Instabilities in a Laboratory-Scale LDI Gas Turbine Engine

DTIC Science & Technology

2013-06-01

combustor by the insertion of a slotted inlet and an exit nozzle , whereas the reduced geometry is acoustically open. Table 2 Summary of Cases Considered... nozzle located at the right-end surface, an outlet condition is imposed by a characteristic back pressure condition. The fuel spray is injected at the...Computational Mesh visualized around the fuel nozzle and swirler III. Decomposition Methods For Combustion Dynamics Diagnostics To understand the

An adjoint-based sensitivity analysis of thermoacoustic network models

NASA Astrophysics Data System (ADS)

Sogaro, Francesca; Morgans, Aimee; Schmid, Peter

2017-11-01

Thermoacoustic instability is a phenomenon that occurs in numerous combustion systems, from rockets to land-based gas turbines. The acoustic oscillations of these systems are of significant importance as they can result in severe vibrations, thrust oscillations, thermal stresses and mechanical loads that lead to fatigue or even failure. In this work we use a low-order network model representation of a combustor system where linear acoustics are solved together with the appropriate boundary conditions, area change jump conditions, acoustic dampers and an appropriate flame transfer function. Special emphasis is directed towards the interaction between acoustically driven instabilities and flame-intrinsic modes. Adjoint methods are used to perform a sensitivity analysis of the spectral properties of the system to changes in the parameters involved. An exchange of modal identity between acoustic and intrinsic modes will be demonstrated and analyzed. The results provide insight into the interplay between various mode types and build a quantitative foundation for the design of combustors.

Analysis of turbojet combustion chamber performances based on flow field simplified mathematical model

NASA Astrophysics Data System (ADS)

Rotaru, Constantin

2017-06-01

In this paper are presented some results about the study of combustion chamber geometrical configurations that are found in aircraft gas turbine engines. The main focus of this paper consists in a study of a new configuration of the aircraft engine combustion chamber with an optimal distribution of gas velocity in front of the turbine. This constructive solution could allow a lower engine rotational speed, a lower temperature in front of the first stage of the turbine and the possibility to increase the turbine pressure ratio. The Arrhenius relationship, which describes the basic dependencies of the reaction rate on pressure, temperature and concentration has been used. and the CFD simulations were made with jet A fuel (which is presented in the Fluent software database) for an annular flame tube with 24 injectors. The temperature profile at the turbine inlet exhibits nonuniformity due to the number of fuel injectors used in the circumferential direction, the spatial nonuniformity in dilution air cooling and mixing characteristics as well as other secondary flow patterns and instabilities that are set up in the flame tube.

Performance and Stability Characteristics of a Uni-Element Swirl Injector for Oxygen-Rich Stage Combustion Cycles

NASA Technical Reports Server (NTRS)

Pal, S.; Kalitan, D.; Woodward, R. D.; Santoro, R. J.

2004-01-01

A uni-element liquid propellant combustion performance and instability study for liquid RP-1 and hot oxygen-rich pre-burner products was conducted, at a chamber pressure of about 1000 psi. using flush and recessed swirl injectors. High-frequency pressure transducer measurements were analyzed to yield the characteristic frequencies which were compared to expected frequencies of the chamber. Modes, which were discovered to be present within the main chamber included, the first longitudinal, detected at approximately 1950 Hz, and the second longitudinal mode at approximately 3800 Hz. An additional first longitudinal quarter wave mode was measured at a frequency of approximately 23000 Hz for the recessed swirl injector configuration. The characteristic instabilities resulting from these experiments were relatively weak averaging 0.2% to 0.3% of the chamber pressure.

Understanding Kelvin-Helmholtz instability in paraffin-based hybrid rocket fuels

NASA Astrophysics Data System (ADS)

Petrarolo, Anna; Kobald, Mario; Schlechtriem, Stefan

2018-04-01

Liquefying fuels show higher regression rates than the classical polymeric ones. They are able to form, along their burning surface, a low viscosity and surface tension liquid layer, which can become unstable (Kelvin-Helmholtz instability) due to the high velocity gas flow in the fuel port. This causes entrainment of liquid droplets from the fuel surface into the oxidizer gas flow. To better understand the droplets entrainment mechanism, optical investigations on the combustion behaviour of paraffin-based hybrid rocket fuels in combination with gaseous oxygen have been conducted in the framework of this research. Combustion tests were performed in a 2D single-slab burner at atmospheric conditions. High speed videos were recorded and analysed with two decomposition techniques. Proper orthogonal decomposition (POD) and independent component analysis (ICA) were applied to the scalar field of the flame luminosity. The most excited frequencies and wavelengths of the wave-like structures characterizing the liquid melt layer were computed. The fuel slab viscosity and the oxidizer mass flow were varied to study their influence on the liquid layer instability process. The combustion is dominated by periodic, wave-like structures for all the analysed fuels. Frequencies and wavelengths characterizing the liquid melt layer depend on the fuel viscosity and oxidizer mass flow. Moreover, for very low mass flows, no wavelength peaks are detected for the higher viscosity fuels. This is important to better understand and predict the onset and development of the entrainment process, which is connected to the amplification of the longitudinal waves.

Combustion Stability Characteristics of the Project Morpheus Liquid Oxygen/Liquid Methane Main Engine

NASA Technical Reports Server (NTRS)

Melcher, J. C.; Morehead, Robert L.

2014-01-01

The Project Morpheus liquid oxygen (LOX) / liquid methane rocket engines demonstrated acousticcoupled combustion instabilities during sea-level ground-based testing at the NASA Johnson Space Center (JSC) and Stennis Space Center (SSC). High-amplitude, 1T, 1R, 1T1R (and higher order) modes appear to be triggered by injector conditions. The instability occurred during the Morpheus-specific engine ignition/start sequence, and did demonstrate the capability to propagate into mainstage. However, the instability was never observed to initiate during mainstage, even at low power levels. The Morpheus main engine is a JSC-designed 5,000 lbf-thrust, 4:1 throttling, pressure-fed cryogenic engine using an impinging element injector design. Two different engine designs, named HD4 and HD5, and two different builds of the HD4 engine all demonstrated similar instability characteristics. Through the analysis of more than 200 hot fire tests on the Morpheus vehicle and SSC test stand, a relationship between ignition stability and injector/chamber pressure was developed. The instability has the distinct characteristic of initiating at high relative injection pressure drop (dP) at low chamber pressure (Pc); i.e., instabilities initiated at high dP/Pc at low Pc during the start sequence. The high dP/Pc during start results during the injector /chamber chill-in, and is enhanced by hydraulic flip in the injector orifice elements. Because of the fixed mixture ratio of the existing engine design (the main valves share a common actuator), it is not currently possible to determine if LOX or methane injector dP/Pc were individual contributors (i.e., LOX and methane dP/Pc typically trend in the same direction within a given test). The instability demonstrated initiation characteristic of starting at or shortly after methane injector chillin. Colder methane (e.g., sub-cooled) at the injector inlet prior to engine start was much more likely to result in an instability. A secondary effect of LOX sub-cooling was also possibly observed; greater LOX sub- cooling improved stability. Some tests demonstrated a low-amplitude 1L-1T instability prior to LOX injector chill-in. The Morpheus main engine also demonstrated chug instabilities during some engine shutdown sequences on the flight vehicle and SSC test stand. The chug instability was also infrequently observed during the startup sequence. The chug instabilities predictably initiated at low dP/Pc at low Pc. The chug instabilities were always self-limiting; startup chug instabilities terminated during throttle-up and shutdown chug instabilities decayed by shutdown termination.

Modeling Turbulent Mixing/Combustion of Bio-Agents Behind Detonations: Effect of Instabilities, Dense Clustering, and Trace Survivability

DTIC Science & Technology

2017-06-01

observed in earlier numerical and experimental investigations. Similar quasi -detonation mode is observed for the case of channel with d/2 = 4 cm. The... control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 1. REPORT DATE (DD-MM-YYYY) 2. REPORT TYPE 3. DATES COVERED (From - To...joule (J) erg 1 × 10 –7 joule (J) kiloton (kt) (TNT equivalent ) 4.184 × 10 12 joule (J) British thermal unit (Btu) (thermochemical) 1.054 350 × 10

Cold-Flow Study of Low Frequency Pressure Instability in Hybrid Rocket Motors

NASA Technical Reports Server (NTRS)

Jenkins, Rhonald M.

1997-01-01

Past experience with hybrid rockets has shown that certain motor operating conditions are conducive to the formation of low frequency pressure oscillations, or flow instabilities, within the motor. Both past and present work in the hybrid propulsion community acknowledges deficiencies in the understanding of such behavior, though it seems probable that the answer lies in an interaction between the flow dynamics and the combustion heat release. Knowledge of the fundamental flow dynamics is essential to the basic understanding of the overall stability problem. A first step in this direction was a study conducted at NASA Marshall Space Flight Center (MSFC), centered around a laboratory-scale two dimensional water flow model of a hybrid rocket motor. Principal objectives included: (1) visualization of flow and measurement of flow velocity distributions: (2) assessment of the importance of shear layer instabilities in driving motor pressure oscillations; (3) determination of the interactions between flow induced shear layers with the mainstream flow, the secondary (wall) throughflow, and solid boundaries; (4) investigation of the interactions between wall flow oscillations and the mainstream flow pressure distribution.

SRB combustion dynamics analysis computer program (CDA-1)

NASA Technical Reports Server (NTRS)

Chung, T. J.; Park, O. Y.

1988-01-01

A two-dimensional numerical model is developed for the unsteady oscillatory combustion of the solid propellant flame zone. Variations of pressure with low and high frequency responses across the long flame, such as in the double-base propellants, are accommodated. The formulation is based on a premixed, laminar flame with a one-step overall chemical reaction and the Arrhenius law of decomposition for the gaseous phase with no condensed phase reaction. Numerical calculations are carried out using the Galerkin finite elements, with perturbations expanded to the zeroth, first, and second orders. The numerical results indicate that amplification of oscillatory motions does indeed prevail in high frequency regions. For the second order system, the trend is similar to the first order system for low frequencies, but instabilities may appear at frequencies lower than those of the first order system. The most significant effect of the second order system is that the admittance is extremely oscillatory between moderately high frequency ranges.

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

NASA Astrophysics Data System (ADS)

Scarborough, David E.

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

Microwave-Assisted Ignition for Improved Internal Combustion Engine Efficiency

NASA Astrophysics Data System (ADS)

DeFilippo, Anthony Cesar

The ever-present need for reducing greenhouse gas emissions associated with transportation motivates this investigation of a novel ignition technology for internal combustion engine applications. Advanced engines can achieve higher efficiencies and reduced emissions by operating in regimes with diluted fuel-air mixtures and higher compression ratios, but the range of stable engine operation is constrained by combustion initiation and flame propagation when dilution levels are high. An advanced ignition technology that reliably extends the operating range of internal combustion engines will aid practical implementation of the next generation of high-efficiency engines. This dissertation contributes to next-generation ignition technology advancement by experimentally analyzing a prototype technology as well as developing a numerical model for the chemical processes governing microwave-assisted ignition. The microwave-assisted spark plug under development by Imagineering, Inc. of Japan has previously been shown to expand the stable operating range of gasoline-fueled engines through plasma-assisted combustion, but the factors limiting its operation were not well characterized. The present experimental study has two main goals. The first goal is to investigate the capability of the microwave-assisted spark plug towards expanding the stable operating range of wet-ethanol-fueled engines. The stability range is investigated by examining the coefficient of variation of indicated mean effective pressure as a metric for instability, and indicated specific ethanol consumption as a metric for efficiency. The second goal is to examine the factors affecting the extent to which microwaves enhance ignition processes. The factors impacting microwave enhancement of ignition processes are individually examined, using flame development behavior as a key metric in determining microwave effectiveness. Further development of practical combustion applications implementing microwave-assisted spark technology will benefit from predictive models which include the plasma processes governing the observed combustion enhancement. This dissertation documents the development of a chemical kinetic mechanism for the plasma-assisted combustion processes relevant to microwave-assisted spark ignition. The mechanism includes an existing mechanism for gas-phase methane oxidation, supplemented with electron impact reactions, cation and anion chemical reactions, and reactions involving vibrationally-excited and electronically-excited species. Calculations using the presently-developed numerical model explain experimentally-observed trends, highlighting the relative importance of pressure, temperature, and mixture composition in determining the effectiveness of microwave-assisted ignition enhancement.

System Detects Vibrational Instabilities

NASA Technical Reports Server (NTRS)

Bozeman, Richard J., Jr.

1990-01-01

Sustained vibrations at two critical frequencies trigger diagnostic response or shutdown. Vibration-analyzing electronic system detects instabilities of combustion in rocket engine. Controls pulse-mode firing of engine and identifies vibrations above threshold amplitude at 5.9 and/or 12kHz. Adapted to other detection and/or control schemes involving simultaneous real-time detection of signals above or below preset amplitudes at two or more specified frequencies. Potential applications include rotating machinery and encoders and decoders in security systems.

Mechanisms Underpinning Degradation of Protective Oxides and Thermal Barrier Coatings in High Hydrogen Content (HHC) - Fueled Turbines

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

Mumm, Daniel

2013-08-31

The overarching goal of this research program has been to evaluate the potential impacts of coal-derived syngas and high-hydrogen content fuels on the degradation of turbine hot-section components through attack of protective oxides and thermal barrier coatings. The primary focus of this research program has been to explore mechanisms underpinning the observed degradation processes, and connections to the combustion environments and characteristic non-combustible constituents. Based on the mechanistic understanding of how these emerging fuel streams affect materials degradation, the ultimate goal of the program is to advance the goals of the Advanced Turbine Program by developing materials design protocols leadingmore » to turbine hot-section components with improved resistance to service lifetime degradation under advanced fuels exposures. This research program has been focused on studying how: (1) differing combustion environments – relative to traditional natural gas fired systems – affect both the growth rate of thermally grown oxide (TGO) layers and the stability of these oxides and of protective thermal barrier coatings (TBCs); and (2) how low levels of fuel impurities and characteristic non-combustibles interact with surface oxides, for instance through the development of molten deposits that lead to hot corrosion of protective TBC coatings. The overall program has been comprised of six inter-related themes, each comprising a research thrust over the program period, including: (i) evaluating the role of syngas and high hydrogen content (HHC) combustion environments in modifying component surface temperatures, heat transfer to the TBC coatings, and thermal gradients within these coatings; (ii) understanding the instability of TBC coatings in the syngas and high hydrogen environment with regards to decomposition, phase changes and sintering; (iii) characterizing ash deposition, molten phase development and infiltration, and associated corrosive/thermo-chemical attack mechanisms; (iv) developing a mechanics-based analysis of the driving forces for crack growth and delamination, based on molten phase infiltration, misfit upon cooling, and loss of compliance; (v) understanding changes in TGO growth mechanisms associated with these emerging combustion product streams; and (vi) identifying degradation resistant alternative materials (including new compositions or bi-layer concepts) for use in mitigating the observed degradation modes. To address the materials stability concerns, this program integrated research thrusts aimed at: (1) Conducting tests in simulated syngas and HHC environments to evaluate materials evolution and degradation mechanisms; assessing thermally grown oxide development unique to HHC environmental exposures; carrying out high-resolution imaging and microanalysis to elucidate the evolution of surface deposits (molten phase formation and infiltration); exploring thermo-chemical instabilities; assessing thermo-mechanical drivers and thermal gradient effects on degradation; and quantitatively measuring stress evolution due to enhanced sintering and thermo-chemical instabilities induced in the coating. (2) Executing experiments to study the melting and infiltration of simulated ash deposits, and identifying reaction products and evolving phases associated with molten phase corrosion mechanisms; utilizing thermal spray techniques to fabricate test coupons with controlled microstructures to study mechanisms of instability and degradation; facilitating thermal gradient testing; and developing new materials systems for laboratory testing; (3) Correlating information on the resulting combustion environments to properly assess materials exposure conditions and guide the development of lab-scale simulations of material exposures; specification of representative syngas and high-hydrogen fuels with realistic levels of impurities and contaminants, to explore differences in heat transfer, surface degradation, and deposit formation; and facilitating combustion rig testing of materials test coupons.« less

Simultaneous identification of transfer functions and combustion noise of a turbulent flame

NASA Astrophysics Data System (ADS)

Merk, M.; Jaensch, S.; Silva, C.; Polifke, W.

2018-05-01

The Large Eddy Simulation/System Identification (LES/SI) approach allows to deduce a flame transfer function (FTF) from LES of turbulent reacting flow: Time series of fluctuations of reference velocity and global heat release rate resulting from broad-band excitation of a simulated turbulent flame are post-processed via SI techniques to derive a low order model of the flame dynamics, from which the FTF is readily deduced. The current work investigates an extension of the established LES/SI approach: In addition to estimation of the FTF, a low order model for the combustion noise source is deduced from the same time series data. By incorporating such a noise model into a linear thermoacoustic model, it is possible to predict the overall level as well as the spectral distribution of sound pressure in confined combustion systems that do not exhibit self-excited thermoacoustic instability. A variety of model structures for estimation of a noise model are tested in the present study. The suitability and quality of these model structures are compared against each other, their sensitivity regarding certain time series properties is studied. The influence of time series length, signal-to-noise ratio as well as acoustic reflection coefficient of the boundary conditions on the identification are examined. It is shown that the Box-Jenkins model structure is superior to simpler approaches for the simultaneous identification of models that describe the FTF as well as the combustion noise source. Subsequent to the question of the most adequate model structure, the choice of optimal model order is addressed, as in particular the optimal parametrization of the noise model is not obvious. Akaike's Information Criterion and a model residual analysis are applied to draw qualitative and quantitative conclusions on the most suitable model order. All investigations are based on a surrogate data model, which allows a Monte Carlo study across a large parameter space with modest computationally effort. The conducted study constitutes a solid basis for the application of advanced SI techniques to actual LES data.

Modeling and simulation of combustion dynamics in lean-premixed swirl-stabilized gas-turbine engines

NASA Astrophysics Data System (ADS)

Huang, Ying

This research focuses on the modeling and simulation of combustion dynamics in lean-premixed gas-turbines engines. The primary objectives are: (1) to establish an efficient and accurate numerical framework for the treatment of unsteady flame dynamics; and (2) to investigate the parameters and mechanisms responsible for driving flow oscillations in a lean-premixed gas-turbine combustor. The energy transfer mechanisms among mean flow motions, periodic motions and background turbulent motions in turbulent reacting flow are first explored using a triple decomposition technique. Then a comprehensive numerical study of the combustion dynamics in a lean-premixed swirl-stabilized combustor is performed. The analysis treats the conservation equations in three dimensions and takes into account finite-rate chemical reactions and variable thermophysical properties. Turbulence closure is achieved using a large-eddy-simulation (LES) technique. The compressible-flow version of the Smagorinsky model is employed to describe subgrid-scale turbulent motions and their effect on large-scale structures. A level-set flamelet library approach is used to simulate premixed turbulent combustion. In this approach, the mean flame location is modeled using a level-set G-equation, where G is defined as a distance function. Thermophysical properties are obtained using a presumed probability density function (PDF) along with a laminar flamelet library. The governing equations and the associated boundary conditions are solved by means of a four-step Runge-Kutta scheme along with the implementation of the message passing interface (MPI) parallel computing architecture. The analysis allows for a detailed investigation into the interaction between turbulent flow motions and oscillatory combustion of a swirl-stabilized injector. Results show good agreement with an analytical solution and experimental data in terms of acoustic properties and flame evolution. A study of flame bifurcation from a stable state to an unstable state indicates that the inlet flow temperature and equivalence ratio are the two most important variables determining the stability characteristics of the combustor. Under unstable operating conditions, several physical processes responsible for driving combustion instabilities in the chamber have been identified and quantified. These processes include vortex shedding and acoustic interaction, coupling between the flame evolution and local flow oscillations, vortex and flame interaction and coupling between heat release and acoustic motions. The effects of inlet swirl number on the flow development and flame dynamics in the chamber are also carefully studied. In the last part of this thesis, an analytical model is developed using triple decomposition techniques to model the combustion response of turbulent premixed flames to acoustic oscillations.

Laminar Dust Flames: A Program of Microgravity and Ground Based Studies at McGill

NASA Technical Reports Server (NTRS)

Goroshin, Sam; Lee, John

1999-01-01

Fundamental knowledge of heterogeneous combustion mechanisms is required to improve utilization of solid fuels (e.g. coal), safe handling of combustible dusts in industry, and solid propulsion systems. The objective of the McGill University research program on dust combustion is to obtain a reliable set of data on basic combustion parameters for dust suspensions (i.e. laminar burning velocity, flame structure, quenching distance, flammability limits, etc.) over a range of particle sizes, dust concentrations, and types of fuel. This set of data then permits theoretical models to be validated and, when necessary, new models to be developed to describe the detailed reaction mechanisms and transport processes. Microgravity is essential to the generation of a uniform dust suspension of arbitrary particle size and concentration. When particles with a characteristic size on the order of tens of microns are suspended, they rapidly settle in a gravitational field. To maintain a particulate in suspension for time duration adequate to carry out combustion experiments invariably requires continuous convective flow in excess of the gravitational settling velocity (which is comparable with and can even exceed the dust laminar burning velocity). This makes the experiments turbulent in nature and thus renders it impossible to study laminar dust flames. Even for small particle sizes on the order of microns, a stable laminar dust flow can be maintained only for relatively low dust concentrations at normal gravity conditions. High dust loading leads to gravitational instability of the dust cloud and to the formation of recirculation cells in the dust suspension in a confined volume, or to the rapid sedimentation of the dense dust cloud, as a whole, in an unconfined volume. Many important solid fuels such as carbon and boron also have low laminar flame speeds (of the order of several centimeters per second). Convection that occurs in combustion products due to buoyancy disrupts the low speed dust flames and makes observation of such flames at normal gravity difficult.

A Study of Premixed, Shock-Induced Combustion With Application to Hypervelocity Flight

NASA Technical Reports Server (NTRS)

Axdahl, Erik L.

2013-01-01

One of the current goals of research in hypersonic, airbreathing propulsion is access to higher Mach numbers. A strong driver of this goal is the desire to integrate a scramjet engine into a transatmospheric vehicle airframe in order to improve performance to low Earth orbit (LEO) or the performance of a semiglobal transport. An engine concept designed to access hypervelocity speeds in excess of Mach 10 is the shock-induced combustion ramjet (i.e. shcramjet). This dissertation presents numerical studies simulating the physics of a shcramjet vehicle traveling at hypervelocity speeds with the goal of understanding the physics of fuel injection, wall autoignition mitigation, and combustion instability in this flow regime.

Annual Research Briefs, 1998

NASA Technical Reports Server (NTRS)

Spinks, Debra (Compiler)

1998-01-01

The topics contained in this progress report are direct numerical simulation of turbulent non-premixed combustion with realistic chemistry; LES of non-premixed turbulent reacting flows with conditional source term estimation; measurements of the three-dimensional scalar dissipation rate in gas-phase planar turbulent jets; direct simulation of a jet diffusion flame; on the use of interpolating wavelets in the direct numerical simulation of combustion; on the use of a dynamically adaptive wavelet collocation algorithm in DNS (direct numerical simulation) of non-premixed turbulent combustion; 2D simulations of Hall thrusters; computation of trailing-edge noise at low mach number using LES and acoustic analogy; weakly nonlinear modeling of the early stages of bypass transition; interactions between freestream turbulence and boundary layers; interfaces at the outer boundaries of turbulent motions; largest scales of turbulent wall flows; the instability of streaks in near-wall turbulence; an implementation of the v(sup 2) - f model with application to transonic flows; heat transfer predictions in cavities; a structure-based model with stropholysis effects; modeling a confined swirling coaxial jet; subgrid-scale models based on incremental unknowns for large eddy simulations; subgrid scale modeling taking the numerical error into consideration; towards a near-wall model for LES of a separated diffuser flow; on the feasibility of merging LES with RANS (Reynolds Averaging Numerical simulation) for the near-wall region of attached turbulent flows; large-eddy simulation of a separated boundary layer; numerical study of a channel flow with variable properties; on the construction of high order finite difference schemes on non-uniform meshes with good conservation properties; development of immersed boundary methods for complex geometries; and particle methods for micro and macroscale flow simulations.

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.

Time varying voltage combustion control and diagnostics sensor

DOEpatents

Chorpening, Benjamin T [Morgantown, WV; Thornton, Jimmy D [Morgantown, WV; Huckaby, E David [Morgantown, WV; Fincham, William [Fairmont, WV

2011-04-19

A time-varying voltage is applied to an electrode, or a pair of electrodes, of a sensor installed in a fuel nozzle disposed adjacent the combustion zone of a continuous combustion system, such as of the gas turbine engine type. The time-varying voltage induces a time-varying current in the flame which is measured and used to determine flame capacitance using AC electrical circuit analysis. Flame capacitance is used to accurately determine the position of the flame from the sensor and the fuel/air ratio. The fuel and/or air flow rate (s) is/are then adjusted to provide reduced flame instability problems such as flashback, combustion dynamics and lean blowout, as well as reduced emissions. The time-varying voltage may be an alternating voltage and the time-varying current may be an alternating current.

Novel Active Combustion Control Valve

NASA Technical Reports Server (NTRS)

Caspermeyer, Matt

2014-01-01

This project presents an innovative solution for active combustion control. Relative to the state of the art, this concept provides frequency modulation (greater than 1,000 Hz) in combination with high-amplitude modulation (in excess of 30 percent flow) and can be adapted to a large range of fuel injector sizes. Existing valves often have low flow modulation strength. To achieve higher flow modulation requires excessively large valves or too much electrical power to be practical. This active combustion control valve (ACCV) has high-frequency and -amplitude modulation, consumes low electrical power, is closely coupled with the fuel injector for modulation strength, and is practical in size and weight. By mitigating combustion instabilities at higher frequencies than have been previously achieved (approximately 1,000 Hz), this new technology enables gas turbines to run at operating points that produce lower emissions and higher performance.

Adjoint-based sensitivity analysis of low-order thermoacoustic networks using a wave-based approach

NASA Astrophysics Data System (ADS)

Aguilar, José G.; Magri, Luca; Juniper, Matthew P.

2017-07-01

Strict pollutant emission regulations are pushing gas turbine manufacturers to develop devices that operate in lean conditions, with the downside that combustion instabilities are more likely to occur. Methods to predict and control unstable modes inside combustion chambers have been developed in the last decades but, in some cases, they are computationally expensive. Sensitivity analysis aided by adjoint methods provides valuable sensitivity information at a low computational cost. This paper introduces adjoint methods and their application in wave-based low order network models, which are used as industrial tools, to predict and control thermoacoustic oscillations. Two thermoacoustic models of interest are analyzed. First, in the zero Mach number limit, a nonlinear eigenvalue problem is derived, and continuous and discrete adjoint methods are used to obtain the sensitivities of the system to small modifications. Sensitivities to base-state modification and feedback devices are presented. Second, a more general case with non-zero Mach number, a moving flame front and choked outlet, is presented. The influence of the entropy waves on the computed sensitivities is shown.

Puffing flame instability - Part II: Predicting the onset and frequency

NASA Astrophysics Data System (ADS)

Boettcher, Philipp; Shepherd, Joseph; Menon, Shyam; Blanquart, Guillaume

2011-11-01

Experiments and simulations have been performed on fuel rich n- hexane air mixtures in a closed vessel. Both experiments and simulations show a distinct cyclic combustion or ``puffing'' mode. The misalignment of buoyancy induced pressure gradients and density gradients across the flame front is responsible for the generation of vorticity and its subsequent roll-up into vortex rings. In the present work, a simplified model is proposed based on the fundamental interactions between fluid mechanical and chemical parameters. This simplified fluid mechanics model is based on dimensional analysis and is used to predict the onset and frequency of the puffing behavior. This work was sponsored by The Boeing Company through CTBA-GTA-1.

The cosmic web and microwave background fossilize the first turbulent combustion

NASA Astrophysics Data System (ADS)

Gibson, Carl H.

2015-09-01

The weblike structure of the cosmic microwave background CMB temperature fluctuations are interpreted as fossils of the first turbulent combustion that drives the big bang1,2,3. Modern turbulence theory3 requires that inertial vortex forces cause turbulence to always cascade from small scales to large, contrary to the standard turbulence model where the cascade is reversed. Assuming that the universe begins at Planck length 10-35 m and temperature 1032 K, the mechanism of the big bang is a powerful turbulent combustion instability, where turbulence forms at the Kolmogorov scale and mass-energy is extracted by < -10113 Pa negative stresses from big bang turbulence working against gravity. Prograde accretion of a Planck antiparticle on a spinning particle-antiparticle pair releases 42% of a particle rest mass from the Kerr metric, producing a spinning gas of turbulent Planck particles that cascades to larger scales at smaller temperatures (10-27 m, 1027 K) retaining the Planck density 1097 kg m-3, where quarks form and gluon viscosity fossilizes the turbulence. Viscous stress powers inflation to ~ 10 m and ~ 10100 kg. The CMB shows signatures of both plasma and big bang turbulence. Direct numerical simulations support the new turbulence theory6.

Validation of an Adaptive Combustion Instability Control Method for Gas-Turbine Engines

NASA Technical Reports Server (NTRS)

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

2004-01-01

This paper describes ongoing testing of an adaptive control method to suppress high frequency thermo-acoustic instabilities like those found in lean-burning, low emission combustors that are being developed for future aircraft gas turbine engines. The method called Adaptive Sliding Phasor Averaged Control, was previously tested in an experimental rig designed to simulate a combustor with an instability of about 530 Hz. Results published earlier, and briefly presented here, demonstrated that this method was effective in suppressing the instability. Because this test rig did not exhibit a well pronounced instability, a question remained regarding the effectiveness of the control methodology when applied to a more coherent instability. To answer this question, a modified combustor rig was assembled at the NASA Glenn Research Center in Cleveland, Ohio. The modified rig exhibited a more coherent, higher amplitude instability, but at a lower frequency of about 315 Hz. Test results show that this control method successfully reduced the instability pressure of the lower frequency test rig. In addition, due to a certain phenomena discovered and reported earlier, the so called Intra-Harmonic Coupling, a dramatic suppression of the instability was achieved by focusing control on the second harmonic of the instability. These results and their implications are discussed, as well as a hypothesis describing the mechanism of intra-harmonic coupling.

Initiation and structures of gaseous detonation

NASA Astrophysics Data System (ADS)

Vasil'ev, A. A.; Vasiliev, V. A.

2018-03-01

The analysis of the initiation of a detonation wave (DW) and the emergence of a multi-front structure of the DW-front are presented. It is shown that the structure of the DW arises spontaneously at the stage of a strong overdriven of the wave. The hypothesis of the gradual enhancement of small perturbations on an initially smooth initiating blast wave, traditionally used in the numerical simulation of multi-front detonation, does not agree with the experimental data. The instability of the DW is due to the chemical energy release of the combustible mixture Q. A technique for determining the Q-value of mixture was proposed, based on reconstruction of the trajectory of the expanding wave from the position of the strong explosion model. The wave trajectory at the critical initiation of a multifront detonation in a combustible mixture is compared with the trajectory of an explosive wave from the same initiator in an inert mixture whose gas-dynamic parameters are equivalent to the parameters of the combustible mixture. The energy release of a mixture is defined as the difference in the joint energy release of the initiator and the fuel mixture during the critical initiation and energy release of the initiator when the blast wave is excited in an inert mixture. Observable deviations of the experimental profile of Q from existing model representations were found.

Solid Rocket Motor Combustion Instability Modeling in COMSOL Multiphysics

NASA Technical Reports Server (NTRS)

Fischbach, S. R.

2015-01-01

Combustion instability modeling of Solid Rocket Motors (SRM) remains a topic of active research. Many rockets display violent fluctuations in pressure, velocity, and temperature originating from the complex interactions between the combustion process, acoustics, and steady-state gas dynamics. Recent advances in defining the energy transport of disturbances within steady flow-fields have been applied by combustion stability modelers to improve the analysis framework. Employing this more accurate global energy balance requires a higher fidelity model of the SRM flow-field and acoustic mode shapes. The current industry standard analysis tool utilizes a one dimensional analysis of the time dependent fluid dynamics along with a quasi-three dimensional propellant grain regression model to determine the SRM ballistics. The code then couples with another application that calculates the eigenvalues of the one dimensional homogenous wave equation. The mean flow parameters and acoustic normal modes are coupled to evaluate the stability theory developed and popularized by Culick. The assumption of a linear, non-dissipative wave in a quiescent fluid remains valid while acoustic amplitudes are small and local gas velocities stay below Mach 0.2. The current study employs the COMSOL Multiphysics finite element framework to model the steady flow-field parameters and acoustic normal modes of a generic SRM. This work builds upon previous efforts to verify the use of the acoustic velocity potential equation (AVPE) laid out by Campos. The acoustic velocity potential (psi) describing the acoustic wave motion in the presence of an inhomogeneous steady high-speed flow is defined by, del squared psi - (lambda/c) squared psi - M x [M x del((del)(psi))] - 2((lambda)(M)/c + M x del(M) x (del)(psi) - 2(lambda)(psi)[M x del(1/c)] = 0. with M as the Mach vector, c as the speed of sound, and ? as the complex eigenvalue. The study requires one way coupling of the CFD High Mach Number Flow (HMNF) and mathematics module. The HMNF module evaluates the gas flow inside of a SRM using St. Robert's law to model the solid propellant burn rate, slip boundary conditions, and the supersonic outflow condition. Results from the HMNF model are verified by comparing the pertinent ballistics parameters with the industry standard code outputs (i.e. pressure drop, axial velocity, exit velocity). These results are then used by the coefficient form of the mathematics module to determine the complex eigenvalues of the AVPE. The mathematics model is truncated at the nozzle sonic line, where a zero flux boundary condition is self-satisfying. The remaining boundaries are modeled with a zero flux boundary condition, assuming zero acoustic absorption on all surfaces. The one way coupled analysis is perform four times utilizing geometries determined through traditional SRM modeling procedures. The results of the steady-state CFD and AVPE analyses are used to calculate the linear acoustic growth rate as is defined by Flandro and Jacob. In order to verify the process implemented within COMSOL we first employ the Culick theory and compare the results with the industry standard. After the process is verified, the Flandro/Jacob energy balance theory is employed and results displayed.

Combustion Stability of the Gas Generator Assembly from J-2X Engine E10001 and Powerpack Tests

NASA Technical Reports Server (NTRS)

Hulka, J. R.; Kenny, R. L.; Casiano, M. J.

2013-01-01

Testing of a powerpack configuration (turbomachinery and gas generator assembly) and the first complete engine system of the liquid oxygen/liquid hydrogen propellant J-2X rocket engine have been completed at the NASA Stennis Space Center. The combustion stability characteristics of the gas generator assemblies on these two systems are of interest for reporting since considerable effort was expended to eliminate combustion instability during early development of the gas generator assembly with workhorse hardware. Comparing the final workhorse gas generator assembly development test data to the powerpack and engine system test data provides an opportunity to investigate how the nearly identical configurations of gas generator assemblies operate with two very different propellant supply systems one the autonomous pressure-fed test configuration on the workhorse development test stand, the other the pump-fed configurations on the powerpack and engine systems. The development of the gas generator assembly and the elimination of the combustion instability on the pressure-fed workhorse test stand have been reported extensively in the two previous Liquid Propulsion Subcommittee meetings 1-7. The powerpack and engine system testing have been conducted from mid-2011 through 2012. All tests of the powerpack and engine system gas generator systems to date have been stable. However, measureable dynamic behavior, similar to that observed on the pressure-fed test stand and reported in Ref. [6] and attributed to an injection-coupled response, has appeared in both powerpack and engine system tests. As discussed in Ref. [6], these injection-coupled responses are influenced by the interaction of the combustion chamber with a branch pipe in the hot gas duct that supplies gaseous helium to pre-spin the turbine during the start transient. This paper presents the powerpack and engine system gas generator test data, compares these data to the development test data, and provides additional combustion stability analyses of the configurations.

A Study of the Impact of Variations on Aerodynamic Flow in Gas Turbine Engines via Monte-Carlo Simulations

NASA Technical Reports Server (NTRS)

Ngo, Khiem Viet; Tumer, Irem Y.

2003-01-01

The unsteady compressible inviscid flow is characterized by the conservations of mass, momentum, and energy; or simply the Euler equations. In this paper, a study of the subsonic one-dimensional Euler equations with local preconditioning is presented with a modal analysis approach. Specifically, this study investigates the behavior of airflow in a gas turbine engine using the specified conditions at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine, under the impact of variations in pressure, velocity, temperature, and density at low Mach numbers. Two main questions that motivate this research are: 1) Is there any aerodynamic problem with the existing gas turbine engines that could impact aircraft performance? 2) If yes, what aspect of a gas turbine engine could be improved via design to alleviate that impact and to optimize aircraft performance. This paper presents an initial attempt to the flow behavior in terms (perturbation) using simulation outputs from a customer-deck model obtained from Pratt&Whitney, (i.e., pressure, temperature, velocity, density) about their mean states at the inflow and outflow boundaries of the compressor, combustion chamber, and turbine. Flow behavior is analyzed for the high pressure compressor and combustion chamber employing the conditions on their left and right boundaries. In the same fashion, similar analyses are carried out for the high and low-pressure turbines. In each case, the eigenfrequencies that are obtained for different boundary conditions are examined closely based on their probabilistic distributions, a result of a Monte Carlo 10,000-sample simulation. Furthermore, the characteristic waves and eave response are analyzed and contrasted among different cases, with and without preconditioners. The results reveal the existence of flow instabilities due to the combined effect of variations and excessive pressures; which are clearly the case in the combustion chamber and high-pressure turbine. Finally a discussion is presented on potential impacts of the instabilities and what can be improved via design to alleviate them for a better aircraft performance.

Experimental analysis of thermo-acoustic instabilities in a generic gas turbine combustor by phase-correlated PIV, chemiluminescence, and laser Raman scattering measurements

NASA Astrophysics Data System (ADS)

Arndt, Christoph M.; Severin, Michael; Dem, Claudiu; Stöhr, Michael; Steinberg, Adam M.; Meier, Wolfgang

2015-04-01

A gas turbine model combustor for partially premixed swirl flames was equipped with an optical combustion chamber and operated with CH4 and air at atmospheric pressure. The burner consisted of two concentric nozzles for separately controlled air flows and a ring of holes 12 mm upstream of the nozzle exits for fuel injection. The flame described here had a thermal power of 25 kW, a global equivalence ratio of 0.7, and exhibited thermo-acoustic instabilities at a frequency of approximately 400 Hz. The phase-dependent variations in the flame shape and relative heat release rate were determined by OH* chemiluminescence imaging; the flow velocities by stereoscopic particle image velocimetry (PIV); and the major species concentrations, mixture fraction, and temperature by laser Raman scattering. The PIV measurements showed that the flow field performed a "pumping" mode with varying inflow velocities and extent of the inner recirculation zone, triggered by the pressure variations in the combustion chamber. The flow field oscillations were accompanied by variations in the mixture fraction in the inflow region and at the flame root, which in turn were mainly caused by the variations in the CH4 concentration. The mean phase-dependent changes in the fluxes of CH4 and N2 through cross-sectional planes of the combustion chamber at different heights above the nozzle were estimated by combining the PIV and Raman data. The results revealed a periodic variation in the CH4 flux by more than 150 % in relation to the mean value, due to the combined influence of the oscillating flow velocity, density variations, and CH4 concentration. Based on the experimental results, the feedback mechanism of the thermo-acoustic pulsations could be identified as a periodic fluctuation of the equivalence ratio and fuel mass flow together with a convective delay for the transport of fuel from the fuel injector to the flame zone. The combustor and the measured data are well suited for the validation of numerical combustion simulations.

Transient flow combustion

NASA Technical Reports Server (NTRS)

Tacina, R. R.

1984-01-01

Non-steady combustion problems can result from engine sources such as accelerations, decelerations, nozzle adjustments, augmentor ignition, and air perturbations into and out of the compressor. Also non-steady combustion can be generated internally from combustion instability or self-induced oscillations. A premixed-prevaporized combustor would be particularly sensitive to flow transients because of its susceptability to flashback-autoignition and blowout. An experimental program, the Transient Flow Combustion Study is in progress to study the effects of air and fuel flow transients on a premixed-prevaporized combustor. Preliminary tests performed at an inlet air temperature of 600 K, a reference velocity of 30 m/s, and a pressure of 700 kPa. The airflow was reduced to 1/3 of its original value in a 40 ms ramp before flashback occurred. Ramping the airflow up has shown that blowout is more sensitive than flashback to flow transients. Blowout occurred with a 25 percent increase in airflow (at a constant fuel-air ratio) in a 20 ms ramp. Combustion resonance was found at some conditions and may be important in determining the effects of flow transients.

Status on the Verification of Combustion Stability for the J-2X Engine Thrust Chamber Assembly

NASA Technical Reports Server (NTRS)

Casiano, Matthew; Hinerman, Tim; Kenny, R. Jeremy; Hulka, Jim; Barnett, Greg; Dodd, Fred; Martin, Tom

2013-01-01

Development is underway of the J -2X engine, a liquid oxygen/liquid hydrogen rocket engine for use on the Space Launch System. The Engine E10001 began hot fire testing in June 2011 and testing will continue with subsequent engines. The J -2X engine main combustion chamber contains both acoustic cavities and baffles. These stability aids are intended to dampen the acoustics in the main combustion chamber. Verification of the engine thrust chamber stability is determined primarily by examining experimental data using a dynamic stability rating technique; however, additional requirements were included to guard against any spontaneous instability or rough combustion. Startup and shutdown chug oscillations are also characterized for this engine. This paper details the stability requirements and verification including low and high frequency dynamics, a discussion on sensor selection and sensor port dynamics, and the process developed to assess combustion stability. A status on the stability results is also provided and discussed.

Periodic equivalence ratio modulation method and apparatus for controlling combustion instability

DOEpatents

Richards, George A.; Janus, Michael C.; Griffith, Richard A.

2000-01-01

The periodic equivalence ratio modulation (PERM) method and apparatus significantly reduces and/or eliminates unstable conditions within a combustion chamber. The method involves modulating the equivalence ratio for the combustion device, such that the combustion device periodically operates outside of an identified unstable oscillation region. The equivalence ratio is modulated between preselected reference points, according to the shape of the oscillation region and operating parameters of the system. Preferably, the equivalence ratio is modulated from a first stable condition to a second stable condition, and, alternatively, the equivalence ratio is modulated from a stable condition to an unstable condition. The method is further applicable to multi-nozzle combustor designs, whereby individual nozzles are alternately modulated from stable to unstable conditions. Periodic equivalence ratio modulation (PERM) is accomplished by active control involving periodic, low frequency fuel modulation, whereby low frequency fuel pulses are injected into the main fuel delivery. Importantly, the fuel pulses are injected at a rate so as not to affect the desired time-average equivalence ratio for the combustion device.

Quantifying Instability Sources in Liquid Rocket Engines

NASA Technical Reports Server (NTRS)

Farmer, Richard C.; Cheng, Gary C.

2000-01-01

Computational fluid dynamics methodology to predict the effects of combusting flows on acoustic pressure oscillations in liquid rocket engines (LREs) is under development. 'Me intent of the investigation is to develop the causal physics of combustion driven acoustic resonances in LREs. The crux of the analysis is the accurate simulation of pressure/density/sound speed in a combustor which when used by the FDNS-RFV CFD code will produce realistic flow phenomena. An analysis of a gas generator considered for the Fastrac engine will be used as a test validation case.

Tomographic reconstruction of heat release rate perturbations induced by helical modes in turbulent swirl flames

NASA Astrophysics Data System (ADS)

Moeck, Jonas P.; Bourgouin, Jean-François; Durox, Daniel; Schuller, Thierry; Candel, Sébastien

2013-04-01

Swirl flows with vortex breakdown are widely used in industrial combustion systems for flame stabilization. This type of flow is known to sustain a hydrodynamic instability with a rotating helical structure, one common manifestation of it being the precessing vortex core. The role of this unsteady flow mode in combustion is not well understood, and its interaction with combustion instabilities and flame stabilization remains unclear. It is therefore important to assess the structure of the perturbation in the flame that is induced by this helical mode. Based on principles of tomographic reconstruction, a method is presented to determine the 3-D distribution of the heat release rate perturbation associated with the helical mode. Since this flow instability is rotating, a phase-resolved sequence of projection images of light emitted from the flame is identical to the Radon transform of the light intensity distribution in the combustor volume and thus can be used for tomographic reconstruction. This is achieved with one stationary camera only, a vast reduction in experimental and hardware requirements compared to a multi-camera setup or camera repositioning, which is typically required for tomographic reconstruction. Different approaches to extract the coherent part of the oscillation from the images are discussed. Two novel tomographic reconstruction algorithms specifically tailored to the structure of the heat release rate perturbations related to the helical mode are derived. The reconstruction techniques are first applied to an artificial field to illustrate the accuracy. High-speed imaging data acquired in a turbulent swirl-stabilized combustor setup with strong helical mode oscillations are then used to reconstruct the 3-D structure of the associated perturbation in the flame.

Reacting flow studies in a dump combustor: Enhanced volumetric heat release rates and flame anchorability

NASA Astrophysics Data System (ADS)

Behrens, Alison Anne

Reacting flow studies in a novel dump combustor facility focused on increasing volumetric heat release rates, under stable burning conditions, and understanding the physical mechanisms governing flame anchoring in an effort to extend range and maneuverability of compact, low drag, air-breathing engines. Countercurrent shear flow was enhanced within the combustor as the primary control variable. Experiments were performed burning premixed JP10/air and methane/air in a dump combustor using reacting flow particle image velocimetry (PIV) and chemiluminescence as the primary diagnostics. Stable combustion studies burning lean mixtures of JP10/air aimed to increase volumetric heat release rates through the implementation of countercurrent shear control. Countercurrent shear flow was produced by creating a suction flow from a low pressure cavity connected to the dump combustor via a gap directly below the trailing edge. Chemiluminescence measurements showed that enhancing countercurrent shear within the combustor doubles volumetric heat release rates. PIV measurements indicate that counterflow acts to increase turbulent kinetic energy while maintaining constant strain rates. This acts to increase flame surface area through flame wrinkling without disrupting the integrity of the flame. Flame anchorability is one of the most important fundamental aspects to understand when trying to enhance turbulent combustion in a high-speed engine without increasing drag. Studies burning methane/air mixtures used reacting flow PIV to study flame anchoring. The operating point with the most stable flame anchor exhibited a correspondingly strong enthalpy flux of products into reactants via a single coherent structure positioned downstream of the step. However, the feature producing a strong flame anchor, i.e. a single coherent structure, also is responsible for combustion instabilities, therefore making this operating point undesirable. Counterflow control was found to create the best flow features for stable, robust, compact combustion. Enhancing countercurrent shear flow within a dump combustor enhances burning rates, provides a consistent pump of reaction-initiating combustion products required for sustained combustion, while maintaining flow three dimensionality needed to disrupt combustion instabilities. Future studies will focus on geometric and control scenarios that further reduce drag penalties while creating these same flow features found with countercurrent shear thus producing robust operating points.

Orbital Maneuvering Engine Feed System Coupled Stability Investigation, Computer User's Manual

NASA Technical Reports Server (NTRS)

Schuman, M. D.; Fertig, K. W.; Hunting, J. K.; Kahn, D. R.

1975-01-01

An operating manual for the feed system coupled stability model was given, in partial fulfillment of a program designed to develop, verify, and document a digital computer model that can be used to analyze and predict engine/feed system coupled instabilities in pressure-fed storable propellant propulsion systems over a frequency range of 10 to 1,000 Hz. The first section describes the analytical approach to modelling the feed system hydrodynamics, combustion dynamics, chamber dynamics, and overall engineering model structure, and presents the governing equations in each of the technical areas. This is followed by the program user's guide, which is a complete description of the structure and operation of the computerized model. Last, appendices provide an alphabetized FORTRAN symbol table, detailed program logic diagrams, computer code listings, and sample case input and output data listings.

A positivity-preserving, implicit defect-correction multigrid method for turbulent combustion

NASA Astrophysics Data System (ADS)

Wasserman, M.; Mor-Yossef, Y.; Greenberg, J. B.

2016-07-01

A novel, robust multigrid method for the simulation of turbulent and chemically reacting flows is developed. A survey of previous attempts at implementing multigrid for the problems at hand indicated extensive use of artificial stabilization to overcome numerical instability arising from non-linearity of turbulence and chemistry model source-terms, small-scale physics of combustion, and loss of positivity. These issues are addressed in the current work. The highly stiff Reynolds-averaged Navier-Stokes (RANS) equations, coupled with turbulence and finite-rate chemical kinetics models, are integrated in time using the unconditionally positive-convergent (UPC) implicit method. The scheme is successfully extended in this work for use with chemical kinetics models, in a fully-coupled multigrid (FC-MG) framework. To tackle the degraded performance of multigrid methods for chemically reacting flows, two major modifications are introduced with respect to the basic, Full Approximation Storage (FAS) approach. First, a novel prolongation operator that is based on logarithmic variables is proposed to prevent loss of positivity due to coarse-grid corrections. Together with the extended UPC implicit scheme, the positivity-preserving prolongation operator guarantees unconditional positivity of turbulence quantities and species mass fractions throughout the multigrid cycle. Second, to improve the coarse-grid-correction obtained in localized regions of high chemical activity, a modified defect correction procedure is devised, and successfully applied for the first time to simulate turbulent, combusting flows. The proposed modifications to the standard multigrid algorithm create a well-rounded and robust numerical method that provides accelerated convergence, while unconditionally preserving the positivity of model equation variables. Numerical simulations of various flows involving premixed combustion demonstrate that the proposed MG method increases the efficiency by a factor of up to eight times with respect to an equivalent single-grid method, and by two times with respect to an artificially-stabilized MG method.

Analysis of Thermo-Diffusive Cellular Instabilities in Continuum Combustion Fronts

NASA Astrophysics Data System (ADS)

Azizi, Hossein; Gurevich, Sebastian; Provatas, Nikolas; Department of Physics, Centre Physics of Materials Team

We explore numerically the morphological patterns of thermo-diffusive instabilities in combustion fronts with a continuum solid fuel source, within a range of Lewis numbers, focusing on the cellular regime. Cellular and dendritic instabilities are found at low Lewis numbers. These are studied using a dynamic adaptive mesh refinement technique that allows very large computational domains, thus allowing us to reduce finite size effects that can affect or even preclude the emergence of these patterns. The distinct types of dynamics found in the vicinity of the critical Lewis number. These types of dynamics are classified as ``quasi-linear'' and characterized by low amplitude cells that may be strongly affected by the mode selection mechanism and growth prescribed by the linear theory. Below this range of Lewis number, highly non-linear effects become prominent and large amplitude, complex cellular and seaweed dendritic morphologies emerge. The cellular patterns simulated in this work are similar to those observed in experiments of flame propagation over a bed of nano-aluminum powder burning with a counter-flowing oxidizer conducted by Malchi et al. It is noteworthy that the physical dimension of our computational domain is roughly close to their experimental setup. This work was supported by a Canadian Space Agency Class Grant ''Percolating Reactive Waves in Particulate Suspensions''. We thank Compute Canada for computing resources.

Feedback control of acoustic disturbance transient growth in triggering thermoacoustic instability

NASA Astrophysics Data System (ADS)

Zhao, Dan; Reyhanoglu, Mahmut

2014-08-01

Transient growth of acoustic disturbances could trigger thermoacoustic instability in a combustion system with non-orthogonal eigenmodes, even with stable eigenvalues. In this work, feedback control of transient growth of flow perturbations in a Rijke-type combustion system is considered. For this, a generalized thermoacoustic model with distributed monopole-like actuators is developed. The model is formulated in state-space to gain insights on the interaction between various eigenmodes and the dynamic response of the system to the actuators. Three critical parameters are identified: (1) the mode number, (2) the number of actuators, and (3) the locations of the actuators. It is shown that in general the number of the actuators K is related to the mode number N as K=N2. For simplicity in illustrating the main results of the paper, two different thermoacoustic systems are considered: system (a) with one mode and system (b) that involves two modes. The actuator location effect is studied in system (a) and it is found that the actuator location plays an important role in determining the control effort. In addition, sensitivity analysis of pressure- and velocity-related control parameters is conducted. In system (b), when the actuators are turned off (i.e., open-loop configuration), it is observed that acoustic energy transfers from the high frequency mode to the lower frequency mode. After some time, the energy is transferred back. Moreover, the high frequency oscillation grows into nonlinear limit cycle with the low frequency oscillation amplified. As a linear-quadratic regulator (LQR) is implemented to tune the actuators, both systems become asymptotically stable. However, the LQR controller fails in eliminating the transient growth, which may potentially trigger thermoacoustic instability. In order to achieve strict dissipativity (i.e., unity maximum transient growth), a transient growth controller is systematically designed and tested in both systems. Comparison is then made between the performance of the LQR controller and that of the transient growth controller. It is found in both systems that the transient growth controller achieves both exponential decay of the flow disturbance energy and unity maximum transient growth.

Assessment of numerical methods for the solution of fluid dynamics equations for nonlinear resonance systems

NASA Technical Reports Server (NTRS)

Przekwas, A. J.; Yang, H. Q.

1989-01-01

The capability of accurate nonlinear flow analysis of resonance systems is essential in many problems, including combustion instability. Classical numerical schemes are either too diffusive or too dispersive especially for transient problems. In the last few years, significant progress has been made in the numerical methods for flows with shocks. The objective was to assess advanced shock capturing schemes on transient flows. Several numerical schemes were tested including TVD, MUSCL, ENO, FCT, and Riemann Solver Godunov type schemes. A systematic assessment was performed on scalar transport, Burgers' and gas dynamic problems. Several shock capturing schemes are compared on fast transient resonant pipe flow problems. A system of 1-D nonlinear hyperbolic gas dynamics equations is solved to predict propagation of finite amplitude waves, the wave steepening, formation, propagation, and reflection of shocks for several hundred wave cycles. It is shown that high accuracy schemes can be used for direct, exact nonlinear analysis of combustion instability problems, preserving high harmonic energy content for long periods of time.

Acoustic Characterization of Compact Jet Engine Simulator Units

NASA Technical Reports Server (NTRS)

Doty, Michael J.; Haskin, Henry H.

2013-01-01

Two dual-stream, heated jet, Compact Jet Engine Simulator (CJES) units are designed for wind tunnel acoustic experiments involving a Hybrid Wing Body (HWB) vehicle. The newly fabricated CJES units are characterized with a series of acoustic and flowfield investigations to ensure successful operation with minimal rig noise. To limit simulator size, consistent with a 5.8% HWB model, the CJES units adapt Ultra Compact Combustor (UCC) technology developed at the Air Force Research Laboratory. Stable and controllable operation of the combustor is demonstrated using passive swirl air injection and backpressuring of the combustion chamber. Combustion instability tones are eliminated using nonuniform flow conditioners in conjunction with upstream screens. Through proper flow conditioning, rig noise is reduced by more than 20 dB over a broad spectral range, but it is not completely eliminated at high frequencies. The low-noise chevron nozzle concept designed for the HWB test shows expected acoustic benefits when installed on the CJES unit, and consistency between CJES units is shown to be within 0.5 dB OASPL.

Numerical techniques for solving nonlinear instability problems in smokeless tactical solid rocket motors. [finite difference technique

NASA Technical Reports Server (NTRS)

Baum, J. D.; Levine, J. N.

1980-01-01

The selection of a satisfactory numerical method for calculating the propagation of steep fronted shock life waveforms in a solid rocket motor combustion chamber is discussed. A number of different numerical schemes were evaluated by comparing the results obtained for three problems: the shock tube problems; the linear wave equation, and nonlinear wave propagation in a closed tube. The most promising method--a combination of the Lax-Wendroff, Hybrid and Artificial Compression techniques, was incorporated into an existing nonlinear instability program. The capability of the modified program to treat steep fronted wave instabilities in low smoke tactical motors was verified by solving a number of motor test cases with disturbance amplitudes as high as 80% of the mean pressure.

CFD Analysis of the 24-inch JIRAD Hybrid Rocket Motor

NASA Technical Reports Server (NTRS)

Liang, Pak-Yan; Ungewitter, Ronald; Claflin, Scott

1996-01-01

A series of multispecies, multiphase computational fluid dynamics (CFD) analyses of the 24-inch diameter joint government industry industrial research and development (JIRAD) hybrid rocket motor is described. The 24-inch JIRAD hybrid motor operates by injection of liquid oxygen (LOX) into a vaporization plenum chamber upstream of ports in the hydroxyl-terminated polybutadiene (HTPB) solid fuel. The injector spray pattern had a strong influence on combustion stability of the JIRAD motor so a CFD study was initiated to define the injector end flow field under different oxidizer spray patterns and operating conditions. By using CFD to gain a clear picture of the flow field and temperature distribution within the JIRAD motor, it is hoped that the fundamental mechanisms of hybrid combustion instability may be identified and then suppressed by simple alterations to the oxidizer injection parameters such as injection angle and velocity. The simulations in this study were carried out using the General Algorithm for Analysis of Combustion SYstems (GALACSY) multiphase combustion codes. GALACSY consists of a comprehensive set of droplet dynamic submodels (atomization, evaporation, etc.) and a computationally efficient hydrocarbon chemistry package built around a robust Navier-Stokes solver optimized for low Mach number flows. Lagrangian tracking of dispersed particles describes a closely coupled spray phase. The CFD cases described in this paper represent various levels of simplification of the problem. They include: (A) gaseous oxygen with combusting fuel vapor blowing off the walls at various oxidizer injection angles and velocities, (B) gaseous oxygen with combusting fuel vapor blowing off the walls, and (C) liquid oxygen with combusting fuel vapor blowing off the walls. The study used an axisymmetric model and the results indicate that the injector design significantly effects the flow field in the injector end of the motor. Markedly different recirculation patterns are observed in the vaporization chamber as the oxygen velocity and/or spray pattern is varied. The ability of these recirculation patterns to stabilize the diffusion flame above the surface of the solid fuel gives a plausible explanation for the experimentally determined combustion stability characteristics of the JIRAD motor, and suggests how combustion stability can be assured by modifications to the injector design.

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, this scheme would be implemented by the adaptive sliding phaser averaged control (ASPAC) algorithm, which requires very little detailed knowledge of the combustor dynamics. In the ASPAC algorithm, the power of the instability signal would be calculated from the wide-bandpass- filtered combustion-pressure signal and averaged over a period of time (typically of the order of a few hundredths of a second) corresponding to the controller updating cycle [not to be confused with the controller sampling cycle, which would be much shorter (typically of the order of 10(exp -4) second)].

On the Experimental and Theoretical Investigations of Lean Partially Premixed Combustion, Burning Speed, Flame Instability and Plasma Formation of Alternative Fuels at High Temperatures and Pressures

NASA Astrophysics Data System (ADS)

Askari, Omid

This dissertation investigates the combustion and injection fundamental characteristics of different alternative fuels both experimentally and theoretically. The subjects such as lean partially premixed combustion of methane/hydrogen/air/diluent, methane high pressure direct-injection, thermal plasma formation, thermodynamic properties of hydrocarbon/air mixtures at high temperatures, laminar flames and flame morphology of synthetic gas (syngas) and Gas-to-Liquid (GTL) fuels were extensively studied in this work. These subjects will be summarized in three following paragraphs. The fundamentals of spray and partially premixed combustion characteristics of directly injected methane in a constant volume combustion chamber have been experimentally studied. The injected fuel jet generates turbulence in the vessel and forms a turbulent heterogeneous fuel-air mixture in the vessel, similar to that in a Compressed Natural Gas (CNG) Direct-Injection (DI) engines. The effect of different characteristics parameters such as spark delay time, stratification ratio, turbulence intensity, fuel injection pressure, chamber pressure, chamber temperature, Exhaust Gas recirculation (EGR) addition, hydrogen addition and equivalence ratio on flame propagation and emission concentrations were analyzed. As a part of this work and for the purpose of control and calibration of high pressure injector, spray development and characteristics including spray tip penetration, spray cone angle and overall equivalence ratio were evaluated under a wide range of fuel injection pressures of 30 to 90 atm and different chamber pressures of 1 to 5 atm. Thermodynamic properties of hydrocarbon/air plasma mixtures at ultra-high temperatures must be precisely calculated due to important influence on the flame kernel formation and propagation in combusting flows and spark discharge applications. A new algorithm based on the statistical thermodynamics was developed to calculate the ultra-high temperature plasma composition and thermodynamic properties. The method was applied to compute the thermodynamic properties of hydrogen/air and methane/air plasma mixtures for a wide range of temperatures (1,000-100,000 K), pressures (10-6-100 atm) and different equivalence ratios within flammability limit. In calculating the individual thermodynamic properties of the atomic species, the Debye-Huckel cutoff criterion has been used for terminating the series expression of the electronic partition function. A new differential-based multi-shell model was developed in conjunction with Schlieren photography to measure laminar burning speed and to study the flame instabilities for different alternative fuels such as syngas and GTL. Flame instabilities such as cracking and wrinkling were observed during flame propagation and discussed in terms of the hydrodynamic and thermo-diffusive effects. Laminar burning speeds were measured using pressure rise data during flame propagation and power law correlations were developed over a wide range of temperatures, pressures and equivalence ratios. As a part of this work, the effect of EGR addition and substitution of nitrogen with helium in air on flame morphology and laminar burning speed were extensively investigated. The effect of cell formation on flame surface area of syngas fuel in terms of a newly defined parameter called cellularity factor was also evaluated. In addition to that the experimental onset of auto-ignition and theoretical ignition delay times of premixed GTL/air mixture were determined at high pressures and low temperatures over a wide range of equivalence ratios.

Flexible Inhibitor Fluid-Structure Interaction Simulation in RSRM.

NASA Astrophysics Data System (ADS)

Wasistho, Bono

2005-11-01

We employ our tightly coupled fluid/structure/combustion simulation code 'Rocstar-3' for solid propellant rocket motors to study 3D flows past rigid and flexible inhibitors in the Reusable Solid Rocket Motor (RSRM). We perform high resolution simulations of a section of the rocket near the center joint slot at 100 seconds after ignition, using inflow conditions based on less detailed 3D simulations of the full RSRM. Our simulations include both inviscid and turbulent flows (using LES dynamic subgrid-scale model), and explore the interaction between the inhibitor and the resulting fluid flow. The response of the solid components is computed by an implicit finite element solver. The internal mesh motion scheme in our block-structured fluid solver enables our code to handle significant changes in geometry. We compute turbulent statistics and determine the compound instabilities originated from the natural hydrodynamic instabilities and the inhibitor motion. The ultimate goal is to studdy the effect of inhibitor flexing on the turbulent field.

Simulation of Interaction of Strong Shocks with Gas Bubbles using the Direct Simulation Monte Carlo Method

NASA Astrophysics Data System (ADS)

Puranik, Bhalchandra; Watvisave, Deepak; Bhandarkar, Upendra

2016-11-01

The interaction of a shock with a density interface is observed in several technological applications such as supersonic combustion, inertial confinement fusion, and shock-induced fragmentation of kidney and gall-stones. The central physical process in this interaction is the mechanism of the Richtmyer-Meshkov Instability (RMI). The specific situation where the density interface is initially an isolated spherical or cylindrical gas bubble presents a relatively simple geometry that exhibits all the essential RMI processes such as reflected and refracted shocks, secondary instabilities, turbulence and mixing of the species. If the incident shocks are strong, the calorically imperfect nature needs to be modelled. In the present work, we have carried out simulations of the shock-bubble interaction using the DSMC method for such situations. Specifically, an investigation of the shock-bubble interaction with diatomic gases involving rotational and vibrational excitations at high temperatures is performed, and the effects of such high temperature phenomena will be presented.

Random sphere packing model of heterogeneous propellants

NASA Astrophysics Data System (ADS)

Kochevets, Sergei Victorovich

It is well recognized that combustion of heterogeneous propellants is strongly dependent on the propellant morphology. Recent developments in computing systems make it possible to start three-dimensional modeling of heterogeneous propellant combustion. A key component of such large scale computations is a realistic model of industrial propellants which retains the true morphology---a goal never achieved before. The research presented develops the Random Sphere Packing Model of heterogeneous propellants and generates numerical samples of actual industrial propellants. This is done by developing a sphere packing algorithm which randomly packs a large number of spheres with a polydisperse size distribution within a rectangular domain. First, the packing code is developed, optimized for performance, and parallelized using the OpenMP shared memory architecture. Second, the morphology and packing fraction of two simple cases of unimodal and bimodal packs are investigated computationally and analytically. It is shown that both the Loose Random Packing and Dense Random Packing limits are not well defined and the growth rate of the spheres is identified as the key parameter controlling the efficiency of the packing. For a properly chosen growth rate, computational results are found to be in excellent agreement with experimental data. Third, two strategies are developed to define numerical samples of polydisperse heterogeneous propellants: the Deterministic Strategy and the Random Selection Strategy. Using these strategies, numerical samples of industrial propellants are generated. The packing fraction is investigated and it is shown that the experimental values of the packing fraction can be achieved computationally. It is strongly believed that this Random Sphere Packing Model of propellants is a major step forward in the realistic computational modeling of heterogeneous propellant of combustion. In addition, a method of analysis of the morphology of heterogeneous propellants is developed which uses the concept of multi-point correlation functions. A set of intrinsic length scales of local density fluctuations in random heterogeneous propellants is identified by performing a Monte-Carlo study of the correlation functions. This method of analysis shows great promise for understanding the origins of the combustion instability of heterogeneous propellants, and is believed to become a valuable tool for the development of safe and reliable rocket engines.

High-temperature earth-storable propellant acoustic cavity technology. [for combustion stability

NASA Technical Reports Server (NTRS)

Oberg, C. L.; Hines, W. S.; Falk, A. Y.

1974-01-01

Design criteria, methods and data, were developed to permit effective design of acoustic cavities for use in regeneratively cooled OME-type engines. This information was developed experimentally from two series of motor firings with high-temperature fuel during which the engine stability was evaluated under various conditions and with various cavity configurations. Supplementary analyses and acoustic model testing were used to aid cavity design and interpretation of results. Results from this program clearly indicate that dynamic stability in regeneratively cooled OME-type engines can be ensured through the use of acoustic cavities. Moreover, multiple modes of instability were successfully suppressed with the cavity.

The crown splash

NASA Astrophysics Data System (ADS)

Deegan, Robert; Brunet, Philippe; Eggers, Jens

2008-11-01

The impact of a drop onto a liquid layer and the subsequent splash has important implications for diverse physical processes such as air-sea gas transfer, cooling, and combustion. In the crown splash parameter regime, the splash pattern is highly regular. We focus on this case as a model for the mechanism that leads to secondary droplets, and thus explain the drop size distribution resulting from the splash. We show that the mean number of secondary droplets is determined by the most unstable wavelength of the Rayleigh-Plateau instability. Variations from this mean are governed by the width of the spectrum. Our results for the crown splash will provide the basis for understanding more complicated splashes.

Flameholder Combustion Instability Study

DTIC Science & Technology

1978-05-01

following sections. 18 -. . . . . 7 -7-- (1) Fuel Veporization Before Flameholder As developed in the preceding analysis, the percentage of the liquid...1 Wide Data Scatter 22 A w90 Cones 1.4 Kerosene 1030 - 1 22 4 Hemispheres 1.1 Kerosene 1030 - 1 22 0 60 V-Gutter 1,73 Kerosene 900 - I Approach 4

Shock-initiated Combustion of a Spherical Density Inhomogeneity

NASA Astrophysics Data System (ADS)

Haehn, Nicholas; Oakley, Jason; Rothamer, David; Anderson, Mark; Ranjan, Devesh; Bonazza, Riccardo

2010-11-01

A spherical density inhomogeneity is prepared using fuel and oxidizer at a stoichiometric ratio and Xe as a diluent that increases the overall density of the bubble mixture (55% Xe, 30% H2, 15% O2). The experiments are performed in the Wisconsin Shock Tube Laboratory in a 9.2 m vertical shock tube with a 25.4 cm x 25.4 cm square cross-section. An injector is used to generate a 5 cm diameter soap film bubble filled with the combustible mixture. The injector retracts flush into the side of the tube releasing the bubble into a state of free fall. The combustible bubble is accelerated by a planar shock wave in N2 (2.0 < M < 2.8). The mismatch of acoustic impedances results in shock-focusing at the downstream pole of the bubble. The shock focusing results in localized temperatures and pressures significantly larger than nominal conditions behind a planar shock wave, resulting in auto-ignition at the focus. Planar Mie scattering and chemiluminescence are used simultaneously to visualize the bubble morphology and combustion characteristics. During the combustion phase, both the span-wise and stream-wise lengths of the bubble are seen to increase compared to the non-combustible scenario. Additionally, smaller instabilities are observed on the upstream surface, which are absent in the non-combustible bubbles.

Propagation of fires along mine workings: criteria and limits

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

Pervushin, Yu.V.

1978-01-01

Underground fires account for over 50% of the accidents occuring in Soviet mines. Their prevention therefore occupies a central place in mine rescue practice and accident prevention. The general features of the physical processes occurring during propagation of a flame have been studied in some detail. Attempts have been made to describe underground fires on the basis of experimental data. However, it is not yet possible to make accurate preductions of the behavior of fires in mine workings: very many factors influence their development. The dynamics of spread of a flame along a working involves such diverse phenomena as heatmore » transfer by thermal conduction, radiation, and convection, transfer of oxygen and combustible gaseous components by draughts and diffusion, various chemical reactions on the surface of combustible materials and within the flames, and finally complex surface effects accompanying heat and mass transfer at interfaces between media. In addition, we must take account of the specific conditions prevailing in a mine - the complex geometrical configuration of the workings, the nonuniformity of the combustible materials, and the role of ventilation and its instability during fires. There can be many approaches to the study of such a many-sided process. The most promising lines seem to be those in which experimental models of the complex of possible phenomena are combined with mathematical models of the process, based on the equations of chemical hydrodynamics, in which the alternative variants are realized on a computer.« less

Development of a computerized analysis for solid propellant combustion instability with turbulence

NASA Technical Reports Server (NTRS)

Chung, T. J.; Park, O. Y.

1988-01-01

A multi-dimensional numerical model has been developed for the unsteady state oscillatory combustion of solid propellants subject to acoustic pressure disturbances. Including the gas phase unsteady effects, the assumption of uniform pressure across the flame zone, which has been conventionally used, is relaxed so that a higher frequency response in the long flame of a double-base propellant can be calculated. The formulation is based on a premixed, laminar flame with a one-step overall chemical reaction and the Arrhenius law of decomposition with no condensed phase reaction. In a given geometry, the Galerkin finite element solution shows the strong resonance and damping effect at the lower frequencies, similar to the result of Denison and Baum. Extended studies deal with the higher frequency region where the pressure varies in the flame thickness. The nonlinear system behavior is investigated by carrying out the second order expansion in wave amplitude when the acoustic pressure oscillations are finite in amplitude. Offset in the burning rate shows a negative sign in the whole frequency region considered, and it verifies the experimental results of Price. Finally, the velocity coupling in the two-dimensional model is discussed.

Combustor flame flashback

NASA Technical Reports Server (NTRS)

Proctor, M. P.; Tien, J. S.

1985-01-01

A stainless steel, two-dimensional (rectangular), center-dump, premixed-prevaporized combustor with quartz window sidewalls for visual access was designed, built, and used to study flashback. A parametric study revealed that the flashback equivalence ratio decreased slightly as the inlet air temperature increased. It also indicated that the average premixer velocity and premixer wall temperature were not governing parameters of flashback. The steady-state velocity balance concept as the flashback mechanism was not supported. From visual observation several stages of burning were identified. High speed photography verified upstream flame propagation with the leading edge of the flame front near the premixer wall. Combustion instabilities (spontaneous pressure oscillations) were discovered during combustion at the dump plane and during flashback. The pressure oscillation frequency ranged from 40 to 80 Hz. The peak-to-peak amplitude (up to 1.4 psi) increased as the fuel/air equivalence ratio was increased attaining a maximum value just before flashback. The amplitude suddenly decreased when the flame stabilized in the premixer. The pressure oscillations were large enough to cause a local flow reversal. A simple test using ceramic fiber tufts indicated flow reversals existed at the premixer exit during flickering. It is suspected that flashback occurs through the premixer wall boundary layer flow reversal caused by combustion instability. A theoretical analysis of periodic flow in the premixing channel has been made. The theory supports the flow reversal mechanism.

J-2X Gas Generator Development Testing at NASA Marshall Space Flight Center

NASA Technical Reports Server (NTRS)

Reynolds, D. C.; Hormonzian, Carlo

2010-01-01

NASA is developing a liquid oxygen/liquid hydrogen rocket engine for upper stage and trans-lunar applications of the Ares vehicles for the Constellation program. This engine, designated the J-2X, is a higher pressure, higher thrust variant of the Apollo-era J-2 engine. Development was contracted to Pratt & Whitney Rocketdyne in 2006. Over the past several years, two phases of testing have been completed on the development of the gas generator for the J-2X engine. The hardware has progressed through a variety of workhorse injector, chamber, and feed system configurations. Several of these configurations have resulted in combustion instability of the gas generator assembly. Development of the final configuration of workhorse hardware (which will ultimately be used to verify critical requirements on a component level) has required a balance between changes in the injector and chamber hardware in order to successfully mitigate the combustion instability without sacrificing other engine system requirements. This paper provides an overview of the two completed test series, performed at NASA s Marshall Space Flight Center. The requirements, facility setup, hardware configurations, and test series progression are detailed. Significant levels of analysis have been performed in order to provide design solutions to mitigate the combustion stability issues, and these are briefly covered. Also discussed are the results of analyses related to either anomalous readings or off-nominal testing throughout the two test series.

Some experiments related to L-star instability in rocket motors

NASA Technical Reports Server (NTRS)

Kumar, R. N.; Mcnamara, R. P.

1973-01-01

The influence of condensed phase heterogeneity on the L-star instability of nonmetallized AP/PBAN propellants is explored using four propellants (with monomodal AP particle distributions having 50 per cent weight average points at 11, 39.5, 175, and 350 microns). An economical firing program is used. One-dimensional nature of the Helmholtz mode and the complex nature of the chuff mode are revealed through color movies. The stability boundary on the L-star pressure plot is found to be parabolic. Frequency correlations and many other features reveal the important role of condensed phase details in propellant combustion.

Damping parameter study of a perforated plate with bias flow

NASA Astrophysics Data System (ADS)

Mazdeh, Alireza

One of the main impediments to successful operation of combustion systems in industrial and aerospace applications including gas turbines, ramjets, rocket motors, afterburners (augmenters) and even large heaters/boilers is the dynamic instability also known as thermo-acoustic instability. Concerns with this ongoing problem have grown with the introduction of Lean Premixed Combustion (LPC) systems developed to address the environmental concerns associated with the conventional combustion systems. The most common way to mitigate thermo-acoustic instability is adding acoustic damping to the combustor using acoustic liners. Recently damping properties of bias flow initially introduced to liners only for cooling purposes have been recognized and proven to be an asset in enhancing the damping effectiveness of liners. Acoustic liners are currently being designed using empirical design rules followed by build-test-improve steps; basically by trial and error. There is growing concerns on the lack of reliability associated with the experimental evaluation of the acoustic liners with small size apertures. The development of physics-based tools in assisting the design of such liners has become of great interest to practitioners recently. This dissertation focuses primarily on how Large-Eddy Simulations (LES) or similar techniques such as Scaled Adaptive Simulation (SAS) can be used to characterize damping properties of bias flow. The dissertation also reviews assumptions made in the existing analytical, semi-empirical, and numerical models, provides a criteria to rank order the existing models, and identifies the best existing theoretical model. Flow field calculations by LES provide good insight into the mechanisms that led to acoustic damping. Comparison of simulation results with empirical and analytical studies shows that LES simulation is a viable alternative to the empirical and analytical methods and can accurately predict the damping behavior of liners. Currently the role of LES for research studies concerned with damping properties of liners is limited to validation of other empirical or theoretical approaches. This research has shown that LES can go beyond that and can be used for performing parametric studies to characterize the sensitivity of acoustic properties of multi--perforated liners to the changes in the geometry and flow conditions and be used as a tool to design acoustic liners. The conducted research provides an insightful understanding about the contribution of different flow and geometry parameters such as perforated plate thickness, aperture radius, porosity factors and bias flow velocity. While the study agrees with previous observations obtained by analytical or experimental methods, it also quantifies the impact from these parameters on the acoustic impedance of perforated plate, a key parameter to determine the acoustic performance of any system. The conducted study has also explored the limitations and capabilities of commercial tool when are applied for performing simulation studies on damping properties of liners. The overall agreement between LES results and previous studies proves that commercial tools can be effectively used for these applications under certain conditions.

Premixed Flames Under Microgravity and Normal Gravity Conditions

NASA Astrophysics Data System (ADS)

Krikunova, Anastasia I.; Son, Eduard E.

2018-03-01

Premixed conical CH4-air flames were studied experimentally and numerically under normal straight, reversed gravity conditions and microgravity. Low-gravity experiments were performed in Drop tower. Classical Bunsen-type burner was used to find out features of gravity influence on the combustion processes. Mixture equivalence ratio was varied from 0.8 to 1.3. Wide range of flow velocity allows to study both laminar and weakly turbulized flames. High-speed flame chemoluminescence video-recording was used as diagnostic. The investigations were performed at atmospheric pressure. As results normalized flame height, laminar flame speed were measured, also features of flame instabilities were shown. Low- and high-frequency flame-instabilities (oscillations) have a various nature as velocity fluctuations, preferential diffusion instability, hydrodynamic and Rayleigh-Taylor ones etc., that was explored and demonstrated.

Some experiments related to L-star instability in rocket motors

NASA Technical Reports Server (NTRS)

Kumar, R. N.; Mcnamara, R. P.

1973-01-01

The role of solid phase heterogeneity on the low-pressure L-star instability of nonmetallized AP/PBAN propellants is explored. Four particle size distributions are employed in propellants that are otherwise identical. Over one hundred test firings were conducted in the 21/2 in. diameter L-star burner. Pressure time histories in the chamber and color movies of two firings constitute the raw data. An economical firing program was used which enables the interesting range of L-star values to be covered during a single firing (at a set mean pressure), through the variations in the depleting propellant volume. Time-independent combustion, Helmholtz mode, chuff mode, and the pressure-burst phenomena are revealed as the principal signatures. Of these, the Helmholtz mode is found to be the most ordered form of instability.

Detailed Multidimensional Simulations of the Structure and Dynamics of Flames

NASA Technical Reports Server (NTRS)

Patnaik, G.; Kailasanath, K.

1999-01-01

Numerical simulations in which the various physical and chemical processes can be independently controlled can significantly advance our understanding of the structure, stability, dynamics and extinction of flames. Therefore, our approach has been to use detailed time-dependent, multidimensional, multispecies numerical models to perform carefully designed computational experiments of flames on Earth and in microgravity environments. Some of these computational experiments are complementary to physical experiments performed under the Microgravity Program while others provide a fundamental understanding that cannot be obtained from physical experiments alone. In this report, we provide a brief summary of our recent research highlighting the contributions since the previous microgravity combustion workshop. There are a number of mechanisms that can cause flame instabilities and result in the formation of dynamic multidimensional structures. In the past, we have used numerical simulations to show that it is the thermo-diffusive instability rather than an instability due to preferential diffusion that is the dominant mechanism for the formation of cellular flames in lean hydrogen-air mixtures. Other studies have explored the role of gravity on flame dynamics and extinguishment, multi-step kinetics and radiative losses on flame instabilities in rich hydrogen-air flames, and heat losses on burner-stabilized flames in microgravity. The recent emphasis of our work has been on exploring flame-vortex interactions and further investigating the structure and dynamics of lean hydrogen-air flames in microgravity. These topics are briefly discussed after a brief discussion of our computational approach for solving these problems.

Investigations of two-phase flame propagation under microgravity conditions

NASA Astrophysics Data System (ADS)

Gokalp, Iskender

2016-07-01

Investigations of two-phase flame propagation under microgravity conditions R. Thimothée, C. Chauveau, F. Halter, I Gökalp Institut de Combustion, Aérothermique, Réactivité et Environnement (ICARE), CNRS, 1C Avenue de la Recherche Scientifique, 45071 Orléans Cedex 2, France This paper presents and discusses recent results on two-phase flame propagation experiments we carried out with mono-sized ethanol droplet aerosols under microgravity conditions. Fundamental studies on the flame propagation in fuel droplet clouds or sprays are essential for a better understanding of the combustion processes in many practical applications including internal combustion engines for cars, modern aircraft and liquid rocket engines. Compared to homogeneous gas phase combustion, the presence of a liquid phase considerably complicates the physico-chemical processes that make up combustion phenomena by coupling liquid atomization, droplet vaporization, mixing and heterogeneous combustion processes giving rise to various combustion regimes where ignition problems and flame instabilities become crucial to understand and control. Almost all applications of spray combustion occur under high pressure conditions. When a high pressure two-phase flame propagation is investigated under normal gravity conditions, sedimentation effects and strong buoyancy flows complicate the picture by inducing additional phenomena and obscuring the proper effect of the presence of the liquid droplets on flame propagation compared to gas phase flame propagation. Conducting such experiments under reduced gravity conditions is therefore helpful for the fundamental understanding of two-phase combustion. We are considering spherically propagating two-phase flames where the fuel aerosol is generated from a gaseous air-fuel mixture using the condensation technique of expansion cooling, based on the Wilson cloud chamber principle. This technique is widely recognized to create well-defined mono-size droplets uniformly distributed. Ethanol-air mixtures are used and the experiments are performed under reduced gravity conditions in the Airbus A310 ZERO-G of the CNES, during which a 10-2g gravity level is achieved. The experiments are conducted in a pressure-release type dual chamber which consists of a spherical combustion chamber of 1 L which is centered in a high pressure chamber of 11 L. Propagating flames under various mixture, droplet size and pressure conditions are investigated with various optical techniques. The collected flame images and the deduced flame propagation velocities enabled to establish various flame propagation and cellular instability regimes, mainly depending on the droplet size and droplet density. The experiments also permitted comparisons with gaseous flames having the same global equivalence ratio as the two-phase flames, therefore allowing analyzing clearly the role of the presence of the droplets in the flame propagation process.

Formation of Sprays From Conical Liquid Sheets

NASA Technical Reports Server (NTRS)

Peck, Bill; Mansour, N. N.; Koga, Dennis (Technical Monitor)

1999-01-01

Our objective is to predict droplet size distributions created by fuel injector nozzles in Jet turbines. These results will be used to determine the initial conditions for numerical simulations of the combustion process in gas turbine combustors. To predict the droplet size distribution, we are currently constructing a numerical model to understand the instability and breakup of thin conical liquid sheets. This geometry serves as a simplified model of the liquid jet emerging from a real nozzle. The physics of this process is difficult to study experimentally as the time and length scales are very short. From existing photographic data, it does seem clear that three-dimensional effects such as the formation of streamwise ligaments and the pulling back of the sheet at its edges under the action of surface tension are important.

Combustion Instability in Solid Propellant Rockets

DTIC Science & Technology

1989-03-21

adverse pressure gradients may arise. As suggested in Figure 5.1, the volume behind a submerged nozzle is especially likely to exhibit recircu- lation...ranges of interest. Therefore, the axial vortical velocity is governed to zeroth order in the mean flow Mach number by auz a2 2 U b+ Mbr -&r2 +O(Mb

Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Alternate low-thrust capability task report

NASA Technical Reports Server (NTRS)

Mellish, J. A.

1980-01-01

The feasibility and design impact of a requirement for the advanced expander cycle engine to be adaptable to extended low thrust operation of approximately 1K to 2K lb is assessed. It is determined that the orbit transfer vehicle point design engine can be reduced in thrust with minor injector modifications from 15K to 1K without significantly affecting combustion performance efficiency or injector face/chamber wall thermal compatibility. Likewise, high frequency transverse mode combustion instability is not expected to be detrimentally affected. Primarily, the operational limitations consist of feed system chugging instabilities and potential coupling of the injector response with the chamber longitudinal mode resonances under certain operating conditions. The recommended injector modification for low thrust operation is a change in the oxidizer injector element orifice size. Analyses also indicate that chamber coolant flow stability may be a concern below 2K 1bF operation and oxidizer pump stability could be a problem below a 2K thrust level although a recirculation flow could alleviate the problem.

Amplification of Reynolds number dependent processes by wave distortion. [acoustic instability of liquid propellant rocket engines

NASA Technical Reports Server (NTRS)

Ventrice, M. B.; Fang, J. C.; Purdy, K. R.

1975-01-01

A system using a hot-wire transducer as an analog of a liquid droplet of propellant was employed to investigate the ingredients of the acoustic instability of liquid-propellant rocket engines. It was assumed that the combustion process was vaporization-limited and that the combustion chamber was acoustically similar to a closed-closed right-circular cylinder. Before studying the hot-wire closed-loop system (the analog system), a microphone closed-loop system, which used the response of a microphone as the source of a linear feedback exciting signal, was investigated to establish the characteristics of self-sustenance of acoustic fields. Self-sustained acoustic fields were found to occur only at resonant frequencies of the chamber. In the hot-wire closed-loop system, the response of hot-wire anemometer was used as the source of the feedback exciting signal. The self-sustained acoustic fields which developed in the system were always found to be harmonically distorted and to have as their fundamental frquency a resonant frequency for which there also existed a second resonant frequency which was approximately twice the fundamental frequency.

Investigation of Multiscale and Multiphase Flow, Transport and Reaction in Heavy Oil Recovery Processes

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

Yortsos, Yanis C.

In this report, the thrust areas include the following: Internal drives, vapor-liquid flows, combustion and reaction processes, fluid displacements and the effect of instabilities and heterogeneities and the flow of fluids with yield stress. These find respective applications in foamy oils, the evolution of dissolved gas, internal steam drives, the mechanics of concurrent and countercurrent vapor-liquid flows, associated with thermal methods and steam injection, such as SAGD, the in-situ combustion, the upscaling of displacements in heterogeneous media and the flow of foams, Bingham plastics and heavy oils in porous media and the development of wormholes during cold production.

Advanced Transportation Systems, Alternate Propulsion Subsystem Concepts

NASA Technical Reports Server (NTRS)

1997-01-01

An understanding of the basic flow of of the subject hybrid model has been gained through this series of testing. Changing injectors (axial vs. radial) and inhibiting the flow between the upstream plenum and the CP section changes the basic flow structure, as evidenced by streamline and velocity contour plots. Numerous shear layer structures were identified in the test configurations; these structures include both standing and traveling vortices which may affect combustion ion stability. Standing vortices may play a role in the heat addition process as the oxidizer enters the motor, while traveling vortices may be instability mechanisms in themselves. Finally, the flow visualization and LVD measurements give insight into determining the effects of flow induced shear layers.

Numerical Simulation of Wall Heat Load in Combustor Flow

NASA Astrophysics Data System (ADS)

Panara, D.; Hase, M.; Krebs, W.; Noll, B.

2007-09-01

Due to the major mechanism of NOx generation, there is generally a temperature trade off between improved cycle efficiency, material constraints and low NOx emission. The cycle efficiency is proportional to the highest cycle temperature, but unfortunately also the NOx production increases with increasing combustion temperature. For this reason, the modern combustion chamber design has been oriented towards lean premixed combustion system and more and more attention must be focused on the cooling air management. The challenge is to ensure sufficiently low temperature of the combustion liner with very low amount of film or effusion cooling air. Correct numerical prediction of temperature fields and wall heat load are therefore of critical interest in the modern combustion chamber design. Moreover, lean combustion technology has shown the appearance of thermo-acoustic instabilities which have to be taken into account in the simulation and, more in general, in the design of reliable combustion systems. In this framework, the present investigation addresses the capability of a commercial multiphysics code (ANSYS CFX) to correctly predict the wall heat load and the core flow temperature field in a scaled power generation combustion chamber with a simplified ceramic liner. Comparison are made with the experimental results from the ITS test rig at the University of Karlsruhe [1] and with a previous numerical campaign from [2]. In addition the effect of flow unsteadyness on the wall heat load is discussed showing some limitations of the traditional steady state flow thermal design.

Acoustic response of Helmholtz dampers in the presence of hot grazing flow

NASA Astrophysics Data System (ADS)

Ćosić, B.; Wassmer, D.; Terhaar, S.; Paschereit, C. O.

2015-01-01

Thermoacoustic instabilities are high amplitude instabilities of premixed gas turbine combustors. Cooled passive dampers are used to attenuate or suppress these instabilities in the combustion chamber. For the first time, the influence of temperature differences between the grazing flow in the combustor and the cross-flow emanating from the Helmholtz damper is comprehensively investigated in the linear and nonlinear amplitude regime. The flow field inside the resonator and in the vicinity of the neck is measured with high-speed particle image velocimetry for various amplitudes and at different momentum-flux ratios of grazing and purging flow. Seeding is used as a tracer to qualitatively assess the mixing of the grazing and purging flow as well as the ingestion into the neck of the resonator. Experimentally, the acoustic response for various temperature differences between grazing and purging flow is investigated. The multi-microphone method, in combination with two microphones flush-mounted in the resonator volume and two microphones in the plane of the resonator entrance, is used to determine the impedance of the Helmholtz resonator in the linear and nonlinear amplitude regime for various temperatures and different momentum-flux ratios. Additionally, a thermocouple was used to measure the temperature in the neck. The acoustic response and the temperature measurements are used to obtain the virtual neck length and the effective area jump from a detailed impedance model. This model is extended to include the observed acoustic energy dissipation caused by the density gradients at the neck vicinity. A clear correlation between temperature differences and changes of the mass end-correction is confirmed. The capabilities of the impedance model are demonstrated.

Forced synchronization and asynchronous quenching in a thermo-acoustic system

NASA Astrophysics Data System (ADS)

Mondal, Sirshendu; Pawar, Samadhan A.; Sujith, Raman

2017-11-01

Forced synchronization, which has been extensively studied in theory and experiments, occurs through two different mechanisms known as phase locking and asynchronous quenching. The latter indicates the suppression of oscillation amplitude. In most practical combustion systems such as gas turbine engines, the main concern is high amplitude pressure oscillations, known as thermo-acoustic instability. Thermo-acoustic instability is undesirable and needs to be suppressed because of its damaging consequences to an engine. In the present study, a systematic experimental investigation of forced synchronization is performed in a prototypical thermo-acoustic system, a Rijke tube, in its limit cycle operation. Further, we show a qualitatively similar behavior using a reduced order model. In the phase locking region, the simultaneous occurrence of synchronization and resonant amplification leads to high amplitude pressure oscillations. However, a reduction in the amplitude of natural oscillations by about 78% of the unforced amplitude is observed when the forcing frequency is far lower than the natural frequency. This shows the possibility of suppression of the oscillation amplitude through asynchronous quenching in thermo-acoustic systems.

Investigation of flame driving and flow turning in axial solid rocket instabilities

NASA Astrophysics Data System (ADS)

Zinn, Ben T.; Daniel, Brady R.; Matta, Lawrence M.

1993-08-01

An understanding of the processes responsible for driving and damping acoustic oscillations in solid rocket motors is necessary for developing practical design methods that eliminate or reduce the occurrence combustion instabilities. While state of the art solid rocket stability prediction methods generally account for the flow turning loss, the magnitude and characteristics of this loss have never been fully investigated. Results of an investigation of the role of the flow turning loss in the stability of solid rockets and its dependence upon motor design and operating parameters are described. A one dimensional acoustic stability equation that verifies that the flow turning loss term is appropriately included in the one dimensional stability formulation was derived for a chamber with a constant mean temperature and pressure by an approach independent from that of Culick. This study was extended providing the background and expressions needed to guide an experimental study of the flow turning loss in the presence of mean temperature and density gradients. This allows the study of combustion systems in which mean temperature gradients and heat losses are significant. The relevant conservation equations were solved numerically for the experimental configuration in order to predict the behavior of the flow turning loss and to assist in the analysis of experimental results. Experiments performed, with and without combustion, showed that the flow turning loss strongly depends upon the propellant burning rate and the location of the flow turning region relative to the standing pressure wave.

Large-eddy simulation of a bluff-body stabilised turbulent premixed flame using the transported flame surface density approach

NASA Astrophysics Data System (ADS)

Lee, Chin Yik; Cant, Stewart

2017-07-01

A premixed propane-air flame stabilised on a triangular bluff body in a model jet-engine afterburner configuration is investigated using large-eddy simulation (LES). The reaction rate source term for turbulent premixed combustion is closed using the transported flame surface density (TFSD) model. In this approach, there is no need to assume local equilibrium between the generation and destruction of subgrid FSD, as commonly done in simple algebraic closure models. Instead, the key processes that create and destroy FSD are accounted for explicitly. This allows the model to capture large-scale unsteady flame propagation in the presence of combustion instabilities, or in situations where the flame encounters progressive wrinkling with time. In this study, comprehensive validation of the numerical method is carried out. For the non-reacting flow, good agreement for both the time-averaged and root-mean-square velocity fields are obtained, and the Karman type vortex shedding behaviour seen in the experiment is well represented. For the reacting flow, two mesh configurations are used to investigate the sensitivity of the LES results to the numerical resolution. Profiles for the velocity and temperature fields exhibit good agreement with the experimental data for both the coarse and dense mesh. This demonstrates the capability of LES coupled with the TFSD approach in representing the highly unsteady premixed combustion observed in this configuration. The instantaneous flow pattern and turbulent flame behaviour are discussed, and the differences between the non-reacting and reacting flow are described through visualisation of vortical structures and their interaction with the flame. Lastly, the generation and destruction of FSD are evaluated by examining the individual terms in the FSD transport equation. Localised regions where straining, curvature and propagation are each dominant are observed, highlighting the importance of non-equilibrium effects of FSD generation and destruction in the model afterburner.

Features of the propagation of laminar spherical flames initiated by a spark discharge in mixtures of methane, pentane, and hydrogen with air at atmospheric pressure

NASA Astrophysics Data System (ADS)

Rubtsov, N. M.; Seplyarskii, B. S.; Troshin, K. Ya.; Chernysh, V. I.; Tsvetkov, G. I.

2011-10-01

Using high-speed digital color cinematography, we studied the propagation of a laminar spherical flame in stoichiometric mixtures of hydrogen, methane, and pentane with air in the presence of additives at atmospheric pressure in constant-volume reactors, and derived quantitative data on the time of formation of a stable flame front. Cellular flames caused by gas-dynamic instability attributable to convective flows arising during the afterburning of gas were observed in hydrocarbon-air stoichiometric mixtures diluted with inert additives. It was found that the effect of additives of carbon dioxide and argon (>10%) and minor additives of CCl4 on the combustion of hydrocarbons, and of propylene on the combustion of hydrogen-rich mixtures, lead to periods of delay in the development of a laminar spherical flame; in addition, additives of propylene promote the combustion of hydrogen poor mixtures.

Combustion Processes in Hybrid Rocket Engines

NASA Technical Reports Server (NTRS)

Venkateswaran,S.; Merkle, C. L.

1996-01-01

In recent years, there has been a resurgence of interest in the development of hybrid rocket engines for advanced launch vehicle applications. Hybrid propulsion systems use a solid fuel such as hydroxyl-terminated polybutadiene (HTPB) along with a gaseous/liquid oxidizer. The performance of hybrid combustors depends on the convective and radiative heat fluxes to the fuel surface, the rate of pyrolysis in the solid phase, and the turbulent combustion processes in the gaseous phases. These processes in combination specify the regression rates of the fuel surface and thereby the utilization efficiency of the fuel. In this paper, we employ computational fluid dynamics (CFD) techniques in order to gain a quantitative understanding of the physical trends in hybrid rocket combustors. The computational modeling is tailored to ongoing experiments at Penn State that employ a two dimensional slab burner configuration. The coordinated computational/experimental effort enables model validation while providing an understanding of the experimental observations. Computations to date have included the full length geometry with and with the aft nozzle section as well as shorter length domains for extensive parametric characterization. HTPB is sed as the fuel with 1,3 butadiene being taken as the gaseous product of the pyrolysis. Pure gaseous oxygen is taken as the oxidizer. The fuel regression rate is specified using an Arrhenius rate reaction, which the fuel surface temperature is given by an energy balance involving gas-phase convection and radiation as well as thermal conduction in the solid-phase. For the gas-phase combustion, a two step global reaction is used. The standard kappa - epsilon model is used for turbulence closure. Radiation is presently treated using a simple diffusion approximation which is valid for large optical path lengths, representative of radiation from soot particles. Computational results are obtained to determine the trends in the fuel burning or regression rates as a function of the head-end oxidizer mass flux, G=rho(e)U(e), and the chamber pressure. Furthermore, computation of the full slab burner configuration has also been obtained for various stages of the burn. Comparisons with available experimental data from small scale tests conducted by General Dynamics-Thiokol-Rocketdyne suggest reasonable agreement in the predicted regression rates. Future work will include: (1) a model for soot generation in the flame for more quantitative radiative transfer modelling, (2) a parametric study of combustion efficiency, and (3) transient calculations to help determine the possible mechanisms responsible for combustion instability in hybrid rocket motors.

Time-dependent computational studies of flames in microgravity

NASA Technical Reports Server (NTRS)

Oran, Elaine S.; Kailasanath, K.

1989-01-01

The research performed at the Center for Reactive Flow and Dynamical Systems in the Laboratory for Computational Physics and Fluid Dynamics, at the Naval Research Laboratory, in support of the NASA Microgravity Science and Applications Program is described. The primary focus was on investigating fundamental questions concerning the propagation and extinction of premixed flames in Earth gravity and in microgravity environments. The approach was to use detailed time-dependent, multispecies, numerical models as tools to simulate flames in different gravity environments. The models include a detailed chemical kinetics mechanism consisting of elementary reactions among the eight reactive species involved in hydrogen combustion, coupled to algorithms for convection, thermal conduction, viscosity, molecular and thermal diffusion, and external forces. The external force, gravity, can be put in any direction relative to flame propagation and can have a range of values. A combination of one-dimensional and two-dimensional simulations was used to investigate the effects of curvature and dilution on ignition and propagation of flames, to help resolve fundamental questions on the existence of flammability limits when there are no external losses or buoyancy forces in the system, to understand the mechanism leading to cellular instability, and to study the effects of gravity on the transition to cellular structure. A flame in a microgravity environment can be extinguished without external losses, and the mechanism leading to cellular structure is not preferential diffusion but a thermo-diffusive instability. The simulations have also lead to a better understanding of the interactions between buoyancy forces and the processes leading to thermo-diffusive instability.

Thermal Ignition

NASA Astrophysics Data System (ADS)

Boettcher, Philipp Andreas

Accidental ignition of flammable gases is a critical safety concern in many industrial applications. Particularly in the aviation industry, the main areas of concern on an aircraft are the fuel tank and adjoining regions, where spilled fuel has a high likelihood of creating a flammable mixture. To this end, a fundamental understanding of the ignition phenomenon is necessary in order to develop more accurate test methods and standards as a means of designing safer air vehicles. The focus of this work is thermal ignition, particularly auto-ignition with emphasis on the effect of heating rate, hot surface ignition and flame propagation, and puffing flames. Combustion of hydrocarbon fuels is traditionally separated into slow reaction, cool flame, and ignition regimes based on pressure and temperature. Standard tests, such as the ASTM E659, are used to determine the lowest temperature required to ignite a specific fuel mixed with air at atmospheric pressure. It is expected that the initial pressure and the rate at which the mixture is heated also influences the limiting temperature and the type of combustion. This study investigates the effect of heating rate, between 4 and 15 K/min, and initial pressure, in the range of 25 to 100 kPa, on ignition of n-hexane air mixtures. Mixtures with equivalence ratio ranging from 0.6 to 1.2 were investigated. The problem is also modeled computationally using an extension of Semenov's classical auto-ignition theory with a detailed chemical mechanism. Experiments and simulations both show that in the same reactor either a slow reaction or an ignition event can take place depending on the heating rate. Analysis of the detailed chemistry demonstrates that a mixture which approaches the ignition region slowly undergoes a significant modification of its composition. This change in composition induces a progressive shift of the explosion limit until the mixture is no longer flammable. A mixture that approaches the ignition region sufficiently rapidly undergoes only a moderate amount of thermal decomposition and explodes quite violently. This behavior can also be captured and analyzed using a one-step reaction model, where the heat release is in competition with the depletion of reactants. Hot surface ignition is examined using a glow plug or heated nickel element in a series of premixed n-hexane air mixtures. High-speed schlieren photography, a thermocouple, and a fast response pressure transducer are used to record flame characteristics such as ignition temperature, flame speed, pressure rises, and combustion mode. The ignition event is captured by considering the dominant balance of diffusion and chemical reaction that occurs near a hot surface. Experiments and models show a dependence of ignition temperature on mixture composition, initial pressure, and hot surface size. The mixtures exhibit the known lower flammability limit where the maximum temperature of the hot surface was insufficient at igniting the mixture. Away from the lower flammability limit, the ignition temperature drops to an almost constant value over a wide range of equivalence ratios (0.7 to 2.8) with large variations as the upper flammability limit is approached. Variations in the initial pressure and equivalence ratio also give rise to different modes of combustion: single flame, re-ignition, and puffing flames. These results are successfully compared to computational results obtained using a flamelet model and a detailed chemical mechanism for n-heptane. These different regimes can be delineated by considering the competition between inertia, i.e., flame propagation, and buoyancy, which can be expressed in the Richardson number. In experiments of hot surface ignition and subsequent flame propagation a 10 Hz puffing flame instability is visible in mixtures that are stagnant and premixed prior to the ignition sequence. By varying the size of the hot surface, power input, and combustion vessel volume, we determined that the instability is a function of the interaction of the flame with the fluid flow induced by the combustion products rather than the initial plume established by the hot surface. The phenomenon is accurately reproduced in numerical simulations and a detailed flow field analysis revealed a competition between the inflow velocity at the base of the flame and the flame propagation speed. The increasing inflow velocity, which exceeds the flame propagation speed, is ultimately responsible for creating a puff. The puff is then accelerated upward, allowing for the creation of the subsequent instabilities. The frequency of the puffing is proportional to the gravitational acceleration and inversely proportional to the flame speed. We propose a relation describing the dependence of the frequency on gravitational acceleration, hot surface diameter, and flame speed. This relation shows good agreement for lean and rich n-hexane-air as well as lean hydrogen-air flames.

Investigation of the flow turning loss in unstable solid propellant rocket motors

NASA Astrophysics Data System (ADS)

Matta, Lawrence Mark

The goal of this study was to improve the understanding of the flow turning loss, which contributes to the damping of axial acoustic instabilities in solid propellant rocket motors. This understanding is needed to develop practical methods for designing motors that do not exhibit such instabilities. The flow turning loss results from the interaction of the flow of combustion products leaving the surface of the propellant with the acoustic field in an unstable motor. While state of the art solid rocket stability models generally account for the flow turning loss, its magnitude and characteristics have never been fully investigated. This thesis describes a combined theoretical, numerical, and experimental investigation of the flow turning loss and its dependence upon various motor design and operating parameters. First, a one dimensional acoustic stability equation that verifies the existence of the flow turning loss was derived for a chamber with constant mean pressure and temperature. The theoretical development was then extended to include the effects of mean temperature gradients to accommodate combustion systems in which mean temperature gradients and heat losses are significant. These analyses provided the background and expressions necessary to guide an experimental study. The relevant equations were then solved for the developed experimental setup to predict the behavior of the flow turning loss and the other terms of the developed acoustic stability equation. This was followed by and experimental study in which the flow turning region of an unstable solid propellant rocket motor was simulated. The setup was used, with and without combustion, to determine the dependence of the flow turning loss upon operating conditions. These studies showed that the flow turning loss strongly depends upon the gas velocity at the propellant surface and the location of the flow turning region relative to the standing acoustic wave. The flow turning loss measured in the experiment was found to be small relative to other mechanisms. This, however, was characteristic of the experimental setup and is not representative of actual rocket motors, in which the flow turning loss is often a significant part of the overall stability.

International Experience in Developing Low-Emission Combustors for Land-Based, Large Gas-Turbine Units: Mitsubishi Heavy Industries' Equipment

NASA Astrophysics Data System (ADS)

Bulysova, L. A.; Vasil'ev, V. D.; Berne, A. L.; Gutnik, M. N.; Ageev, A. V.

2018-05-01

This is the second paper in a series of publications summarizing the international experience in the development of low-emission combustors (LEC) for land-based, large (above 250 MW) gas-turbine units (GTU). The purpose of this series is to generalize and analyze the approaches used by various manufacturers in designing flowpaths for fuel and air in LECs, managing fuel combustion, and controlling the fuel flow. The efficiency of advanced GTUs can be as high as 43% (with an output of 350-500 MW) while the efficiency of 600-800 MW combined-cycle units with these GTUs can attain 63.5%. These high efficiencies require a compression ratio of 20-24 and a temperature as high as 1600°C at the combustor outlet. Accordingly, the temperature in the combustion zone also rises. All the requirements for the control of harmful emissions from these GTUs are met. All the manufacturers and designers of LECs for modern GTUs encounter similar problems, such as emissions control, combustion instability, and reliable cooling of hot path parts. Methods of their elimination are different and interesting from the standpoint of science and practice. One more essential requirement is that the efficiency and environmental performance indices must be maintained irrespective of the fuel composition or heating value and also in operation at part loads below 40% of rated. This paper deals with Mitsubishi Series M701 GTUs, F, G, or J class, which have gained a good reputation in the power equipment market. A design of a burner for LECs and a control method providing stable low-emission fuel combustion are presented. The advantages and disadvantages of the use of air bypass valves installed in each liner to maintain a nearly constant air to fuel ratio within a wide range of GTU loads are described. Methods for controlling low- and high-frequency combustion instabilities are outlined. Upgrading of the cooling system for the wall of a liner and a transition piece is of great interest. Change over from effusion (or film) cooling to convective steam cooling and convective air cooling has considerably increased the GTU efficiency.

The Fluids and Combustion Facility

NASA Technical Reports Server (NTRS)

Kundu, Sampa

2004-01-01

Microgravity is an environment with very weak gravitational effects. The Fluids and Combustion Facility (FCF) on the International Space Station (ISS) will support the study of fluid physics and combustion science in a long-duration microgravity environment. The Fluid Combustion Facility's design will permit both independent and remote control operations from the Telescience Support Center. The crew of the International Space Station will continue to insert and remove the experiment module, store and reload removable data storage and media data tapes, and reconfigure diagnostics on either side of the optics benches. Upon completion of the Fluids Combustion Facility, about ten experiments will be conducted within a ten-year period. Several different areas of fluid physics will be studied in the Fluids Combustion Facility. These areas include complex fluids, interfacial phenomena, dynamics and instabilities, and multiphase flows and phase change. Recently, emphasis has been placed in areas that relate directly to NASA missions including life support, power, propulsion, and thermal control systems. By 2006 or 2007, a Fluids Integrated Rack (FIR) and a Combustion Integrated Rack (CIR) will be installed inside the International Space Station. The Fluids Integrated Rack will contain all the hardware and software necessary to perform experiments in fluid physics. A wide range of experiments that meet the requirements of the international space station, including research from other specialties, will be considered. Experiments will be contained in subsystems such as the international standard payload rack, the active rack isolation system, the optics bench, environmental subsystem, electrical power control unit, the gas interface subsystem, and the command and data management subsystem. In conclusion, the Fluids and Combustion Facility will allow researchers to study fluid physics and combustion science in a long-duration microgravity environment. Additional information is included in the original extended abstract.

Onset of Darrieus-Landau Instability in Expanding Flames

NASA Astrophysics Data System (ADS)

Mohan, Shikhar; Matalon, Moshe

2017-11-01

The effect of small amplitude perturbations on the propagation of circular flames in unconfined domains is investigated, computationally and analytically, within the context of the hydrodynamic theory. The flame, treated as a surface of density discontinuity separating fresh combustible mixture from the burnt gas, propagates at a speed dependent upon local curvature and hydrodynamic strain. For mixtures with Lewis numbers above criticality, thermodiffusive effects have stabilizing influences which largely affect the flame at small radii. The amplitude of these disturbances initially decay and only begin to grow once a critical radius is reached. This instability is hydrodynamic in nature and is a consequence of thermal expansion. Through linear stability analysis, predictions of critical flame radius at the onset of instability are obtained as functions of Markstein length and thermal expansion coefficients. The flame evolution is also examined numerically where the motion of the interface is tracked via a level-set method. Consistent with linear stability results, simulations show the flame initially remaining stable and the existence of a particular mode that will be first to grow and later determine the cellular structure observed experimentally at the onset of instability.

Current and Future Critical Issues in Rocket Propulsion Systems

NASA Technical Reports Server (NTRS)

Navaz, Homayun K.; Dix, Jeff C.

1998-01-01

The objective of this research was to tackle several problems that are currently of great importance to NASA. In a liquid rocket engine several complex processes take place that are not thoroughly understood. Droplet evaporation, turbulence, finite rate chemistry, instability, and injection/atomization phenomena are some of the critical issues being encountered in a liquid rocket engine environment. Pulse Detonation Engines (PDE) performance, combustion chamber instability analysis, 60K motor flowfield pattern from hydrocarbon fuel combustion, and 3D flowfield analysis for the Combined Cycle engine were of special interest to NASA. During the summer of 1997, we made an attempt to generate computational results for all of the above problems and shed some light on understanding some of the complex physical phenomena. For this purpose, the Liquid Thrust Chamber Performance (LTCP) code, mainly designed for liquid rocket engine applications, was utilized. The following test cases were considered: (1) Characterization of a detonation wave in a Pulse Detonation Tube; (2) 60K Motor wall temperature studies; (3) Propagation of a pressure pulse in a combustion chamber (under single and two-phase flow conditions); (4) Transonic region flowfield analysis affected by viscous effects; (5) Exploring the viscous differences between a smooth and a corrugated wall; and (6) 3D thrust chamber flowfield analysis of the Combined Cycle engine. It was shown that the LTCP-2D and LTCP-3D codes are capable of solving complex and stiff conservation equations for gaseous and droplet phases in a very robust and efficient manner. These codes can be run on a workstation and personal computers (PC's).

Suresh K. AggarwalQuantified Analysis of a Production Diesel Injector Using X-Ray Radiography and Engine Diagnostics

NASA Astrophysics Data System (ADS)

Ramirez, Anita I.

The work presented in this thesis pursues further the understanding of fuel spray, combustion, performance, and emissions in an internal combustion engine. Various experimental techniques including x-ray radiography, injection rate measurement, and in-cylinder endoscopy are employed in this work to characterize the effects of various upstream conditions such as injection rate profile and fuel physical properties. A single non-evaporating spray from a 6-hole full-production Hydraulically Actuated Electronically Controlled Unit Injector (HEUI) nozzle is studied under engine-like ambient densities with x-ray radiography at the Advanced Photon Source (APS) of Argonne National Laboratory (ANL). Two different injection pressures were investigated and parameters such as fuel mass distribution, spray penetration, cone angle, and spray velocity were obtained. The data acquired with x-ray radiography is used for the development and validation of improved Computational Fluid Dynamic (CFD) models. Rate of injection is studied using the same HEUI in a single cylinder Caterpillar test engine. The injection rate profile is altered to have three levels of initial injection pressure rise. Combustion behavior, engine performance, and emissions information was acquired for three rate profile variations. It is found that NOx emission reduction is achieved when the SOI timing is constant at the penalty of lower power generated in the cycle. However, if CA50 is aligned amongst the three profiles, the NOx emissions and power are constant with a slight penalty in CO emissions. The influence of physical and chemical parameters of fuel is examined in a study of the heavy alcohol, phytol (C20H40O), in internal combustion engine application. Phytol is blended with diesel in 5%, 10%, and 20% by volume. Combustion behavior is similar between pure diesel and the phytol/diesel blends with small differences noted in peak cylinder pressure, ignition delay, and heat release rate in the premix burn phase. Diesel/phytol blends yield marginally lower power values. In-cylinder soot radiation images show combustion instability at the start of the event for the 20% phytol/diesel blend. Overall, NOx emissions are comparable across the different fuels used and no discernible trend is found in CO emissions.

Active Control of Engine Dynamics (Le controle actif pour la dynamique des moteurs)

DTIC Science & Technology

2002-11-01

optimum operating conditions, avoiding, for example, inadvertent operation when the pulsations can cause unacceptable rates of surface heat transfer or...such as shipboard incineration, and power and heat generation in the field. Because the practical problem of suppressing combustion instabilities has...aforementioned physical processes are essentially completed prior to entering the combustor. One consequence of fuel-air premixing is that the heat

Climate change in the age of humans

Treesearch

J. Curt Stager

2014-01-01

The Anthropocene epoch presents a mix of old and new challenges for the world’s forests. Climatic instability has typified most of the Cenozoic Era but today’s situation is unique due to the presence of billions of humans on the planet. The potential rate and magnitude of future warming driven by continued fossil fuel combustion could be unprecedented during the last...

Electroacoustic control of Rijke tube instability

NASA Astrophysics Data System (ADS)

Zhang, Yumin; Huang, Lixi

2017-11-01

Unsteady heat release coupled with pressure fluctuation triggers the thermoacoustic instability which may damage a combustion chamber severely. This study demonstrates an electroacoustic control approach of suppressing the thermoacoustic instability in a Rijke tube by altering the wall boundary condition. An electrically shunted loudspeaker driver device is connected as a side-branch to the main tube via a small aperture. Tests in an impedance tube show that this device has sound absorption coefficient up to 40% under normal incidence from 100 Hz to 400 Hz, namely over two octaves. Experimental result demonstrates that such a broadband acoustic performance can effectively eliminate the Rijke-tube instability from 94 Hz to 378 Hz (when the tube length varies from 1.8 m to 0.9 m, the first mode frequency for the former is 94 Hz and the second mode frequency for the latter is 378 Hz). Theoretical investigation reveals that the devices act as a damper draining out sound energy through a tiny hole to eliminate the instability. Finally, it is also estimated based on the experimental data that small amount of sound energy is actually absorbed when the system undergoes a transition from the unstable to stable state if the contrpaol is activated. When the system is actually stabilized, no sound is radiated so no sound energy needs to be absorbed by the control device.

Development of a solvent processed insensitive propellant

NASA Technical Reports Server (NTRS)

Trask, R.; Costa, E.; Beardell, A. J.

1980-01-01

Two types of low vulnerability propellants are studied which are distinguished by whether the binder is a rubber, such as polyurethane or CTBN, or a plasticizable polymer such as ethyl cellulose or cellulose acetate. The former propellants are made by a partial cure extrusion process while the latter are made by the conventional solvent process. Emphasis is given to a cellulose binder (plasticizer) RDX composition. The type of binder used, the particle size of the RDX and the presence of small quantities of nitrocellulose in the solvent processed compositions have important influences on the mechanical and combustion characteristics of the propellant. The low temperature combustion is of particular concern because of potential breakup of the grains that can lead to instability.

Phase-locked two-line OH planar laser-induced fluorescence thermometry in a pulsating gas turbine model combustor at atmospheric pressure.

PubMed

Giezendanner-Thoben, Robert; Meier, Ulrich; Meier, Wolfgang; Heinze, Johannes; Aigner, Manfred

2005-11-01

Two-line OH planar laser-induced fluorescence (PLIF) thermometry was applied to a swirling CH4/air flame in a gas turbine (GT) model combustor at atmospheric pressure, which exhibited self-excited combustion instability. The potential and limitations of the method are discussed with respect to applications in GT-like flames. A major drawback of using OH as a temperature indicator is that no temperature information can be obtained from regions where OH radicals are missing or present in insufficient concentration. The resulting bias in the average temperature is addressed and quantified for one operating condition by a comparison with results from laser Raman measurements applied in the same flame. Care was taken to minimize saturation effects by decreasing the spectral laser power density to a minimum while keeping an acceptable spatial resolution and signal-to-noise ratio. In order to correct for the influence of laser light attenuation, absorption measurements were performed on a single-shot basis and a correction procedure was applied. The accuracy was determined to 4%-7% depending on the location within the flame and on the temperature level. A GT model combustor with an optical combustion chamber is described, and phase-locked 2D temperature distributions from a pulsating flame are presented. The temperature variations during an oscillation cycle are specified, and the general flame behavior is described. Our main goals are the evaluation of the OH PLIF thermometry and the characterization of a pulsating GT-like flame.

Uncertainty Quantification of Non-linear Oscillation Triggering in a Multi-injector Liquid-propellant Rocket Combustion Chamber

NASA Astrophysics Data System (ADS)

Popov, Pavel; Sideris, Athanasios; Sirignano, William

2014-11-01

We examine the non-linear dynamics of the transverse modes of combustion-driven acoustic instability in a liquid-propellant rocket engine. Triggering can occur, whereby small perturbations from mean conditions decay, while larger disturbances grow to a limit-cycle of amplitude that may compare to the mean pressure. For a deterministic perturbation, the system is also deterministic, computed by coupled finite-volume solvers at low computational cost for a single realization. The randomness of the triggering disturbance is captured by treating the injector flow rates, local pressure disturbances, and sudden acceleration of the entire combustion chamber as random variables. The combustor chamber with its many sub-fields resulting from many injector ports may be viewed as a multi-scale complex system wherein the developing acoustic oscillation is the emergent structure. Numerical simulation of the resulting stochastic PDE system is performed using the polynomial chaos expansion method. The overall probability of unstable growth is assessed in different regions of the parameter space. We address, in particular, the seven-injector, rectangular Purdue University experimental combustion chamber. In addition to the novel geometry, new features include disturbances caused by engine acceleration and unsteady thruster nozzle flow.

Compressible instability of rapidly expanding spherical material interfaces

NASA Astrophysics Data System (ADS)

Mankbadi, Mina Reda

The focus herein is on the instability of a material interface formed during an abrupt release of concentrated energy as in detonative combustion, explosive dispersals, and inertial-confinement fusion. These applications are modeled as a spherical shock-tube in which high-pressure gas initially contained in a small spherical shell is suddenly released. A forward-moving shock and an inward-moving secondary shock are formed, and between them a material interface develops that separates high-density fluid from the low-density one. The wrinkling of this interface controls mixing and energy release. The interface's stability is studied with and without the inclusion of metalized particulates. A numerical scheme is developed to discretize the full nonlinear equations of the base flow, and the 3D linearized perturbed flow equations. Linearization is followed by spherical harmonic decomposition of the disturbances, thereby reducing the 3D computational domain to one-dimensional radial domain. The 3D physical nature of the disturbances is maintained throughout the procedure. An extended Roe-Pike scheme coupled with a WENO scheme is developed to capture the discontinuities and accurately predict the disturbances. In Chapter 2, the contact interface's stability is analyzed in the inviscid single-phase. The disturbances grow exponentially and the growth rate is insensitive to the radial initial-disturbance profile. For wave numbers less than 100, the results are in accordance with previous theories but clarify that compressibility reduces the growth rate. Unlike the classical RTI, the growth rate reaches saturation at high wavenumbers. The parametric studies show that for specific ratios of initial pressure and temperature, the instability can be eliminated altogether. Chapter 3 discusses the full effects of viscosity and thermal diffusivity. Although Prandtl number effects are minimal, viscous effects dampen the high-wave numbers. For a given Reynolds number there is a peak wave number at which the disturbances are most amplified. In Chapter 4, the multiphase case with metalized particles is investigated. The quasi steady gas-particle interaction forces and heat transfer decelerate the contact interface and reduce its Atwood number, which results in reducing the growth of the interfacial instabilities. A parametric study of the multiphase instability is presented to assist in controlling the instability.

Fluid-dynamically coupled solid propellant combustion instability - cold flow simulation

NASA Astrophysics Data System (ADS)

Ben-Reuven, M.

1983-10-01

The near-wall processes in an injected, axisymmetric, viscous flow is examined. Solid propellant rocket instability, in which cold flow simulation is evaluated as a tool to elucidate possible instability driving mechanisms is studied. One such prominent mechanism seems to be visco-acoustic coupling. The formulation is presented in terms of a singular boundary layer problem, with detail (up to second order) given only to the near wall region. The injection Reynolds number is assumed large, and its inverse square root serves as an appropriate small perturbation quantity. The injected Mach number is also small, and taken of the same order as the aforesaid small quantity. The radial-dependence of the inner solutions up to second order is solved, in polynominal form. This leaves the (x,t) dependence to much simpler partial differential equations. Particular results demonstrate the existence of a first order pressure perturbation, which arises due to the dissipative near wall processes. This pressure and the associated viscous friction coefficient are shown to agree very well with experimental injected flow data.

Predictive Evaluations of Oxygen-Rich Hydrocarbon Combustion Gas-Centered Swirl Coaxial Injectors using a Flamelet-Based 3-D CFD Simulation Approach

NASA Technical Reports Server (NTRS)

Richardson, Brian R.; Braman, Kalem; West, Jeff

2016-01-01

NASA Marshall Space Flight Center (MSFC) has embarked upon a joint project with the Air Force to improve the state-of-the-art of space application combustion device design and operational understanding. One goal of the project is to design, build and hot-fire test a 40,000 pound-thrust Oxygen/Rocket Propellant-2 (RP-2) Oxygen-Rich staged engine at MSFC. The overall project goals afford the opportunity to test multiple different injector designs and experimentally evaluate the any effect on the engine performance and combustion dynamics. To maximize the available test resources and benefits, pre-test, combusting flow, Computational Fluid Dynamics (CFD) analysis was performed on the individual injectors to guide the design. The results of the CFD analysis were used to design the injectors for specific, targeted fluid dynamic features and the analysis results also provided some predictive input for acoustic and thermal analysis of the main Thrust Chamber Assembly (TCA). MSFC has developed and demonstrated the ability to utilize a computationally efficient, flamelet-based combustion model to guide the pre-test design of single-element Gas Centered Swirl Coaxial (GCSC) injectors. Previous, Oxygen/RP-2 simulation models utilizing the Loci-STREAM flow solver, were validated using single injector test data from the EC-1 Air Force test facility. The simulation effort herein is an extension of the validated, CFD driven, single-injector design approach applied to single injectors which will be part of a larger engine array. Time-accurate, Three-Dimensional, CFD simulations were performed for five different classes of injector geometries. Simulations were performed to guide the design of the injector to achieve a variety of intended performance goals. For example, two GCSC injectors were designed to achieve stable hydrodynamic behavior of the propellant circuits while providing the largest thermal margin possible within the design envelope. While another injector was designed to purposefully create a hydrodynamic instability in the fuel supply circuit as predicted by the CFD analysis. Future multi-injector analysis and testing will indicate what if any changes occur in the predicted behavior for the single-element injector when the same injector geometry is placed in a multi-element array.

Flame Dynamics and Chemistry in LRE Combustion Instability

DTIC Science & Technology

2016-12-22

simulation conditions are as follows: the upper boundary consists of a mixture of DME, oxygen and nitrogen at a fixed temperature of 300 K, while the lower...Fig. 11 a. However, the reduction effect of increased oxygen con- centration on the cool flame extinction temperature is again over- predicted by... temperature chemistry and extends the hysteresis between ignition and Fig. 11. Ignition and extinction temperatures at various strain rates and oxygen

Acoustically Forced Coaxial Hydrogen / Liquid Oxygen Jet Flames

DTIC Science & Technology

2016-05-15

serious problems in the development of liquid rocket engines. In order to understand and predict them, it is necessary to understand how representative...liquid rocket injector flames react to acoustic waves. In this study, a representative coaxial gaseous hydrogen / liquid oxygen (LOX) jet flame is...Combustion instabilities can pose serious problems in the development of liquid rocket engines. In order to under- stand and predict them, it is

Influences of the Darrieus-Landau instability on premixed turbulent flames

NASA Astrophysics Data System (ADS)

Patyal, Advitya; Matalon, Moshe

2017-11-01

The propagation of turbulent flames in three-dimensional turbulent flows is studied within the context of the hydrodynamic theory. The flame is treated as a surface of density discontinuity with the flow modified by gas expansion resulting from heat released during combustion. The flame is tracked using a level-set method with a propagation speed that depends on the local flame stretch, modulated by a Markstein length. Impact of the Darrieus-Landau instability on the topology of the flame surface is studied. It is shown that similar to passive interfaces, flames under the influence of the hydrodynamic instability resort to cylindrical structures with increasing turbulence intensity, even in 3D. The mechanism of modification of vortical structures in the burned gas is identified in terms of the alignments between the vorticity vector, flame surface normal and eigenvectors of the strain rate tensor. The results indicate that the strain rate tensor is intricately coupled with the normal to the flame surface and creates anisotropy in the orientation of vortical structures, which begins to weaken as the turbulent intensity increases. Furthermore, vorticity budgets are used to highlight the relative importance of baroclinic torque due to Darrieus-Landau instability.

Laser Absorption Measurements of Equivalence Ratios Studied Along With Their Coupling to Pressure Fluctuations in Lean Premixed Prevaporized (LPP) Combustion

NASA Technical Reports Server (NTRS)

Nguyen, Quang-Viet

2001-01-01

Concerns about damaging the Earth's ozone layer as a result of high levels of nitrogen oxides (known collectively as NOx) from high-altitude, high-speed aircraft have prompted the study of lean premixed prevaporized (LPP) combustion in aircraft engines. LPP combustion reduces NOx emissions principally by reducing the peak flame temperatures inside an engine. Recent advances in LPP technologies have realized exceptional reductions in pollutant emissions (single-digit ppm NOx for example). However, LPP combustion also presents major challenges: combustion instability and dynamic coupling effects between fluctuations in heat-release rate, dynamic pressure, and fuel pressure. These challenges are formidable and can literally shake an engine apart if uncontrolled. To better understand this phenomenon so that it can be controlled, we obtained real-time laser absorption measurements of the fuel vapor concentration (and equivalence ratio) simultaneously with the dynamic pressure, flame luminosity, and time-averaged gaseous emissions measurements in a research-type jet-A-fueled LPP combustor. The measurements were obtained in NASA Glenn Research Center's CE-5B optically accessible flame tube facility. The CE-5B facility provides inlet air temperatures and pressures similar to the actual operating conditions of real aircraft engines. The laser absorption measurements were performed using an infrared 3.39 micron HeNe laser in conjunction with a visible HeNe laser for liquid droplet scattering compensation.

A new method for the prediction of combustion instability

NASA Astrophysics Data System (ADS)

Flanagan, Steven Meville

This dissertation presents a new approach to the prediction of combustion instability in solid rocket motors. Previous attempts at developing computational tools to solve this problem have been largely unsuccessful, showing very poor agreement with experimental results and having little or no predictive capability. This is due primarily to deficiencies in the linear stability theory upon which these efforts have been based. Recent advances in linear instability theory by Flandro have demonstrated the importance of including unsteady rotational effects, previously considered negligible. Previous versions of the theory also neglected corrections to the unsteady flow field of the first order in the mean flow Mach number. This research explores the stability implications of extending the solution to include these corrections. Also, the corrected linear stability theory based upon a rotational unsteady flow field extended to first order in mean flow Mach number has been implemented in two computer programs developed for the Macintosh platform. A quasi one-dimensional version of the program has been developed which is based upon an approximate solution to the cavity acoustics problem. The three-dimensional program applies Greens's Function Discretization (GFD) to the solution for the acoustic mode shapes and frequency. GFD is a recently developed numerical method for finding fully three dimensional solutions for this class of problems. The analysis of complex motor geometries, previously a tedious and time consuming task, has also been greatly simplified through the development of a drawing package designed specifically to facilitate the specification of typical motor geometries. The combination of the drawing package, improved acoustic solutions, and new analysis, results in a tool which is capable of producing more accurate and meaningful predictions than have been possible in the past.

The Rocket Engine Advancement Program 2 (REAP2)

NASA Technical Reports Server (NTRS)

Harper, Brent (Technical Monitor); Hawk, Clark W.

2004-01-01

The Rocket Engine Advancement Program (REAP) 2 program is being conducted by a university propulsion consortium consisting of the University of Alabama in Huntsville, Penn State University, Purdue University, Tuskegee University and Auburn University. It has been created to bring their combined skills to bear on liquid rocket combustion stability and thrust chamber cooling. The research team involves well established and known researchers in the propulsion community. The cure team provides the knowledge base, research skills, and commitment to achieve an immediate and continuing impact on present and future propulsion issues. through integrated research teams composed of analysts, diagnosticians, and experimentalists working together in an integrated multi-disciplinary program. This paper provides an overview of the program, its objectives and technical approaches. Research on combustion instability and thrust chamber cooling are being accomplished

High-Speed Turbulent Reacting Flows: Intrinsic Flame Instability and its Effects on the Turbulent Cascade

NASA Astrophysics Data System (ADS)

Poludnenko, Alexei

2016-11-01

Turbulent reacting flows are pervasive both in our daily lives on Earth and in the Universe. They power modern society being at the heart of many energy generation and propulsion systems, such as gas turbines, internal combustion and jet engines. On astronomical scales, thermonuclear turbulent flames are the driver of some of the most powerful explosions in the Universe, knows as Type Ia supernovae. Despite this ubiquity in Nature, turbulent reacting flows still pose a number of fundamental questions often exhibiting surprising and unexpected behavior. In this talk, we will discuss several such phenomena observed in direct numerical simulations of high-speed, premixed, turbulent flames. We show that turbulent flames in certain regimes are intrinsically unstable even in the absence of the surrounding combustor walls or obstacles, which can support the thermoacoustic feedback. Such instability can fundamentally change the structure and dynamics of the turbulent cascade, resulting in a significant (and anisotropic) redistribution of kinetic energy from small to large scales. In particular, three effects are observed. 1) The turbulent burning velocity can develop pulsations with significant peak-to-peak amplitudes. 2) Unstable burning can result in pressure build-up and the formation of pressure waves or shocks when the flame speed approaches or exceeds the speed of a Chapman-Jouguet deflagration. 3) Coupling of pressure and density gradients across the flame can lead to the anisotropic generation of turbulence inside the flame volume and flame acceleration. We extend our earlier analysis, which relied on a simplified single-step reaction model, by demonstrating existence of these effects in realistic chemical flames (hydrogen and methane) and in thermonuclear flames in degenerate, relativistic plasmas found in stellar interiors. Finally, we discuss the implications of these results for subgrid-scale LES combustion models. This work was supported by the Air Force Office of Scientific Research (AFOSR) under Award No. F4FGA06055G001, and the Department of Defense (DoD) High Performance Computing Modernization Program (HPCMP) under a Frontier project award.

Numerical approaches to combustion modeling. Progress in Astronautics and Aeronautics. Vol. 135

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

Oran, E.S.; Boris, J.P.

1991-01-01

Various papers on numerical approaches to combustion modeling are presented. The topics addressed include; ab initio quantum chemistry for combustion; rate coefficient calculations for combustion modeling; numerical modeling of combustion of complex hydrocarbons; combustion kinetics and sensitivity analysis computations; reduction of chemical reaction models; length scales in laminar and turbulent flames; numerical modeling of laminar diffusion flames; laminar flames in premixed gases; spectral simulations of turbulent reacting flows; vortex simulation of reacting shear flow; combustion modeling using PDF methods. Also considered are: supersonic reacting internal flow fields; studies of detonation initiation, propagation, and quenching; numerical modeling of heterogeneous detonations, deflagration-to-detonationmore » transition to reactive granular materials; toward a microscopic theory of detonations in energetic crystals; overview of spray modeling; liquid drop behavior in dense and dilute clusters; spray combustion in idealized configurations: parallel drop streams; comparisons of deterministic and stochastic computations of drop collisions in dense sprays; ignition and flame spread across solid fuels; numerical study of pulse combustor dynamics; mathematical modeling of enclosure fires; nuclear systems.« less

A review of acoustic dampers applied to combustion chambers in aerospace industry

NASA Astrophysics Data System (ADS)

Zhao, Dan; Li, X. Y.

2015-04-01

In engine combustion systems such as rockets, aero-engines and gas turbines, pressure fluctuations are always present, even during normal operation. One of design prerequisites for the engine combustors is stable operation, since large-amplitude self-sustained pressure fluctuations (also known as combustion instability) have the potential to cause serious structural damage and catastrophic engine failure. To dampen pressure fluctuations and to reduce noise, acoustic dampers are widely applied as a passive control means to stabilize combustion/engine systems. However, they cannot respond to the dynamic changes of operating conditions and tend to be effective over certain narrow range of frequencies. To maintain their optimum damping performance over a broad frequency range, extensive researches have been conducted during the past four decades. The present work is to summarize the status, challenges and progress of implementing such acoustic dampers on engine systems. The damping effect and mechanism of various acoustic dampers, such as Helmholtz resonators, perforated liners, baffles, half- and quarter-wave tube are introduced first. A summary of numerical, experimental and theoretical studies are then presented to review the progress made so far. Finally, as an alternative means, ';tunable acoustic dampers' are discussed. Potential, challenges and issues associated with the dampers practical implementation are highlighted.

Deposition and material response from Mach 0.3 burner rig combustion of SRC 2 fuels

NASA Technical Reports Server (NTRS)

Santoro, G. J.; Kohl, F. J.; Stearns, C. A.; Fryburg, G. C.; Johnson, J. R.

1980-01-01

Collectors at 1173K (900 C) were exposed to the combustion products of a Mach 0.3 burner rig fueled with various industrial turbine liquid fuels from solvent refined coals. Four fuels were employed: a naphtha, a light oil, a wash solvent and a mid-heavy distillate blend. The response of four superalloys (IN-100, U 700, IN 792 and M-509) to exposure to the combustion gases from the SRC-2 naphtha and resultant deposits was also determined. The SRC-2 fuel analysis and insights obtained during the combustion experience are discussed. Particular problems encountered were fuel instability and reactions of the fuel with hardware components. The major metallic elements which contributed to the deposits were copper, iron, chromium, calcium, aluminum, nickel, silicon, titanium, zinc, and sodium. The deposits were found to be mainly metal oxides. An equilibrium thermodynamic analysis was employed to predict the chemical composition of the deposits. The agreement between the predicted and observed compounds was excellent. No hot corrosion was observed. This was expected because the deposits contained very little sodium or potassium and consisted mainly of the unreactive oxides. However, the amounts of deposits formed indicated that fouling is a potential problem with the use of these fuels.

Vortex pairing and reverse cascade in a simulated two-dimensional rocket motor-like flow field

NASA Astrophysics Data System (ADS)

Chakravarthy, Kalyana; Chakraborty, Debasis

2017-07-01

Two-dimensional large eddy simulation of a flow experiment intended for studying and understanding transition and parietal vortex shedding has brought to light some interesting features that have never been seen in previous similar simulations and have implications for future computational work on combustion instabilities in rocket motors. The frequency spectrum of pressure at head end shows a peak at the expected value associated with parietal vortex shedding but an additional peak at half this frequency emerges at downstream location. Using vorticity spectra at various distances away from the wall, it is shown that the frequency halving is due to vortex pairing as hypothesized by Dunlap et al. ["Internal flow field studies in a simulated cylindrical port rocket chamber," J. Propul. Power 6(6), 690-704 (1990)] for a similar experiment. As the flow transitions to turbulence towards the nozzle end, inertial range with Kolmogorov scaling becomes evident in the velocity spectrum. Given that the simulation is two-dimensional, such a scaling could be associated with a reverse energy cascade as per Kraichnan-Leith-Bachelor theory. By filtering the simulated flow field and identifying where the energy backscatters into the filtered scales, the regions with a reverse cascade are identified. The implications of this finding on combustion modeling are discussed.

Combustion Sensors: Gas Turbine Applications

NASA Technical Reports Server (NTRS)

Human, Mel

2002-01-01

This report documents efforts to survey the current research directions in sensor technology for gas turbine systems. The work is driven by the current and future requirements on system performance and optimization. Accurate real time measurements of velocities, pressure, temperatures, and species concentrations will be required for objectives such as combustion instability attenuation, pollutant reduction, engine health management, exhaust profile control via active control, etc. Changing combustor conditions - engine aging, flow path slagging, or rapid maneuvering - will require adaptive responses; the effectiveness of such will be only as good as the dynamic information available for processing. All of these issues point toward the importance of continued sensor development. For adequate control of the combustion process, sensor data must include information about the above mentioned quantities along with equivalence ratios and radical concentrations, and also include both temporal and spatial velocity resolution. Ultimately these devices must transfer from the laboratory to field installations, and thus must become low weight and cost, reliable and maintainable. A primary conclusion from this study is that the optics-based sensor science will be the primary diagnostic in future gas turbine technologies.

38th JANNAF Combustion Subcommittee Meeting. Volume 1

NASA Technical Reports Server (NTRS)

Fry, Ronald S. (Editor); Eggleston, Debra S. (Editor); Gannaway, Mary T. (Editor)

2002-01-01

This volume, the first of two volumes, is a collection of 55 unclassified/unlimited-distribution papers which were presented at the Joint Army-Navy-NASA-Air Force (JANNAF) 38th Combustion Subcommittee (CS), 26 th Airbreathing Propulsion Subcommittee (APS), 20th Propulsion Systems Hazards Subcommittee (PSHS), and 21 Modeling and Simulation Subcommittee. The meeting was held 8-12 April 2002 at the Bayside Inn at The Sandestin Golf & Beach Resort and Eglin Air Force Base, Destin, Florida. Topics cover five major technology areas including: 1) Combustion - Propellant Combustion, Ingredient Kinetics, Metal Combustion, Decomposition Processes and Material Characterization, Rocket Motor Combustion, and Liquid & Hybrid Combustion; 2) Liquid Rocket Engines - Low Cost Hydrocarbon Liquid Rocket Engines, Liquid Propulsion Turbines, Liquid Propulsion Pumps, and Staged Combustion Injector Technology; 3) Modeling & Simulation - Development of Multi- Disciplinary RBCC Modeling, Gun Modeling, and Computational Modeling for Liquid Propellant Combustion; 4) Guns Gun Propelling Charge Design, and ETC Gun Propulsion; and 5) Airbreathing - Scramjet an Ramjet- S&T Program Overviews.

Computational Combustion

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

Westbrook, C K; Mizobuchi, Y; Poinsot, T J

2004-08-26

Progress in the field of computational combustion over the past 50 years is reviewed. Particular attention is given to those classes of models that are common to most system modeling efforts, including fluid dynamics, chemical kinetics, liquid sprays, and turbulent flame models. The developments in combustion modeling are placed into the time-dependent context of the accompanying exponential growth in computer capabilities and Moore's Law. Superimposed on this steady growth, the occasional sudden advances in modeling capabilities are identified and their impacts are discussed. Integration of submodels into system models for spark ignition, diesel and homogeneous charge, compression ignition engines, surfacemore » and catalytic combustion, pulse combustion, and detonations are described. Finally, the current state of combustion modeling is illustrated by descriptions of a very large jet lifted 3D turbulent hydrogen flame with direct numerical simulation and 3D large eddy simulations of practical gas burner combustion devices.« less

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

NASA Astrophysics Data System (ADS)

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

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

Prediction of Combustion Instability with Detailed Chemical Kinetics

DTIC Science & Technology

2014-12-01

global reaction where the fuel and oxidizer react to form water and carbon dioxide . The production of carbon monoxide is a known intermediate step in... radially inward on a sleeve which turns the flow in the axial direction with minimal swirl. The oxidizer is decomposed hydrogen peroxide, which is...unburnt mixture in the vicinity of the recirculating hot products. This also causes the hot gas near the back step to move radially inwards, close to the

Analytical Prediction of Motor Component Vibrations Driven by Acoustic Combustion Instability

DTIC Science & Technology

1976-02-01

27"V 1Sy 1 2 oiihedr41 Symmetry .. .. . ., . . C-28 3 SPCD Bulk Data Card Format ......... . . .. .- 29 4 CYJOIN Bulk Data Card Format...analysis, the loads, the values of enforced displacements, and the temperatures may vary from element to element. The SPCD bulk data card (Figure 3) is...Static loads for vech suhc’: -e are spcified with LOAD, ’TEMPERATURE (LOAD), or DE .-I, oiectic.-,!•. Enforced deformations may be specified on SPCD

Experimental investigation of the limits of ethanol combustion in the boundary layer behind an obstacle

NASA Astrophysics Data System (ADS)

Boyarshinov, B. F.

2018-01-01

Experimental data on the flow structure and mass transfer near the boundaries of the region existence of the laminar and turbulent boundary layers with combustion are considered. These data include the results of in-vestigation on reacting flow stability at mixed convection, mass transfer during ethanol evaporation "on the floor" and "on the ceiling", when the flame surface curves to form the large-scale cellular structures. It is shown with the help of the PIV equipment that when Rayleigh-Taylor instability manifests, the mushroom-like structures are formed, where the motion from the flame front to the wall and back alternates. The cellular flame exists in a narrow range of velocities from 0.55 to 0.65 m/s, and mass transfer is three times higher than its level in the standard laminar boundary layer.

Numerical Simulation and Chaotic Analysis of an Aluminum Holding Furnace

NASA Astrophysics Data System (ADS)

Wang, Ji-min; Zhou, Yuan-yuan; Lan, Shen; Chen, Tao; Li, Jie; Yan, Hong-jie; Zhou, Jie-min; Tian, Rui-jiao; Tu, Yan-wu; Li, Wen-ke

2014-12-01

To achieve high heat efficiency, low pollutant emission and homogeneous melt temperature during thermal process of secondary aluminum, taking into account the features of aluminum alloying process, a CFD process model was developed and integrated with heat load and aluminum temperature control model. This paper presented numerical simulation of aluminum holding furnaces using the customized code based on FLUENT packages. Thermal behaviors of aluminum holding furnaces were investigated by probing into main physical fields such as flue gas temperature, velocity, and concentration, and combustion instability of aluminum holding process was represented by chaos theory. The results show that aluminum temperature uniform coefficient firstly decreases during heating phase, then increases and reduces alternately during holding phase, lastly rises during standing phase. Correlation dimension drops with fuel velocity. Maximal Lyapunov exponent reaches to a maximum when air-fuel ratio is close to 1. It would be a clear comprehension about each phase of aluminum holding furnaces to find new technology, retrofit furnace design, and optimize parameters combination.

DWPF Melter Off-Gas Flammability Assessment for Sludge Batch 9

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

Choi, A. S.

2016-07-11

The slurry feed to the Defense Waste Processing Facility (DWPF) melter contains several organic carbon species that decompose in the cold cap and produce flammable gases that could accumulate in the off-gas system and create potential flammability hazard. To mitigate such a hazard, DWPF has implemented a strategy to impose the Technical Safety Requirement (TSR) limits on all key operating variables affecting off-gas flammability and operate the melter within those limits using both hardwired/software interlocks and administrative controls. The operating variables that are currently being controlled include; (1) total organic carbon (TOC), (2) air purges for combustion and dilution, (3)more » melter vapor space temperature, and (4) feed rate. The safety basis limits for these operating variables are determined using two computer models, 4-stage cold cap and Melter Off-Gas (MOG) dynamics models, under the baseline upset scenario - a surge in off-gas flow due to the inherent cold cap instabilities in the slurry-fed melter.« less

Laser-based investigations in gas turbine model combustors

NASA Astrophysics Data System (ADS)

Meier, W.; Boxx, I.; Stöhr, M.; Carter, C. D.

2010-10-01

Dynamic processes in gas turbine (GT) combustors play a key role in flame stabilization and extinction, combustion instabilities and pollutant formation, and present a challenge for experimental as well as numerical investigations. These phenomena were investigated in two gas turbine model combustors for premixed and partially premixed CH4/air swirl flames at atmospheric pressure. Optical access through large quartz windows enabled the application of laser Raman scattering, planar laser-induced fluorescence (PLIF) of OH, particle image velocimetry (PIV) at repetition rates up to 10 kHz and the simultaneous application of OH PLIF and PIV at a repetition rate of 5 kHz. Effects of unmixedness and reaction progress in lean premixed GT flames were revealed and quantified by Raman scattering. In a thermo-acoustically unstable flame, the cyclic variation in mixture fraction and its role for the feedback mechanism of the instability are addressed. In a partially premixed oscillating swirl flame, the cyclic variations of the heat release and the flow field were characterized by chemiluminescence imaging and PIV, respectively. Using phase-correlated Raman scattering measurements, significant phase-dependent variations of the mixture fraction and fuel distributions were revealed. The flame structures and the shape of the reaction zones were visualized by planar imaging of OH distribution. The simultaneous OH PLIF/PIV high-speed measurements revealed the time history of the flow field-flame interaction and demonstrated the development of a local flame extinction event. Further, the influence of a precessing vortex core on the flame topology and its dynamics is discussed.

A Two-Zone Multigrid Model for SI Engine Combustion Simulation Using Detailed Chemistry

DOE PAGES

Ge, Hai-Wen; Juneja, Harmit; Shi, Yu; ...

2010-01-01

An efficient multigrid (MG) model was implemented for spark-ignited (SI) engine combustion modeling using detailed chemistry. The model is designed to be coupled with a level-set-G-equation model for flame propagation (GAMUT combustion model) for highly efficient engine simulation. The model was explored for a gasoline direct-injection SI engine with knocking combustion. The numerical results using the MG model were compared with the results of the original GAMUT combustion model. A simpler one-zone MG model was found to be unable to reproduce the results of the original GAMUT model. However, a two-zone MG model, which treats the burned and unburned regionsmore » separately, was found to provide much better accuracy and efficiency than the one-zone MG model. Without loss in accuracy, an order of magnitude speedup was achieved in terms of CPU and wall times. To reproduce the results of the original GAMUT combustion model, either a low searching level or a procedure to exclude high-temperature computational cells from the grouping should be applied to the unburned region, which was found to be more sensitive to the combustion model details.« less

A factor involved in efficient breakdown of supersonic streamwise vortices

NASA Astrophysics Data System (ADS)

Hiejima, Toshihiko

2015-03-01

Spatially developing processes in supersonic streamwise vortices were numerically simulated at Mach number 5.0. The vortex evolution largely depended on the azimuthal vorticity thickness of the vortices, which governs the negative helicity profile. Large vorticity thickness greatly enhanced the centrifugal instability, with consequent development of perturbations with competing wavenumbers outside the vortex core. During the transition process, supersonic streamwise vortices could generate large-scale spiral structures and a number of hairpin like vortices. Remarkably, the transition caused a dramatic increase in the total fluctuation energy of hypersonic flows, because the negative helicity profile destabilizes the flows due to helicity instability. Unstable growth might also relate to the correlation length between the axial and azimuthal vorticities of the streamwise vortices. The knowledge gained in this study is important for realizing effective fuel-oxidizer mixing in supersonic combustion engines.

A combustion model for studying the effects of ideal gas properties on jet noise

NASA Astrophysics Data System (ADS)

Jacobs, Jerin; Tinney, Charles

2016-11-01

A theoretical combustion model is developed to simulate the influence of ideal gas effects on various aeroacoustic parameters over a range of equivalence ratios. The motivation is to narrow the gap between laboratory and full-scale jet noise testing. The combustion model is used to model propane combustion in air and kerosene combustion in air. Gas properties from the combustion model are compared to real lab data acquired at the National Center for Physical Acoustics at the University of Mississippi as well as outputs from NASA's Chemical Equilibrium Analysis code. Different jet properties are then studied over a range of equivalence ratios and pressure ratios for propane combustion in air, kerosene combustion in air and heated air. The findings reveal negligible differences between the three constituents where the density and sound speed ratios are concerned. Albeit, the area ratio required for perfectly expanded flow is shown to be more sensitive to gas properties, relative to changes in the temperature ratio.

Oscillations of Accretion Disks in Cataclysmic Variable Stars

NASA Astrophysics Data System (ADS)

Osaki, Y.

2013-12-01

The disk instability model for the outbursts of dwarf novae is reviewed, with particular attention given to the superoutburst of SU UMa stars. Two intrinsic instabilities in accretion disks of dwarf novae are known; the thermal instability and the tidal instability. The thermal-tidal instability model (abbreviated the TTI model), which combines these two instabilities, was first proposed in 1989 by Osaki (1989) to explain the superoutburst phenomenon of SU UMa stars. Recent Kepler observations of one SU UMa star, V1504 Cyg, have dramatically demonstrated that the superoutburst phenomenon of the SU UMa stars is explained by the thermal-tidal instability model.

Combustion Fundamentals Research

NASA Technical Reports Server (NTRS)

1983-01-01

Increased emphasis is placed on fundamental and generic research at Lewis Research Center with less systems development efforts. This is especially true in combustion research, where the study of combustion fundamentals has grown significantly in order to better address the perceived long term technical needs of the aerospace industry. The main thrusts for this combustion fundamentals program area are as follows: analytical models of combustion processes, model verification experiments, fundamental combustion experiments, and advanced numeric techniques.

Assessment of Turbulence-Chemistry Interaction Models in the National Combustion Code (NCC) - Part I

NASA Technical Reports Server (NTRS)

Wey, Thomas Changju; Liu, Nan-suey

2011-01-01

This paper describes the implementations of the linear-eddy model (LEM) and an Eulerian FDF/PDF model in the National Combustion Code (NCC) for the simulation of turbulent combustion. The impacts of these two models, along with the so called laminar chemistry model, are then illustrated via the preliminary results from two combustion systems: a nine-element gas fueled combustor and a single-element liquid fueled combustor.

Combustion Instabilities in Liquid-Fuelled Propulsion Systems: Conference Proceedings of the Propulsion and Energetics Panel (72nd) B specialists Meeting Held in Bath (England) on 6-7 October 1988

DTIC Science & Technology

1989-04-01

Recommending effective ways for the member nations to use their research and development capabilities for the common benefit of the NATO community...is effectively swept under the rug and presented in global approximation. When the details of, say, droplet dynamics are treated, only numerical...applicable theory for passive control; development is always a costly trial and error proces; and the effectiveness of a particular design is

A study of pressure-based methodology for resonant flows in non-linear combustion instabilities

NASA Technical Reports Server (NTRS)

Yang, H. Q.; Pindera, M. Z.; Przekwas, A. J.; Tucker, K.

1992-01-01

This paper presents a systematic assessment of a large variety of spatial and temporal differencing schemes on nonstaggered grids by the pressure-based methods for the problems of fast transient flows. The observation from the present study is that for steady state flow problems, pressure-based methods can be very competitive with the density-based methods. For transient flow problems, pressure-based methods utilizing the same differencing scheme are less accurate, even though the wave speeds are correctly predicted.

Investigation of the Flame-Acoustic Wave Interaction during Axial Solid Rocket Instabilities

DTIC Science & Technology

1991-04-30

acoustic exergy by the mean flow was neglected as small with respect to the mean flow independent energy flux. The relative magnitudes of the terms in the...34Laser Rayleigh Thermometry in Turbulent Flames", 18th Symposium ( International ) on Combustion, 1980. 5. T. Chen, Ph.D. Thesis Proposal, G.I.T., 1989. 6...Cantrell, R. H. and Hart, R. W., "Interactions Between Sound and Flow in Acoustic Cavities: Mass, Momentum, and Energy Considerations," Journal of the

Diode Laser-Based Detection of Combustor Instabilities with Application to a Scramjet Engine

DTIC Science & Technology

2010-02-01

Exhibit, Reno, NV, January 9–12, 2006 . [16] G. Rieker, H. Li, X. Liu, et al ., Proc. Combust. Inst. 31 (2007) 3041–3049. [17] H. Li, G.B. Rieker, X...Cincinnati, OH, July 2006 . [24] M. Gruber, J. Donbar, K. Jackson, et al ., J. Propul. Power 17 (2001) 1296–1304. ...in inlet unstart, which causes a significant decrease in captured air massute. Published by Elsevier Inc. All rights reserved. 832 G.B. Rieker et al

Mixing in Shear Coaxial Jets with and without Acoustics (Briefing Charts)

DTIC Science & Technology

2012-05-21

and heat transfer fluctuations in a rocket engine – Irreparable damage can occur in əs • Combustion Instability caused a 4-yr delay in the...common choice for cryogenic liquid rocket engines • Interactions of transverse acoustics with injector’s own modes and mixing needs to be understood...Pr = 0.44 • LAR-thin , Pr = 0.44, J = 0.5 POM 2 POM 1 Average Snapshot Power Spectral Densities (PSD) of Temporal Coefficients of POMs 1 and 2

Lessons Learned in Solid Rocket Combustion Instability

DTIC Science & Technology

2006-11-14

Gary Flandro . Also I wish to thank James Crump and H.B. Mathes who provided guidance during my first ten years at China Lake. VIII. References 1 L...It is also a form of analysis to examine the acoustic boundary with flow normal to the surface. It is sometimes known as the " Flandro boundary layer...Tso, "Flow Turning Losses in Solid Rocket Motors," AFAL-TR-87-095, March 1988. 30 G.A. Flandro , "Solid Propellant Acoustic Admittance Corrections

Application of neural network in the study of combustion rate of natural gas/diesel dual fuel engine.

PubMed

Yan, Zhao-Da; Zhou, Chong-Guang; Su, Shi-Chuan; Liu, Zhen-Tao; Wang, Xi-Zhen

2003-01-01

In order to predict and improve the performance of natural gas/diesel dual fuel engine (DFE), a combustion rate model based on forward neural network was built to study the combustion process of the DFE. The effect of the operating parameters on combustion rate was also studied by means of this model. The study showed that the predicted results were good agreement with the experimental data. It was proved that the developed combustion rate model could be used to successfully predict and optimize the combustion process of dual fuel engine.

LOX/hydrocarbon rocket engine analytical design methodology development and validation. Volume 2: Appendices

NASA Technical Reports Server (NTRS)

Niiya, Karen E.; Walker, Richard E.; Pieper, Jerry L.; Nguyen, Thong V.

1993-01-01

This final report includes a discussion of the work accomplished during the period from Dec. 1988 through Nov. 1991. The objective of the program was to assemble existing performance and combustion stability models into a usable design methodology capable of designing and analyzing high-performance and stable LOX/hydrocarbon booster engines. The methodology was then used to design a validation engine. The capabilities and validity of the methodology were demonstrated using this engine in an extensive hot fire test program. The engine used LOX/RP-1 propellants and was tested over a range of mixture ratios, chamber pressures, and acoustic damping device configurations. This volume contains time domain and frequency domain stability plots which indicate the pressure perturbation amplitudes and frequencies from approximately 30 tests of a 50K thrust rocket engine using LOX/RP-1 propellants over a range of chamber pressures from 240 to 1750 psia with mixture ratios of from 1.2 to 7.5. The data is from test configurations which used both bitune and monotune acoustic cavities and from tests with no acoustic cavities. The engine had a length of 14 inches and a contraction ratio of 2.0 using a 7.68 inch diameter injector. The data was taken from both stable and unstable tests. All combustion instabilities were spontaneous in the first tangential mode. Although stability bombs were used and generated overpressures of approximately 20 percent, no tests were driven unstable by the bombs. The stability instrumentation included six high-frequency Kistler transducers in the combustion chamber, a high-frequency Kistler transducer in each propellant manifold, and tri-axial accelerometers. Performance data is presented, both characteristic velocity efficiencies and energy release efficiencies, for those tests of sufficient duration to record steady state values.

Advanced Chemical Modeling for Turbulent Combustion Simulations

DTIC Science & Technology

2012-05-03

premixed combustion. The chemistry work proposes a method for defining jet fuel surrogates, describes how different sub- mechanisms can be incorporated...Chemical Modeling For Turbulent Combustion Simulations Final Report submitted by: Heinz Pitsch (PI) Stanford University Mechanical Engineering Flow Physics...predict the combustion characteristics of fuel oxidation and pollutant emissions from engines . The relevant fuel chemistry must be accurately modeled

Symposium on Combustion /International/, 16th, Massachusetts Institute of Technology, Cambridge, Mass., August 15-20, 1976, Proceedings

NASA Technical Reports Server (NTRS)

1977-01-01

Aspects of combustion technology in power systems are considered, taking into account a combustion in large boilers, the control of over-all thermal efficiency of combustion heating systems, a comparison of mathematical models of the radiative behavior of a large-scale experimental furnace, a concentric multiannular swirl burner, and the effects of water introduction on diesel engine combustion and emissions. Attention is also given to combustion and related processes in energy production from coal, spray and droplet combustion, soot formation and growth, the kinetics of elementary reactions, flame structure and chemistry, propellant ignition and combustion, fire and explosion research, mathematical modeling, high output combustion systems, turbulent flames and combustion, and ignition, optical, and electrical properties.

Sub-grid scale combustion models for large eddy simulation of unsteady premixed flame propagation around obstacles.

PubMed

Di Sarli, Valeria; Di Benedetto, Almerinda; Russo, Gennaro

2010-08-15

In this work, an assessment of different sub-grid scale (sgs) combustion models proposed for large eddy simulation (LES) of steady turbulent premixed combustion (Colin et al., Phys. Fluids 12 (2000) 1843-1863; Flohr and Pitsch, Proc. CTR Summer Program, 2000, pp. 61-82; Kim and Menon, Combust. Sci. Technol. 160 (2000) 119-150; Charlette et al., Combust. Flame 131 (2002) 159-180; Pitsch and Duchamp de Lageneste, Proc. Combust. Inst. 29 (2002) 2001-2008) was performed to identify the model that best predicts unsteady flame propagation in gas explosions. Numerical results were compared to the experimental data by Patel et al. (Proc. Combust. Inst. 29 (2002) 1849-1854) for premixed deflagrating flame in a vented chamber in the presence of three sequential obstacles. It is found that all sgs combustion models are able to reproduce qualitatively the experiment in terms of step of flame acceleration and deceleration around each obstacle, and shape of the propagating flame. Without adjusting any constants and parameters, the sgs model by Charlette et al. also provides satisfactory quantitative predictions for flame speed and pressure peak. Conversely, the sgs combustion models other than Charlette et al. give correct predictions only after an ad hoc tuning of constants and parameters. Copyright 2010 Elsevier B.V. All rights reserved.

Modeling the combustion behavior of hazardous waste in a rotary kiln incinerator.

PubMed

Yang, Yongxiang; Pijnenborg, Marc J A; Reuter, Markus A; Verwoerd, Joep

2005-01-01

Hazardous wastes have complex physical forms and chemical compositions and are normally incinerated in rotary kilns for safe disposal and energy recovery. In the rotary kiln, the multifeed stream and wide variation of thermal, physical, and chemical properties of the wastes cause the incineration system to be highly heterogeneous, with severe temperature fluctuations and unsteady combustion chemistry. Incomplete combustion is often the consequence, and the process is difficult to control. In this article, modeling of the waste combustion is described by using computational fluid dynamics (CFD). Through CFD simulation, gas flow and mixing, turbulent combustion, and heat transfer inside the incinerator were predicted and visualized. As the first step, the waste in various forms was modeled to a hydrocarbon-based virtual fuel mixture. The combustion of the simplified waste was then simulated with a seven-gas combustion model within a CFD framework. Comparison was made with previous global three-gas combustion model with which no chemical behavior can be derived. The distribution of temperature and chemical species has been investigated. The waste combustion model was validated with temperature measurements. Various operating conditions and the influence on the incineration performance were then simulated. Through this research, a better process understanding and potential optimization of the design were attained.

Research in Supercritical Fuel Properties and Combustion Modeling

DTIC Science & Technology

2015-09-18

AFRL-AFOSR-VA-TR-2015-0296 RESEARCH IN SUPERCRITICAL FUEL PROPERTIES AND COMBUSTION MODELING Gregory Faris SRI INTERNATIONAL MENLO PARK CA Final...Properties and Combustion Modeling 5a. CONTRACT NUMBER 5b. GRANT NUMBER FA9550-13-1-0177 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Gregory W...carbon atom species for combustion modeling and optimization. On the stimulated scattering task, we have tested new methods for rapidly scanning

Simulation of air pollution due to marine engines

NASA Astrophysics Data System (ADS)

Stan, L. C.

2017-08-01

This paperwork tried to simulate the combustion inside the marine engines using the newest computer methods and technologies with the result of a diverse and rich palette of solutions, extremely useful for the study and prediction of complex phenomena of the fuel combustion. The paperwork is contributing to the theoretical systematization of the area of interest bringing into attention a thoroughly inventory of the thermodynamic description of the phenomena which take place in the combustion process into the marine diesel engines; to the in depth multidimensional combustion models description along with the interdisciplinary phenomenology taking place in the combustion models; to the FEA (Finite Elements Method) modelling for the combustion chemistry in the nonpremixed mixtures approach considered too; the CFD (Computational Fluid Dynamics) model was issued for the combustion area and a rich palette of results interesting for any researcher of the process.

Combustion behaviors of GO2/GH2 swirl-coaxial injector using non-intrusive optical diagnostics

NASA Astrophysics Data System (ADS)

GuoBiao, Cai; Jian, Dai; Yang, Zhang; NanJia, Yu

2016-06-01

This research evaluates the combustion behaviors of a single-element, swirl-coaxial injector in an atmospheric combustion chamber with gaseous oxygen and gaseous hydrogen (GO2/GH2) as the propellants. A brief simulated flow field schematic comparison between a shear-coaxial injector and the swirl-coaxial injector reveals the distribution characteristics of the temperature field and streamline patterns. Advanced optical diagnostics, i.e., OH planar laser-induced fluorescence and high-speed imaging, are simultaneously employed to determine the OH radical spatial distribution and flame fluctuations, respectively. The present study focuses on the flame structures under varying O/F mixing ratios and center oxygen swirl intensities. The combined use of several image-processing methods aimed at OH instantaneous images, including time-averaged, root-mean-square, and gradient transformation, provides detailed information regarding the distribution of the flow field. The results indicate that the shear layers anchored on the oxygen injector lip are the main zones of chemical heat release and that the O/F mixing ratio significantly affects the flame shape. Furthermore, with high-speed imaging, an intuitionistic ignition process and several consecutive steady-state images reveal that lean conditions make it easy to drive the combustion instabilities and that the center swirl intensity has a moderate influence on the flame oscillation strength. The results of this study provide a visualized analysis for future optimal swirl-coaxial injector designs.

1990 Fuel oil utilization workshop

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

McDonald, B.L.; Lange, H.B.; Miller, M.N.

1992-01-01

Following a 1983 EPRI-sponsored workshop on utility boiler problems (EPRI report AP-3753), the Institute has responded to the need for better information on fuel utilization by sponsoring annual utility-focused workshops. This workshop is the sixth in a series of annual events designed to address this need. The objective was to provide utility personnel with an opportunity to exchange information on residual oil use in fossil steam plants. Participants at the 1990 workshop, held in Arlington, Virginia, October 31-November 1, 1990, included 37 representatives from 19 electric utilities, including representatives from Mexico, Canada, and Spain, as well as the Institute demore » Investigaciones Electricas in Mexico. The workshop comprised formal presentations followed by question-and-answer sessions and three 2-hour discussion group sessions. Topics included a water/oil emulsion test summary, a NO{sub x} reduction program, particulate and unburned carbon emissions reductions from oil-fired boilers using combustion promoters, a utility perspective on oil spills, and size distribution and opacity of particulate matter emissions from combustion of residual fuel oils. In addition, participants discussed the development of a coke formation index, instability and compatibility of residual fuel oils, the clean combustion of heavy liquid fuels, toxic air emissions from the combustion of residual fuel oils, H{sub 2}S release from residual fuel oils, and increased reliability of superheater and reheater tubes and headers by optimization of steam-side and gas-side temperatures.« less

The present state and future directions of PDF methods

NASA Technical Reports Server (NTRS)

Pope, S. B.

1992-01-01

The objectives of the workshop are presented in viewgraph format, as is this entire article. The objectives are to discuss the present status and the future direction of various levels of engineering turbulence modeling related to Computational Fluid Dynamics (CFD) computations for propulsion; to assure that combustion is an essential part of propulsion; and to discuss Probability Density Function (PDF) methods for turbulent combustion. Essential to the integration of turbulent combustion models is the development of turbulent model, chemical kinetics, and numerical method. Some turbulent combustion models typically used in industry are the k-epsilon turbulent model, the equilibrium/mixing limited combustion, and the finite volume codes.

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).

PREMIXED FLAME PROPAGATION AND MORPHOLOGY IN A CONSTANT VOLUME COMBUSTION CHAMBER

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

Hariharan, A; Wichman, IS

2014-06-04

This work presents an experimental and numerical investigation of premixed flame propagation in a constant volume rectangular channel with an aspect ratio of six (6) that serves as a combustion chamber. Ignition is followed by an accelerating cusped finger-shaped flame-front. A deceleration of the flame is followed by the formation of a "tulip"-shaped flame-front. Eventually, the flame is extinguished when it collides with the cold wall on the opposite channel end. Numerical computations are performed to understand the influence of pressure waves, instabilities, and flow field effects causing changes to the flame structure and morphology. The transient 2D numerical simulationmore » results are compared with transient 3D experimental results. Issues discussed are the appearance of oscillatory motions along the flame front and the influences of gravity on flame structure. An explanation is provided for the formation of the "tulip" shape of the premixed flame front.« less

Characterization of flow disturbances in a coal fired combustion flow train

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

Winkleman, B.C.; Giel, T.V.; Lineberry, J.T.

1990-01-01

Audible rumbles are known to accompany operation of the CFFF low mass flow train and visual/aural observations indicate simultaneous dropouts in the diffuser light emission. Three hypotheses, coal flow disturbances, combustion instabilities, and slag entrainment into the flow, are presented as possible causes of the rumbles. Wideband instrumentation including line reversals, luminosities, and dynamic pressures were used to investigate the rumble phenomena. The observational evidence implies that briefly before the rumble sound, the vitation heater pressure rises and a cold opaque structure moves from upstream to downstream through the aerodynamic duct, diffuser, and radiant furnace. Steady state thermodynamic analysis ofmore » the flow train at conditions corresponding to measured rumble phenomena are presented. It is concluded that a dispersed structure of slag particles entrained from the combustor is the most viable hypothesis. 8 refs., 23 figs., 2 tabs.« less

Review on pressure swirl injector in liquid rocket engine

NASA Astrophysics Data System (ADS)

Kang, Zhongtao; Wang, Zhen-guo; Li, Qinglian; Cheng, Peng

2018-04-01

The pressure swirl injector with tangential inlet ports is widely used in liquid rocket engine. Commonly, this type of pressure swirl injector consists of tangential inlet ports, a swirl chamber, a converging spin chamber, and a discharge orifice. The atomization of the liquid propellants includes the formation of liquid film, primary breakup and secondary atomization. And the back pressure and temperature in the combustion chamber could have great influence on the atomization of the injector. What's more, when the combustion instability occurs, the pressure oscillation could further affects the atomization process. This paper reviewed the primary atomization and the performance of the pressure swirl injector, which include the formation of the conical liquid film, the breakup and atomization characteristics of the conical liquid film, the effects of the rocket engine environment, and the response of the injector and atomization on the pressure oscillation.

Thermal Model of the Promoted Combustion Test

NASA Technical Reports Server (NTRS)

Jones, Peter D.

1996-01-01

Flammability of metals in high pressure, pure oxygen environments, such as rocket engine turbopumps, is commonly evaluated using the Promoted Combustion Test (PCT). The PCT emphasizes the ability of an ignited material to sustain combustion, as opposed to evaluating the sample's propensity to ignite in the first place. A common arrangement is a rod of the sample material hanging in a chamber in which a high pressure, pure oxygen environment is maintained. An igniter of some energetically combusting material is fixed to the bottom of the rod and fired. This initiates combustion, and the sample burns and melts at its bottom tip. A ball of molten material forms, and this ball detaches when it grows too large to be supported by surface tension with the rod. In materials which do not sustain combustion, the combustion then extinguishes. In materials which do sustain combustion, combustion re-initiates from molten residue left on the bottom of the rod, and the melt ball burns and grows until it detaches again. The purpose of this work is development of a PCT thermal simulation model, detailing phase change, melt detachment, and the several heat transfer modes. Combustion is modeled by a summary rate equation, whose parameters are identified by comparison to PCT results. The sensitivity of PCT results to various physical and geometrical parameters is evaluated. The identified combustion parameters may be used in design of new PCT arrangements, as might be used for flammability assessment in flow-dominated environments. The Haynes 214 nickel-based superalloy, whose PCT results are applied here, burns heterogeneously (fuel and oxidizer are of different phases; combustion takes place on the fuel surface). Heterogeneous combustion is not well understood. (In homogeneous combustion, the metal vaporizes, and combustion takes place in an analytically treatable cloud above the surface). Thermal modeling in heterogeneous combustion settings provides a means for linking test results more directly to detailed combustion mechanics, leading to improved data analysis, and improved understanding of heterogeneous combustion phenomena.

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

Hicks, E. P.; Rosner, R., E-mail: eph2001@columbia.edu

In this paper, we provide support for the Rayleigh-Taylor-(RT)-based subgrid model used in full-star simulations of deflagrations in Type Ia supernovae explosions. We use the results of a parameter study of two-dimensional direct numerical simulations of an RT unstable model flame to distinguish between the two main types of subgrid models (RT or turbulence dominated) in the flamelet regime. First, we give scalings for the turbulent flame speed, the Reynolds number, the viscous scale, and the size of the burning region as the non-dimensional gravity (G) is varied. The flame speed is well predicted by an RT-based flame speed model.more » Next, the above scalings are used to calculate the Karlovitz number (Ka) and to discuss appropriate combustion regimes. No transition to thin reaction zones is seen at Ka = 1, although such a transition is expected by turbulence-dominated subgrid models. Finally, we confirm a basic physical premise of the RT subgrid model, namely, that the flame is fractal, and thus self-similar. By modeling the turbulent flame speed, we demonstrate that it is affected more by large-scale RT stretching than by small-scale turbulent wrinkling. In this way, the RT instability controls the flame directly from the large scales. Overall, these results support the RT subgrid model.« less

Fundamentals of Gas Turbine combustion

NASA Technical Reports Server (NTRS)

Gerstein, M.

1979-01-01

Combustion problems and research recommendations are discussed in the areas of atomization and vaporization, combustion chemistry, combustion dynamics, and combustion modelling. The recommendations considered of highest priority in these areas are presented.

High-Temperature Piezoelectric Ceramic Developed

NASA Technical Reports Server (NTRS)

Sayir, Ali; Farmer, Serene C.; Dynys, Frederick W.

2005-01-01

Active combustion control of spatial and temporal variations in the local fuel-to-air ratio is of considerable interest for suppressing combustion instabilities in lean gas turbine combustors and, thereby, achieving lower NOx levels. The actuator for fuel modulation in gas turbine combustors must meet several requirements: (1) bandwidth capability of 1000 Hz, (2) operating temperature compatible with the fuel temperature, which is in the vicinity of 400 F, (3) stroke of approximately 4 mils (100 m), and (4) force of 300 lb-force. Piezoelectric actuators offer the fastest response time (microsecond time constants) and can generate forces in excess of 2000 lb-force. The state-of-the-art piezoceramic material in industry today is Pb(Zr,Ti)O3, called PZT. This class of piezoelectric ceramic is currently used in diesel fuel injectors and in the development of high-response fuel modulation valves. PZT materials are generally limited to operating temperatures of 250 F, which is 150 F lower than the desired operating temperature for gas turbine combustor fuel-modulation injection valves. Thus, there is a clear need to increase the operating temperature range of piezoceramic devices for active combustion control in gas turbine engines.

Dynamical behavior of lean swirling premixed flame generated by change in gravitational orientation

NASA Astrophysics Data System (ADS)

Gotoda, Hiroshi; Miyano, Takaya; Shepherd, Ian

2010-11-01

The dynamic behavior of flame front instability in lean swirling premixed flame generated by the effect of gravitational orientation has been experimentally investigated in this work. When the gravitational direction is changed relative to the flame front, i.e., in inverted gravity, an unstably fluctuating flame (unstable flame) is formed in a limited domain of equivalence ratio and swirl number (Gotoda. H et al., Physical Review E, vol. 81, 026211, 2010). The time history of flame front fluctuations show that in the buoyancy-dominated region, chaotic irregular fluctuation with low frequencies is superimposed on the dominant periodic oscillation of the unstable flame. This periodic oscillation is produced by unstable large-scale vortex motion in combustion products generated by a change in the buoyancy/swirl interaction due to the inversion of gravitational orientation. As a result, the dynamic behavior of the unstable flame becomes low-dimensional deterministic chaos. Its dynamics maintains low-dimensional deterministic chaos even in the momentum-dominated region, in which vortex breakdown in the combustion products clearly occurs. These results were clearly demonstrated by the use of nonlinear time series analysis based on chaos theory, which has not been widely applied to the investigation of combustion phenomena.

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).

Numerical Simulation on Hydrodynamics and Combustion in a Circulating Fluidized Bed under O2/CO2 and Air Atmospheres

NASA Astrophysics Data System (ADS)

Zhou, W.; Zhao, C. S.; Duan, L. B.; Qu, C. R.; Lu, J. Y.; Chen, X. P.

Oxy-fuel circulating fluidized bed (CFB) combustion technology is in the stage of initial development for carbon capture and storage (CCS). Numerical simulation is helpful to better understanding the combustion process and will be significant for CFB scale-up. In this paper, a computational fluid dynamics (CFD) model was employed to simulate the hydrodynamics of gas-solid flow in a CFB riser based on the Eulerian-Granular multiphase model. The cold model predicted the main features of the complex gas-solid flow, including the cluster formation of the solid phase along the walls, the flow structure of up-flow in the core and downward flow in the annular region. Furthermore, coal devolatilization, char combustion and heat transfer were considered by coupling semi-empirical sub-models with CFD model to establish a comprehensive model. The gas compositions and temperature profiles were predicted and the outflow gas fractions are validated with the experimental data in air combustion. With the experimentally validated model being applied, the concentration and temperature distributions in O2/CO2 combustion were predicted. The model is useful for the further development of a comprehensive model including more sub-models, such as pollutant emissions, and better understanding the combustion process in furnace.

A study of the current group evaporation/combustion theories

NASA Technical Reports Server (NTRS)

Shen, Hayley H.

1990-01-01

Liquid fuel combustion can be greatly enhanced by disintegrating the liquid fuel into droplets, an effect achieved by various configurations. A number of experiments carried out in the seventies showed that combustion of droplet arrays and sprays do not form individual flames. Moreover, the rate of burning in spray combustion greatly deviates from that of the single combustion rate. Such observations naturally challenge its applicability to spray combustion. A number of mathematical models were developed to evaluate 'group combustion' and the related 'group evaporation' phenomena. This study investigates the similarity and difference of these models and their applicability to spray combustion. Future work that should be carried out in this area is indicated.

Longitudinal Tobacco Use Transitions Among Adolescents and Young Adults: 2014-2016.

PubMed

Hair, Elizabeth C; Romberg, Alexa R; Niaura, Raymond; Abrams, David B; Bennett, Morgane A; Xiao, Haijun; Rath, Jessica M; Pitzer, Lindsay; Vallone, Donna

2018-02-13

Among youth, the frequency and prevalence of using more than one tobacco (small cigar, cigarette, and hookah) or nicotine-containing product (e-cigarettes-ENDS) are changing. These shifts pose challenges for regulation, intervention, and prevention campaigns because of scant longitudinal data on the stability of use patterns in this changing product landscape. A nationally representative longitudinal survey of 15- to 21-year olds (n = 15,275) was used to describe transitions between never use, noncurrent use, and past 30-day use of combustible tobacco, e-cigarettes (ENDS), and dual use of both kinds of products. A multistate model was fit to observations collected every 6 months across 2.5 years to estimate the probability of transitions between states (TPs), the average time in state (sojourn time), and the effect of age on transitions. Current state strongly predicted future state over time intervals of 1 year or less, but only weakly predicted future state at longer intervals: TP to noncurrent use was higher for ENDS-only than combustible-only users over a 6-month interval but was similar for both groups over a 2-year interval. Sojourn time was significantly longer for combustible-only (0.52 years) and dual use (0.55 years) than ENDS-only use (0.27 years); older youth were more likely than younger youth to stay combustible tobacco users or noncurrent users. The dynamics of transitions between combustible tobacco products and ENDS in a population of youth and young adults suggest that policy and prevention efforts must consider the frequent changes and instability over a 1-year or less time period in use patterns among young people. The study addresses an urgent need in public health for timely information on how youth and young adults use tobacco and nicotine products. We found that youth, particularly adolescents, moved frequently between using ENDS and combustible tobacco products either alone or together. Importantly, the utility of current-use states for predicting future use states declined for time horizons longer than 1 year. Our results demonstrate a need for caution in interpreting product transitions. Longitudinal data with frequent observations and coverage of a wide range of possible product types is required to fully characterize usage patterns in youth. © The Author 2018. Published by Oxford University Press on behalf of the Society for Research on Nicotine and Tobacco. All rights reserved. For permissions, please e-mail: journals.permissions@oup.com.

Modeling complex chemical effects in turbulent nonpremixed combustion

NASA Technical Reports Server (NTRS)

Smith, Nigel S. A.

1995-01-01

Virtually all of the energy derived from the consumption of combustibles occurs in systems which utilize turbulent fluid motion. Since combustion is largely related to the mixing of fluids and mixing processes are orders of magnitude more rapid when enhanced by turbulent motion, efficiency criteria dictate that chemically powered devices necessarily involve fluid turbulence. Where combustion occurs concurrently with mixing at an interface between two reactive fluid bodies, this mode of combustion is called nonpremixed combustion. This is distinct from premixed combustion where flame-fronts propagate into a homogeneous mixture of reactants. These two modes are limiting cases in the range of temporal lag between mixing of reactants and the onset of reaction. Nonpremixed combustion occurs where this lag tends to zero, while premixed combustion occurs where this lag tends to infinity. Many combustion processes are hybrids of these two extremes with finite non-zero lag times. Turbulent nonpremixed combustion is important from a practical standpoint because it occurs in gas fired boilers, furnaces, waste incinerators, diesel engines, gas turbine combustors, and afterburners etc. To a large extent, past development of these practical systems involved an empirical methodology. Presently, efficiency standards and emission regulations are being further tightened (Correa 1993), and empiricism has had to give way to more fundamental research in order to understand and effectively model practical combustion processes (Pope 1991). A key element in effective modeling of turbulent combustion is making use of a sufficiently detailed chemical kinetic mechanism. The prediction of pollutant emission such as oxides of nitrogen (NO(x)) and sulphur (SO(x)) unburned hydrocarbons, and particulates demands the use of detailed chemical mechanisms. It is essential that practical models for turbulent nonpremixed combustion are capable of handling large numbers of 'stiff' chemical species equations.

Investigation of HCCI Combustion of Diethyl Ether and Ethanol Mixtures Using Carbon 14 Tracing and Numerical Simulations

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

Mack, J H; Dibble, R W; Buchholz, B A

2004-01-16

Despite the rapid combustion typically experienced in Homogeneous Charge Compression Ignition (HCCI), components in fuel mixtures do not ignite in unison or burn equally. In our experiments and modeling of blends of diethyl ether (DEE) and ethanol (EtOH), the DEE led combustion and proceeded further toward completion, as indicated by {sup 14}C isotope tracing. A numerical model of HCCI combustion of DEE and EtOH mixtures supports the isotopic findings. Although both approaches lacked information on incompletely combusted intermediates plentiful in HCCI emissions, the numerical model and {sup 14}C tracing data agreed within the limitations of the single zone model. Despitemore » the fact that DEE is more reactive than EtOH in HCCI engines, they are sufficiently similar that we did not observe a large elongation of energy release or significant reduction in inlet temperature required for light-off, both desired effects for the combustion event. This finding suggests that, in general, HCCI combustion of fuel blends may have preferential combustion of some of the blend components.« less

Combustion system CFD modeling at GE Aircraft Engines

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

Combustion system CFD modeling at GE Aircraft Engines

NASA Astrophysics Data System (ADS)

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

1995-03-01

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

Predictive modeling and reducing cyclic variability in autoignition engines

DOEpatents

Hellstrom, Erik; Stefanopoulou, Anna; Jiang, Li; Larimore, Jacob

2016-08-30

Methods and systems are provided for controlling a vehicle engine to reduce cycle-to-cycle combustion variation. A predictive model is applied to predict cycle-to-cycle combustion behavior of an engine based on observed engine performance variables. Conditions are identified, based on the predicted cycle-to-cycle combustion behavior, that indicate high cycle-to-cycle combustion variation. Corrective measures are then applied to prevent the predicted high cycle-to-cycle combustion variation.

Investigation of critical burning of fuel droplets

NASA Technical Reports Server (NTRS)

Faeth, G. M.

1979-01-01

The general problem of spray combustion was investigated. The combustion of bipropellent droplets; combustion of hydrozine fuels; and combustion of sprays were studied. A model was developed to predict mean velocities and temperatures in a combusting gas jet.

Investigating co-combustion characteristics of bamboo and wood.

PubMed

Liang, Fang; Wang, Ruijuan; Jiang, Changle; Yang, Xiaomeng; Zhang, Tao; Hu, Wanhe; Mi, Bingbing; Liu, Zhijia

2017-11-01

To investigate co-combustion characteristics of bamboo and wood, moso bamboo and masson pine were torrefied and mixed with different blend ratios. The combustion process was examined by thermogravimetric analyzer (TGA). The results showed the combustion process of samples included volatile emission and oxidation combustion as well as char combustion. The main mass loss of biomass blends occurred at volatile emission and oxidation combustion stage, while that of torrefied biomass occurred at char combustion stage. With the increase of bamboo content, characteristic temperatures decreased. Compared with untreated biomass, torrefied biomass had a higher initial and burnout temperature. With the increase of heating rates, combustion process of samples shifted to higher temperatures. Compared with non-isothermal models, activation energy obtained from isothermal model was lower. The result is helpful to promote development of co-combustion of bamboo and masson pine wastes. Copyright © 2017 Elsevier Ltd. All rights reserved.

Research on optimization of combustion efficiency of thermal power unit based on genetic algorithm

NASA Astrophysics Data System (ADS)

Zhou, Qiongyang

2018-04-01

In order to improve the economic performance and reduce pollutant emissions of thermal power units, the characteristics of neural network in establishing boiler combustion model are analyzed based on the analysis of the main factors affecting boiler efficiency by using orthogonal method. In addition, on the basis of this model, the genetic algorithm is used to find the best control amount of the furnace combustion in a certain working condition. Through the genetic algorithm based on real number encoding and roulette selection is concluded: the best control quantity at a condition of furnace combustion can be combined with the boiler combustion system model for neural network training. The precision of the neural network model is further improved, and the basic work is laid for the research of the whole boiler combustion optimization system.

Critical evaluation of Jet-A spray combustion using propane chemical kinetics in gas turbine combustion simulated by KIVA-2

NASA Technical Reports Server (NTRS)

Nguyen, H. L.; Ying, S.-J.

1990-01-01

Jet-A spray combustion has been evaluated in gas turbine combustion with the use of propane chemical kinetics as the first approximation for the chemical reactions. Here, the numerical solutions are obtained by using the KIVA-2 computer code. The KIVA-2 code is the most developed of the available multidimensional combustion computer programs for application of the in-cylinder combustion dynamics of internal combustion engines. The released version of KIVA-2 assumes that 12 chemical species are present; the code uses an Arrhenius kinetic-controlled combustion model governed by a four-step global chemical reaction and six equilibrium reactions. Researchers efforts involve the addition of Jet-A thermophysical properties and the implementation of detailed reaction mechanisms for propane oxidation. Three different detailed reaction mechanism models are considered. The first model consists of 131 reactions and 45 species. This is considered as the full mechanism which is developed through the study of chemical kinetics of propane combustion in an enclosed chamber. The full mechanism is evaluated by comparing calculated ignition delay times with available shock tube data. However, these detailed reactions occupy too much computer memory and CPU time for the computation. Therefore, it only serves as a benchmark case by which to evaluate other simplified models. Two possible simplified models were tested in the existing computer code KIVA-2 for the same conditions as used with the full mechanism. One model is obtained through a sensitivity analysis using LSENS, the general kinetics and sensitivity analysis program code of D. A. Bittker and K. Radhakrishnan. This model consists of 45 chemical reactions and 27 species. The other model is based on the work published by C. K. Westbrook and F. L. Dryer.

Effects of fuel cetane number on the structure of diesel spray combustion: An accelerated Eulerian stochastic fields method

NASA Astrophysics Data System (ADS)

Jangi, Mehdi; Lucchini, Tommaso; Gong, Cheng; Bai, Xue-Song

2015-09-01

An Eulerian stochastic fields (ESF) method accelerated with the chemistry coordinate mapping (CCM) approach for modelling spray combustion is formulated, and applied to model diesel combustion in a constant volume vessel. In ESF-CCM, the thermodynamic states of the discretised stochastic fields are mapped into a low-dimensional phase space. Integration of the chemical stiff ODEs is performed in the phase space and the results are mapped back to the physical domain. After validating the ESF-CCM, the method is used to investigate the effects of fuel cetane number on the structure of diesel spray combustion. It is shown that, depending of the fuel cetane number, liftoff length is varied, which can lead to a change in combustion mode from classical diesel spray combustion to fuel-lean premixed burned combustion. Spray combustion with a shorter liftoff length exhibits the characteristics of the classical conceptual diesel combustion model proposed by Dec in 1997 (http://dx.doi.org/10.4271/970873), whereas in a case with a lower cetane number the liftoff length is much larger and the spray combustion probably occurs in a fuel-lean-premixed mode of combustion. Nevertheless, the transport budget at the liftoff location shows that stabilisation at all cetane numbers is governed primarily by the auto-ignition process.

The Diesel Combustion Collaboratory: Combustion Researchers Collaborating over the Internet

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

C. M. Pancerella; L. A. Rahn; C. Yang

2000-02-01

The Diesel Combustion Collaborator (DCC) is a pilot project to develop and deploy collaborative technologies to combustion researchers distributed throughout the DOE national laboratories, academia, and industry. The result is a problem-solving environment for combustion research. Researchers collaborate over the Internet using DCC tools, which include: a distributed execution management system for running combustion models on widely distributed computers, including supercomputers; web-accessible data archiving capabilities for sharing graphical experimental or modeling data; electronic notebooks and shared workspaces for facilitating collaboration; visualization of combustion data; and video-conferencing and data-conferencing among researchers at remote sites. Security is a key aspect of themore » collaborative tools. In many cases, the authors have integrated these tools to allow data, including large combustion data sets, to flow seamlessly, for example, from modeling tools to data archives. In this paper the authors describe the work of a larger collaborative effort to design, implement and deploy the DCC.« less

Recent advances in large-eddy simulation of spray and coal combustion

NASA Astrophysics Data System (ADS)

Zhou, L. X.

2013-07-01

Large-eddy simulation (LES) is under its rapid development and is recognized as a possible second generation of CFD methods used in engineering. Spray and coal combustion is widely used in power, transportation, chemical and metallurgical, iron and steel making, aeronautical and astronautical engineering, hence LES of spray and coal two-phase combustion is particularly important for engineering application. LES of two-phase combustion attracts more and more attention; since it can give the detailed instantaneous flow and flame structures and more exact statistical results than those given by the Reynolds averaged modeling (RANS modeling). One of the key problems in LES is to develop sub-grid scale (SGS) models, including SGS stress models and combustion models. Different investigators proposed or adopted various SGS models. In this paper the present author attempts to review the advances in studies on LES of spray and coal combustion, including the studies done by the present author and his colleagues. Different SGS models adopted by different investigators are described, some of their main results are summarized, and finally some research needs are discussed.

Suppression of combustion oscillations with mechanical damping devices

NASA Technical Reports Server (NTRS)

1971-01-01

Nonarray absorbing devices were investigated for use in rocket thrust chambers as instability suppressors. A theory for designing absorbing devices suitable for rocket application is derived, and a nonarray computer program is developed. The experimental program used to verify the theory is discussed. It is concluded that individual acoustical devices can be designed for maximum energy absorption, and it is recommended that single resonators be designed so that the ratio of the aperture diameter to the product of the quarter-wave length and cavity backing depth is less than one.

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

NASA Technical Reports Server (NTRS)

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

1982-01-01

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

Extended lattice Boltzmann scheme for droplet combustion.

PubMed

Ashna, Mostafa; Rahimian, Mohammad Hassan; Fakhari, Abbas

2017-05-01

The available lattice Boltzmann (LB) models for combustion or phase change are focused on either single-phase flow combustion or two-phase flow with evaporation assuming a constant density for both liquid and gas phases. To pave the way towards simulation of spray combustion, we propose a two-phase LB method for modeling combustion of liquid fuel droplets. We develop an LB scheme to model phase change and combustion by taking into account the density variation in the gas phase and accounting for the chemical reaction based on the Cahn-Hilliard free-energy approach. Evaporation of liquid fuel is modeled by adding a source term, which is due to the divergence of the velocity field being nontrivial, in the continuity equation. The low-Mach-number approximation in the governing Navier-Stokes and energy equations is used to incorporate source terms due to heat release from chemical reactions, density variation, and nonluminous radiative heat loss. Additionally, the conservation equation for chemical species is formulated by including a source term due to chemical reaction. To validate the model, we consider the combustion of n-heptane and n-butanol droplets in stagnant air using overall single-step reactions. The diameter history and flame standoff ratio obtained from the proposed LB method are found to be in good agreement with available numerical and experimental data. The present LB scheme is believed to be a promising approach for modeling spray combustion.

40 CFR Table 4 to Subpart Bbbb of... - Model Rule-Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a

Code of Federal Regulations, 2011 CFR

2011-07-01

... Existing Small Municipal Waste Combustion Unit a 4 Table 4 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a For...

40 CFR Table 2 to Subpart Bbbb of... - Model Rule-Class I Emission Limits for Existing Small Municipal Waste Combustion Units a

Code of Federal Regulations, 2014 CFR

2014-07-01

... Existing Small Municipal Waste Combustion Units a 2 Table 2 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class I Emission Limits for Existing Small Municipal Waste Combustion Units a For...

40 CFR Table 4 to Subpart Bbbb of... - Model Rule-Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a

Code of Federal Regulations, 2014 CFR

2014-07-01

... Existing Small Municipal Waste Combustion Unit a 4 Table 4 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a For...

40 CFR Table 4 to Subpart Bbbb of... - Model Rule-Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a

Code of Federal Regulations, 2010 CFR

2010-07-01

... Existing Small Municipal Waste Combustion Unit a 4 Table 4 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a For...

40 CFR Table 2 to Subpart Bbbb of... - Model Rule-Class I Emission Limits for Existing Small Municipal Waste Combustion Units a

Code of Federal Regulations, 2012 CFR

2012-07-01

... Existing Small Municipal Waste Combustion Units a 2 Table 2 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class I Emission Limits for Existing Small Municipal Waste Combustion Units a For...

40 CFR Table 2 to Subpart Bbbb of... - Model Rule-Class I Emission Limits for Existing Small Municipal Waste Combustion Units a

Code of Federal Regulations, 2013 CFR

2013-07-01

... Existing Small Municipal Waste Combustion Units a 2 Table 2 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class I Emission Limits for Existing Small Municipal Waste Combustion Units a For...

40 CFR Table 2 to Subpart Bbbb of... - Model Rule-Class I Emission Limits for Existing Small Municipal Waste Combustion Units a

Code of Federal Regulations, 2010 CFR

2010-07-01

... Existing Small Municipal Waste Combustion Units a 2 Table 2 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class I Emission Limits for Existing Small Municipal Waste Combustion Units a For...

40 CFR Table 2 to Subpart Bbbb of... - Model Rule-Class I Emission Limits for Existing Small Municipal Waste Combustion Units a

Code of Federal Regulations, 2011 CFR

2011-07-01

... Existing Small Municipal Waste Combustion Units a 2 Table 2 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class I Emission Limits for Existing Small Municipal Waste Combustion Units a For...

40 CFR Table 4 to Subpart Bbbb of... - Model Rule-Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a

Code of Federal Regulations, 2012 CFR

2012-07-01

... Existing Small Municipal Waste Combustion Unit a 4 Table 4 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a For...

40 CFR Table 4 to Subpart Bbbb of... - Model Rule-Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a

Code of Federal Regulations, 2013 CFR

2013-07-01

... Existing Small Municipal Waste Combustion Unit a 4 Table 4 to Subpart BBBB of Part 60 Protection of... NEW STATIONARY SOURCES Emission Guidelines and Compliance Times for Small Municipal Waste Combustion... Part 60—Model Rule—Class II Emission Limits for Existing Small Municipal Waste Combustion Unit a For...

Fuel properties to enable lifted-flame combustion

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

Kurtz, Eric

The Fuel Properties to Enable Lifted-Flame Combustion project responded directly to solicitation DE-FOA-0000239 AOI 1A, Fuels and Lubricants for Advanced Combustion Regimes. This subtopic was intended to encompass clean and highly-efficient, liquid-fueled combustion engines to achieve extremely low engine-out nitrogen oxides (NOx) and particulate matter (PM) as a target and similar efficiency as state-of-the-art direct injection diesel engines. The intent of this project was to identify how fuel properties can be used to achieve controllable Leaner Lifted Flame Combustion (LLFC) with low NOx and PM emissions. Specifically, this project was expected to identify and test key fuel properties to enablemore » LLFC and their compatibility with current fuel systems and to enhance combustion models to capture the effect of fuel properties on advanced combustion. Successful demonstration of LLFC may reduce the need for after treatment devices, thereby reducing costs and improving thermal efficiency. The project team consisted of key technical personnel from Ford Motor Company (FMC), the University of Wisconsin-Madison (UW), Sandia National Laboratories (SNL) and Lawrence Livermore National Laboratories (LLNL). Each partner had key roles in achieving project objectives. FMC investigated fuel properties relating to LLFC and sooting tendency. Together, FMC and UW developed and integrated 3D combustion models to capture fuel property combustion effects. FMC used these modeling results to develop a combustion system and define fuel properties to support a single-cylinder demonstration of fuel-enabled LLFC. UW investigated modeling the flame characteristics and emissions behavior of different fuels, including those with different cetane number and oxygen content. SNL led spray combustion experiments to quantify the effect of key fuel properties on combustion characteristics critical for LLFC, as well as single cylinder optical engine experiments to improve fundamental understanding of flame lift-off, generate model validation data, and demonstrate LLFC concurrent with FMC efforts. Additionally, LLNL was added to the project during the second year to develop a detailed kinetic mechanism for a key oxygenate to support CFD modeling. Successful completion of this project allowed the team to enhance fundamental understanding of LLFC, improve the state of current combustion models and increase understanding of desired fuel properties. This knowledge also improves our knowledge of how cost effective and environmentally friendly renewable fuels can assist in helping meet future emission and greenhouse gas regulations.« less

A comprehensive evaluation of different radiation models in a gas turbine combustor under conditions of oxy-fuel combustion with dry recycle

NASA Astrophysics Data System (ADS)

Kez, V.; Liu, F.; Consalvi, J. L.; Ströhle, J.; Epple, B.

2016-03-01

The oxy-fuel combustion is a promising CO2 capture technology from combustion systems. This process is characterized by much higher CO2 concentrations in the combustion system compared to that of the conventional air-fuel combustion. To accurately predict the enhanced thermal radiation in oxy-fuel combustion, it is essential to take into account the non-gray nature of gas radiation. In this study, radiation heat transfer in a 3D model gas turbine combustor under two test cases at 20 atm total pressure was calculated by various non-gray gas radiation models, including the statistical narrow-band (SNB) model, the statistical narrow-band correlated-k (SNBCK) model, the wide-band correlated-k (WBCK) model, the full spectrum correlated-k (FSCK) model, and several weighted sum of gray gases (WSGG) models. Calculations of SNB, SNBCK, and FSCK were conducted using the updated EM2C SNB model parameters. Results of the SNB model are considered as the benchmark solution to evaluate the accuracy of the other models considered. Results of SNBCK and FSCK are in good agreement with the benchmark solution. The WBCK model is less accurate than SNBCK or FSCK. Considering the three formulations of the WBCK model, the multiple gases formulation is the best choice regarding the accuracy and computational cost. The WSGG model with the parameters of Bordbar et al. (2014) [20] is the most accurate of the three investigated WSGG models. Use of the gray WSSG formulation leads to significant deviations from the benchmark data and should not be applied to predict radiation heat transfer in oxy-fuel combustion systems. A best practice to incorporate the state-of-the-art gas radiation models for high accuracy of radiation heat transfer calculations at minimal increase in computational cost in CFD simulation of oxy-fuel combustion systems for pressure path lengths up to about 10 bar m is suggested.

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

NASA Astrophysics Data System (ADS)

Mehanna Ismail, Mohammed Ali

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

Numerical modelling of biomass combustion: Solid conversion processes in a fixed bed furnace

NASA Astrophysics Data System (ADS)

Karim, Md. Rezwanul; Naser, Jamal

2017-06-01

Increasing demand for energy and rising concerns over global warming has urged the use of renewable energy sources to carry a sustainable development of the world. Bio mass is a renewable energy which has become an important fuel to produce thermal energy or electricity. It is an eco-friendly source of energy as it reduces carbon dioxide emissions. Combustion of solid biomass is a complex phenomenon due to its large varieties and physical structures. Among various systems, fixed bed combustion is the most commonly used technique for thermal conversion of solid biomass. But inadequate knowledge on complex solid conversion processes has limited the development of such combustion system. Numerical modelling of this combustion system has some advantages over experimental analysis. Many important system parameters (e.g. temperature, density, solid fraction) can be estimated inside the entire domain under different working conditions. In this work, a complete numerical model is used for solid conversion processes of biomass combustion in a fixed bed furnace. The combustion system is divided in to solid and gas phase. This model includes several sub models to characterize the solid phase of the combustion with several variables. User defined subroutines are used to introduce solid phase variables in commercial CFD code. Gas phase of combustion is resolved using built-in module of CFD code. Heat transfer model is modified to predict the temperature of solid and gas phases with special radiation heat transfer solution for considering the high absorptivity of the medium. Considering all solid conversion processes the solid phase variables are evaluated. Results obtained are discussed with reference from an experimental burner.

Numerical Simulation of Combustion and Rotor-Stator Interaction in a Turbine Combustor

DOE PAGES

Isvoranu, Dragos D.; Cizmas, Paul G. A.

2003-01-01

This article presents the development of a numerical algorithm for the computation of flow and combustion in a turbine combustor. The flow and combustion are modeled by the Reynolds-averaged Navier-Stokes equations coupled with the species-conservation equations. The chemistry model used herein is a two-step, global, finite-rate combustion model for methane and combustion gases. The governing equations are written in the strong conservation form and solved using a fully implicit, finite-difference approximation. The gas dynamics and chemistry equations are fully decoupled. A correction technique has been developed to enforce the conservation of mass fractions. The numerical algorithm developed herein has beenmore » used to investigate the flow and combustion in a one-stage turbine combustor.« less

Parallel distributed, reciprocal Monte Carlo radiation in coupled, large eddy combustion simulations

NASA Astrophysics Data System (ADS)

Hunsaker, Isaac L.

Radiation is the dominant mode of heat transfer in high temperature combustion environments. Radiative heat transfer affects the gas and particle phases, including all the associated combustion chemistry. The radiative properties are in turn affected by the turbulent flow field. This bi-directional coupling of radiation turbulence interactions poses a major challenge in creating parallel-capable, high-fidelity combustion simulations. In this work, a new model was developed in which reciprocal monte carlo radiation was coupled with a turbulent, large-eddy simulation combustion model. A technique wherein domain patches are stitched together was implemented to allow for scalable parallelism. The combustion model runs in parallel on a decomposed domain. The radiation model runs in parallel on a recomposed domain. The recomposed domain is stored on each processor after information sharing of the decomposed domain is handled via the message passing interface. Verification and validation testing of the new radiation model were favorable. Strong scaling analyses were performed on the Ember cluster and the Titan cluster for the CPU-radiation model and GPU-radiation model, respectively. The model demonstrated strong scaling to over 1,700 and 16,000 processing cores on Ember and Titan, respectively.

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.

Simulating Hadronic-to-Quark-Matter with Burn-UD: Recent work and astrophysical applications

NASA Astrophysics Data System (ADS)

Welbanks, Luis; Ouyed, Amir; Koning, Nico; Ouyed, Rachid

2017-06-01

We present the new developments in Burn-UD, our in-house hydrodynamic combustion code used to model the phase transition of hadronic-to-quark matter. Our two new modules add neutrino transport and the time evolution of a (u, d, s) quark star (QS). Preliminary simulations show that the inclusion of neutrino transport points towards new hydrodynamic instabilities that increase the burning speed. A higher burning speed could elicit the deflagration to detonation of a neutron star (NS) into a QS. We propose that a Quark-Nova (QN: the explosive transition of a NS to a QS) could help us explain the most energetic astronomical events to this day: superluminous supernovae (SLSNe). Our models consider a QN occurring in a massive binary, experiencing two common envelope stages and a QN occurring after the supernova explosion of a Wolf-Rayet (WO) star. Both models have been successful in explaining the double humped light curves of over half a dozen SLSNe. We also introduce SiRop our r-process simulation code and propose that a QN site has the hot temperatures and neutron densities required to make it an ideal site for the r-process.

Analysis of pressure spectra measurements in a ducted combustion system. Ph.D. Thesis - Toledo Univ.

NASA Technical Reports Server (NTRS)

Miles, J. H.

1980-01-01

Combustion noise propagation in an operating ducted liquid fuel combustion system is studied in relation to the development of combustion noise prediction and suppression techniques. The presence of combustor emissions in the duct is proposed as the primary mechanism producing the attenuation and dispersion of combustion noise propagating in an operating liquid fuel combustion system. First, a complex mathematical model for calculating attenuation and dispersion taking into account mass transfer, heat transfer, and viscosity effects due to the presence of liquid fuel droplets or solid soot particles is discussed. Next, a simpler single parameter model for calculating pressure auto-spectra and cross-spectra which takes into account dispersion and attenuation due to heat transfer between solid soot particles and air is developed. Then, auto-spectra and cross-spectra obtained from internal pressure measurements in a combustion system consisting of a J-47 combustor can, a spool piece, and a long duct are presented. Last, analytical results obtained with the single parameter model are compared with the experimental measurements. The single parameter model results are shown to be in excellent agreement with the measurements.

Numerical study of influence of molecular diffusion in the Mild combustion regime

NASA Astrophysics Data System (ADS)

Mardani, Amir; Tabejamaat, Sadegh; Ghamari, Mohsen

2010-09-01

In this paper, the importance of molecular diffusion versus turbulent transport in the moderate or intense low-oxygen dilution (Mild) combustion mode has been numerically studied. The experimental conditions of Dally et al. [Proc. Combust. Inst. 29 (2002) 1147-1154] were used for modelling. The EDC model was used to describe the turbulence-chemistry interaction. The DRM-22 reduced mechanism and the GRI 2.11 full mechanism were used to represent the chemical reactions of an H2/methane jet flame. The importance of molecular diffusion for various O2 levels, jet Reynolds numbers and H2 fuel contents was investigated. Results show that the molecular diffusion in Mild combustion cannot be ignored in comparison with the turbulent transport. Also, the method of inclusion of molecular diffusion in combustion modelling has a considerable effect on the accuracy of numerical modelling of Mild combustion. By decreasing the jet Reynolds number, decreasing the oxygen concentration in the airflow or increasing H2 in the fuel mixture, the influence of molecular diffusion on Mild combustion increases.

Analysis of pressure spectra measurements in a ducted combustion system

NASA Astrophysics Data System (ADS)

Miles, J. H.

1980-11-01

Combustion noise propagation in an operating ducted liquid fuel combustion system is studied in relation to the development of combustion noise prediction and suppression techniques. The presence of combustor emissions in the duct is proposed as the primary mechanism producing the attenuation and dispersion of combustion noise propagating in an operating liquid fuel combustion system. First, a complex mathematical model for calculating attenuation and dispersion taking into account mass transfer, heat transfer, and viscosity effects due to the presence of liquid fuel droplets or solid soot particles is discussed. Next, a simpler single parameter model for calculating pressure auto-spectra and cross-spectra which takes into account dispersion and attenuation due to heat transfer between solid soot particles and air is developed. Then, auto-spectra and cross-spectra obtained from internal pressure measurements in a combustion system consisting of a J-47 combustor can, a spool piece, and a long duct are presented. Last, analytical results obtained with the single parameter model are compared with the experimental measurements. The single parameter model results are shown to be in excellent agreement with the measurements.

Numerical Investigation Into Effect of Fuel Injection Timing on CAI/HCCI Combustion in a Four-Stroke GDI Engine

NASA Astrophysics Data System (ADS)

Cao, Li; Zhao, Hua; Jiang, Xi; Kalian, Navin

2006-02-01

The Controlled Auto-Ignition (CAI) combustion, also known as Homogeneous Charge Compression Ignition (HCCI), was achieved by trapping residuals with early exhaust valve closure in conjunction with direct injection. Multi-cycle 3D engine simulations have been carried out for parametric study on four different injection timings in order to better understand the effects of injection timings on in-cylinder mixing and CAI combustion. The full engine cycle simulation including complete gas exchange and combustion processes was carried out over several cycles in order to obtain the stable cycle for analysis. The combustion models used in the present study are the Shell auto-ignition model and the characteristic-time combustion model, which were modified to take the high level of EGR into consideration. A liquid sheet breakup spray model was used for the droplet breakup processes. The analyses show that the injection timing plays an important role in affecting the in-cylinder air/fuel mixing and mixture temperature, which in turn affects the CAI combustion and engine performance.

Analysis of the performance, emission and combustion characteristics of a turbocharged diesel engine fuelled with Jatropha curcas biodiesel-diesel blends using kernel-based extreme learning machine.

PubMed

Silitonga, Arridina Susan; Hassan, Masjuki Haji; Ong, Hwai Chyuan; Kusumo, Fitranto

2017-11-01

The purpose of this study is to investigate the performance, emission and combustion characteristics of a four-cylinder common-rail turbocharged diesel engine fuelled with Jatropha curcas biodiesel-diesel blends. A kernel-based extreme learning machine (KELM) model is developed in this study using MATLAB software in order to predict the performance, combustion and emission characteristics of the engine. To acquire the data for training and testing the KELM model, the engine speed was selected as the input parameter, whereas the performance, exhaust emissions and combustion characteristics were chosen as the output parameters of the KELM model. The performance, emissions and combustion characteristics predicted by the KELM model were validated by comparing the predicted data with the experimental data. The results show that the coefficient of determination of the parameters is within a range of 0.9805-0.9991 for both the KELM model and the experimental data. The mean absolute percentage error is within a range of 0.1259-2.3838. This study shows that KELM modelling is a useful technique in biodiesel production since it facilitates scientists and researchers to predict the performance, exhaust emissions and combustion characteristics of internal combustion engines with high accuracy.

Comparing kinetic Monte Carlo and thin-film modeling of transversal instabilities of ridges on patterned substrates

NASA Astrophysics Data System (ADS)

Tewes, Walter; Buller, Oleg; Heuer, Andreas; Thiele, Uwe; Gurevich, Svetlana V.

2017-03-01

We employ kinetic Monte Carlo (KMC) simulations and a thin-film continuum model to comparatively study the transversal (i.e., Plateau-Rayleigh) instability of ridges formed by molecules on pre-patterned substrates. It is demonstrated that the evolution of the occurring instability qualitatively agrees between the two models for a single ridge as well as for two weakly interacting ridges. In particular, it is shown for both models that the instability occurs on well defined length and time scales which are, for the KMC model, significantly larger than the intrinsic scales of thermodynamic fluctuations. This is further evidenced by the similarity of dispersion relations characterizing the linear instability modes.

Development and Validation of a 3-Dimensional CFB Furnace Model

NASA Astrophysics Data System (ADS)

Vepsäläinen, Arl; Myöhänen, Karl; Hyppäneni, Timo; Leino, Timo; Tourunen, Antti

At Foster Wheeler, a three-dimensional CFB furnace model is essential part of knowledge development of CFB furnace process regarding solid mixing, combustion, emission formation and heat transfer. Results of laboratory and pilot scale phenomenon research are utilized in development of sub-models. Analyses of field-test results in industrial-scale CFB boilers including furnace profile measurements are simultaneously carried out with development of 3-dimensional process modeling, which provides a chain of knowledge that is utilized as feedback for phenomenon research. Knowledge gathered by model validation studies and up-to-date parameter databases are utilized in performance prediction and design development of CFB boiler furnaces. This paper reports recent development steps related to modeling of combustion and formation of char and volatiles of various fuel types in CFB conditions. Also a new model for predicting the formation of nitrogen oxides is presented. Validation of mixing and combustion parameters for solids and gases are based on test balances at several large-scale CFB boilers combusting coal, peat and bio-fuels. Field-tests including lateral and vertical furnace profile measurements and characterization of solid materials provides a window for characterization of fuel specific mixing and combustion behavior in CFB furnace at different loads and operation conditions. Measured horizontal gas profiles are projection of balance between fuel mixing and reactions at lower part of furnace and are used together with both lateral temperature profiles at bed and upper parts of furnace for determination of solid mixing and combustion model parameters. Modeling of char and volatile based formation of NO profiles is followed by analysis of oxidizing and reducing regions formed due lower furnace design and mixing characteristics of fuel and combustion airs effecting to formation ofNO furnace profile by reduction and volatile-nitrogen reactions. This paper presents CFB process analysis focused on combustion and NO profiles in pilot and industrial scale bituminous coal combustion.

Low-Emission combustion of fuel in aeroderivative gas turbines

NASA Astrophysics Data System (ADS)

Bulysova, L. A.; Vasil'ev, V. D.; Berne, A. L.

2017-12-01

The paper is the first of a planned set of papers devoted to the world experience in development of Low Emission combustors (LEC) for industrial Gas Turbines (GT). The purpose of the article is to summarize and analyze the most successful experience of introducing the principles of low-emission combustion of the so-called "poor" (low fuel concentration in air when the excess air ratio is about 1.9-2.1) well mixed fuelair mixtures in the LEC for GTs and ways to reduce the instability of combustion. The consideration examples are the most successful and widely used aero-derivative GT. The GT development meets problems related to the difference in requirements and operation conditions between the aero, industrial, and power production GT. One of the main problems to be solved is the LEC development to mitigate emissions of the harmful products first of all the Nitrogen oxides NOx. The ways to modify or convert the initial combustors to the LEC are shown. This development may follow location of multiburner mixers within the initial axial envelope dimensions or conversion of circular combustor to the can type one. The most interesting are Natural Gas firing GT without water injection into the operating process or Dry Low emission (DLE) combustors. The current GT efficiency requirement may be satisfied at compressor exit pressure above 3 MPa and Turbine Entry temperature (TET) above 1500°C. The paper describes LEC examples based on the concept of preliminary prepared air-fuel mixtures' combustion. Each combustor employs its own fuel supply control concept based on the fuel flow-power output relation. In the case of multiburner combustors, the burners are started subsequently under a specific scheme. The can type combustors have combustion zones gradually ignited following the GT power change. The combustion noise problem experienced in lean mixtures' combustion is also considered, and the problem solutions are described. The GT test results show wide ranges of stable operation at needed levels of NOx and CO emissions. The world experience analysis and generalization and investigation of the further development directions for the high performance GT will assist development of domestic LEC for prospective GTs.

Unsteady numerical simulations of the stability and dynamics of flames

NASA Technical Reports Server (NTRS)

Kailasanath, K.; Patnaik, G.; Oran, E. S.

1995-01-01

In this report we describe the research performed at the Naval Research Laboratory in support of the NASA Microgravity Science and Applications Program over the past three years (from Feb. 1992) with emphasis on the work performed since the last microgravity combustion workshop. The primary objective of our research is to develop an understanding of the differences in the structure, stability, dynamics and extinction of flames in earth gravity and in microgravity environments. Numerical simulations, in which the various physical and chemical processes can be independently controlled, can significantly advance our understanding of these differences. Therefore, our approach is to use detailed time-dependent, multi-dimensional, multispecies numerical models to perform carefully designed computational experiments. The basic issues we have addressed, a general description of the numerical approach, and a summary of the results are described in this report. More detailed discussions are available in the papers published which are referenced herein. Some of the basic issues we have addressed recently are (1) the relative importance of wall losses and gravity on the extinguishment of downward-propagating flames; (2) the role of hydrodynamic instabilities in the formation of cellular flames; (3) effects of gravity on burner-stabilized flames, and (4) effects of radiative losses and chemical-kinetics on flames near flammability limits. We have also expanded our efforts to include hydrocarbon flames in addition to hydrogen flames and to perform simulations in support of other on-going efforts in the microgravity combustion sciences program. Modeling hydrocarbon flames typically involves a larger number of species and a much larger number of reactions when compared to hydrogen. In addition, more complex radiation models may also be needed. In order to efficiently compute such complex flames recent developments in parallel computing have been utilized to develop a state-of-the-art parallel flame code. This is discussed below in some detail after a brief discussion of the numerical models.

Non-equilibrium diffusion combustion of a fuel droplet

NASA Astrophysics Data System (ADS)

Tyurenkova, Veronika V.

2012-06-01

A mathematical model for the non-equilibrium combustion of droplets in rocket engines is developed. This model allows to determine the divergence of combustion rate for the equilibrium and non-equilibrium model. Criterion for droplet combustion deviation from equilibrium is introduced. It grows decreasing droplet radius, accommodation coefficient, temperature and decreases on decreasing diffusion coefficient. Also divergence from equilibrium increases on reduction of droplet radius. Droplet burning time essentially increases under non-equilibrium conditions. Comparison of theoretical and experimental data shows that to have adequate solution for small droplets it is necessary to use the non-equilibrium model.

Methods of the working processes modelling of an internal combustion engine by an ANSYS IC Engine module

NASA Astrophysics Data System (ADS)

Kurchatkin, I. V.; Gorshkalev, A. A.; Blagin, E. V.

2017-01-01

This article deals with developed methods of the working processes modelling in the combustion chamber of an internal combustion engine (ICE). Methods includes description of the preparation of a combustion chamber 3-d model, setting of the finite-element mesh, boundary condition setting and solution customization. Aircraft radial engine M-14 was selected for modelling. The cycle of cold blowdown in the ANSYS IC Engine software was carried out. The obtained data were compared to results of known calculation methods. A method of engine’s induction port improvement was suggested.

Evaluation of char combustion models: measurement and analysis of variability in char particle size and density

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

Maloney, Daniel J; Monazam, Esmail R; Casleton, Kent H

Char samples representing a range of combustion conditions and extents of burnout were obtained from a well-characterized laminar flow combustion experiment. Individual particles from the parent coal and char samples were characterized to determine distributions in particle volume, mass, and density at different extent of burnout. The data were then compared with predictions from a comprehensive char combustion model referred to as the char burnout kinetics model (CBK). The data clearly reflect the particle- to-particle heterogeneity of the parent coal and show a significant broadening in the size and density distributions of the chars resulting from both devolatilization and combustion.more » Data for chars prepared in a lower oxygen content environment (6% oxygen by vol.) are consistent with zone II type combustion behavior where most of the combustion is occurring near the particle surface. At higher oxygen contents (12% by vol.), the data show indications of more burning occurring in the particle interior. The CBK model does a good job of predicting the general nature of the development of size and density distributions during burning but the input distribution of particle size and density is critical to obtaining good predictions. A significant reduction in particle size was observed to occur as a result of devolatilization. For comprehensive combustion models to provide accurate predictions, this size reduction phenomenon needs to be included in devolatilization models so that representative char distributions are carried through the calculations.« less

Evaluation of a hybrid kinetics/mixing-controlled combustion model for turbulent premixed and diffusion combustion using KIVA-II

NASA Technical Reports Server (NTRS)

Nguyen, H. Lee; Wey, Ming-Jyh

1990-01-01

Two-dimensional calculations were made of spark ignited premixed-charge combustion and direct injection stratified-charge combustion in gasoline fueled piston engines. Results are obtained using kinetic-controlled combustion submodel governed by a four-step global chemical reaction or a hybrid laminar kinetics/mixing-controlled combustion submodel that accounts for laminar kinetics and turbulent mixing effects. The numerical solutions are obtained by using KIVA-2 computer code which uses a kinetic-controlled combustion submodel governed by a four-step global chemical reaction (i.e., it assumes that the mixing time is smaller than the chemistry). A hybrid laminar/mixing-controlled combustion submodel was implemented into KIVA-2. In this model, chemical species approach their thermodynamics equilibrium with a rate that is a combination of the turbulent-mixing time and the chemical-kinetics time. The combination is formed in such a way that the longer of the two times has more influence on the conversion rate and the energy release. An additional element of the model is that the laminar-flame kinetics strongly influence the early flame development following ignition.

Evaluation of a hybrid kinetics/mixing-controlled combustion model for turbulent premixed and diffusion combustion using KIVA-2

NASA Technical Reports Server (NTRS)

Nguyen, H. Lee; Wey, Ming-Jyh

1990-01-01

Two dimensional calculations were made of spark ignited premixed-charge combustion and direct injection stratified-charge combustion in gasoline fueled piston engines. Results are obtained using kinetic-controlled combustion submodel governed by a four-step global chemical reaction or a hybrid laminar kinetics/mixing-controlled combustion submodel that accounts for laminar kinetics and turbulent mixing effects. The numerical solutions are obtained by using KIVA-2 computer code which uses a kinetic-controlled combustion submodel governed by a four-step global chemical reaction (i.e., it assumes that the mixing time is smaller than the chemistry). A hybrid laminar/mixing-controlled combustion submodel was implemented into KIVA-2. In this model, chemical species approach their thermodynamics equilibrium with a rate that is a combination of the turbulent-mixing time and the chemical-kinetics time. The combination is formed in such a way that the longer of the two times has more influence on the conversion rate and the energy release. An additional element of the model is that the laminar-flame kinetics strongly influence the early flame development following ignition.

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

Large eddy simulation modelling of combustion for propulsion applications.

PubMed

Fureby, C

2009-07-28

Predictive modelling of turbulent combustion is important for the development of air-breathing engines, internal combustion engines, furnaces and for power generation. Significant advances in modelling non-reactive turbulent flows are now possible with the development of large eddy simulation (LES), in which the large energetic scales of the flow are resolved on the grid while modelling the effects of the small scales. Here, we discuss the use of combustion LES in predictive modelling of propulsion applications such as gas turbine, ramjet and scramjet engines. The LES models used are described in some detail and are validated against laboratory data-of which results from two cases are presented. These validated LES models are then applied to an annular multi-burner gas turbine combustor and a simplified scramjet combustor, for which some additional experimental data are available. For these cases, good agreement with the available reference data is obtained, and the LES predictions are used to elucidate the flow physics in such devices to further enhance our knowledge of these propulsion systems. Particular attention is focused on the influence of the combustion chemistry, turbulence-chemistry interaction, self-ignition, flame holding burner-to-burner interactions and combustion oscillations.

Flex Fuel Optimized SI and HCCI Engine

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

Zhu, Guoming; Schock, Harold; Yang, Xiaojian

The central objective of the proposed work is to demonstrate an HCCI (homogeneous charge compression ignition) capable SI (spark ignited) engine that is capable of fast and smooth mode transition between SI and HCCI combustion modes. The model-based control technique was used to develop and validate the proposed control strategy for the fast and smooth combustion mode transition based upon the developed control-oriented engine; and an HCCI capable SI engine was designed and constructed using production ready two-step valve-train with electrical variable valve timing actuating system. Finally, smooth combustion mode transition was demonstrated on a metal engine within eight enginemore » cycles. The Chrysler turbocharged 2.0L I4 direct injection engine was selected as the base engine for the project and the engine was modified to fit the two-step valve with electrical variable valve timing actuating system. To develop the model-based control strategy for stable HCCI combustion and smooth combustion mode transition between SI and HCCI combustion, a control-oriented real-time engine model was developed and implemented into the MSU HIL (hardware-in-the-loop) simulation environment. The developed model was used to study the engine actuating system requirement for the smooth and fast combustion mode transition and to develop the proposed mode transition control strategy. Finally, a single cylinder optical engine was designed and fabricated for studying the HCCI combustion characteristics. Optical engine combustion tests were conducted in both SI and HCCI combustion modes and the test results were used to calibrate the developed control-oriented engine model. Intensive GT-Power simulations were conducted to determine the optimal valve lift (high and low) and the cam phasing range. Delphi was selected to be the supplier for the two-step valve-train and Denso to be the electrical variable valve timing system supplier. A test bench was constructed to develop control strategies for the electrical variable valve timing (VVT) actuating system and satisfactory electrical VVT responses were obtained. Target engine control system was designed and fabricated at MSU for both single-cylinder optical and multi-cylinder metal engines. Finally, the developed control-oriented engine model was successfully implemented into the HIL simulation environment. The Chrysler 2.0L I4 DI engine was modified to fit the two-step vale with electrical variable valve timing actuating system. A used prototype engine was used as the base engine and the cylinder head was modified for the two-step valve with electrical VVT actuating system. Engine validation tests indicated that cylinder #3 has very high blow-by and it cannot be reduced with new pistons and rings. Due to the time constraint, it was decided to convert the four-cylinder engine into a single cylinder engine by blocking both intake and exhaust ports of the unused cylinders. The model-based combustion mode transition control algorithm was developed in the MSU HIL simulation environment and the Simulink based control strategy was implemented into the target engine controller. With both single-cylinder metal engine and control strategy ready, stable HCCI combustion was achived with COV of 2.1% Motoring tests were conducted to validate the actuator transient operations including valve lift, electrical variable valve timing, electronic throttle, multiple spark and injection controls. After the actuator operations were confirmed, 15-cycle smooth combustion mode transition from SI to HCCI combustion was achieved; and fast 8-cycle smooth combustion mode transition followed. With a fast electrical variable valve timing actuator, the number of engine cycles required for mode transition can be reduced down to five. It was also found that the combustion mode transition is sensitive to the charge air and engine coolant temperatures and regulating the corresponding temperatures to the target levels during the combustion mode transition is the key for a smooth combustion mode transition. As a summary, the proposed combustion mode transition strategy using the hybrid combustion mode that starts with the SI combustion and ends with the HCCI combustion was experimentally validated on a metal engine. The proposed model-based control approach made it possible to complete the SI-HCCI combustion mode transition within eight engine cycles utilizing the well controlled hybrid combustion mode. Without intensive control-oriented engine modeling and HIL simulation study of using the hybrid combustion mode during the mode transition, it would be impossible to validate the proposed combustion mode transition strategy in a very short period.« less

Low-gravity fluid dynamics and transport phenomena. Progress in Astronautics and Aeronautics. Vol. 130

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

Koster, J.N.; Sani, R.L.

1990-01-01

Various papers on low-gravity fluid dynamics and transport phenomena are presented. Individual topics addressed include: fluid management in low gravity, nucleate pool boiling in variable gravity, application of energy-stability theory to problems in crystal growth, thermosolutal convection in liquid HgCdTe near the liquidus temperature, capillary surfaces in microgravity, thermohydrodynamic instabilities and capillary flows, interfacial oscillators, effects of gravity jitter on typical fluid science experiments and on natural convection in a vertical cylinder. Also discussed are: double-diffusive convection and its effects under reduced gravity, segregation and convection in dendritic alloys, fluid flow and microstructure development, analysis of convective situations with themore » Soret effect, complex natural convection in low Prandtl number metals, separation physics, phase partitioning in reduced gravity, separation of binary alloys with miscibility gap in the melt, Ostwald ripening in liquids, particle cloud combustion in reduced gravity, opposed-flow flame spread with implications for combustion at microgravity.« less

Turbulent Mixing and Afterburn in Post-Detonation Flow with Dense Particle Clouds

NASA Astrophysics Data System (ADS)

Menon, Suresh

2015-06-01

Reactive metal particles are used as additives in most explosives to enhance afterburn and augment the impact of the explosive. The afterburn is highly dependent on the particle dispersal and mixing in the post-detonation flow. The post-detonation flow is generally characterized by hydrodynamic instabilities emanating from the interaction of the blast waves with the detonation product gases and the ambient air. Further, influenced by the particles, the flow evolves and develops turbulent structures, which play vital role in determining mixing and combustion. Past studies in the field in open literature are reviewed along with some recent studies conducted using three dimensional numerical simulations of particle dispersal and combustion in the post-detonation flow. Spherical nitromethane charges enveloped by particle shells of varying thickness are considered along with dense loading effects. In dense flows, the particles block the flow of the gases and therefore, the role of the inter-particle interactions on particle dispersal cannot be ignored. Thus, both dense and dilute effects must be modeled simultaneously to simulate the post-detonation flow. A hybrid equation of state is employed to study the evolution of flow from detonation initiation till the late time mixing and afterburn. The particle dispersal pattern in each case is compared with the available experimental results. The burn rate and the energy release in each case is quantified and the effect of total mass of the particles and the particle size is analyzed in detail. Strengths and limitations of the various methods used for such studies as well as the uncertainties in the modeling strategies are also highlighted. Supported by Defense Threat Reduction Agency.

Thermophysics Characterization of Kerosene Combustion

NASA Technical Reports Server (NTRS)

Wang, Ten-See

2001-01-01

A one-formula surrogate fuel formulation and its quasi-global combustion kinetics model are developed to support the design of injectors and thrust chambers of kerosene-fueled rocket engines. This surrogate fuel model depicts a fuel blend that properly represents the general physical and chemical properties of kerosene. The accompanying gaseous-phase thermodynamics of the surrogate fuel is anchored with the heat of formation of kerosene and verified by comparing a series of one-dimensional rocket thrust chamber calculations. The quasi-global combustion kinetics model consists of several global steps for parent fuel decomposition, soot formation, and soot oxidation and a detailed wet-CO mechanism to complete the combustion process. The final thermophysics formulations are incorporated with a computational fluid dynamics model for prediction of the combustion efficiency of an unielement, tripropellant combustor and the radiation of a kerosene-fueled thruster plume. The model predictions agreed reasonably well with those of the tests.

Combustion and flow modelling applied to the OMV VTE

NASA Technical Reports Server (NTRS)

Larosiliere, Louis M.; Jeng, San-Mou

1990-01-01

A predictive tool for hypergolic bipropellant spray combustion and flow evolution in the OMV VTE (orbital maneuvering vehicle variable thrust engine) is described. It encompasses a computational technique for the gas phase governing equations, a discrete particle method for liquid bipropellant sprays, and constitutive models for combustion chemistry, interphase exchanges, and unlike impinging liquid hypergolic stream interactions. Emphasis is placed on the phenomenological modelling of the hypergolic liquid bipropellant gasification processes. An application to the OMV VTE combustion chamber is given in order to show some of the capabilities and inadequacies of this tool.

40 CFR 60.1635 - What must I do if I close my municipal waste combustion unit and then restart my municipal waste...

Code of Federal Regulations, 2011 CFR

2011-07-01

... waste combustion unit and then restart my municipal waste combustion unit? 60.1635 Section 60.1635... Combustion Units Constructed on or Before August 30, 1999 Model Rule-Increments of Progress § 60.1635 What must I do if I close my municipal waste combustion unit and then restart my municipal waste combustion...

40 CFR 60.1635 - What must I do if I close my municipal waste combustion unit and then restart my municipal waste...

Code of Federal Regulations, 2013 CFR

2013-07-01

... waste combustion unit and then restart my municipal waste combustion unit? 60.1635 Section 60.1635... Combustion Units Constructed on or Before August 30, 1999 Model Rule-Increments of Progress § 60.1635 What must I do if I close my municipal waste combustion unit and then restart my municipal waste combustion...

40 CFR 60.1635 - What must I do if I close my municipal waste combustion unit and then restart my municipal waste...

Code of Federal Regulations, 2014 CFR

2014-07-01

... waste combustion unit and then restart my municipal waste combustion unit? 60.1635 Section 60.1635... Combustion Units Constructed on or Before August 30, 1999 Model Rule-Increments of Progress § 60.1635 What must I do if I close my municipal waste combustion unit and then restart my municipal waste combustion...

40 CFR 60.1635 - What must I do if I close my municipal waste combustion unit and then restart my municipal waste...

Code of Federal Regulations, 2012 CFR

2012-07-01

... waste combustion unit and then restart my municipal waste combustion unit? 60.1635 Section 60.1635... Combustion Units Constructed on or Before August 30, 1999 Model Rule-Increments of Progress § 60.1635 What must I do if I close my municipal waste combustion unit and then restart my municipal waste combustion...

40 CFR 60.1635 - What must I do if I close my municipal waste combustion unit and then restart my municipal waste...

Code of Federal Regulations, 2010 CFR

2010-07-01

... waste combustion unit and then restart my municipal waste combustion unit? 60.1635 Section 60.1635... Combustion Units Constructed on or Before August 30, 1999 Model Rule-Increments of Progress § 60.1635 What must I do if I close my municipal waste combustion unit and then restart my municipal waste combustion...

Apollo Contour Rocket Nozzle in the Propulsion Systems Laboratory

NASA Image and Video Library

1964-07-21

Bill Harrison and Bud Meilander check the setup of an Apollo Contour rocket nozzle in the Propulsion Systems Laboratory at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The Propulsion Systems Laboratory contained two 14-foot diameter test chambers that could simulate conditions found at very high altitudes. The facility was used in the 1960s to study complex rocket engines such as the Pratt and Whitney RL-10 and rocket components such as the Apollo Contour nozzle, seen here. Meilander oversaw the facility’s mechanics and the installation of test articles into the chambers. Harrison was head of the Supersonic Tunnels Branch in the Test Installations Division. Researchers sought to determine the impulse value of the storable propellant mix, classify and improve the internal engine performance, and compare the results with analytical tools. A special setup was installed in the chamber that included a device to measure the thrust load and a calibration stand. Both cylindrical and conical combustion chambers were examined with the conical large area ratio nozzles. In addition, two contour nozzles were tested, one based on the Apollo Service Propulsion System and the other on the Air Force’s Titan transtage engine. Three types of injectors were investigated, including a Lewis-designed model that produced 98-percent efficiency. It was determined that combustion instability did not affect the nozzle performance. Although much valuable information was obtained during the tests, attempts to improve the engine performance were not successful.

Turbulent Combustion in SDF Explosions

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

Kuhl, A L; Bell, J B; Beckner, V E

2009-11-12

A heterogeneous continuum model is proposed to describe the dispersion and combustion of an aluminum particle cloud in an explosion. It combines the gas-dynamic conservation laws for the gas phase with a continuum model for the dispersed phase, as formulated by Nigmatulin. Inter-phase mass, momentum and energy exchange are prescribed by phenomenological models. It incorporates a combustion model based on the mass conservation laws for fuel, air and products; source/sink terms are treated in the fast-chemistry limit appropriate for such gasdynamic fields, along with a model for mass transfer from the particle phase to the gas. The model takes intomore » account both the afterburning of the detonation products of the C-4 booster with air, and the combustion of the Al particles with air. The model equations were integrated by high-order Godunov schemes for both the gas and particle phases. Numerical simulations of the explosion fields from 1.5-g Shock-Dispersed-Fuel (SDF) charge in a 6.6 liter calorimeter were used to validate the combustion model. Then the model was applied to 10-kg Al-SDF explosions in a an unconfined height-of-burst explosion. Computed pressure histories are compared with measured waveforms. Differences are caused by physical-chemical kinetic effects of particle combustion which induce ignition delays in the initial reactive blast wave and quenching of reactions at late times. Current simulations give initial insights into such modeling issues.« less

Azimuthally spinning wave modes and heat release in an annular combustor

NASA Astrophysics Data System (ADS)

Nygard, Hakon; Mazur, Marek; Dawson, James R.; Worth, Nicholas A.

2017-11-01

In order to reduce NOx emissions from aeroengines and stationary gas turbines the fuel-air mixture can be made leaner, at the risk of introducing potentially damaging thermo-acoustic instabilities. At present this phenomenon is not understood well enough to eliminate these instabilities at the design stage. Recently, the presence of different azimuthal modes in annular combustors has been demonstrated both experimentally and numerically. These naturally occurring instabilities in annular geometry have been observed to constantly switch between spinning and standing modes, making it more difficult to analyse the flame structure and dynamics. Very recently this issue was partially addressed using novel acoustic forcing to generate a standing mode. In the present study this concept has been developed further by creating an azimuthal array of loud speakers, which for the first time permits predominantly spinning modes to be set up inside the combustion chamber. The use of pressure and high speed OH* measurements enables the study of the flame dynamics and heat release rate oscillations of the combustor, which will be reported in the current paper. The ability to precisely control the azimuthal mode of oscillation greatly enhances our further understanding of the phenomenon. This project has received funding from the European Research Council (ERC) under the European Union's Horizon 2020 research and innovation programme (Grant Agreement No 677931 TAIAC).

A statistical model for combustion resonance from a DI diesel engine with applications

NASA Astrophysics Data System (ADS)

Bodisco, Timothy; Low Choy, Samantha; Masri, Assaad; Brown, Richard J.

2015-08-01

Introduced in this paper is a Bayesian model for isolating the resonant frequency from combustion chamber resonance. The model shown in this paper focused on characterising the initial rise in the resonant frequency to investigate the rise of in-cylinder bulk temperature associated with combustion. By resolving the model parameters, it is possible to determine: the start of pre-mixed combustion, the start of diffusion combustion, the initial resonant frequency, the resonant frequency as a function of crank angle, the in-cylinder bulk temperature as a function of crank angle and the trapped mass as a function of crank angle. The Bayesian method allows for individual cycles to be examined without cycle-averaging-allowing inter-cycle variability studies. Results are shown for a turbo-charged, common-rail compression ignition engine run at 2000 rpm and full load.

Discrete model of gas-free spin combustion of a powder mixture

NASA Astrophysics Data System (ADS)

Klimenok, Kirill L.; Rashkovskiy, Sergey A.

2015-01-01

We propose a discrete model of gas-free combustion of a cylindrical sample which reproduces in detail a spin combustion mode. It is shown that a spin combustion, in its classical sense as a continuous spiral motion of heat release zones on the surface of the sample, does not exist. Such a concept has arisen due to the misinterpretation of the experimental data. This study shows that in fact a spinlike combustion is realized, at which two energy release zones appear on the lateral surface of the sample and propagate circumferentially in the opposite directions. After some time two new heat release zones are formed on the next layer of the cylinder surface and make the same counter-circular motion. This process continues periodically and from a certain angle it looks like a spiral movement of the luminous zone along the lateral surface of the sample. The model shows that on approaching the combustion limit the process becomes more complicated and the spinlike combustion mode shifts to a more complex mode with multiple zones of heat release moving in different directions along the lateral surface. It is shown that the spin combustion mode appears due to asymmetry of initial conditions and always transforms into a layer-by-layer combustion mode with time.

Discrete model of gas-free spin combustion of a powder mixture.

PubMed

Klimenok, Kirill L; Rashkovskiy, Sergey A

2015-01-01

We propose a discrete model of gas-free combustion of a cylindrical sample which reproduces in detail a spin combustion mode. It is shown that a spin combustion, in its classical sense as a continuous spiral motion of heat release zones on the surface of the sample, does not exist. Such a concept has arisen due to the misinterpretation of the experimental data. This study shows that in fact a spinlike combustion is realized, at which two energy release zones appear on the lateral surface of the sample and propagate circumferentially in the opposite directions. After some time two new heat release zones are formed on the next layer of the cylinder surface and make the same counter-circular motion. This process continues periodically and from a certain angle it looks like a spiral movement of the luminous zone along the lateral surface of the sample. The model shows that on approaching the combustion limit the process becomes more complicated and the spinlike combustion mode shifts to a more complex mode with multiple zones of heat release moving in different directions along the lateral surface. It is shown that the spin combustion mode appears due to asymmetry of initial conditions and always transforms into a layer-by-layer combustion mode with time.

Numerical Investigation of the Acoustic Damping of Plane Acoustic Waves by Perforated Liners with Bias Flow

NASA Astrophysics Data System (ADS)

Zhao, Dan; Zhong, Zhi Yuan

Perforated liners are extensively used in aero-engines and gas turbine combustors to suppress combustion instabilities. These liners, typically subjected to a low Mach number bias flow (a cooling flow through perforated holes), are fitted along the bounding walls of a combustor to convert acoustic energy into flow energy by generating vorticity at the rims of the perforated apertures. To investigate the acoustic damping of such liners with bias flow on plane acoustic waves, a time-domain numerical model is developed to compute acoustic wave propagation in a cylindrical duct with a single-layer liner attached. The damping mechanism of the liner is characterized in real-time by using a 'compliance', developed especially for this work. It is a rational function representation of the frequency-domain homogeneous compliance adapted from the Rayleigh conductivity of a single aperture with mean bias flow in the z-domain. The liner 'compliance' model is then incorporated into partial differential equations of the duct system, which are solved by using the method of lines. The numerical results are then evaluated by comparing with the numerical results of Eldredge and Dowling's frequency-domain model. Good agreement is observed. This confirms that the model and the approach developed are suitable for real-time characterizing the acoustic damping of perforated liners.

Modeling of a Sequential Two-Stage Combustor

NASA Technical Reports Server (NTRS)

Hendricks, R. C.; Liu, N.-S.; Gallagher, J. R.; Ryder, R. C.; Brankovic, A.; Hendricks, J. A.

2005-01-01

A sequential two-stage, natural gas fueled power generation combustion system is modeled to examine the fundamental aerodynamic and combustion characteristics of the system. The modeling methodology includes CAD-based geometry definition, and combustion computational fluid dynamics analysis. Graphical analysis is used to examine the complex vortical patterns in each component, identifying sources of pressure loss. The simulations demonstrate the importance of including the rotating high-pressure turbine blades in the computation, as this results in direct computation of combustion within the first turbine stage, and accurate simulation of the flow in the second combustion stage. The direct computation of hot-streaks through the rotating high-pressure turbine stage leads to improved understanding of the aerodynamic relationships between the primary and secondary combustors and the turbomachinery.

Combustion Of Metals In Reduced Gravity And Extraterrestrial Environments

NASA Technical Reports Server (NTRS)

Abbud-Madrid, A.; Modak, A.; Branch, M. C.

2003-01-01

The recent focus of this research project has been to model the combustion of isolated metal droplets and, in particular, to couple the existing theories and formulations of phenomena such as condensation, reaction kinetics, radiation, and surface reactions to formulate a more complete combustion model. A fully transient, one-dimensional (spherical symmetry) numerical model that uses detailed chemical kinetics, multi-component molecular transport mechanisms, condensation kinetics, and gas phase radiation heat transfer was developed. A coagulation model was used to simulate the particulate formation of MgO. The model was used to simulate the combustion of an Mg droplet in pure O2 and CO2. Methanol droplet combustion is considered as a test case for the solution method for both quasi-steady and fully transient simulations. Although some important processes unique to methanol combustion, such as water absorption at the surface, are not included in the model, the results are in sufficient agreement with the published data. Since the major part of the heat released in combustion of Mg, and in combustion of metals in general, is due to the condensation of the metal oxide, it is very important to capture the condensation processes correctly. Using the modified nucleation theory, an Arrhenius type rate expression is derived to calculate the condensation rate of MgO. This expression can be easily included in the CHEMKIN reaction mechanism format. Although very little property data is available for MgO, the condensation rate expression derived using the existing data is able to capture the condensation of MgO. An appropriate choice of the reference temperature to calculate the rate coefficients allows the model to correctly predict the subsequent heat release and hence the flame temperature.

Modeling of natural acoustic frequencies of a gas-turbine plant combustion chamber

NASA Astrophysics Data System (ADS)

Zubrilin, I. A.; Gurakov, N. I.; Zubrilin, R. A.; Matveev, S. G.

2017-05-01

The paper presents results of determination of natural acoustic frequencies of a gas-turbine plant annular combustion chamber model using 3D-simulation. At the beginning, a calculation procedure for determining natural acoustic frequencies of the gas-turbine plant combustion chamber was worked out. The effect of spatial inhomogeneity of the flow parameters (fluid composition, pressure, temperature) arising in combustion and some geometrical parameters (cooling holes of the flame tube walls) on the calculation results is studied. It is found that the change of the fluid composition in combustion affects the acoustic velocity not more than 5%; therefore, the air with a volume variable temperature can be taken as a working fluid in the calculation of natural acoustic frequencies. It is also shown that the cooling holes of the flame tube walls with diameter less than 2 mm can be neglected in the determination of the acoustic modes in the frequency range of up to 1000 Hz. This reduces the number of the grid-model elements by a factor of six in comparison with a model that considers all of the holes. Furthermore, a method of export of spatial inhomogeneity of the flow parameters from a CFD solver sector model to the annular combustion chamber model in a modal solver is presented. As a result of the obtained model calculation, acoustic modes of the combustion chamber in the frequency range of up to 1000 Hz are determined. For a standard engine condition, a potentially dangerous acoustic mode with a frequency close to the ripple frequency of the precessing vortex core, which is formed behind the burner device of this combustion chamber, is detected.

NASA Tech Briefs, July 2005

NASA Technical Reports Server (NTRS)

2005-01-01

Thin-Film Resistance Heat-Flux Sensors Circuit Indicates that Voice-Recording Disks are Nearly Full Optical Sensing of Combustion Instabilities in Gas Turbines Topics include: Crane-Load Contact Sensor; Hexagonal and Pentagonal Fractal Multiband Antennas; Multifunctional Logic Gate Controlled by Temperature; Multifunctional Logic Gate Controlled by Supply Voltage; Power Divider for Waveforms Rich in Harmonics; SCB Quantum Computers Using iSWAP and 1-Qubit Rotations; CSAM Metrology Software Tool; Update on Rover Sequencing and Visualization Program; Selecting Data from a Star Catalog; Rotating Desk for Collaboration by Two Computer Programmers; Variable-Pressure Washer; Magnetically Attached Multifunction Maintenance Rover; Improvements in Fabrication of Sand/Binder Cores for Casting; Solid Freeform Fabrication of Composite-Material Objects; Efficient Computational Model of Hysteresis; Gauges for Highly Precise Metrology of a Compound Mirror; Improved Electrolytic Hydrogen Peroxide Generator; High-Power Fiber Lasers Using Photonic Band Gap Materials; Ontology-Driven Information Integration; Quantifying Traversability of Terrain for a Mobile Robot; More About Arc-Welding Process for Making Carbon Nanotubes; Controlling Laser Spot Size in Outer Space; or Software-Reconfigurable Processors for Spacecraft.

A CFD model for biomass combustion in a packed bed furnace

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

Karim, Md. Rezwanul; Department of Mechanical & Chemical Engineering, Islamic University of Technology, Gazipur 1704; Ovi, Ifat Rabbil Qudrat

Climate change has now become an important issue which is affecting environment and people around the world. Global warming is the main reason of climate change which is increasing day by day due to the growing demand of energy in developed countries. Use of renewable energy is now an established technique to decrease the adverse effect of global warming. Biomass is a widely accessible renewable energy source which reduces CO{sub 2} emissions for producing thermal energy or electricity. But the combustion of biomass is complex due its large variations and physical structures. Packed bed or fixed bed combustion is themore » most common method for the energy conversion of biomass. Experimental investigation of packed bed biomass combustion is difficult as the data collection inside the bed is challenging. CFD simulation of these combustion systems can be helpful to investigate different operational conditions and to evaluate the local values inside the investigation area. Available CFD codes can model the gas phase combustion but it can’t model the solid phase of biomass conversion. In this work, a complete three-dimensional CFD model is presented for numerical investigation of packed bed biomass combustion. The model describes the solid phase along with the interface between solid and gas phase. It also includes the bed shrinkage due to the continuous movement of the bed during solid fuel combustion. Several variables are employed to represent different parameters of solid mass. Packed bed is considered as a porous bed and User Defined Functions (UDFs) platform is used to introduce solid phase user defined variables in the CFD. Modified standard discrete transfer radiation method (DTRM) is applied to model the radiation heat transfer. Preliminary results of gas phase velocity and pressure drop over packed bed have been shown. The model can be useful for investigation of movement of the packed bed during solid fuel combustion.« less

Investigation of Ignition and Combustion Processes of Diesel Engines Operating with Turbulence and Air-storage Chambers

NASA Technical Reports Server (NTRS)

Petersen, Hans

1938-01-01

The flame photographs obtained with combustion-chamber models of engines operating respectively, with turbulence chamber and air-storage chambers or cells, provide an insight into the air and fuel movements that take place before and during combustion in the combustion chamber. The relation between air velocity, start of injection, and time of combustion was determined for the combustion process employing a turbulence chamber.

Analysis of unsteady reacting flows and impact of chemistry description in Large Eddy Simulations of side-dump ramjet combustors

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

Roux, A.; Gicquel, L.Y.M.; Staffelbach, G.

2010-01-15

Among all the undesired phenomena observed in ramjet combustors, combustion instabilities are of foremost importance and predicting them using Large Eddy Simulation (LES) is an active research field. While acoustics are naturally captured by compressible LES provided that the proper boundary conditions are applied, combustion/chemistry modelling remains a critical issue and its impact on numerical predictions must still be assessed for complex applications. To do so, two different ramjet LES's are compared here. The first simulation is based on a standard one-step chemistry known to over-estimate the laminar flame speed in fuel rich conditions. The second simulation uses the samemore » scheme but introduces a correction of reaction rates for rich flames to match a detailed mechanism provided by Peters (1993). Even though the two chemical schemes are very similar and very few points burn in rich regimes, distinct limit-cycles are obtained with LES depending on which scheme is used. Results obtained with the standard one-step chemistry exhibit high frequency self-sustained oscillations. Multiple flame fronts are stabilized in the vicinity of the shear layer developing at the exit of the air inlets. When compared to the experiment, the fitted one-step scheme yields better predictions than the standard scheme. With the fitted scheme, the flame is detached from the air inlets and stabilizes in the regions identified in the experiment (Ristori et al. (2005), Heid and Ristori (2003), Heid and Ristori (2005), Ristori et al. (1999)). LES and experiments exhibit all main low-frequency modes including the first longitudinal acoustic mode. The high frequencies excited with the standard scheme are damped with the fitted scheme. The chemical scheme is found, for this ramjet burner, to have a strong impact on the predicted stability: approximate chemical schemes even in a limited range of equivalence ratio can lead to the occurence of non-physical combustion oscillations. (author)« less

The effect of fuel inlet turbulence intensity on H2/CH4 flame structure of MILD combustion using the LES method

NASA Astrophysics Data System (ADS)

Afarin, Yashar; Tabejamaat, Sadegh

2013-06-01

Large eddy simulations (LES) are employed to investigate the effect of the inlet turbulence intensity on the H2/CH4 flame structure in a hot and diluted co-flow stream which emulates the (Moderate or Intense Low-oxygen Dilution) MILD combustion regime. In this regard, three fuel inlet turbulence intensity profiles with the values of 4%, 7% and 10% are superimposed on the annular mixing layer. The effects of these changes on the flame structure under the MILD condition are studied for two oxygen concentrations of 3% and 9% (by mass) in the oxidiser stream and three hot co-flow temperatures 1300, 1500 and 1750 K. The turbulence-chemistry interaction of the numerically unresolved scales is modelled using the (Partially Stirred Reactor) PaSR method, where the full mechanism of GRI-2.11 represents the chemical reactions. The influences of the turbulence intensity on the flame structure under the MILD condition are studied by using the profile of temperature, CO and OH mass fractions in both physical and mixture fraction spaces at two downstream locations. Also, the effects of this parameter are investigated by contours of OH, HCO and CH2O radicals in an area near the nozzle exit zone. Results show that increasing the fuel inlet turbulence intensity has a profound effect on the flame structure particularly at low oxygen mass fraction. This increment weakens the combustion zone and results in a decrease in the peak values of the flame temperature and OH and CO mass fractions. Furthermore, increasing the inlet turbulence intensity decreases the flame thickness, and increases the MILD flame instability and diffusion of un-burnt fuel through the flame front. These effects are reduced by increasing the hot co-flow temperature which reinforces the reaction zone.

High temperature high velocity direct power extraction using an open-cycle oxy-combustion system

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

Love, Norman

The implementation of oxy-fuel technology in fossil-fuel power plants may contribute to increased system efficiencies and a reduction of pollutant emissions. One technology that has potential to utilize the temperature of undiluted oxy-combustion flames is open-cycle magnetohydrodynamic (MHD) power generators. These systems can be configured as a topping cycle and provide high enthalpy, electrically conductive flows for direct conversion of electricity. This report presents the design and modeling strategies of a MHD combustor operating at temperatures exceeding 3000 K. Throughout the study, computational fluid dynamics (CFD) models were extensively used as a design and optimization tool. A lab-scale 60 kWthmore » model was designed, manufactured and tested as part of this project. A fully-coupled numerical method was developed in ANSYS FLUENT to characterize the heat transfer in the system. This study revealed that nozzle heat transfer may be predicted through a 40% reduction of the semi-empirical Bartz correlation. Experimental results showed good agreement with the numerical evaluation, with the combustor exhibiting a favorable performance when tested during extended time periods. A transient numerical method was employed to analyze fuel injector geometries for the 60-kW combustor. The ANSYS FLUENT study revealed that counter-swirl inlets achieve a uniform pressure and velocity ratio when the ports of the injector length to diameter ratio (L/D) is 4. An angle of 115 degrees was found to increase distribution efficiency. The findings show that this oxy-combustion concept is capable of providing a high-enthalpy environment for seeding, in order to render the flow to be conductive. Based on previous findings, temperatures in the range of 2800-3000 K may enable magnetohydrodynamic power extraction. The heat loss fraction in this oxy-combustion system, based on CFD and analytical calculations, at optimal operating conditions, was estimated to be less than 10 percent. Furthermore, the heat transfer design removed approximately 7 MW/m2. The results observed in the lab-scale system were employed to develop a 1-MW scaled prototype. Scaling methods were based on critical design criteria found in similar systems, aimed at replicating combustion flow fields and reducing possible instabilities. A numerical simulation of the combustor wall was developed for a combined thermal steady model and static structural model. This combined model was developed predict combined stress parameters within the wall during testing conditions. Both models were developed within ANSYS FEA software package. The relative accuracy presented as well major performance parameters are discussed to assess the design's validity and ensure safety. The scaled prototype was manufactured through selective laser melting (SLM)-based additive manufacturing to reduce lead times and increase geometrical complexity. Additional CFD models were developed to optimize coolant manifold system parameters and perform a parametric study on channel geometry. An investigation on coolant manifold geometry demonstrated improvements in channel flow distribution when enlarging manifold lengths and increasing the number of tubes feeding into the flow. A three-dimensional model based on a single channel was developed to capture the effect of variable properties and thermal stratification. All cases in the simulation exhibited higher wall temperatures and lower convective coefficients than those determined through 1-D analytical equations. This implies that pressure and velocity safety factors must be implemented during system operation. Overall, the findings made in this investigation are thought to be of value to researchers and industrial practitioners when designing oxy-fuel direct power extraction systems operating at temperatures exceeding 3000 K. In addition to this, the implementation of the developed technology at pilot and commercial scales could result in a significant improvement in the efficiencies of heritage and next-generation power cycles.« less

Spray combustion model improvement study, 1

NASA Technical Reports Server (NTRS)

Chen, C. P.; Kim, Y. M.; Shang, H. M.

1993-01-01

This study involves the development of numerical and physical modeling in spray combustion. These modeling efforts are mainly motivated to improve the physical submodels of turbulence, combustion, atomization, dense spray effects, and group vaporization. The present mathematical formulation can be easily implemented in any time-marching multiple pressure correction methodologies such as MAST code. A sequence of validation cases includes the nonevaporating, evaporating and_burnin dense_sprays.

Two-Phase Model of Combustion in Explosions

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

Kuhl, A L; Khasainov, B; Bell, J

2006-06-19

A two-phase model for Aluminum particle combustion in explosions is proposed. It combines the gas-dynamic conservation laws for the gas phase with the continuum mechanics laws of multi-phase media, as formulated by Nigmatulin. Inter-phase mass, momentum and energy exchange are prescribed by the Khasainov model. Combustion is specified as material transformations in the Le Chatelier diagram which depicts the locus of thermodynamic states in the internal energy-temperature plane according to Kuhl. Numerical simulations are used to show the evolution of two-phase combustion fields generated by the explosive dissemination of a powdered Al fuel.

The numerical modelling and process simulation for the fault diagnosis of rotary kiln incinerator.

PubMed

Roh, S D; Kim, S W; Cho, W S

2001-10-01

The numerical modelling and process simulation for the fault diagnosis of rotary kiln incinerator were accomplished. In the numerical modelling, two models applied to the modelling within the kiln are the combustion chamber model including the mass and energy balance equations for two combustion chambers and 3D thermal model. The combustion chamber model predicts temperature within the kiln, flue gas composition, flux and heat of combustion. Using the combustion chamber model and 3D thermal model, the production-rules for the process simulation can be obtained through interrelation analysis between control and operation variables. The process simulation of the kiln is operated with the production-rules for automatic operation. The process simulation aims to provide fundamental solutions to the problems in incineration process by introducing an online expert control system to provide an integrity in process control and management. Knowledge-based expert control systems use symbolic logic and heuristic rules to find solutions for various types of problems. It was implemented to be a hybrid intelligent expert control system by mutually connecting with the process control systems which has the capability of process diagnosis, analysis and control.

Global burden of mortalities due to chronic exposure to ambient PM2.5 from open combustion of domestic waste

NASA Astrophysics Data System (ADS)

Kodros, John K.; Wiedinmyer, Christine; Ford, Bonne; Cucinotta, Rachel; Gan, Ryan; Magzamen, Sheryl; Pierce, Jeffrey R.

2016-12-01

Uncontrolled combustion of domestic waste has been observed in many countries, creating concerns for air quality; however, the health implications have not yet been quantified. We incorporate the Wiedinmyer et al (2014 Environ. Sci. Technol. 48 9523-30) emissions inventory into the global chemical-transport model, GEOS-Chem, and provide a first estimate of premature adult mortalities from chronic exposure to ambient PM2.5 from uncontrolled combustion of domestic waste. Using the concentration-response functions (CRFs) of Burnett et al (2014 Environ. Health Perspect. 122 397-403), we estimate that waste-combustion emissions result in 270 000 (5th-95th: 213 000-328 000) premature adult mortalities per year. The confidence interval results only from uncertainty in the CRFs and assumes equal toxicity of waste-combustion PM2.5 to all other PM2.5 sources. We acknowledge that this result is likely sensitive to choice of chemical-transport model, CRFs, and emission inventories. Our central estimate equates to 9% of adult mortalities from exposure to ambient PM2.5 reported in the Global Burden of Disease Study 2010. Exposure to PM2.5 from waste combustion increases the risk of premature mortality by more than 0.5% for greater than 50% of the population. We consider sensitivity simulations to uncertainty in waste-combustion emission mass, the removal of waste-combustion emissions, and model resolution. A factor-of-2 uncertainty in waste-combustion PM2.5 leads to central estimates ranging from 138 000 to 518 000 mortalities per year for factors-of-2 reductions and increases, respectively. Complete removal of waste combustion would only avoid 191 000 (5th-95th: 151 000-224 000) mortalities per year (smaller than the total contributed premature mortalities due to nonlinear CRFs). Decreasing model resolution from 2° × 2.5° to 4° × 5° results in 16% fewer mortalities attributed to waste-combustion PM2.5, and over Asia, decreasing resolution from 0.5° × 0.666° to 2° × 2.5° results in 21% fewer mortalities attributed to waste-combustion PM2.5. Owing to coarse model resolution, our global estimates of premature mortality from waste-combustion PM2.5 are likely a lower bound.

Symposium /International/ on Combustion, 18th, University of Waterloo, Waterloo, Ontario, Canada, August 17-22, 1980, Proceedings

NASA Technical Reports Server (NTRS)

1981-01-01

Problems related to combustion generated pollution are explored, taking into account the mechanism of NO formation from nitrogen compounds in hydrogen flames studied by laser fluorescence, the structure and similarity of nitric oxide production in turbulent diffusion flames, the effect of steam addition on NO formation, and the formation of NO2 by laminar flames. Other topics considered are concerned with propellant combustion, fluidized bed combustion, the combustion of droplets and sprays, premixed flame studies, fire studies, and flame stabilization. Attention is also given to coal flammability, chemical kinetics, turbulent combustion, soot, coal combustion, the modeling of combustion processes, combustion diagnostics, detonations and explosions, ignition, internal combustion engines, combustion studies, and furnaces.

Convective Electrokinetic Instability With Conductivity Gradients

NASA Astrophysics Data System (ADS)

Chen, Chuan-Hua; Lin, Hao; Lele, Sanjiva; Santiago, Juan

2003-11-01

Electrokinetic flow instability has been experimentally identified and quantified in a glass T-junction microchannel system with a cross section of 11 um x 155 um. In this system, buffers of different conductivities were electrokinetically driven into a common mixing channel by a DC electric field. A convective instability was observed with a threshold electric field of 0.45 kV/cm for a 10:1 conductivity ratio. A physical model has been developed which consists of a modified Ohmic model formulation for electrolyte solutions and the Navier-Stokes equations with an electric body force term. The model and experiments show that bulk charge accumulation in regions of conductivity gradients is the key mechanism of such instabilities. A linear stability analysis was performed in a convective framework, and Briggs-Bers criteria were applied to determine the nature of instability. The analysis shows the instability is governed by two key parameters: the ratio of molecular diffusion to electroviscous time scale which governs the onset of instability, and the ratio of electroviscous to electroosmotic velocity which governs whether the instability is convective or absolute. The model predicted critical electric field, growth rate, wavelength, and phase speed which were comparable to experimental data.

Dynamical instabilities in axisymmetric stellar systems. I - Oblate E6 models

NASA Technical Reports Server (NTRS)

Levison, Harold F.; Duncan, Martin J.; Smith, Bruce F.

1990-01-01

The stability of a set of models based on isothermal oblate E6 elliptical galaxies is studied using N-body techniques. The only stable models are those that are near the isotropic model and have nearly equal number of stars in retrograde and prograde orbits. Fast rotators are unstable to modes that appear to be analogous to the classical streaming instability seen in many disk systems. Systems with a large velocity dispersion in the direction of the cylindrical radius are unstable to modes that appear to be similar to the radial orbit instability observed in some spherical systems. However, evidence is presented that these two instabilities may be related, and an instability criterion that applies to both is constructed.

Combustion properties of Kraft Black Liquors

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

Frederick, W.J. Jr.; Hupa, M.

1993-04-01

In a previous study of the phenomena involved in the combustion of black liquor droplets a numerical model was developed. The model required certain black liquor specific combustion information which was then not currently available, and additional data were needed for evaluating the model. The overall objectives of the project reported here was to provide experimental data on key aspects of black liquor combustion, to interpret the data, and to put it into a form which would be useful for computational models for recovery boilers. The specific topics to be investigated were the volatiles and char carbon yields from pyrolysismore » of single black liquor droplets; a criterion for the onset of devolatilization and the accompanying rapid swelling; and the surface temperature of black liquor droplets during pyrolysis, combustion, and gasification. Additional information on the swelling characteristics of black liquor droplets was also obtained as part of the experiments conducted.« less

Experiments on the Richtmyer-Meshkov Instability of Incompressible Fluids

NASA Technical Reports Server (NTRS)

Jacobs, J.; Niederhaus, C.

2000-01-01

Richtmyer-Meshkov (R-M) instability occurs when two different density fluids are impulsively accelerated in the direction normal to their nearly planar interface. The instability causes small perturbations on the interface to grow and possibly become turbulent given the proper initial conditions. R-M instability is similar to the Rayleigh-Taylor (R-T) instability, which is generated when the two fluids undergo a constant acceleration. R-M instability is a fundamental fluid instability that is important to fields ranging from astrophysics to high-speed combustion. For example, R-M instability is currently the limiting factor in achieving a net positive yield with inertial confinement fusion. The experiments described here utilize a novel technique that circumvents many of the experimental difficulties previously limiting the study of the R-M instability. A Plexiglas tank contains two unequal density liquids and is gently oscillated horizontally to produce a controlled initial fluid interface shape. The tank is mounted to a sled on a high speed, low friction linear rail system, constraining the main motion to the vertical direction. The sled is released from an initial height and falls vertically until it bounces off of a movable spring, imparting an impulsive acceleration in the upward direction. As the sled travels up and down the rails, the spring retracts out of the way, allowing the instability to evolve in free-fall until impacting a shock absorber at the end of the rails. The impulsive acceleration provided to the system is measured by a piezoelectric accelerometer mounted on the tank, and a capacitive accelerometer measures the low-level drag of the bearings. Planar Laser-Induced Fluorescence is used for flow visualization, which uses an Argon ion laser to illuminate the flow and a CCD camera, mounted to the sled, to capture images of the interface. This experimental study investigates the instability of an interface between incompressible, miscible liquids with an initial sinusoidal perturbation. The amplitude of the disturbance during the experiment is measured and compared to theory. The results show good agreement (within 10%) with linear stability theory up to nondimensional amplitude ka = 0.7 (wavenumber x amplitude). These results hold true for an initial ka (before acceleration) of -0.7 less than ka less than -0.06, while the linear theory was developed for absolute value of ka much less than 1. In addition, a third order weakly nonlinear perturbation theory is shown to be accurate for amplitudes as large as ka = 1.3, even though the interface becomes double-valued at ka = 1.1. As time progresses, the vorticity on the interface concentrates, and the interface spirals around the alternating sign vortex centers to form a mushroom pattern. At higher Reynolds Number (based on circulation), an instability of the vortex cores has been observed. While time limitations of the apparatus prevent determination of a critical Reynolds Number, the lowest Reynolds Number this vortex instability has been observed at is 5000.

Saw-tooth instability in storage rings: simulations and dynamical model

NASA Astrophysics Data System (ADS)

Migliorati, M.; Palumbo, L.; Dattoli, G.; Mezi, L.

1999-11-01

The saw-tooth instability in storage rings is studied by means of a time-domain simulation code which takes into account the self-induced wake fields. The results are compared with those from a dynamical heuristic model exploiting two coupled non-linear differential equations, accounting for the time behavior of the instability growth rate and for the anomalous growth of the energy spread. This model is shown to reproduce the characteristic features of the instability in a fairly satisfactory way.

An assessment of thermodynamic merits for current and potential future engine operating strategies

DOE PAGES

Wissink, Martin L.; Splitter, Derek A.; Dempsey, Adam B.; ...

2017-02-01

The present work compares the fundamental thermodynamic underpinnings (i.e., working fluid properties and heat release profile) of various combustion strategies with engine measurements. The approach employs a model that separately tracks the impacts on efficiency due to differences in rate of heat addition, volume change, mass addition, and molecular weight change for a given combination of working fluid, heat release profile, and engine geometry. Comparative analysis between measured and modeled efficiencies illustrates fundamental sources of efficiency reductions or opportunities inherent to various combustion regimes. Engine operating regimes chosen for analysis include stoichiometric spark-ignited combustion and lean compression-ignited combustion including HCCI,more » SA-HCCI, RCCI, GCI, and CDC. Within each combustion regime, effects such as engine load, combustion duration, combustion phasing, combustion chamber geometry, fuel properties, and charge dilution are explored. Model findings illustrate that even in the absence of losses such as heat transfer or incomplete combustion, the maximum possible thermal efficiency inherent to each operating strategy varies to a significant degree. Additionally, the experimentally measured losses are observed to be unique within a given operating strategy. The findings highlight the fact that in order to create a roadmap for future directions in ICE technologies, it is important to not only compare the absolute real-world efficiency of a given combustion strategy, but to also examine the measured efficiency in context of what is thermodynamically possible with the working fluid and boundary conditions prescribed by a strategy.« less

An assessment of thermodynamic merits for current and potential future engine operating strategies

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

Wissink, Martin L.; Splitter, Derek A.; Dempsey, Adam B.

The present work compares the fundamental thermodynamic underpinnings (i.e., working fluid properties and heat release profile) of various combustion strategies with engine measurements. The approach employs a model that separately tracks the impacts on efficiency due to differences in rate of heat addition, volume change, mass addition, and molecular weight change for a given combination of working fluid, heat release profile, and engine geometry. Comparative analysis between measured and modeled efficiencies illustrates fundamental sources of efficiency reductions or opportunities inherent to various combustion regimes. Engine operating regimes chosen for analysis include stoichiometric spark-ignited combustion and lean compression-ignited combustion including HCCI,more » SA-HCCI, RCCI, GCI, and CDC. Within each combustion regime, effects such as engine load, combustion duration, combustion phasing, combustion chamber geometry, fuel properties, and charge dilution are explored. Model findings illustrate that even in the absence of losses such as heat transfer or incomplete combustion, the maximum possible thermal efficiency inherent to each operating strategy varies to a significant degree. Additionally, the experimentally measured losses are observed to be unique within a given operating strategy. The findings highlight the fact that in order to create a roadmap for future directions in ICE technologies, it is important to not only compare the absolute real-world efficiency of a given combustion strategy, but to also examine the measured efficiency in context of what is thermodynamically possible with the working fluid and boundary conditions prescribed by a strategy.« less

Kinetics of devolatilization and oxidation of a pulverized biomass in an entrained flow reactor under realistic combustion conditions

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

Jimenez, Santiago; Remacha, Pilar; Ballester, Javier

2008-03-15

In this paper the results of a complete set of devolatilization and combustion experiments performed with pulverized ({proportional_to}500 {mu}m) biomass in an entrained flow reactor under realistic combustion conditions are presented. The data obtained are used to derive the kinetic parameters that best fit the observed behaviors, according to a simple model of particle combustion (one-step devolatilization, apparent oxidation kinetics, thermally thin particles). The model is found to adequately reproduce the experimental trends regarding both volatile release and char oxidation rates for the range of particle sizes and combustion conditions explored. The experimental and numerical procedures, similar to those recentlymore » proposed for the combustion of pulverized coal [J. Ballester, S. Jimenez, Combust. Flame 142 (2005) 210-222], have been designed to derive the parameters required for the analysis of biomass combustion in practical pulverized fuel configurations and allow a reliable characterization of any finely pulverized biomass. Additionally, the results of a limited study on the release rate of nitrogen from the biomass particle along combustion are shown. (author)« less

Computational Analysis of End-of-Injection Transients and Combustion Recession

NASA Astrophysics Data System (ADS)

Jarrahbashi, Dorrin; Kim, Sayop; Knox, Benjamin W.; Genzale, Caroline L.; Georgia Institute of Technology Team

2016-11-01

Mixing and combustion of ECN Spray A after end of injection are modeled with different chemical kinetics models to evaluate the impact of mechanism formulation and low-temperature chemistry on predictions of combustion recession. Simulations qualitatively agreed with the past experimental observations of combustion recession. Simulations with the Cai mechanism show second-stage ignition in distinct regions near the nozzle, initially spatially separated from the lifted diffusion flame, but then rapidly merge with flame. By contrast, the Yao mechanism fails to predict sufficient low-temperature chemistry in mixtures upstream of the diffusion flame and combustion recession. The effects of the shape and duration of the EOI transient on the entrainment wave near the nozzle, the likelihood of combustion recession, and the spatiotemporal development of mixing and chemistry in near-nozzle mixtures are also investigated. With a more rapid ramp-down injection profile, a weaker combustion recession occurs. For extremely fast ramp-down, the entrainment flux varies rapidly near the nozzle and over-leaning of the mixture completely suppresses combustion recession. For a slower ramp-down profile complete combustion recession back toward the nozzle is observed.

Instability in interacting dark sector: an appropriate holographic Ricci dark energy model

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

Herrera, Ramón; Hipólito-Ricaldi, W.S.; Videla, Nelson, E-mail: ramon.herrera@pucv.cl, E-mail: wiliam.ricaldi@ufes.br, E-mail: nelson.videla@ing.uchile.cl

In this paper we investigate the consequences of phantom crossing considering the perturbative dynamics in models with interaction in their dark sector. By mean of a general study of gauge-invariant variables in comoving gauge, we relate the sources of instabilities in the structure formation process with the phantom crossing. In order to illustrate these relations and its consequences in more detail, we consider a specific case of an holographic dark energy interacting with dark matter. We find that in spite of the model is in excellent agreement with observational data at background level, however it is plagued of instabilities inmore » its perturbative dynamics. We reconstruct the model in order to avoid these undesirable instabilities, and we show that this implies a modification of the concordance model at background. Also we find drastic changes on the parameters space in our model when instabilities are avoided.« less

Numerical Simulation of the Thermal Process in a W-Shape Radiant Tube Burner

NASA Astrophysics Data System (ADS)

Wang, Yi; Li, Jiyong; Zhang, Lifeng; Ling, Haitao; Li, Yanlong

2014-07-01

In the current work, three-dimensional mathematical models were developed for the heat transfer and combustion in a W-shape radiant tube burner (RTB) and were solved using Fluent software (ANSYS Inc., Canonsburg, PA). The standard k- ɛ model, nonpremixed combustion model, and the discrete ordinate model were used for the modeling of turbulence, combustion, and radiant heat transfer, respectively. In addition, the NO x postprocessor was used for the prediction of the NO emission. A corresponding experiment was performed for the validation of mathematical models. The details of fluid flow, heat transfer, and combustion in the RTB were investigated. Moreover, the effect of the air/fuel ratio (A/F) and air staging on the performance of RTB was studied with the reference indexes including heat efficiency, maximum temperature difference on shell wall, and NO emission at the outlet. The results indicated that a low speed zone formed in the vicinity of the combustion chamber outlet, and there were two relative high-temperature zones in the RTB, one in combustion chamber that favored the flame stability and the other from the main flame in the RTB. The maximum temperature difference was 95.48 K. As the A/F increased, the temperature increased first and then decreased. As the ratio of the primary to secondary air increased, the recirculation zone at the outlet of combustion chamber shrank gradually to disappear, and the flame length was longer and the temperature in flame decreased correspondingly.

The influence of fuel type to combustion characteristic in diffusion flame drying by computational fluid dynamics simulation

NASA Astrophysics Data System (ADS)

Septiani, Eka Lutfi; Widiyastuti, W.; Machmudah, Siti; Nurtono, Tantular; Winardi, Sugeng

2017-05-01

Diffusion flame spray drying has become promising method in nanoparticles synthesis giving several advantages and low operation cost. In order to scale up the process which needs high experimentation time and cost, Computational Fluid Dynamics (CFD) by Ansys Fluent 15.0 software has been used. Combustion characteristic in diffusion flame reactor may affects particle size distribution. This study aims to observe influence of fuel type to combustion characteristic in the reactor. Large Eddy Simulation (LES) and non-premixed combustion model are selected for the turbulence and combustion model respectively. Methane, propane, and LPG in 0.5 L/min were used as type of fuel. While the oxidizer is air with 200% excess of O2. Simulation result shown that the maximum temperature was obtained from propane-air combustion in 2268 K. However, the stable temperature contour was achieved by methane-air combustion.

Advancing predictive models for particulate formation in turbulent flames via massively parallel direct numerical simulations

PubMed Central

Bisetti, Fabrizio; Attili, Antonio; Pitsch, Heinz

2014-01-01

Combustion of fossil fuels is likely to continue for the near future due to the growing trends in energy consumption worldwide. The increase in efficiency and the reduction of pollutant emissions from combustion devices are pivotal to achieving meaningful levels of carbon abatement as part of the ongoing climate change efforts. Computational fluid dynamics featuring adequate combustion models will play an increasingly important role in the design of more efficient and cleaner industrial burners, internal combustion engines, and combustors for stationary power generation and aircraft propulsion. Today, turbulent combustion modelling is hindered severely by the lack of data that are accurate and sufficiently complete to assess and remedy model deficiencies effectively. In particular, the formation of pollutants is a complex, nonlinear and multi-scale process characterized by the interaction of molecular and turbulent mixing with a multitude of chemical reactions with disparate time scales. The use of direct numerical simulation (DNS) featuring a state of the art description of the underlying chemistry and physical processes has contributed greatly to combustion model development in recent years. In this paper, the analysis of the intricate evolution of soot formation in turbulent flames demonstrates how DNS databases are used to illuminate relevant physico-chemical mechanisms and to identify modelling needs. PMID:25024412

Gaseous emissions from the combustion of a waste mixture containing a high concentration of N2O.

PubMed

Dong, Changqing; Yang, Yongping; Zhang, Junjiao; Lu, Xuefeng

2009-01-01

This paper is focused on reducing the emissions from the combustion of a waste mixture containing a high concentration of N2O. A rate model and an equilibrium model were used to predict gaseous emissions from the combustion of the mixture. The influences of temperature and methane were considered, and the experimental research was carried out in a tabular reactor and a pilot combustion furnace. The results showed that for the waste mixture, the combustion temperature should be in the range of 950-1100 degrees C and the gas residence time should be 2s or higher to reduce emissions.

Numerical model of two-dimensional heterogeneous combustion in porous media under natural convection or forced filtration

NASA Astrophysics Data System (ADS)

Lutsenko, Nickolay A.

2018-03-01

A novel mathematical model and original numerical method for investigating the two-dimensional waves of heterogeneous combustion in porous media are proposed and described in detail. The mathematical model is constructed within the framework of the model of interacting interpenetrating continua and includes equations of state, continuity, momentum conservation and energy for solid and gas phases. Combustion, considered in the paper, is due to the exothermic reaction between fuel in the porous solid medium and oxidiser contained in the gas flowing through the porous object. The original numerical method is based on a combination of explicit and implicit finite-difference schemes. A distinctive feature of the proposed model is that the gas velocity at the open boundaries (inlet and outlet) of the porous object is unknown and has to be found from the solution of the problem, i.e. the flow rate of the gas regulates itself. This approach allows processes to be modelled not only under forced filtration, but also under free convection, when there is no forced gas input in porous objects, which is typical for many natural or anthropogenic disasters (burning of peatlands, coal dumps, landfills, grain elevators). Some two-dimensional time-dependent problems of heterogeneous combustion in porous objects have been solved using the proposed numerical method. It is shown that two-dimensional waves of heterogeneous combustion in porous media can propagate in two modes with different characteristics, as in the case of one-dimensional combustion, but the combustion front can move in a complex manner, and gas dynamics within the porous objects can be complicated. When natural convection takes place, self-sustaining combustion waves can go through the all parts of the object regardless of where an ignition zone was located, so the all combustible material in each part of the object is burned out, in contrast to forced filtration.

Revised users manual, Pulverized Coal Gasification or Combustion: 2-dimensional (87-PCGC-2): Final report, Volume 2. [87-PCGC-2

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

Smith, P.J.; Smoot, L.D.; Brewster, B.S.

1987-12-01

A two-dimensional, steady-state model for describing a variety of reactive and non-reactive flows, including pulverized coal combustion and gasification, is presented. Recent code revisions and additions are described. The model, referred to as 87-PCGC-2, is applicable to cylindrical axi-symmetric systems. Turbulence is accounted for in both the fluid mechanics equations and the combustion scheme. Radiation from gases, walls, and particles is taken into account using either a flux method or discrete ordinates method. The particle phase is modeled in a Lagrangian framework, such that mean paths of particle groups are followed. Several multi-step coal devolatilization schemes are included along withmore » a heterogeneous reaction scheme that allows for both diffusion and chemical reaction. Major gas-phase reactions are modeled assuming local instantaneous equilibrium, and thus the reaction rates are limited by the turbulent rate mixing. A NO/sub x/ finite rate chemistry submodel is included which integrates chemical kinetics and the statistics of the turbulence. The gas phase is described by elliptic partial differential equations that are solved by an iterative line-by-line technique. Under-relaxation is used to achieve numerical stability. The generalized nature of the model allows for calculation of isothermal fluid mechanicsgaseous combustion, droplet combustion, particulate combustion and various mixtures of the above, including combustion of coal-water and coal-oil slurries. Both combustion and gasification environments are permissible. User information and theory are presented, along with sample problems. 106 refs.« less

Studies in nonlinear problems of energy. Progress report, October 1, 1993--September 30, 1994

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

Matkowsky, B.J.

1994-09-01

The authors concentrate on modeling, analysis and large scale scientific computation of combustion and flame propagation phenomena, with emphasis on the transition from laminar to turbulent combustion. In the transition process a flame passed through a stages exhibiting increasingly complex spatial and temporal patterns which serve as signatures identifying each stage. Often the transitions arise via bifurcation. The authors investigate nonlinear dynamics, bifurcation and pattern formation in the successive stage of transition. They describe the stability of combustion waves, and transitions to combustion waves exhibiting progressively higher degrees of spatio-temporal complexity. One aspect of this research program is the systematicmore » derivation of appropriate, approximate models from the original models governing combustion. The approximate models are then analyzed. The authors are particularly interested in understanding the basic mechanisms affecting combustion, which is a prerequisite to effective control of the process. They are interested in determining the effects of varying various control parameters, such as Nusselt number, Lewis number, heat release, activation energy, Damkohler number, Reynolds number, Prandtl number, Peclet number, etc. The authors have also considered a number of problems in self-propagating high-temperature synthesis (SHS), in which combustion waves are employed to synthesize advanced materials. Efforts are directed toward understanding fundamental mechanisms. 167 refs.« less

Tabulated Combustion Model Development For Non-Premixed Flames

NASA Astrophysics Data System (ADS)

Kundu, Prithwish

Turbulent non-premixed flames play a very important role in the field of engineering ranging from power generation to propulsion. The coupling of fluid mechanics and complicated combustion chemistry of fuels pose a challenge for the numerical modeling of these type of problems. Combustion modeling in Computational Fluid Dynamics (CFD) is one of the most important tools used for predictive modeling of complex systems and to understand the basic fundamentals of combustion. Traditional combustion models solve a transport equation of each species with a source term. In order to resolve the complex chemistry accurately it is important to include a large number of species. However, the computational cost is generally proportional to the cube of number of species. The presence of a large number of species in a flame makes the use of CFD computationally expensive and beyond reach for some applications or inaccurate when solved with simplified chemistry. For highly turbulent flows, it also becomes important to incorporate the effects of turbulence chemistry interaction (TCI). The aim of this work is to develop high fidelity combustion models based on the flamelet concept and to significantly advance the existing capabilities. A thorough investigation of existing models (Finite-rate chemistry and Representative Interactive Flamelet (RIF)) and comparative study of combustion models was done initially on a constant volume combustion chamber with diesel fuel injection. The CFD modeling was validated with experimental results and was also successfully applied to a single cylinder diesel engine. The effect of number of flamelets on the RIF model and flamelet initialization strategies were studied. The RIF model with multiple flamelets is computationally expensive and a model was proposed on the frame work of RIF. The new model was based on tabulated chemistry and incorporated TCI effects. A multidimensional tabulated chemistry database generation code was developed based on the 1D diffusion flame solver. The proposed model did not use progress variables like the traditional chemistry tabulation methods. The resulting model demonstrated an order of magnitude computational speed up over the RIF model. The results were validated across a wide range of operating conditions for diesel injections and the results were in close agreement to those of the experimental data. History of scalar dissipation rates plays a very important role in non premixed flames. However, tabulated methods have not been able to incorporate this physics in their models. A comparative approach is developed that can quantify these effects and find correlations with flow variables. A new model is proposed to include these effects in tabulated combustion models. The model is initially validated for 1D counterflow diffusion flame problems at engine conditions. The model is further implemented and validated in a 3D RANS code across a range of operating conditions for spray flames.

Technology for Transient Simulation of Vibration during Combustion Process in Rocket Thruster

NASA Astrophysics Data System (ADS)

Zubanov, V. M.; Stepanov, D. V.; Shabliy, L. S.

2018-01-01

The article describes the technology for simulation of transient combustion processes in the rocket thruster for determination of vibration frequency occurs during combustion. The engine operates on gaseous propellant: oxygen and hydrogen. Combustion simulation was performed using the ANSYS CFX software. Three reaction mechanisms for the stationary mode were considered and described in detail. The way for obtaining quick CFD-results with intermediate combustion components using an EDM model was found. The way to generate the Flamelet library with CFX-RIF was described. A technique for modeling transient combustion processes in the rocket thruster was proposed based on the Flamelet library. A cyclic irregularity of the temperature field like vortex core precession was detected in the chamber. Frequency of flame precession was obtained with the proposed simulation technique.

DNS and LES/FMDF of turbulent jet ignition and combustion

NASA Astrophysics Data System (ADS)

Validi, Abdoulahad; Jaberi, Farhad

2014-11-01

The ignition and combustion of lean fuel-air mixtures by a turbulent jet flow of hot combustion products injected into various geometries are studied by high fidelity numerical models. Turbulent jet ignition (TJI) is an efficient method for starting and controlling the combustion in complex propulsion systems and engines. The TJI and combustion of hydrogen and propane in various flow configurations are simulated with the direct numerical simulation (DNS) and the hybrid large eddy simulation/filtered mass density function (LES/FMDF) models. In the LES/FMDF model, the filtered form of the compressible Navier-Stokes equations are solved with a high-order finite difference scheme for the turbulent velocity and the FMDF transport equation is solved with a Lagrangian stochastic method to obtain the scalar field. The DNS and LES/FMDF data are used to study the physics of TJI and combustion for different turbulent jet igniter and gas mixture conditions. The results show the very complex and different behavior of the turbulence and the flame structure at different jet equivalence ratios.

Investigation on combustion characteristics and NO formation of methane with swirling and non-swirling high temperature air

NASA Astrophysics Data System (ADS)

Li, Xing; Jia, Li

2014-10-01

Combustion characteristics of methane jet flames in an industrial burner working in high temperature combustion regime were investigated experimentally and numerically to clarify the effects of swirling high temperature air on combustion. Speziale-Sarkar-Gatski (SSG) Reynolds stress model, Eddy-Dissipation Model (EDM), Discrete Ordinates Method (DTM) combined with Weighted-Sum-of-Grey Gases Model (WSGG) were employed for the numerical simulation. Both Thermal-NO and Prompt-NO mechanism were considered to evaluate the NO formation. Temperature distribution, NO emissions by experiment and computation in swirling and non-swirling patterns show combustion characteristics of methane jet flames are totally different. Non-swirling high temperature air made high NO formation while significant NO prohibition were achieved by swirling high temperature air. Furthermore, velocity fields, dimensionless major species mole fraction distributions and Thermal-NO molar reaction rate profiles by computation interpret an inner exhaust gas recirculation formed in the combustion zone in swirling case.

Modelling and simulation of wood chip combustion in a hot air generator system.

PubMed

Rajika, J K A T; Narayana, Mahinsasa

2016-01-01

This study focuses on modelling and simulation of horizontal moving bed/grate wood chip combustor. A standalone finite volume based 2-D steady state Euler-Euler Computational Fluid Dynamics (CFD) model was developed for packed bed combustion. Packed bed combustion of a medium scale biomass combustor, which was retrofitted from wood log to wood chip feeding for Tea drying in Sri Lanka, was evaluated by a CFD simulation study. The model was validated by the experimental results of an industrial biomass combustor for a hot air generation system in tea industry. Open-source CFD tool; OpenFOAM was used to generate CFD model source code for the packed bed combustion and simulated along with an available solver for free board region modelling in the CFD tool. Height of the packed bed is about 20 cm and biomass particles are assumed to be spherical shape with constant surface area to volume ratio. Temperature measurements of the combustor are well agreed with simulation results while gas phase compositions have discrepancies. Combustion efficiency of the validated hot air generator is around 52.2 %.

Indirect combustion noise of auxiliary power units

NASA Astrophysics Data System (ADS)

Tam, Christopher K. W.; Parrish, Sarah A.; Xu, Jun; Schuster, Bill

2013-08-01

Recent advances in noise suppression technology have significantly reduced jet and fan noise from commercial jet engines. This leads many investigators in the aeroacoustics community to suggest that core noise could well be the next aircraft noise barrier. Core noise consists of turbine noise and combustion noise. There is direct combustion noise generated by the combustion processes, and there is indirect combustion noise generated by the passage of combustion hot spots, or entropy waves, through constrictions in an engine. The present work focuses on indirect combustion noise. Indirect combustion noise has now been found in laboratory experiments. The primary objective of this work is to investigate whether indirect combustion noise is also generated in jet and other engines. In a jet engine, there are numerous noise sources. This makes the identification of indirect combustion noise a formidable task. Here, our effort concentrates exclusively on auxiliary power units (APUs). This choice is motivated by the fact that APUs are relatively simple engines with only a few noise sources. It is, therefore, expected that the chance of success is higher. Accordingly, a theoretical model study of the generation of indirect combustion noise in an Auxiliary Power Unit (APU) is carried out. The cross-sectional areas of an APU from the combustor to the turbine exit are scaled off to form an equivalent nozzle. A principal function of a turbine in an APU is to extract mechanical energy from the flow stream through the exertion of a resistive force. Therefore, the turbine is modeled by adding a negative body force to the momentum equation. This model is used to predict the ranges of frequencies over which there is a high probability for indirect combustion noise generation. Experimental spectra of internal pressure fluctuations and far-field noise of an RE220 APU are examined to identify anomalous peaks. These peaks are possible indirection combustion noise. In the case of the APU RE220, such peaks are identified. The frequency ranges of these peaks are found to overlap those predicted by the model theory. Based on this agreement, a tentative conclusion is drawn that there is good reason to believe that APUs do generate measurable indirect combustion noise. This paper is dedicated to the memory of Prof. Phil Doak for his numerous contributions to Aeroacoustics and the Journal of Sound and Vibration.

Spectral optimization and uncertainty quantification in combustion modeling

NASA Astrophysics Data System (ADS)

Sheen, David Allan

Reliable simulations of reacting flow systems require a well-characterized, detailed chemical model as a foundation. Accuracy of such a model can be assured, in principle, by a multi-parameter optimization against a set of experimental data. However, the inherent uncertainties in the rate evaluations and experimental data leave a model still characterized by some finite kinetic rate parameter space. Without a careful analysis of how this uncertainty space propagates into the model's predictions, those predictions can at best be trusted only qualitatively. In this work, the Method of Uncertainty Minimization using Polynomial Chaos Expansions is proposed to quantify these uncertainties. In this method, the uncertainty in the rate parameters of the as-compiled model is quantified. Then, the model is subjected to a rigorous multi-parameter optimization, as well as a consistency-screening process. Lastly, the uncertainty of the optimized model is calculated using an inverse spectral optimization technique, and then propagated into a range of simulation conditions. An as-compiled, detailed H2/CO/C1-C4 kinetic model is combined with a set of ethylene combustion data to serve as an example. The idea that the hydrocarbon oxidation model should be understood and developed in a hierarchical fashion has been a major driving force in kinetics research for decades. How this hierarchical strategy works at a quantitative level, however, has never been addressed. In this work, we use ethylene and propane combustion as examples and explore the question of hierarchical model development quantitatively. The Method of Uncertainty Minimization using Polynomial Chaos Expansions is utilized to quantify the amount of information that a particular combustion experiment, and thereby each data set, contributes to the model. This knowledge is applied to explore the relationships among the combustion chemistry of hydrogen/carbon monoxide, ethylene, and larger alkanes. Frequently, new data will become available, and it will be desirable to know the effect that inclusion of these data has on the optimized model. Two cases are considered here. In the first, a study of H2/CO mass burning rates has recently been published, wherein the experimentally-obtained results could not be reconciled with any extant H2/CO oxidation model. It is shown in that an optimized H2/CO model can be developed that will reproduce the results of the new experimental measurements. In addition, the high precision of the new experiments provide a strong constraint on the reaction rate parameters of the chemistry model, manifested in a significant improvement in the precision of simulations. In the second case, species time histories were measured during n-heptane oxidation behind reflected shock waves. The highly precise nature of these measurements is expected to impose critical constraints on chemical kinetic models of hydrocarbon combustion. The results show that while an as-compiled, prior reaction model of n-alkane combustion can be accurate in its prediction of the detailed species profiles, the kinetic parameter uncertainty in the model remains to be too large to obtain a precise prediction of the data. Constraining the prior model against the species time histories within the measurement uncertainties led to notable improvements in the precision of model predictions against the species data as well as the global combustion properties considered. Lastly, we show that while the capability of the multispecies measurement presents a step-change in our precise knowledge of the chemical processes in hydrocarbon combustion, accurate data of global combustion properties are still necessary to predict fuel combustion.

Turbulent transport measurements in a model of GT-combustor

NASA Astrophysics Data System (ADS)

Chikishev, L. M.; Gobyzov, O. A.; Sharaborin, D. K.; Lobasov, A. S.; Dulin, V. M.; Markovich, D. M.; Tsatiashvili, V. V.

2016-10-01

To reduce NOx formation modern industrial power gas-turbines utilizes lean premixed combustion of natural gas. The uniform distribution of local fuel/air ratio in the combustion chamber plays one of the key roles in the field of lean combustion to prevent thermo-acoustic pulsations. Present paper reports on simultaneous Particle Image Velocimetry and acetone Planar Laser Induced Fluorescence measurements in a cold model of GT-combustor to investigate mixing processes which are relevant to the organization of lean premixed combustion. Velocity and passive admixture pulsations correlations were measured to verify gradient closer model, which is often used in Reynolds-Averaged Navier-Stokes (RANS) simulation of turbulent mixing.

Analysis of rocket engine injection combustion processes

NASA Technical Reports Server (NTRS)

Salmon, J. W.

1976-01-01

A critique is given of the JANNAF sub-critical propellant injection/combustion process analysis computer models and application of the models to correlation of well documented hot fire engine data bases. These programs are the distributed energy release (DER) model for conventional liquid propellants injectors and the coaxial injection combustion model (CICM) for gaseous annulus/liquid core coaxial injectors. The critique identifies model inconsistencies while the computer analyses provide quantitative data on predictive accuracy. The program is comprised of three tasks: (1) computer program review and operations; (2) analysis and data correlations; and (3) documentation.

The hybrid RANS/LES of partially premixed supersonic combustion using G/Z flamelet model

NASA Astrophysics Data System (ADS)

Wu, Jinshui; Wang, Zhenguo; Bai, Xuesong; Sun, Mingbo; Wang, Hongbo

2016-10-01

In order to describe partially premixed supersonic combustion numerically, G/Z flamelet model is developed and compared with finite rate model in hybrid RANS/LES simulation to study the strut-injection supersonic combustion flow field designed by the German Aerospace Center. A new temperature calculation method based on time-splitting method of total energy is introduced in G/Z flamelet model. Simulation results show that temperature predictions in partially premixed zone by G/Z flamelet model are more consistent with experiment than finite rate model. It is worth mentioning that low temperature reaction zone behind the strut is well reproduced. Other quantities such as average velocity and average velocity fluctuation obtained by developed G/Z flamelet model are also in good agreement with experiment. Besides, simulation results by G/Z flamelet also reveal the mechanism of partially premixed supersonic combustion by the analyses of the interaction between turbulent burning velocity and flow field.

40 CFR Table 5 to Subpart Bbbb of... - Model Rule-Carbon Monoxide Emission Limits for Existing Small Municipal Waste Combustion Units

Code of Federal Regulations, 2012 CFR

2012-07-01

... Limits for Existing Small Municipal Waste Combustion Units 5 Table 5 to Subpart BBBB of Part 60... Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 5 Table 5 to Subpart... Combustion Units For the following municipal waste combustion units You must meet the following carbon...

40 CFR Table 3 to Subpart Bbbb of... - Model Rule-Class I Nitrogen Oxides Emission Limits for Existing Small Municipal Waste Combustion...

Code of Federal Regulations, 2011 CFR

2011-07-01

... Emission Limits for Existing Small Municipal Waste Combustion Unitsa,b,c 3 Table 3 to Subpart BBBB of Part... Municipal Waste Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 3 Table... Municipal Waste Combustion Unitsa,b,c Municipal waste combustion technology Limits for class I municipal...

40 CFR Table 5 to Subpart Bbbb of... - Model Rule-Carbon Monoxide Emission Limits for Existing Small Municipal Waste Combustion Units

Code of Federal Regulations, 2014 CFR

2014-07-01

... Limits for Existing Small Municipal Waste Combustion Units 5 Table 5 to Subpart BBBB of Part 60... Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 5 Table 5 to Subpart... Combustion Units For the following municipal waste combustion units You must meet the following carbon...

40 CFR Table 5 to Subpart Bbbb of... - Model Rule-Carbon Monoxide Emission Limits for Existing Small Municipal Waste Combustion Units

Code of Federal Regulations, 2010 CFR

2010-07-01

... Limits for Existing Small Municipal Waste Combustion Units 5 Table 5 to Subpart BBBB of Part 60... Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 5 Table 5 to Subpart... Combustion Units For the following municipal waste combustion units You must meet the following carbon...

40 CFR Table 5 to Subpart Bbbb of... - Model Rule-Carbon Monoxide Emission Limits for Existing Small Municipal Waste Combustion Units

Code of Federal Regulations, 2013 CFR

2013-07-01

... Limits for Existing Small Municipal Waste Combustion Units 5 Table 5 to Subpart BBBB of Part 60... Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 5 Table 5 to Subpart... Combustion Units For the following municipal waste combustion units You must meet the following carbon...

40 CFR Table 3 to Subpart Bbbb of... - Model Rule-Class I Nitrogen Oxides Emission Limits for Existing Small Municipal Waste Combustion...

Code of Federal Regulations, 2013 CFR

2013-07-01

... Emission Limits for Existing Small Municipal Waste Combustion Unitsa,b,c 3 Table 3 to Subpart BBBB of Part... Municipal Waste Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 3 Table... Municipal Waste Combustion Unitsa,b,c Municipal waste combustion technology Limits for class I municipal...

40 CFR Table 3 to Subpart Bbbb of... - Model Rule-Class I Nitrogen Oxides Emission Limits for Existing Small Municipal Waste Combustion...

Code of Federal Regulations, 2010 CFR

2010-07-01

... Emission Limits for Existing Small Municipal Waste Combustion Unitsa,b,c 3 Table 3 to Subpart BBBB of Part... Municipal Waste Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 3 Table... Municipal Waste Combustion Unitsa,b,c Municipal waste combustion technology Limits for class I municipal...

40 CFR Table 5 to Subpart Bbbb of... - Model Rule-Carbon Monoxide Emission Limits for Existing Small Municipal Waste Combustion Units

Code of Federal Regulations, 2011 CFR

2011-07-01

... Limits for Existing Small Municipal Waste Combustion Units 5 Table 5 to Subpart BBBB of Part 60... Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 5 Table 5 to Subpart... Combustion Units For the following municipal waste combustion units You must meet the following carbon...

40 CFR Table 3 to Subpart Bbbb of... - Model Rule-Class I Nitrogen Oxides Emission Limits for Existing Small Municipal Waste Combustion...

Code of Federal Regulations, 2012 CFR

2012-07-01

... Emission Limits for Existing Small Municipal Waste Combustion Unitsa,b,c 3 Table 3 to Subpart BBBB of Part... Municipal Waste Combustion Units Constructed on or Before August 30, 1999 Pt. 60, Subpt. BBBB, Table 3 Table... Municipal Waste Combustion Unitsa,b,c Municipal waste combustion technology Limits for class I municipal...

Taylor instability in rhyolite lava flows

NASA Technical Reports Server (NTRS)

Baum, B. A.; Krantz, W. B.; Fink, J. H.; Dickinson, R. E.

1989-01-01

A refined Taylor instability model is developed to describe the surface morphology of rhyolite lava flows. The effect of the downslope flow of the lava on the structures resulting from the Taylor instability mechanism is considered. Squire's (1933) transformation is developed for this flow in order to extend the results to three-dimensional modes. This permits assessing why ridges thought to arise from the Taylor instability mechanism are preferentially oriented transverse to the direction of lava flow. Measured diapir and ridge spacings for the Little and Big Glass Mountain rhyolite flows in northern California are used in conjunction with the model in order to explore the implications of the Taylor instability for flow emplacement. The model suggests additional lava flow features that can be measured in order to test whether the Taylor instability mechanism has influenced the flows surface morphology.

Thermo-chemical modelling of a village cookstove for design improvement

NASA Astrophysics Data System (ADS)

Honkalaskar, Vijay H.; Sohoni, Milind; Bhandarkar, Upendra V.

2014-05-01

Cookstove operation comprises three basic processes, namely combustion of firewood, natural air draft due to the buoyancy induced by the temperature difference between the hearth and its surroundings, and heat transfer to the pot, stove body and surrounding atmosphere. Owing to the heterogenous and unsteady burning of solid fuel, there exist nonlinear and dynamic interrelationships among these process parameters. A steady-state analytical model of the cookstove operation is developed for its design improvement by splitting the hearth into three zones to study char combustion, volatile combustion and heat transfer to the pot bottom separately. It comprises a total of seven relations corresponding to a thorough analysis of the three basic processes. A novel method is proposed to model the combustion of wood to mimic the realities closely. Combustion space above the fuel bed is split into 1000 discrete parts to study the combustion of volatiles by considering a set of representative volatile gases. Model results are validated by comparing them with a set of water boiling tests carried on a traditional cookstove in the laboratory. It is found that the major thrust areas to improve the thermal performance are combustion of volatiles and the heat transfer to the pot. It is revealed that the existing design dimensions of the traditional cookstove are close to their optimal values. Addition of twisted-tape inserts in the hearth of the cookstove shows an improvement in the thermal performance due to increase in the heat transfer coefficient to the pot bottom and improved combustion of volatiles.