Combustion instability mechanism of a lean premixed gas turbine combustor

Lean premixed combustion has been considered as one of the promising solutions for the reduction of NOx emissions from gas turbines. However, unstable combustion of lean premixed flow becomes a real challenge on the way to design a reliable, highly efficient dry low NOx gas turbine combustor. Contrary to a conventional diffusion type combustion system, characteristics of premixed combustion significantly depend on a premixing degree of combusting flow. Combustion behavior in terms of stability has been studied in a model gas turbine combustor burning natural gas and air. Incompleteness of premixing is identified as significant perturbation source for inducing unstable combustion. Application of a simple convection time lag theory can only predict instability modes but cannot determine whether instability occurs or not. Low frequency perturbations are observed at the onset of instability and believed to initiate the coupling between heat release rate and pressure fluctuations.     

Linear stability analysis of acoustically driven pressure oscillations in a lean premixed gas turbine combustor

The dynamic response of a turbulent premixed flame to acoustic velocity perturbations was experimentally determined in a swirl-stabilized lean-premixed gas turbine combustor. CH* chemiluminescence intensity and the twomicrophone method were used to measure heat release rates and inlet velocity fluctuations, respectively. Using the n-τ formulation, gain and phase of flame transfer functions were incorporated into an analytic thermoacoustic model to predict instability frequencies and modal structures. Self-excited instability measurements were performed to verify eigenfrequencies predicted by the thermoacoustic model. Instability frequency predicted by the model is supported by experimental results. Results show that the self-excited instability frequency of ～ 220 Hz results from the fact that the flames amplify flow perturbations with f = 150 ～ 250 Hz. The other instability frequency of ～ 350 Hz occurs because the whole combustion system has an eigenfrequency corresponding to the ¼-wave eigenmode of the mixing section.     

Analysis of the combustion oscillation in a silo-type gas turbine combustor and its suppression

The characteristics of combustion oscillation of a silo-type 79.5 MW gas turbine combustor in commercial operation and its suppression have been investigated. The oscillation of the lean premixed gas turbine combustor resulting from the combustion instability occurred at near full load operation. An FFT analysis of the combustion dynamics showed that the dominant frequency of the oscillation would be that of the 1st longitudinal acoustic resonance mode of the combustor. To suppress the combustion oscillation, a passive control technique for reducing the combustion instability was employed; that is, the fuel to the combustor was redistributed by adjusting the operational schedule of one of six fuel control valves, which would lead the increase of the local operational equivalence ratio near the central recirculation zone of the combustor. By doing so, the oscillation was successfully reduced to the permissible level while the amount of NO_x emission met proper regulatory level set by the local government.     

Hot-fire injector test for determination of combustion stability boundaries using model chamber

This study realizes the conceptual method to predict combustion instability in actual full-scale combustion chamber of rocket engines by experimental tests with model (sub-scale) chamber. The model chamber was designed based on the methodologies proposed in the previous work regarding geometrical dimensions and operating conditions, and hot-fire test procedures were followed to obtain stability boundaries. From the experimental tests, two instability regions are presented by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio, which are customary for combustor designers. It is found that instability characteristics in the chamber with the adopted jet injectors can be explained by the correlation between the characteristic burning or mixing time and the characteristic acoustic time. In each instability region, dynamic behaviors of flames are investigated to verify the hydrodynamically-derived characteristic lengths of the jet injectors. Large-amplitude pressure oscillation observed in upper instability region is found to be generated by lifted-off flames.     

On the method for hot-fire modeling of high-frequency combustion instability in liquid rocket engines

This study presents the methodological aspects of combustion instability modeling and pro-vides the numerical results of the model (sub-scale) combustion chamber, regarding geometrical dimensions and operating conditions, which are for determining the combustion stability boundaries using the model chamber. An approach to determine the stability limits and acoustic characteristics of injectors is described intensively. Procedures for extrapolation of the model operating parameters to the actual conditions are presented, which allow the hot-fire test data to be presented by parameters of the combustion chamber pressure and mixture (oxidizer/fuel) ratio, which are customary for designers. Tests with the model chamber, based on the suggested scaling method, are far more cost-effective than with the actual (full-scale) chamber and useful for injector screening at the initial stage of the combustor development in a viewpoint of combustion instabilities.     

Numerical simulation of the unsteady combustion of solid rocket propellants at a harmonic pressure change

This study focuses on a numerical investigation of the unsteady burning rate of solid propellants at a harmonic pressure change in the combustion chamber of a solid propellant rocket engine. The physico-mathematical model includes the equations of heat transfer and decomposition of the oxidizer in the solid phase and two phases, the dual velocity, and the two-temperature reaction flow of gasification products. The boundary conditions on the solid fuel surface implement the conservation of energy fluxes and the mass of components. We numerically calculate the unsteady burning rate of metallized solid propellant and nitroglycerin powder under a harmonic pressure change in the combustion chamber of a solid propellant rocket engine and determine the dependence of the burning rate amplitude on the frequency of pressure oscillations. The amplitude of the burning rate depends nonmonotonously on the oscillation frequency. With increasing frequency, the amplitude first rises and then declines.

     

Experimental study of the role of gas-liquid scheme injector as an acoustic resonator in a combustion chamber

In a liquid rocket engine, the role of gas-liquid scheme injector as an acoustic resonator or absorber is studied experimentally for combustion stability by adopting linear acoustic test. The acoustic-pressure signals or responses from the chamber are monitored by acoustic amplitude. Acoustic behavior in a rocket combustor with a single injector is investigated and the acoustic-damping effect of the injector is evaluated for cold condition by the quantitative parameter of damping factor as a function of injector length. From the experimental data, it is found that the injector can play a significant role in acoustic damping when it is tuned finely. The optimum tuning-length of the injector to maximize the damping capacity is near half of a full wavelength of the first longitudinal overtone mode traveling in the injector with the acoustic frequency intended for damping in the chamber. When the injector has large diameter, the phenomenon of the mode split is observed near the optimum injector length and thereby, the acoustic-damping effect of the tuned injectors can be degraded.     

Characteristics of self-excited combustion oscillation and combustion control by forced pulsating mixture supply

Characteristics of self-excited combustion oscillation are experimentally studied using confined premixed flames stabilized by a rearward-facing step. A new idea to suppress combustion oscillation was applied to the flames. The characteristics of unsteady combustion were examined, which is driven by forced pulsating mixture supply that can modulate its amplitude and frequency. The self-excited combustion oscillation having weaker flow velocity fluctuation intensity than that of the forced pulsating supply can be suppressed by the method. The effects of the forced pulsation amplitude and frequency on controlling self-excited combustion oscillations were also investigated comparing with the steady mixture supply. The unsteady combustion used in this experiment plays an important role in controlling self-excited combustion oscillations, and it also exhibits desirable performances, from a practical point of view, such as high combustion load and reduced pollutant emissions of nitric oxide.     

Experimental study on combustion instability and attenuation characteristics in the lab-scale gas turbine combustor with a sponge-like porous medium

The present study has experimentally investigated the combustion instability and its attenuation characteristics in the lab-scale swirlstabilized premixed combustor with a sponge-like porous medium. Unlike the conventional premixed burners, this model combustor has the unique features including a porous dump plane and an acoustic cavity, which was devised for attenuating the combustion instability. When replacing the dump plane with a non-porous medium, the burner becomes the conventional design. In order to evaluate the damping effects of the porous medium on the unstable flame dynamics, various acoustic and photonic measurements and flame visualization were made. Special emphasis is given to the effects of the acoustic cavity length on the stabilization characteristics. Results showed that the model combustor with the porous dump plane and the acoustic cavity exhibited dramatic attenuation of the pressure oscillation intensity by up to about 40 %. The attenuation was increased with increasing the acoustic cavity length. It was also found that the attenuation is effective not only for the fundamental resonance but also for its higher harmonics. Based on the experimental results, detailed discussions are made for the combustion instability and its attenuation characteristics in the model gas turbine combustor with porous and nonporous media.

     

Transient analysis of hybrid rocket combustion by the zeldovich-novozhilov method

Hybrid rocket combustion has a manifestation of stable response to the perturbations compared to solid propellant combustion. Recently, it has revealed that the low frequency combustion instability about 10 Hz was occurred mainly due to thermal inertia of solid fuel. In this paper. the combustion response function was theoretically derived by use of ZN (Zeldovich-Novozhilov) method. The result with HTPB/LOX combination showed a quite good agreement in response function with previous works and could predict the low frequency oscillations with a peak around 10 Hz which was observed experimentally. Also, it was found that the amplification region in the frequency domain is independent of the regression rate exponentn but showed the dependence of activation energy. Moreover, the response function has shown that the hybrid combustion system was stable due to negative heat release of solid fuel for vaporization, even though the addition of energetic ingredients such as AP and Al could lead to increase heat release at the fuel surface.     

Improvement of flame stability and NOx reduction in hydrogen-added ultra lean premixed combustion

Lean premixed combustion is a well known method in gas turbine combustors that can reduce fuel consumption and decrease flame temperature. In lean premixed flames, flame instabilities can occur because the combustion takes place near the lean flammable limit. For the purpose of increasing flame stability, a small amount of hydrogen was added into a fuel, which has ultra low lean flammable limit. The extinction stretch rate increased and total equivalence ratio at extinction decreased with hydrogen addition; consequently, ultra lean premixed combustion was possible and flame stability could be achieved at low temperature conditions. The NO_x emission increased with hydrogen addition for the same stretch rate and equivalence ratio, but the extinction stretch rate and lean flammability limit was enlarged. Consequently, NO_x emission decreased with hydrogen addition in the near extinction conditions. Hydrogen addition could improve flame stability and reduce NO_x emission in ultra lean premixed combustion. This paper was recommended for publication in revised form by Associate Editor Ohchae Kwon Dr. Eun-Seong Cho received his B.S. and M.S. degrees in Mechanical Engineering from Hanyang University, Korea, in 1996 and 1998, respectively. He then received his Ph.D. degree from Seoul National University, Korea, in 2005. He was a principal engineer of KD Navien research center and currently a research associate at Delft University of Technology, The Netherlands. His research interests include eco-friendly clean combustion technology, new and renewable energy systems. Prof. Suk Ho Chung received his B.S. degree from Seoul National University, Korea, in 1976 and Ph.D. degree in Mechanical Engineering from Northwestern University, USA, in 1983. He is a Professor since 1984 in the School of Mechanical and Aerospace Engineering at Seoul National University in Seoul, Korea. His research interests cover combustion fundamentals, pollutant formation, laser diagnostics, and plasma-assisted combustion.     

An experiment analysis of MILD combustion with liquid fuel spray in a combustion vessel

Mild combustion is characterized by its distinguished features, such as suppressed pollutant emission, homogeneous temperature distribution, reduced noise, and thermal stress. Recently, many studies have revealed the potential of MILD combustion in various power systems but most studies have been focused on gas phase fuel MILD combustion. Therefore, further study on MILD combustion using liquid fuel is needed for the application to a liquid-fueled gas turbine especially. In this work, we studied experimentally on the formation of liquid fuel MILD combustion under the condition of high dilution by burnt gas generated from a first premixed flame in two stages combustor which consists of the first premixed burner and secondary combustor. In particular, the effects of burnt gas velocity and oxygen level of burnt gas on the formation of liquid fuel MILD combustion were investigated. The results show that as the burnt gas velocity through the nozzle becomes higher, the color of flames was changed from yellow to pale blue and flames became very short. The OH radical measured by ICCD camera was uniformly distributed on the pale blue flame surface and its intensity was very low compared to conventional liquid diffusion flame. As burnt gas velocity is increased, local high-temperature region appeared to be diminished and the flame temperature became spatially uniform. And CO emission was sampled around 1 ppm and NOx emission was measured around 10 ppm under the overall equivalence ratio of 0.8 to 0.98 for 40 mm or less diameter of velocity control nozzle. This low NOx emission seems to be attributed to maintaining the average temperature in secondary combustor below the threshold temperature of thermal NOx formation. In view of the uniform temperature distribution, low OH radical intensity and low NOx emission data in the secondary combustor, formation of stable MILD combustion using kerosene liquid fuel could be verified at high burnt gas velocity.

     

Biparametric assessment of the combustion stability in an industrial gas turbine combustor

Here we propose a generalized procedure for a two-parameter assessment of the Combustion stability (CS) of industrial gas turbines. In evaluating the CS, this procedure employs two parameters of measured dynamic pressure data: the Root-mean-square (RMS) pressure as the primary parameter and the damping ratio as a secondary parameter. The former tells the time-averaged level of the dynamic pressure, and, the latter, the degree of acoustic energy loss. A data point pairing the two parameters, which are evaluated at a specific instance of a combustion process, identifies Instantaneous combustion stability (ICS) by its location on a 2-D domain of both parameters. Collective representation of the ICS points on the domain produces a CS map of the combustion process. The locus of the ICS point on the map represents the temporal variation of CS during the combustion process. The biparametric assessment procedure divides the CS map into three regimes (i.e., stable, transitional and unstable) by utilizing two threshold values for the RMS pressure and one for the damping ratio. The feasibility of the proposed procedure was tested with the dynamic pressure data from a model gas turbine combustor burning synthetic natural gas. Then the technique was applied experimental data obtained from a laboratory-scale lean premixed combustor to identify the three regimes of the combustion process of a reported case. We found that the procedure is able to provide gas turbine operators with valuable information on CS during a combustion process, especially on the transitional regime.

     

基于CFD的氢气／空气当量比对微尺度燃烧室燃烧特性影响的分析

以美国麻省理工学院（MIT）研制的硅基六晶片微燃烧室为研究对象,提出利用二维CFD（计算流体动力学）数值模拟的方法,研究在保持微尺度燃烧室进口氢气／空气流量不变的情况下,改变氢气／空气当量比对燃烧室燃烧特性的影响。整个模拟计算主要包括氢气／空气的流动路径、微燃烧室的内部区域以及整个燃烧室的墙壁面;同时在计算过程中我们考虑了氢气／空气的流体动力学特性、传热学特性和详细的基元反应机理。结果表明,利用二维CFD数值模拟的方法研究微尺度燃烧室燃烧特性可行,与国外实际测量结果较为相似,为今后微型燃气轮机燃烧室的研制及改进提供了一定的参考依据。     

Effects of inlet blockage of gas-liquid scheme injector on acoustic tuning for acoustic damping in a combustion chamber

In a rocket combustor, purely acoustic tuning of gas-liquid scheme injector is studied numerically for acoustic absorption by adopting linear acoustic analysis. Acoustic behavior in the combustor with a single injector is investigated to assure the optimum injector length. Acoustic-absorption effect of the injector is evaluated for cold condition by the quantitative parameter of damping factor as a function of injector length in the chamber for several boundary absorption coefficients. Irrespective of boundary absorption at the chamber wall, it is assured that the optimum tuning-length of the injector corresponds to half of a full wavelength of the first longitudinal overtone mode traveling in the injector with the acoustic frequency intended for damping in the chamber. Although boundary absorption affects little the tuning length of the injector, it appreciably affects damping capacity. Acoustic absorption at the wall increases with boundary absorption coefficient, but purely acoustic-damping effect induced by the tuned injector decreases with the coefficient. As another design parameter, effects of blockage at the injector inlet on acoustic tuning are investigated. It is found that the optimum injector length is shifted depending on the blockage ratio. Suitable combination of injector length and blockage should be made for maximum damping.     

Experimental study on flame structure and temperature characteristics in a lean premixed model gas turbine combustor

Experimental study was carried out in an atmospheric pressure, laboratory-scale dump combustor showing features of combustion instabilities. Flame structure and heat release rates were obtained from OH emission spectroscopy. Qualitative comparisons were made between line —integrated OH chemiluminescence image and Abel—transformed one. Local Rayleigh index distributions were also examined. Mean temperature, normalized standard deviation and temperature fluctuations were measured by coherent anti-Stokes Raman spectroscopy (CARS). To see the periodic behavior of oscillating flames, phase—resolved measurements were performed with respect to the pressure wave in the combustor. Results on system damping and driving characteristics were provided as a function of equivalence ratio. It also could be observed that phase resolved temperatures have been changed in a well—defined manner, while its difference between maximum and minimum reached up to 280K. These results would be expected to play an important role in better understanding of driving mechanisms and thermo-acoustic interactions.     

贫油直喷燃烧室回火的数值研究

针对燃烧稳定性中的回火问题,对贫油直喷燃烧室的回火特性进行了研究.燃烧中的回火指的是火焰从燃烧室传入到了预混区中的这样一种现象,它包含了声波、湍流、燃烧之间的复杂的相互作用,是燃烧研究中的关键问题之一.利用Fluent中的混合分数/PDF平衡化学反应模型对贫油直喷燃烧室进行了大涡模拟,分别计算了冷态和热态下的流场,捕捉了贫油直喷燃烧室内回火的动态过程,其中亚格子模型采用WALE模型,燃油射流采用离散相模型.研究结果表明,大涡模拟能够较好地反映流场分布情况,贫油直喷燃烧室内存在中心回流区,模拟结果与实验结果吻合较好;贫油直喷燃烧室中的回火发生在中心流区域,回火时,中心回流区向上游移动,这促使了火焰的向上游传播,并最终导致了回火的发生.     

A study on the characteristics of combustion variations of compression ignition engine

This paper describes the results of a study of the variation of combustion characteristics in a precombusion chamber type water-cooled diesel engine. Statiscal analysis on cycle-by-cycle variation of combustion characteristics such as rate of pressure rise, heat release rate, and mass burning rate from combustion pressure-crank angle data of one thousand cycles were made under several operating conditions. The influence of engine speed and coolant temperature upon maximum frequency of combustion characteristics are discussed also.     

Flow instability due to cryogenic cavitation in the downstream of orifice

Flow instability in LRE (liquid rocket engine) occurs due to various reasons such as flow interactions with valve, orifice and venturi, etc. The inception of cavitation, especially in the propellant feeding system, is the primary cause of mass and pressure oscillations because of the cyclic formation and depletion of cavitation. Meanwhile, the main propellant in a liquid rocket engine is the cryogenic fluid, which properties are very sensitive to temperature variation. And the change of propellant properties to temperature variation by thermodynamic effect needs to be properly taken into account in the flow analysis in order to understand basic mechanisms for cryogenic cavitation. The present study focuses on the formation of cryogenic cavitation by using the IDM model suggested by Shyy and coworkers. The flow instability was also numerically investigated in the downstream of orifice with a developed numerical code. Calculation results show that cryogenic cavitation can be a primary source of flow instability, leading to mass fluctuations accompanied by pressure oscillations. The prediction of cavitation in cryogenic fluid is of vital importance in designing a feeding system of an LRE. This paper was recommended for publication in revised form by Associate Editor Jun Sang Park Changjin Lee received his B.S. and M.S. degrees in Aeronautical Engineering from Seoul National University in 1983 and 1985. He then went on to receive his Ph.D. degree from University of Illinois at Urbana- Champaign in 1992. Dr. Lee is currently a Professor at the department of Aerospace Engineering at Konkuk University in SEOUL, Korea. His research interests are in the area of combustion instabilities of hybrid, liquid rocket and jet propulsions. Tae-Seong Roh received his B.S. and M.S. degrees in Aeronautical Engineering from Seoul National University in 1984 and 1986. He then went on to receive his Ph.D. degree from Pennsylvania State University in 1995. Dr. Roh is currently a Professor at the department of Aerospace Engineering at Inha University in Incheon, Korea. His research interests are in the area of combustion instabilities, rocket and jet propulsions, interior ballistics, and gas turbine engine defect diagnostics.     

Combustion stability characteristics of the model chamber with various configurations of triplet impinging-jet injectors

Combustion stability characteristics in actual full-scale combustion chamber of a rocket engine are investigated by experimental tests with the model (sub-scale) chamber. The present hot-fire tests adopt the combustion chamber with three configurations of triplet impinging-jet injectors such as F-O-O-F, F-O-F, and O-F-O configurations. Combustion stability boundaries are obtained and presented by the parameters of combustion-chamber pressure and mixture (oxidizer/fuel) ratio. From the experimental tests, two instability regions are observed and the pressure oscillations have the similar patterns irrespective of injector configuration. But, the O-F-O injector configuration shows broader upper-instability region than the other configurations. To verify the instability mechanism for the lower and upper instability regions, air-purge acoustic test is conducted and the photograph of the flames is taken. As a result, it is found that the pressure oscillations in the two regions can be characterized by the first impinging point of hydraulic jets and pre-blowout combustion, respectively.