航空发动机加力燃烧室不稳定燃烧机理与控制方法研究
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摘要
航空发动机加力燃烧室中的振荡燃烧一直以来都是研究者们关注的问题,引起燃烧不稳定的详细机理和激励过程到目前为止还不十分清楚,如何发展有效的手段来抑制燃烧室中不稳定的发生也仍然是需要研究的问题。
本文围绕某型航空发动机加力蜂鸣问题,利用模型燃烧室预混燃烧及Rijke管热声振荡的一系列实验研究,探讨引起燃烧不稳定的机理、影响因素和抑制方法,并建立了能够适用于航空发动机加力燃烧室的热声振荡理论模型,在CFD数值模拟研究的基础上,对该型发动机加力燃烧室中热声振荡现象及其控制规律进行机理性的研究和分析,并试图找出一种主被动相结合的控制热声振荡新方法。
首先是在已有的三维圆管热声振荡理论模型的基础上,发展了一个适合于航空发动机加力燃烧室的不稳定燃烧预测模型。该模型利用了模态匹配的方法来处理变截面、声衬(隔热屏)加入等复杂问题,并考虑了热源、边界条件等因素的影响。该模型的建立为预测航空发动机加力燃烧室振荡燃烧特性及控制规律提供了一种理论探索途径。
其次为研究航空发动机加力燃烧室燃烧不稳定激励及控制方法,开展了Rijke管热声振荡研究及模型燃烧室预混燃烧不稳定性研究系列实验等基础工作。在Rijke管实验中,对热源位置、穿孔板穿孔率、声衬长度、偏流等因素对热声振荡特性的影响进行了研究,结果表明:热源位置、声衬长度、穿孔率以及偏流等因素的改变能够引起Rijke管中的有效声压、振荡频率及不稳定区域的变化。并能够从中找到抑制热声振荡的控制方法和规律。对模型燃烧室预混燃烧不稳定性研究实验中,着重研究了当量比、旋流和扩张段结构等对于振荡燃烧的影响,实验表明燃烧室发生的是纵向自激式振荡。振荡模式是燃烧室与预混段之间的耦合的共振现象,振荡频率由整个系统决定,预混段的振荡是由燃烧室引起的受迫振荡。另外当量比对燃烧的稳定性有重要影响,扩张段结构会与旋流会共同影响燃烧不稳定频率和脉动压力。这些工作为进一步研究影响热声振荡特性的诸因素及其控制规律,同时也为航空发动机加力燃烧室的振荡燃烧控制积累了有价值的经验。
第三是利用CFD方法对模型燃烧室及某型航空发动机加力燃烧室中的不稳定燃烧进行了数值模拟。对模型燃烧室预混燃烧的数值模拟解释了实验中的现象,并发现在燃烧室进口突扩面处形成以中心回流区和角回流区与混气射流形成的内外两个剪切层决定火焰的结构和运动。应用不同的数值模型对航空发动机加力燃烧室在最大加力状态时出现的振荡燃烧的数值模拟表明:火焰稳定器后的涡脱落现象是引起加力蜂鸣的重要原因,而燃油分布、雾化等因素会影响到火焰稳定器唇口附近的剪切层形态,从而对火焰面的脉动产生影响。数值模拟的结果还表明采用的数值方法对对勾画燃烧室中的涡系结构、燃烧形态以及对振荡燃烧的判断有很大的影响,因此对模型的选择仍然是利用CFD方法研究燃烧室流场的难点之一。通过CFD的模拟得到了加力燃烧室的平均流动参数,为圆管热声振荡理论模型在加力燃烧室中的具体应用提供了条件。非定常数值模拟的结果也从一个侧面考察了加力燃烧室中发生振荡燃烧的机理和影响因素。
第四是利用圆管热声振荡理论模型计算了航空发动机加力燃烧室中的热声振荡现象,并探讨抑制热声振荡的机理及控制规律。这项工作的结果表明该模型在研究这类问题时的适用性,预测结果具用一定的准确性。同样对Rijke管的计算结果也证明了偏流的引入对于主被动结合控制热声振荡方法的重要性。在对航空发动机加力燃烧室中的热源位置、隔热屏等结构的模拟研究中,揭示了这些结构对热声振荡的影响规律,从而为发现新的控制方法提供理论依据和思路。
Oscillation combustion in the afterburner has been one of main concerns in the research and development of modern aero-engine. The relevant mechanism and physical process is neither totally clear for now, nor are the effective ways to suppress the occurrence of instability in the chamber.
For the buzz of the aero-engine afterburner, this paper conducted a series of experimental study on thermoacoustic oscillation and combustion instability, by making use of the lean premixed gas combustor experiment for model chamber and thermoacoustic oscillation experiment of Rijke tube. The mechanism of combustion instability and control approaches were discussed, and the thermoacoustic theory method for afterburner of aero-engine was established. Based on the CFD numerical simulation analysis, the aero-engine was studied and analyzed on its thermoacoustic oscillation phenomenon and control mechanism, whose aim is to explore a new way of active and passive combination to control thermoacoustic oscillation.
First, the three-dimensional theory method of thermoacoustic oscillation in a tube was developed. The model can be used to describe the sound propagation in a variable cross-section tube with complex acoustic lining structure, such as perforation plate. The method deals with the complex problems of variable cross-section or adding acoustic lining in use of mode matching approach, which includes the influence of heat release and boundary conditions. The model provides theoretical way for the research of the thermoacoustic oscillation in the aero-engine afterburner.
Secondly, a series of experiments were conducted on thermoacoustic oscillation of Rijke tube and lean premixed gas combustor. The influence of heat release position on thermoacoustic oscillation character of such factors, open area ratio of perforated plate, acoustic lining length, and bias flow were investigated by means of the Rijke tube. The research results show that the factors of heat release position, open area ratio of perforated plate, acoustic lining length, and bias flow can introduce the variation of effective acoustic pressure, oscillating frequency and unstable region. From the results the control approaches and laws of suppressing thermoacoustic oscillation can be explored. In the experiments of lean premixed gas combustor instability research, the influence of equivalent rate, rotational flow, and expansion structure on oscillation combustion was emphasized. The experiments indicate that self-induced oscillation occurs in the chamber. The oscillation modal is a resonance phenomenon coupled between the chamber and premixed part. The oscillating frequency was determined by the whole system, and oscillation in premixed part was forced oscillation induced by the chamber. Besides, the equivalent rate has an important influence on combustion instability, and expansion with rotational flow together influences combustion instability frequency and oscillation pressure.
Thirdly, the numerical simulation was performed on the unstable combustion in lean premixed gas combustor and the afterburner of aero-engine by CFD. The numerical simulation of lean premixed gas combustor explains the phenomenon in the experiments. It is found that the two shear layers of jets with central recirculation zone and corner recirculation zone at the sudden expansion inlet of chamber determine the structure and movement of flame. By applying different numerical models, the numerical simulation on the unstable combustion existing in the afterburner of the aero-engine at the maximum thrust state indicates that: vortex shedding after the flame holders is the important reason of inducing boost buzzing, and the fuel contribution and atomization can influence the configuration of shear layer near the lips of the flame holders as to influence the pulse of flame surface. The numerical simulation results also indicate that the numerical methods have remarkable influence on the structure and configuration of vortex; therefore it is one of difficulty to choose models for CFD research. Besides, the average flow parameters of the afterburner were obtained by CFD simulation, which provide the conditions for the application of thermoacoustic oscillation model of the afterburner.
Fourthly, the thermoacoustic oscillation in the afterburner of aero-engine was calculated using the present thermoacoustic model, and the mechanism and control laws of suppressing thermoacoustic oscillation were discussed. The results of the research indicate that the method has applicability in studying such problems and the predicted thermoacoustic frequency has enough accuracy. The calculation results for Rijke tube indicate show the importance of bias flow injection on thermoacoustic oscillation control method of active and passive combination. The review of heat release positions and heat shield structures indicates the influence laws for such structures, which provides the theory evidence for finding new control ways to supress the oscillation in afterburner of aero-engine.
本文围绕某型航空发动机加力蜂鸣问题,利用模型燃烧室预混燃烧及Rijke管热声振荡的一系列实验研究,探讨引起燃烧不稳定的机理、影响因素和抑制方法,并建立了能够适用于航空发动机加力燃烧室的热声振荡理论模型,在CFD数值模拟研究的基础上,对该型发动机加力燃烧室中热声振荡现象及其控制规律进行机理性的研究和分析,并试图找出一种主被动相结合的控制热声振荡新方法。
首先是在已有的三维圆管热声振荡理论模型的基础上,发展了一个适合于航空发动机加力燃烧室的不稳定燃烧预测模型。该模型利用了模态匹配的方法来处理变截面、声衬(隔热屏)加入等复杂问题,并考虑了热源、边界条件等因素的影响。该模型的建立为预测航空发动机加力燃烧室振荡燃烧特性及控制规律提供了一种理论探索途径。
其次为研究航空发动机加力燃烧室燃烧不稳定激励及控制方法,开展了Rijke管热声振荡研究及模型燃烧室预混燃烧不稳定性研究系列实验等基础工作。在Rijke管实验中,对热源位置、穿孔板穿孔率、声衬长度、偏流等因素对热声振荡特性的影响进行了研究,结果表明:热源位置、声衬长度、穿孔率以及偏流等因素的改变能够引起Rijke管中的有效声压、振荡频率及不稳定区域的变化。并能够从中找到抑制热声振荡的控制方法和规律。对模型燃烧室预混燃烧不稳定性研究实验中,着重研究了当量比、旋流和扩张段结构等对于振荡燃烧的影响,实验表明燃烧室发生的是纵向自激式振荡。振荡模式是燃烧室与预混段之间的耦合的共振现象,振荡频率由整个系统决定,预混段的振荡是由燃烧室引起的受迫振荡。另外当量比对燃烧的稳定性有重要影响,扩张段结构会与旋流会共同影响燃烧不稳定频率和脉动压力。这些工作为进一步研究影响热声振荡特性的诸因素及其控制规律,同时也为航空发动机加力燃烧室的振荡燃烧控制积累了有价值的经验。
第三是利用CFD方法对模型燃烧室及某型航空发动机加力燃烧室中的不稳定燃烧进行了数值模拟。对模型燃烧室预混燃烧的数值模拟解释了实验中的现象,并发现在燃烧室进口突扩面处形成以中心回流区和角回流区与混气射流形成的内外两个剪切层决定火焰的结构和运动。应用不同的数值模型对航空发动机加力燃烧室在最大加力状态时出现的振荡燃烧的数值模拟表明:火焰稳定器后的涡脱落现象是引起加力蜂鸣的重要原因,而燃油分布、雾化等因素会影响到火焰稳定器唇口附近的剪切层形态,从而对火焰面的脉动产生影响。数值模拟的结果还表明采用的数值方法对对勾画燃烧室中的涡系结构、燃烧形态以及对振荡燃烧的判断有很大的影响,因此对模型的选择仍然是利用CFD方法研究燃烧室流场的难点之一。通过CFD的模拟得到了加力燃烧室的平均流动参数,为圆管热声振荡理论模型在加力燃烧室中的具体应用提供了条件。非定常数值模拟的结果也从一个侧面考察了加力燃烧室中发生振荡燃烧的机理和影响因素。
第四是利用圆管热声振荡理论模型计算了航空发动机加力燃烧室中的热声振荡现象,并探讨抑制热声振荡的机理及控制规律。这项工作的结果表明该模型在研究这类问题时的适用性,预测结果具用一定的准确性。同样对Rijke管的计算结果也证明了偏流的引入对于主被动结合控制热声振荡方法的重要性。在对航空发动机加力燃烧室中的热源位置、隔热屏等结构的模拟研究中,揭示了这些结构对热声振荡的影响规律,从而为发现新的控制方法提供理论依据和思路。
Oscillation combustion in the afterburner has been one of main concerns in the research and development of modern aero-engine. The relevant mechanism and physical process is neither totally clear for now, nor are the effective ways to suppress the occurrence of instability in the chamber.
For the buzz of the aero-engine afterburner, this paper conducted a series of experimental study on thermoacoustic oscillation and combustion instability, by making use of the lean premixed gas combustor experiment for model chamber and thermoacoustic oscillation experiment of Rijke tube. The mechanism of combustion instability and control approaches were discussed, and the thermoacoustic theory method for afterburner of aero-engine was established. Based on the CFD numerical simulation analysis, the aero-engine was studied and analyzed on its thermoacoustic oscillation phenomenon and control mechanism, whose aim is to explore a new way of active and passive combination to control thermoacoustic oscillation.
First, the three-dimensional theory method of thermoacoustic oscillation in a tube was developed. The model can be used to describe the sound propagation in a variable cross-section tube with complex acoustic lining structure, such as perforation plate. The method deals with the complex problems of variable cross-section or adding acoustic lining in use of mode matching approach, which includes the influence of heat release and boundary conditions. The model provides theoretical way for the research of the thermoacoustic oscillation in the aero-engine afterburner.
Secondly, a series of experiments were conducted on thermoacoustic oscillation of Rijke tube and lean premixed gas combustor. The influence of heat release position on thermoacoustic oscillation character of such factors, open area ratio of perforated plate, acoustic lining length, and bias flow were investigated by means of the Rijke tube. The research results show that the factors of heat release position, open area ratio of perforated plate, acoustic lining length, and bias flow can introduce the variation of effective acoustic pressure, oscillating frequency and unstable region. From the results the control approaches and laws of suppressing thermoacoustic oscillation can be explored. In the experiments of lean premixed gas combustor instability research, the influence of equivalent rate, rotational flow, and expansion structure on oscillation combustion was emphasized. The experiments indicate that self-induced oscillation occurs in the chamber. The oscillation modal is a resonance phenomenon coupled between the chamber and premixed part. The oscillating frequency was determined by the whole system, and oscillation in premixed part was forced oscillation induced by the chamber. Besides, the equivalent rate has an important influence on combustion instability, and expansion with rotational flow together influences combustion instability frequency and oscillation pressure.
Thirdly, the numerical simulation was performed on the unstable combustion in lean premixed gas combustor and the afterburner of aero-engine by CFD. The numerical simulation of lean premixed gas combustor explains the phenomenon in the experiments. It is found that the two shear layers of jets with central recirculation zone and corner recirculation zone at the sudden expansion inlet of chamber determine the structure and movement of flame. By applying different numerical models, the numerical simulation on the unstable combustion existing in the afterburner of the aero-engine at the maximum thrust state indicates that: vortex shedding after the flame holders is the important reason of inducing boost buzzing, and the fuel contribution and atomization can influence the configuration of shear layer near the lips of the flame holders as to influence the pulse of flame surface. The numerical simulation results also indicate that the numerical methods have remarkable influence on the structure and configuration of vortex; therefore it is one of difficulty to choose models for CFD research. Besides, the average flow parameters of the afterburner were obtained by CFD simulation, which provide the conditions for the application of thermoacoustic oscillation model of the afterburner.
Fourthly, the thermoacoustic oscillation in the afterburner of aero-engine was calculated using the present thermoacoustic model, and the mechanism and control laws of suppressing thermoacoustic oscillation were discussed. The results of the research indicate that the method has applicability in studying such problems and the predicted thermoacoustic frequency has enough accuracy. The calculation results for Rijke tube indicate show the importance of bias flow injection on thermoacoustic oscillation control method of active and passive combination. The review of heat release positions and heat shield structures indicates the influence laws for such structures, which provides the theory evidence for finding new control ways to supress the oscillation in afterburner of aero-engine.
引文
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