燃气透平叶片冷却通道及叶顶间隙流动和换热研究
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摘要
燃气透平常被喻为航空发动机的“心脏”。纵观航空技术的发展历程,每一次重大进展都离不开燃气透平设计及制造技术的革新。由于航空发动机应用的特殊性,它无疑集中体现了现代动力机械设计和制造的最先进水平,并代表着新一代动力机械的发展方向。正因如此,航空发动机技术历来是世界发达国家优先发展、高度垄断、严密封锁的关键技术,也是一个国家军事装备水平、科技发展、工业实力及综合国力的重要标志之一。在我国,为了满足高性能、大推力战斗机的发展需求,新一代燃气透平的研发刻不容缓。
在保持战斗机重量和尺寸不变的前提下,提高透平入口燃气温度是提高发动机性能的重要途径之一。然而这是一柄双刃剑,因为提高燃气入口温度在获得更高的发动机推重比的同时,给发动机高温部件的可靠性和使用寿命也带来巨大威胁,特别是透平叶片。因此必须采取有效的透平叶片冷却技术来应对这一威胁,以保障高温、高压、高转速工作条件下发动机的可靠性和使用寿命。从早期的叶片内部简单直接冷却技术,到如今先进的气膜冷却、层板冷却、内通道肋片与外表面气膜组合冷却技术,燃气透平发动机技术伴随着叶片冷却技术的改进而迅速发展。尽管如此,人们对透平叶片冷却技术仍然不断提出更高的要求：如何设计最佳的冷却结构在达到更好冷却效果的同时降低二次流气动损失?如何用最少的冷却介质达到更合理的冷却效率分布?自航空发动机问世以来,围绕诸如此类伴随着发动机技术发展而衍生的课题,无数科研工作者做了大量的研究工作。尽管如此,时至今日,这些研究课题依旧非常热门。
本文以实验研究为基础,结合数值模拟计算,针对航空发动机透平叶片内部肋化冷却通道以及叶片顶部间隙区的流动和换热特性,展开了一系列基础科学问题研究。研究内容主要可分为以下三个部分：
第一部分,透平叶片内部肋化通道换热及流阻特性实验。搭建实验平台,设计了两种不同几何结构的矩形肋化冷却通道：单面斜肋通道和双面交错排列斜肋通道。测量两种冷却通道的局部和平均换热系数以及肋片导致的二次流气动压力损失,综合比较了两种冷却通道的综合换热效率。所展示的测量数据不仅可为肋化通道结构设计提供参考,还可为数值模拟时选择湍流模型及检验计算方法提供验证数据源。针对附加肋片可以强化换热但同时导致流动压力损失增大这一问题,根据流场的相似性原理,本文首先提出一种肋化通道的优化设计方法,即：先固定肋片高度e,改变肋片间距p,通过实验得到综合换热效率最高的肋片间距与高度比p/e；然后固定所得到的最佳p/e,改变肋片高度e,寻求固定几何通道中的最佳肋片高度。通过上述两步优化过程,可得到固定几何形状的最佳肋片高度和间距设计参数。此优化方法的特点在于能较大地减少肋片优化设计工作量,提高肋化通道设计工作效率。
第二部分,叶顶间隙区流动和换热的实验和数值研究。利用学校工程实验中心先进的示踪粒子图像测速技术(PIV),捕捉旋转叶片顶部间隙区特征截面上的速度分布。通过不同截面上的速度分布,观察叶片顶部泄漏涡的产生和发展过程,分析泄漏流和叶栅通道流的混合特性。在实验研究的同时,本文利用商业软件,开展叶片顶部间隙流动的换热数值研究,并以实验数据为依据,比较利用五种湍流模型所得数值计算结果。数值计算结果与实验测量数据比较的表明：使用RNGk—ε湍流模型得到的数值结果与实验测量数据最接近。此外,本文通过实验还观测到叶片顶部不同位置气膜孔冷气射流对泄漏流的阻挡作用。为了深入研究气膜冷却对泄漏流的影响,本文设计了不同气膜孔排布方式和气膜孔角度,用经过实验检验的湍流模型和数值方法,比较了不同气膜孔设计方式的差异,并在此基础上提出改进叶顶气膜孔设计的建议。
第三部分,叶片旋转和间隙高度对叶顶间隙区换热的影响。在研究旋转叶片顶部的流动和换热特性时,前人采用的实验状态有三种：1)叶片和端壁都处于静止状态,2)叶片静止而端壁运动的状态,3)叶片旋转而端壁静止的状态。真实叶片的工作状态是第三种,然而受实验测量手段的制约,当前大多数实验研究是在第一种和第二种状态下进行的,第三种状态下进行的很少,而且都是针对低温低压低转速条件的。由于前两种近似实验方案与实际叶片的工况相差很大,人们不断地提出疑问：近似方案究竟能否准确模拟叶顶的流动和换热情况?不同实验研究中方案如何选取?针对前人研究叶片顶部间隙泄漏流采用不同实验方案的争议,本文以GE-E3旋转叶片为模型,在旋转速度8450r/min的条件下,选取三种间隙高度0.3,0.75,1.2mm(分别对应于叶片高度的1%,2.5%,4%),采用前人推荐并用实验数据验证过的旋转问题湍流模型,用数值模拟方法,综合分析了三种实验状态下间隙高度对叶片顶部间隙区的流动和换热特性影响。本文通过数值分析,提出不同间隙高度下与真实叶片工作状态比较接近的实验方案,从而为转子叶片顶部间隙泄漏流及冷却特性实验研究装置设计提供有力的理论依据。
Gas turbine as aero-engine is usually likened as the aircraft's heart. Throughout the development of aeronautics and astronautics technologies, every major progress was inseparable from the development of aero-engine design and manufacture technology. Due to the particularity of the aero-engine applications, there is no doubt that the design and manufacture of the aero-engine represents the most advanced technique and key development direction of modern power machinery. For this reason, the aero-engine techniques in every world power are very crucial, developed priority, monopolized highly, blockaded tightly, because the correlation techniques are seen to be one of the most important symbols of national military level, industry equipment and overall national strength. In our country, to meet the demands of supersonic and high performance aircraft, it's imperative to investigate and improve a new generation of aero-engine.
Under the premise of maintaining the same engine weight and dimension scale, one important approach to improve the engine efficiency is to increase turbine inlet temperature, but this approach is a double-edged sword. One hand side, a higher gas temperature at the inlet can bring a higher ratio of thrust to weight, but on the other hand side, the higher inlet temperature will result in a higher thermal load to the components exposed directly to extreme high temperature environment, especially to the turbine blade. Hence, more effective cooling methods must be developed to protect these components, and ensure the reliability and service life of the engine under long-term high temperature, high pressure conditions. At the beginning of turbine blade cooling design, only simple straight internal cooling channel was applied. Through several generations of cooling technique development, such as impingement cooling, film cooling and advanced film cooling, today the combination cooling techniques of internal ribbed channel and external film cooling have been widely used in modern gas turbine designs. However, more requirements of turbine blade cooling technique are still popular research topics, such as, how get optimal design of cooling structures to reduce the aerodynamic loss of the second flow? How achieve a reasonable cooling efficiency distribution with the least coolant consumption? Around these scientific questions deriving from the development of aero-engine gas turbine, numerous researchers have been carried out a lot of work in this field, since the invention of aero-engine. However, these topics are still very popular even nowadays.
This dissertation will present a series of experimental and numerical investigations on the fluid flow and heat transfer characteristics of the internal cooling air channel and tip leakage of a typical turbine blade through the following three chapters.
In the first section, the cooling air flow and heat transfer characteristics of the ribbed channels within a turbine blade were studied through experiments. Experimental platform was set up, and two different ribbed channels were designed, one channel is ribbed by inclined ribs installed on one wall, and the other channel by inclined ribs staggered installed on two opposite walls. Experiments were performed, and the aerodynamic loss, local and average heat transfer coefficients of the two channels was measured. The synthetical heat transfer efficiency of two structures was calculated and compared. The experimental data provide a reference for the designers of the ribbed channels, and can be used to validate turbulence models and calculation strategies in the numerical simulations carried out by our research group. Based on the experiments, according to the flow similar principles, a simplified optimization approach of the ribbed channels was proposed:during the first step of the optimization approach, the rib's height e is fixed, but the rib's pitch p is changed, the structure with the highest synthetical heat transfer coefficient was obtained by this step; in the second step, at the fixed p/e obtained by the first step, the rib height e is changed (but p/e maintains the value obtained by the previous step). Through the optimization of two steps, geometry parameters of a ribbed channel with the highest synthetical heat transfer coefficient can be obtained. The advantage of this optimization method is that it significantly reduces the work load of the ribbed channel design and improves work efficiency thereby.
In the second section, an experimental and numerical investigation on tip leakage flow and blade tip heat transfer is presented. Two-dimension flow fields at the certain cross sections of turbine blade tip gap were captured by the PIV (Particle Image Velocimetry) system of USTC Engineering Experimental Center. Through the two-dimensional fluid flow fields, the generation and development of the tip leakage vortex of the turbine blade was analyzed. At the same time, a numerical simulation was carried out by commercial software. The numerical results obtained by five turbulence models were compared with the experimental data measured by the PIV. Through the comparison, the RNGk-ε model was suggested as the most close to the experimental data. Based on the experimental and numerical investigations, the effect of tip film holes injection to the tip leakage flow was numerical analyzed. To reduce the tip leakage flow using the cooling air injection through film holes, different arrangements and angles of the film holes were designed, and the corresponding cooling effectiveness and reduction effect of the leakage flow were compared through numerical simulations. Based on the numerical comparisons, this chapter provides the investigators and designers with a relatively comprehensive reference regarding turbulence model choice and tip film holes design to reduce leakage flow.
In the third section, the effect of rotation and tip gap on blade tip heat transfer was numerically investigated. When investigating the characteristics of the fluid flow and heat transfer of the blade tip, three working states were used:1) both blade and endwall are stationary;2) the blade is stationary, but the endwall is rotating;3) the blade is rotating, but the endwall is stationary. The actual turbine blade worked under the third state, while resticted by experimental survey methods, most of the current experimental investigations were carried out under the first and second state. Great difference existed between the first two approximate experimental programs and actual turbine blade working condition, people doubted that whether the approximate program can accurately simulate the characteristics of flow and heat transfer on blade tip? How to select experimental program appropriately? To solve the controversy concerned with the experimental scheme of tip leakage flow, a typical GE-E3turbine blade was used as specimen, and numerical simulations were carried out at a rotating speed of8450r/min, three tip gap values0.3,0.75,1.2mm (1%,2.5%and4%of blade height respectively). The characteristics of the fluid flow and heat transfer of the blade tip gap were investigated under the three working states. Based on the numerical analysis, the better experimental schemes, which are more close to the real operation state of gas turbine rotor blade under different tip gaps, are suggested to investigators and designers of the experimental schemes of blade tip leakage flow and cooling performances.
在保持战斗机重量和尺寸不变的前提下,提高透平入口燃气温度是提高发动机性能的重要途径之一。然而这是一柄双刃剑,因为提高燃气入口温度在获得更高的发动机推重比的同时,给发动机高温部件的可靠性和使用寿命也带来巨大威胁,特别是透平叶片。因此必须采取有效的透平叶片冷却技术来应对这一威胁,以保障高温、高压、高转速工作条件下发动机的可靠性和使用寿命。从早期的叶片内部简单直接冷却技术,到如今先进的气膜冷却、层板冷却、内通道肋片与外表面气膜组合冷却技术,燃气透平发动机技术伴随着叶片冷却技术的改进而迅速发展。尽管如此,人们对透平叶片冷却技术仍然不断提出更高的要求：如何设计最佳的冷却结构在达到更好冷却效果的同时降低二次流气动损失?如何用最少的冷却介质达到更合理的冷却效率分布?自航空发动机问世以来,围绕诸如此类伴随着发动机技术发展而衍生的课题,无数科研工作者做了大量的研究工作。尽管如此,时至今日,这些研究课题依旧非常热门。
本文以实验研究为基础,结合数值模拟计算,针对航空发动机透平叶片内部肋化冷却通道以及叶片顶部间隙区的流动和换热特性,展开了一系列基础科学问题研究。研究内容主要可分为以下三个部分：
第一部分,透平叶片内部肋化通道换热及流阻特性实验。搭建实验平台,设计了两种不同几何结构的矩形肋化冷却通道：单面斜肋通道和双面交错排列斜肋通道。测量两种冷却通道的局部和平均换热系数以及肋片导致的二次流气动压力损失,综合比较了两种冷却通道的综合换热效率。所展示的测量数据不仅可为肋化通道结构设计提供参考,还可为数值模拟时选择湍流模型及检验计算方法提供验证数据源。针对附加肋片可以强化换热但同时导致流动压力损失增大这一问题,根据流场的相似性原理,本文首先提出一种肋化通道的优化设计方法,即：先固定肋片高度e,改变肋片间距p,通过实验得到综合换热效率最高的肋片间距与高度比p/e；然后固定所得到的最佳p/e,改变肋片高度e,寻求固定几何通道中的最佳肋片高度。通过上述两步优化过程,可得到固定几何形状的最佳肋片高度和间距设计参数。此优化方法的特点在于能较大地减少肋片优化设计工作量,提高肋化通道设计工作效率。
第二部分,叶顶间隙区流动和换热的实验和数值研究。利用学校工程实验中心先进的示踪粒子图像测速技术(PIV),捕捉旋转叶片顶部间隙区特征截面上的速度分布。通过不同截面上的速度分布,观察叶片顶部泄漏涡的产生和发展过程,分析泄漏流和叶栅通道流的混合特性。在实验研究的同时,本文利用商业软件,开展叶片顶部间隙流动的换热数值研究,并以实验数据为依据,比较利用五种湍流模型所得数值计算结果。数值计算结果与实验测量数据比较的表明：使用RNGk—ε湍流模型得到的数值结果与实验测量数据最接近。此外,本文通过实验还观测到叶片顶部不同位置气膜孔冷气射流对泄漏流的阻挡作用。为了深入研究气膜冷却对泄漏流的影响,本文设计了不同气膜孔排布方式和气膜孔角度,用经过实验检验的湍流模型和数值方法,比较了不同气膜孔设计方式的差异,并在此基础上提出改进叶顶气膜孔设计的建议。
第三部分,叶片旋转和间隙高度对叶顶间隙区换热的影响。在研究旋转叶片顶部的流动和换热特性时,前人采用的实验状态有三种：1)叶片和端壁都处于静止状态,2)叶片静止而端壁运动的状态,3)叶片旋转而端壁静止的状态。真实叶片的工作状态是第三种,然而受实验测量手段的制约,当前大多数实验研究是在第一种和第二种状态下进行的,第三种状态下进行的很少,而且都是针对低温低压低转速条件的。由于前两种近似实验方案与实际叶片的工况相差很大,人们不断地提出疑问：近似方案究竟能否准确模拟叶顶的流动和换热情况?不同实验研究中方案如何选取?针对前人研究叶片顶部间隙泄漏流采用不同实验方案的争议,本文以GE-E3旋转叶片为模型,在旋转速度8450r/min的条件下,选取三种间隙高度0.3,0.75,1.2mm(分别对应于叶片高度的1%,2.5%,4%),采用前人推荐并用实验数据验证过的旋转问题湍流模型,用数值模拟方法,综合分析了三种实验状态下间隙高度对叶片顶部间隙区的流动和换热特性影响。本文通过数值分析,提出不同间隙高度下与真实叶片工作状态比较接近的实验方案,从而为转子叶片顶部间隙泄漏流及冷却特性实验研究装置设计提供有力的理论依据。
Gas turbine as aero-engine is usually likened as the aircraft's heart. Throughout the development of aeronautics and astronautics technologies, every major progress was inseparable from the development of aero-engine design and manufacture technology. Due to the particularity of the aero-engine applications, there is no doubt that the design and manufacture of the aero-engine represents the most advanced technique and key development direction of modern power machinery. For this reason, the aero-engine techniques in every world power are very crucial, developed priority, monopolized highly, blockaded tightly, because the correlation techniques are seen to be one of the most important symbols of national military level, industry equipment and overall national strength. In our country, to meet the demands of supersonic and high performance aircraft, it's imperative to investigate and improve a new generation of aero-engine.
Under the premise of maintaining the same engine weight and dimension scale, one important approach to improve the engine efficiency is to increase turbine inlet temperature, but this approach is a double-edged sword. One hand side, a higher gas temperature at the inlet can bring a higher ratio of thrust to weight, but on the other hand side, the higher inlet temperature will result in a higher thermal load to the components exposed directly to extreme high temperature environment, especially to the turbine blade. Hence, more effective cooling methods must be developed to protect these components, and ensure the reliability and service life of the engine under long-term high temperature, high pressure conditions. At the beginning of turbine blade cooling design, only simple straight internal cooling channel was applied. Through several generations of cooling technique development, such as impingement cooling, film cooling and advanced film cooling, today the combination cooling techniques of internal ribbed channel and external film cooling have been widely used in modern gas turbine designs. However, more requirements of turbine blade cooling technique are still popular research topics, such as, how get optimal design of cooling structures to reduce the aerodynamic loss of the second flow? How achieve a reasonable cooling efficiency distribution with the least coolant consumption? Around these scientific questions deriving from the development of aero-engine gas turbine, numerous researchers have been carried out a lot of work in this field, since the invention of aero-engine. However, these topics are still very popular even nowadays.
This dissertation will present a series of experimental and numerical investigations on the fluid flow and heat transfer characteristics of the internal cooling air channel and tip leakage of a typical turbine blade through the following three chapters.
In the first section, the cooling air flow and heat transfer characteristics of the ribbed channels within a turbine blade were studied through experiments. Experimental platform was set up, and two different ribbed channels were designed, one channel is ribbed by inclined ribs installed on one wall, and the other channel by inclined ribs staggered installed on two opposite walls. Experiments were performed, and the aerodynamic loss, local and average heat transfer coefficients of the two channels was measured. The synthetical heat transfer efficiency of two structures was calculated and compared. The experimental data provide a reference for the designers of the ribbed channels, and can be used to validate turbulence models and calculation strategies in the numerical simulations carried out by our research group. Based on the experiments, according to the flow similar principles, a simplified optimization approach of the ribbed channels was proposed:during the first step of the optimization approach, the rib's height e is fixed, but the rib's pitch p is changed, the structure with the highest synthetical heat transfer coefficient was obtained by this step; in the second step, at the fixed p/e obtained by the first step, the rib height e is changed (but p/e maintains the value obtained by the previous step). Through the optimization of two steps, geometry parameters of a ribbed channel with the highest synthetical heat transfer coefficient can be obtained. The advantage of this optimization method is that it significantly reduces the work load of the ribbed channel design and improves work efficiency thereby.
In the second section, an experimental and numerical investigation on tip leakage flow and blade tip heat transfer is presented. Two-dimension flow fields at the certain cross sections of turbine blade tip gap were captured by the PIV (Particle Image Velocimetry) system of USTC Engineering Experimental Center. Through the two-dimensional fluid flow fields, the generation and development of the tip leakage vortex of the turbine blade was analyzed. At the same time, a numerical simulation was carried out by commercial software. The numerical results obtained by five turbulence models were compared with the experimental data measured by the PIV. Through the comparison, the RNGk-ε model was suggested as the most close to the experimental data. Based on the experimental and numerical investigations, the effect of tip film holes injection to the tip leakage flow was numerical analyzed. To reduce the tip leakage flow using the cooling air injection through film holes, different arrangements and angles of the film holes were designed, and the corresponding cooling effectiveness and reduction effect of the leakage flow were compared through numerical simulations. Based on the numerical comparisons, this chapter provides the investigators and designers with a relatively comprehensive reference regarding turbulence model choice and tip film holes design to reduce leakage flow.
In the third section, the effect of rotation and tip gap on blade tip heat transfer was numerically investigated. When investigating the characteristics of the fluid flow and heat transfer of the blade tip, three working states were used:1) both blade and endwall are stationary;2) the blade is stationary, but the endwall is rotating;3) the blade is rotating, but the endwall is stationary. The actual turbine blade worked under the third state, while resticted by experimental survey methods, most of the current experimental investigations were carried out under the first and second state. Great difference existed between the first two approximate experimental programs and actual turbine blade working condition, people doubted that whether the approximate program can accurately simulate the characteristics of flow and heat transfer on blade tip? How to select experimental program appropriately? To solve the controversy concerned with the experimental scheme of tip leakage flow, a typical GE-E3turbine blade was used as specimen, and numerical simulations were carried out at a rotating speed of8450r/min, three tip gap values0.3,0.75,1.2mm (1%,2.5%and4%of blade height respectively). The characteristics of the fluid flow and heat transfer of the blade tip gap were investigated under the three working states. Based on the numerical analysis, the better experimental schemes, which are more close to the real operation state of gas turbine rotor blade under different tip gaps, are suggested to investigators and designers of the experimental schemes of blade tip leakage flow and cooling performances.
引文
[1]张伟,2008,航空发动机.北京：航空工业出版社.
[2]王云,2009,航空发动机原理.北京：北京航空航天大学出版社.
[3]国外航空技术,发动机类第16号,1973.
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