跨音速透平叶栅尾缘劈缝射流的数值研究

采用Spalart-Allmaras(S-A)湍流模型对跨音速导叶尾缘劈缝射流的定常流动结构进行了模拟分析,研究不同尾缘射流压比对尾缘激波结构与强度、尾迹形态、各种能量损失的影响规律.结果表明:劈缝射流可以减小尾迹宽度与低速峰值,降低尾缘燕尾波的强度,射流对压力面侧激波的削弱作用更大;射流使燕尾波的形成位置更接近尾缘,导致燕尾波张角增大;射流可以降低叶栅的总动能损失,压比对激波损失和尾迹损失的影响更明显,但对边界层损失的影响较小;根据叶栅出口的状态可知,存在一个最佳的射流压比.     

跨声速涡轮叶片半劈缝长度及冷气量对叶栅流场的影响

通过在不同尾部劈缝结构以及冷气量情况下对某型跨声速叶栅数值模拟,得出了劈缝长度及冷气量对叶栅流道内及尾缘附近流场结构影响的规律。主要表现为:尾缘劈缝结构使叶片尾缘内伸波变为两道;长尾缘劈缝以及大尾缘冷气量不仅能够减小尾迹宽度、降低尾缘损失,也能够使叶片吸力面分离泡减小,黏性损失降低。     

基于强化换热的偏斜尾缘设计

采用SST湍流模型对偏斜尾劈缝射流的流动结构与传热特点进行模拟分析,研究不同尾缘射流速比,入口攻角及偏折角对于尾缘位置边界层状态、涡流结构、尾迹形态及流动损失的影响规律.结果表明:偏斜尾缘劈缝射流的使用能够有效改善尾缘位置的换热性能,射流对于尾缘折转区的冲刷作用使得吸力面侧得到较好的冷却;偏斜尾缘的采用也能够有效抑制大...     

尾缘造型及冷气喷射对跨声速涡轮叶栅损失的影响

通过对不同尾缘造型、不同尾缘冷气喷射量下某跨声速涡轮叶栅的数值模拟,初步得出了尾缘劈缝冷却对尾缘损失以及叶栅能量损失影响的规律。其主要表现为:从减小叶栅能量损失角度来讲,尾缘冷气喷射流量存在最佳值,且随劈缝长度增加,此最佳冷气喷射流量越小;从减小尾缘激波强度角度来讲,较大的冷气流量以及较长的尾缘劈缝有利于减小激波损失,但会消耗过多冷气并增加掺混损失,导致总损失增加。     

表面粗糙度对压气机叶栅流动特性的影响

在低速平面叶栅风洞中,实验研究了表面粗糙度对高负荷压气机流动特性的影响,并对叶片吸力面不同位置布置的表面粗糙度进行了对比分析。通过墨迹流场显示法对叶栅壁面流场进行了测量,利用五孔气动探针对叶栅出口截面进行了扫掠,给出了不同方案出口截面马赫数、二次流速度矢量的分布以及叶栅的流场特征,以分析和探讨表面粗糙度对叶栅流动特性的影响。结果表明,吸力面局部表面粗糙度的增加使得角区分离范围减小;且随着粗糙带向尾缘移动,角区分离范围的减小程度也逐渐增加。     

高亚音速大折转角叶型内激波结构与损失分析

本文采用组合多项式曲线构造了具有高亚音进口条件的大折转角压气机静叶叶型,探索了进一步提高跨音速压气机负荷时,静叶根部区域可能存在的激波结构和损失特征.采用数值模拟方法对不同条件下的高亚音速大折转角压气机叶栅流场进行了数值模拟,结果表明:叶栅内的激波结构与进口马赫数、攻角以及叶型的转角等参数密切相关.通过对叶栅出口的损失分析发现,激波与附面层相互作用改变了原有附面层内的损失分布规律,形成了由激波强度和位置所决定的沿叶片表面法线方向大小基本不变的高损失区域,叶型损失的大小和激波与吸力面最低压力点之间的相对位置密切相关.     

叶栅内凝结激波与气动激波干涉的数值研究

对自发凝结两相流动中凝结激波与气动激波之间的干涉进行了数值研究,气相采用N-S方程,液相凝结过程应用构造的多阶复合参数进行积分求解.模拟得到的数值纹影图显示了不同出口马赫数情况下汽轮机叶栅中过热蒸汽流动和自发凝结流动激波系的分布.结果表明:凝结激波会影响气动激波的强度和出口气流角,消弱气动激波在吸力面反射引起的边界层分离现象,增强尾迹的强度,并影响相关损失的产生.     

燃气轮机高压涡轮动静叶非定常干涉数值研究

采用数值模拟方法对燃气轮机跨声速涡轮级内动静叶的非定常气动干扰进行了深入的研究,分析了静叶尾迹输运、尾缘激波和尾迹涡的变化以及下游动叶静压的非定常扰动的影响。结果表明:燃气轮机跨音速涡轮级中静叶的非定常特定区域集中在吸力面扩压段到尾缘区域;尾缘激波的压力面分支在吸力面的反射激波对有下游动叶的影响非常大;尾迹传播方向随着下游势流的变化而产生相应的变化。     

基于LES的压气机叶栅通道非定常流动结构研究

为了进一步理解压气机叶栅通道内的非定常流动结构,采用大涡模拟(LES)方法研究了来流附面层厚度和稠度变化对叶栅通道内涡系结构及总压损失系数的影响。研究表明：来流附面层增厚将导致端壁处流体的轴向动能降低,使得马蹄涡压力面分支更早地流向相邻叶片吸力面;来流附面层越厚,通道涡在叶栅尾缘沿展向抬升的高度越高,角区分离的范围也越大;叶栅的总压损失随附面层增厚而增加,附面层损失增加显著,二次流损失有所增大;稠度较低时叶栅吸力面表面存在分离,会对通道涡及角区分离产生影响;稠度增大,横向压力梯度减小,叶栅流道的速度分布更均匀,通道涡的强度和尺度减小,角区分离的范围减小;稠度增大使叶表不再分离时,总压损失显著降低,但稠度继续增大会使气流与叶片表面的摩擦损失增加。     

吹风比和唇板厚度对叶片尾缘气膜冷却的影响研究

为了探究吹风比、唇板厚度对叶片尾缘半劈缝冷却结构气膜冷却特性的影响,采用数值模拟方法对比唇板厚度为4,5和3 mm,吹风比Br为0.5,0.8,1.0和1.5条件下叶片尾缘后台阶上的气膜冷却效率。结果表明：在吹风比Br为0.5时,叶片尾缘后台阶上产生的回流区大,冷气向展向扩散范围广,冷气在近劈缝一端向展向覆盖的较好,由于吹风比小,冷气流速慢,动量小,在后台阶远端燃气与冷气掺混量大,导致冷气冷却能力降低;在大吹风比下（Br=1.5）,冷气流速快,冷气从劈缝射出集中覆盖在劈缝下游处,而肋下游冷气覆盖效果差。唇板厚度影响着唇板出口处形成的回流区,增大唇板厚度将导致半劈缝出口气流分离所产生的涡强度变大,促进燃气与冷气的掺混,降低冷却效率,薄唇板会使尾缘气膜冷却效率显著提高。     

空心静叶缝隙抽吸对蒸汽流场影响的数值研究

在缝隙抽吸条件下,采用Fluent软件对某600MW汽轮机末级空心静叶栅内的蒸汽流场进行了三维数值模拟,讨论了缝隙位置与结构参数对主蒸汽流场的影响。结果表明:随着缝隙位置从叶片前缘向叶片尾缘的移动,汽流总压损失系数逐渐减小。缝隙抽吸使叶栅进口马赫数有所升高,且内弧上的缝隙抽吸对马赫数的影响要大于背弧上的影响;当缝隙角度α=45°时,缝隙抽吸对叶栅进口马赫数的影响较大;缝隙宽度的增大使叶栅进口马赫数升高的幅度也越大。缝隙抽吸对叶片表面压力分布总体影响不大,仅在缝隙附近的汽流静压有较大的降低;随着缝隙宽度的增加,缝隙处的汽流静压降低越大。另外,缝隙抽吸使叶片表面的汽流边界层有所减薄。     

Numerical investigation on the effects of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade

This paper presents a detailed numerical investigation of the influence of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade. The boundary layer suction (BLS) was used to control the location and strength of shock waves, with the aspirated slot locating at 49% chord, where is just downstream of the impingement point of shock wave at the leading edge. The numerical simulation is based on NUMECA, a commercial software, where the cell-centered control volume approach with third-order spatial accuracy is used to solve the 3-D Reynolds-averaged Navier-Stokes equations under the Cartesian coordinate system. Several conclusions can be made through the observation of the numerical results. (1) Multiple shock waves in cascade passage leaded the velocity deficits of boundary layer on suction surface downstream of shock wave, resulting in seriously separated flow on the suction side of blade, especially when the front shock wave is much stronger than the rest of the shocks. (2) BLS with small mass flow rate can not effectively improve the boundary layer. When the impingement point of oblique shock wave coming from cascade leading edge is bled to downstream of the passage shock wave by BLS, the boundary layer flow is greatly improved. However, if the BLS mass flow rate exceeds a critical value (1.2%), the boundary layer downstream of shock wave would separate from suction surface. (3) At the blade mid-span, the aerodynamic performance of compressor blade is improved as BLS mass flow rate increases. The optimum BLS is about 1.2%. Compared with the baseline case, the BLS with flow rate of 1.2% increases the total pressure recovery coefficient by 12%, and decreases diffusion factor by 18% and deviation angle to 7 ° while keeping the pressure rise constant. (4) The three dimensional flow structure of the compressor cascade ranged from 25% span to 75% span was improved greatly with the 1.2% BLS flow rate. However it could not control the development of the corner boundary layer effectively.     

The influence of synthetic jet excitation on secondary flow in compressor cascade

The detailed numerical simulation has been carried out to investigate the effect of synthetic jet excitation on the secondary flow at 5° incidence in a compressor cascade, in which the synthetic jet actuation is equipped on the suction surface. The influence of excitation position, one fixed near the trailing edge and the other fixed a little far from the trailing edge, has also been studied. The results show that unsteady disturbance of desirable synthetic jet effectively enhances the mixing of the fluid inside the separation region, which reduces the vortex intensity and the energy loss, improves the flow status in the cascade, and also suppresses velocity fluctuation near the trailing edge. Additionally, the actuation fixed near the separation region proves to be more effective and exit load distribution is more uniform due to the employment of the synthetic jet.     

An improvement on the efficiency of a single rotor transonic compressor by reducing the shock wave strength on the blade suction surfaces

It is well known that increasing the rotational velocity is an effective way to increase the total pressure ratio. With increasing flow velocity especially under the condition of transonic flow in the supersonic region, where exist strong shock waves, the shock wave loss becomes main and important. Simultaneously, there occurs boundary layer separation due to the shock wave / boundary layer interaction. In the present paper the transonic compressor blades were studied and analyzed to find a proper and simple way to reduce the shock wave loss by optimizing the suction surface configuration or controlling the gradient of isentropic Mach number on the suction surface. A Navier-Stokes solver combined with a modified design algorithm was developed and used. The NASA single rotor for transonic flow compressor was served as a numerical example to show the effectiveness of this method. Two cases for both original and modified rotors were analyzed and compared.     

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.     

超微涡轮动叶栅叶顶间隙对流场影响的数值模拟

通过数值求解基于雷诺时均的三维定常粘性N-S方程,结合RNGk-ε湍流模型和非平衡壁面函数,对一种超微型向心涡轮动叶栅内的流动情况进行了数值模拟。揭示了具有极低展弦比动叶栅叶顶间隙对流场参数分布和气动损失的影响,为超微涡轮的设计和改进提供了理论依据。模拟结果表明,叶顶间隙的大小对通道内马赫数分布有重要影响,其中顶部间隙射流所引发的泄漏涡与主流的掺混是主流马赫数降低的重要原因;叶顶间隙的存在使得总压损失系数均匀化,即近壁区和主流区的总压损失都较高;动叶栅在叶展方向上的载荷分布均匀,弦向载荷主要由接近尾缘的弧段承担;模拟中还解析出三维的尾迹涡,这主要是动叶栅尾缘过厚所导致,应进行叶型改进。     

导叶约化中心位置对涡轮非定常气动计算的影响

针对工程上常用的基于叶片数约化的涡轮非定常计算方法,为了掌握导叶约化中心位置对于涡轮内部流动非定常计算结果的影响规律,对导叶前缘约化、导叶尾缘约化、不约化三个算例进行了非定常计算与对比分析.研究结果表明:导叶前缘约化方式对于涡轮气动性能时均值影响量级在1％以上,而导叶尾缘约化方式的影响不到前者的1/2;两种约化方式均能...     

Aerodynamic Performance Improvement of a Highly Loaded Compressor Airfoil with Coanda Jet Flap

Coanda jet flap is an effective flow control technique,which offers pressurized high streamwise velocity to eliminate the boundary layer flow separation and increase the aerodynamic loading of compressor blades.Traditionally,there is only single-jet flap on the blade suction side.A novel Coanda double-jet flap configuration combining the front-jet slot near the blade leading edge and the rear-jet slot near the blade trailing edge is proposed and investigated in this paper.The reference highly loaded compressor profile is the Zierke&Deutsch double-circular-arc airfoil with the diffusion factor of 0.66.Firstly,three types of Coanda jet flap configurations including front-jet,rear-jet and the novel double-jet flaps are designed based on the 2D flow fields in the highly loaded compressor blade passage.The Back Propagation Neural Network(BPNN)combined with the genetic algorithm(GA)is adopted to obtain the optimal geometry for each type of Coanda jet flap configuration.Numerical simulations are then performed to understand the effects of the three optimal Coanda jet flaps on the compressor airfoil performance.Results indicate all the three types of Coanda jet flaps effectively improve the aerodynamic performance of the highly loaded airfoil,and the Coanda double-jet flap behaves best in controlling the boundary layer flow separation.At the inlet flow condition with incidence angle of 5°,the total pressure loss coefficient is reduced by 52.5%and the static pressure rise coefficient is increased by 25.7%with Coanda double-jet flap when the normalized jet mass flow ratio of the front jet and the rear jet is equal to 1.5%and 0.5%,respectively.The impacts of geometric parameters and jet mass flow ratios on the airfoil aerodynamic performance are further analyzed.It is observed that the geometric design parameters of Coanda double-jet flap determine airfoil thickness and jet slot position,which plays the key role in supressing flow separation on the airfoil suction side.Furthermore,there exists an optimal combination of front-jet and rear-jet mass flow ratios to achieve the minimum flow loss at each incidence angle of incoming flow.These results indicate that Coanda double-jet flap combining the adjust of jet mass flow rate varying with the incidence angle of incoming flow would be a promising adaptive flow control technique.     

导流板抽吸对扇形叶栅实验件流场的周期性影响

为保证导叶-静叶结构的扇形叶栅实验顺利开展,针对已有扇形叶栅实验件,设计了两种导流板抽吸方案:两侧导流板流道转接的位置设置抽吸缝和在前一方案的基础上增加左侧抽吸缝的抽吸量、同时不在右侧设抽吸缝。两种方案中都于导叶与静叶转接位置开设矩形抽吸缝,宽度均为2 mm。对不同方案的实验件流场进行数值模拟,通过对比分析导叶及静叶栅出口气动参数和流场结构,确定了能够大幅提高实验件流场分布周向均匀性的导流板抽吸和结构改进方案。研究表明:导流板结构改进和设置抽吸缝都可以在一定程度上改善流场的周期性;导流板抽吸缝开设在气流分离区,可减小分离强度范围,改善实验件整体周期性;第2种方案可大幅提高实验件流场分布的周向均匀性,使可测量流道数增加到7个。     

变几何涡轮可调叶栅过渡态特性研究

变几何涡轮使发动机在变工况下的性能得到提升,为了更透彻地了解变几何涡轮导叶转动过程中参数的变化情况,通过数值模拟及试验方法探究可调叶栅过渡态特性。将变几何涡轮导叶进行调节,导叶调大范围为0°~6°,导叶调小范围为0°~-5°,观察过渡态参数变化规律。试验研究表明：导叶在调大及调小过程中,导叶出口质量流量、绝对气流角和绝对马赫数随转角接近线性变化,导叶出口总压损失系数和熵增接近抛物线变化;导叶从0°向-5°转动过程绝对出口马赫数减小了2.2%,总压损失系数增加了37.3%;导叶从0°向6°转动过程中,导叶出口马赫数增加了1.5%,导叶出口总压损失系数减小了15.8%;在导叶转角和二次流改变的影响下,吸力侧和压力侧来流在导叶尾缘后掺混改变,沿叶高分布的出口绝对气流角不同程度地偏离几何出口角;导叶转角调大,上部通道涡沿叶高上移,泄漏涡和通道涡相互削弱,总压损失系数减小。