Study on the influence of pneumatic probe on the wake flow field of transonic turbine cascade

The pneumatic probe is widely used for contact measurements in turbomachinery flow field research.However,it inevitably interferes with the original flow field,leading to additional errors,particularly in wake flow fields or transonic regions with significant pressure gradients.This study employed Reynolds-Averaged Navier-Stokes delete and high-fidelity numerical simulation to investigate the impact of an inserted pneumatic probe on the wake flow field of a transonic turbine blade and compared it to the baseline flow field.Results indicate that the probe causes the shock waves premature occurrence in the high subsonic wake region near the turbine blade trailing edge.These shock waves affect vortex shedding by thickening the boundary layer near the trailing edge and changing the shedding pattern from high-frequency-low-energy to low-frequencyhigh-energy.In addition,the extra flow loss is incurred,and the blade's heat transfer characteristic is changed.This research provides a reference for testing experiments in complex transonic flow fields,guiding experimental researchers to minimize instrument interference with the original flow field.

EXPERIMENTAL MEASUREMENT AND NUMERICAL SIMULATION FOR FLOW FIELD AND FILM COOLING EFFECTIVENESS IN FILM-COOLED TURBINE

Numerical simulation of three-dimensional flow field and film cooling effectiveness in film-cooled turbine rotor and stationary turbine cascade were carried out by using the k- ε turbulence model, and the predictions of the three-dimensional velocities were compared with the measured results by Laser-Doppler Velocimetry （LDV）. Results reveal the secondary flow near the blade surface in the wake region behind the jet hole. Compared with the stationary cascade, there are the centrifugal force and Coriolis force existing in the flow field of the turbine rotor, and these forces make the three-dimensional flow field change in the turbine rotor, especially for the radial velocity. The effect of rotation on the flow field and the film cooling effectiveness on the pressure side is more apparent than that on the suction side as is shown in the computational and measured results, and the low film cooling effectiveness appears on the pressure surface of the turbine rotor blade compared with that of the stationary cascade.

Numerical Simulation on Unsteady Flow Mechanism of a 1.5-Stage Axial Transonic Compressor

In this paper,a numerical simulation method is used to calculate a 1.5-stage axial transonic compressor to explore its unsteady flow mechanism.The performance curve is compared with the experimental data to verify the calculation method with a high numerical accuracy,which shows that the unsteady calculation has good reliability.According to the analysis of the data from the monitoring points under the near-stall condition,the unsteady disturbances originate from the tip region of blade and perform the strongest at the blade pressure surface with a broadband characteristic.Further analysis is conducted by combining with the characteristics of the transient flow field at the tip of blade.The results show that the unsteady pressure fluctuations are caused by the migration of the new vortex cores.These new vortex cores are generated by the breakdown of leakage vortex in the downstream,which is induced by the leakage vortex and shock wave interference.Moreover,the relationship between the unsteady flow characteristics and the working conditions is also studied.The leakage vortex intensity and the shock wave strength gradually increase with the decrease of flow rate.When the combination of the leakage vortex intensity and shock wave strength reaches the first threshold,a single frequency of unsteady disturbances appears at the blade tip.When the combination of the leakage vortex intensity and shock wave strength reaches the second threshold,the frequency of unsteady disturbances changes to a broadband.

Flow Control Effect of Spanwise Distributed Pulsed Arc Discharge Plasma Actuation on Supersonic Compressor Cascade Flow

To achieve efficient control of supersonic compressor cascade flow,a type of spanwise distributed pulsed arc discharge plasma actuation(PADPA)was designed.To simulate the influences of PADPA on the flow field,a phenomenological model was established.Then,the flow control effects of PADPA on supersonic compressor cascade flow were researched numerically.The results show that under low static pressure ratio condition,the compressive wave induced by PADPA reduced the intensity of the passage shock wave,which eventually reduced shock wave loss.It was also found that PADPA produced an adverse pressure gradient(pre-compression effect)around the actuation location,which reduced the strength of the high adverse pressure gradient induced by the passage shock wave.The airflow on both sides of the actuation location was accelerated by PADPA owing to the spanwise distributed layout.Thus,it improved the ability of the boundary layer to resist the effect of the adverse pressure gradient and reduced the separation zone.Consequently,the total pressure loss was reduced by 6.8%.Under high pressure ratio condition,the effect of PADPA on the suction side controlling the large separation of the boundary layer was insignificant.The total pressure loss also increased slightly.

Computations of unsteady transonic flow over airfoil at low Reynolds number

Transonic flow over a thin airfoil at low Reynolds number was studied numerically by directly solving two-dimensional full Navier-Stokes equations through 5th order weighted essentially non-oscillatory(WENO) scheme without using any turbulence model.A series of distinguished unsteady phenomena for a thin 2-D transonic airfoil flow were presented.Due to continuous adverse pressure gradient in the subsonic flow downstream of the sonic line,the unsteady separated boundary layer with main vortex and secondary vortex was developed at the rear of the airfoil.At the trailing edge,the vortex-shedding was characterized by periodical connection of the main vortex and secondary vortex on the other side of the airfoil.The unsteady separation and vortex-shedding occurred with the same period.On the airfoil surface,the average pulse pressure related to the unsteady supersonic region was obviously smaller than that related to the vortex-shedding at the trailing edge.With the attack angle increasing from 0°to 2°,the frequency of vortex-shedding decreases about 4.2%.At last,the turbulence intensity and many second-order statistics in the wake region were investigated.

A Study of 2D Separated Cascade Flow at Large Angles of Attack Using Euler Equations

This paper presents a numerical simulation method developed for separated flow in cascades using the Euler equations and demonstrates the feasibility of this method.MacCormack's two-steps explicit finite difference scheme is used to discretize the equations in conservation form,and the artificial viscosity is added to the dis- cretized inviscid equations by means of the self-adapted filter technique.The initial separation boundary is given according to simple experimental results.The numerical simulation results including subsonic and transonic turbine cascades flow with or without separation show that the fundamental idea of this numerical method is reasonable and simple.The present study indicates that for solving certain engineering problems it is a simple and effective tool for adding some viscosity corrections to inviscid flow model,especially the current when the Navier-Stokes equations have not been solved very effectively for various complicated flows in turbomachinery.

叶片破损对压气机性能的影响

为了研究叶片破损对压气机内部流场的影响机理,本文针对带有破损的跨音速压气机叶片进行气动性能变化规律和机理研究。通过NASA rotor 37实验结果验证了计算方法的准确性,利用数值仿真研究了叶片破损对压气性能的影响规律,并对2种典型工况下叶片破损前后的流场结构进行对比分析,探讨引起性能变化的具体原因。研究表明:叶片破损会导致压气机的压比降低0.6631%、绝热效率降低0.7877%,且压气机的稳定裕度降低更为明显。叶尖破损主要对叶顶流场结构产生影响,使叶顶区域泄漏流增强,当其与激波相互作用时,通道内产生了流动堵塞和流动损失。本文研究成果可为实际服役中的航空发动机特性评定提供理论参考。

高负荷涡轮叶栅风洞流动周期性研究

为设计符合要求的平面叶栅试验段以开展跨音速高负荷平面叶栅气动试验,以高负荷涡轮转子叶栅为研究对象,采用数值模拟方法研究了不同来流攻角和不同出口等熵马赫数对高负荷涡轮平面叶栅进口流场的均匀性与叶栅通道流场周期性的影响.结果表明:平面叶栅进口流场沿周向呈现出不均匀的波动分布,靠近吸力面侧壁的三个通道的来流均匀性和准确性普遍较差,叶栅中部和靠近压力面侧壁的通道进口均匀性和准确性较好;进口流场品质直接影响着叶栅通道内流场的周期性,而来流攻角对叶栅进口流场品质的影响比出口等熵马赫数更明显,在正攻角下,攻角偏差极限值约为0.7°,不同叶栅通道进口流场均匀性和准确性较好;从亚音速流动到超音速流动,靠近试验件吸力面侧壁的叶栅通道产生强激波,致使叶片80%轴向弦长处压力误差值增大,叶栅通道流场周期性下降.

A combined application of micro-vortex generator and boundary layer suction in a high-load compressor cascade

In the current study, the effects of a combined application between micro-vortex generator and boundary layer suction on the flow characteristics of a high-load compressor cascade are investigated. The micro-vortex generator with a special configuration and the longitudinal suction slot are adopted. The calculated results show that a reverse flow region, which is considered the main reason for occurring stall at 7.9° incidence, grows and collapses rapidly near the leading edge and leads to two critical points occurring on the end-wall with the increasing incidence in the baseline. As the micro-vortex generator is introduced in the baseline cascade, the corner separation is switched to a trailing edge separation by the thrust from the induced vortex. Meanwhile, the occurrence of failure is delayed due to the mixed low energy fluid and main flow. The synergistic effects between the micro-vortex generator and the boundary layer suction on the performance of the cascade are superior to the baseline at all the incidence conditions before the occurrence of failure, and the sudden deterioration of the cascade occurs at 10.3° incidence. The optimal results show that the farther upstream suction position, the lower total pressure loss of the cascade with vortex generator at the near stall condition. Moreover, the induced vortex with a leg can migrate the accumulated low energy fluid backward to delay the occurrence of stall.

超声速差法平面叶栅流场试验和数值研究

针对平面叶栅狭小通道中流场非接触测量难题,提出一种基于超声速差原理的二维流场测量方法。设计一种2×4超声波斜、平口换能器组合阵列,研制一套多通道超声法测量和气流速度场重建系统。开展平面叶栅流场测量试验,成功获取了3种不同进口速度条件下流道二维速度场。利用COMSOL Multiphysics中CFD模块同步开展叶栅流场数值模拟,并与试验结果对比。结果表明:超声测量与数值模拟结果较为接近,速度大小和偏转角误差均在5%以内,较好地反映叶栅通道的流动特点。提出的超声测量技术有望进一步用于高马赫数流场研究。

进口雷诺数和湍流度对压气机叶栅流动特性影响

针对高空条件下,无人机发动机的压气机气动性能明显下降的问题,本文以某高亚声速压气机叶型V103为研究对象,采用数值模拟方法研究了3种雷诺数(Re=1.5×10^(5),4.5×10^(5),1.2×106)条件下,不同来流湍流度(Tu=1%,3%,5%,6%,7.5%,10%)对压气机叶栅内部流动特性的影响,分析了流动分离和转捩之间的变化关系,以及由此引起的叶型损失变化规律。研究结果表明:当湍流度不变时,随着雷诺数减小,分离泡长度增加,但分离点位置不变;当雷诺数不变,湍流度减小,层流分离点位置会提前,并且前移现象在低雷诺数下更明显;当湍流度增加,叶片表面分离泡消失,但损失仍然呈现出随着湍流度增加而增大的趋势;在不同雷诺数条件下,都存在一个对应于最小叶型损失的来流湍流度,且该湍流度随着雷诺数的减小而略有增大。本文为修正低雷诺数条件下压气机叶型损失模型和发展预测压气机高空低雷诺数效应的数值模型提供理论指导。

一种多级降压调节阀的降压能力研究

通过理论分析得到多级降压串式调节阀的阀芯与阀座之间的流通面积、相对流阻系数与压降比的关系,若多级降压串式调节阀的各级流通面积和流道形式完全相同,则流阻系数相同,进而各级分压相同。以公称通径DN50,公称压力900Lb的三级降压串式调节阀为例,采用Fluent数值模拟分析,验证不同阀芯与阀座的流通面积分布对压降比的影响,给出阀座级间径向流通面积与轴向流通面积的推荐值;并推导出各串级间的相对流阻系数关系,便于多级降压串式调节阀的压降计算;结论同样适用于同规格不同降压级数串式调节阀的压降分析与计算,对工程设计具有指导或借鉴意义。

串列静子叶片流动机理研究

提高级负荷水平一直是压气机气动设计领域不断追求的目标。串列叶片可以突破常规布局气动负荷的限制,且具有良好的工程应用前景。为了探索串列叶片三维流动机理及前后排相互影响机制,利用数值模拟方法对来流马赫数为0.8的串列静子叶片近端壁流动特性展开研究,分析了前后排相互影响机制,主要是针对前/后排叶片单排工作与串列条件下进行对比。通过研究表明,以后排叶片势作用、缝隙射流、尾迹扩散为代表的典型流动现象显著地影响了前后排叶片的扩压过程,进而影响了气动负荷沿流向的分配及端壁损失,串列静子近端壁的气动设计需要考虑前后排相互干涉的影响。

基于木星大气环境的降落伞系统气动特性研究

木星探测是未来行星探测的重要发展方向,而降落伞是进入木星大气探测必不可少的气动减速装置。文章基于“伽利略号”探测任务,设计了满足未来木星探测需求的降落伞系统简化模型,并针对该降落伞系统进行了数值模拟,研究了木星大气和地球风洞实验环境中不同来流马赫数下降落伞系统的复杂流动现象及气动力变化规律。在木星大气环境中,降落伞的阻力系数和横向力系数大小以及横向力系数波动幅度均高于风洞试验环境,阻力系数波动幅度均低于风洞实验环境。此外,还研究了木星大气环境中不同来流攻角下降落伞系统的气动特性。研究表明,木星大气环境中降落伞系统气动特性与风洞实验结果有差异,因此未来在设计用于木星探测的降落伞系统时,应考虑由于木星大气环境对降落伞系统气动特性的影响。

端壁非定常脉冲射流对高速扩压叶栅性能的影响

为探究非定常脉冲振荡射流对高速平面扩压叶栅气动性能、分离流动控制以及流场结构的影响,基于CFX数值模拟方法对平面扩压叶栅进行端壁非定常脉冲射流研究,分析射流效果随射流位置、角度和强度的变化规律。结果表明,通过角区脉冲射流可以显著提高叶栅气动性能,仅采用不足叶栅主流0.3%的射流流量,就能使叶栅出口总压损失系数降低28.66%。当射流位于吸力面侧分离起始位置稍下游时控制效果最佳;射流角度、射流强度和射流频率的最佳值分别为α=20°,C_(u)=110%和F^(+)=0.80;脉冲射流具有较好的适应性,在来流冲角i=-8°~+4°内均能降低叶栅损失。脉冲射流主要通过抑制和推迟通道涡和集中脱落涡的发展,减小其影响范围来改善叶栅内的涡系结构。

压气机平面叶栅试验台中尾板调节的数值研究

针对压气机平面叶栅试验台中尾板影响叶栅出口流场周期性的特征,以包含高亚音静子叶栅的全尺寸叶栅试验台为研究对象,数值研究了尾板长度和角度改变时叶栅出口流场的周期性变化情况。结果表明:相同叶栅进出口压比情况下,尾板越长,叶栅进口马赫数越高;在高亚音马赫数进口情况下,不同进气攻角下的压气机叶栅流场在不同尾板长度下都保持较好的周期性;在较低亚音马赫数进口情况下,选取较长尾板并进行角度调节有利于改进不同攻角下叶栅流场的周期性。

大偏转角时超声速五孔探针激波图谱数值研究

激波是超声速流动的关键性问题,而超声速来流条件下气动探针的激波图谱的研究有待进一步丰富。本文以适用于超声速来流条件下的五孔气动探针为研究对象,提出一种复合型五孔压力-温度探针的结构设计方法,并采用经过实验校核的数值方法分析跨、超声速来流绕流探针而形成的激波图谱的变化规律,其中着重关注在正对来流和相对于来流存在较大偏转角度时探针头部及杆体的激波结构。研究结果表明:跨声速来流条件下,探针头部前端会形成一道强度较大的正激波结构;超声速来流条件下,探针与来流存在较大偏转角度时,探针背风面杆体后端会形成“λ”型激波,随着来流马赫数增大最终变为正激波,且激波与附面层相互作用导致探针表面的附面层发生分离,产生了低能流体团。本文的研究成果实现探针头部和杆体激波图谱的再现,充实了超声速探针的基础数据库。

使用PAFV湍流模型对叶栅流动的数值研究

为进一步发展更有效的湍流模型,提高对高负荷跨声速压气机流动的数值计算精度,以侧偏平均脉动速度和应变率张量为基础构造了新的侧偏平均脉动速度(partial average fluctuation velocity, PAFV)湍流模型。在使用Python计算机语言进行编程计算过程中,将流动控制方程在一般曲线坐标系下变换后,采用有限差分方法进行空间离散。方程对流项采用Steger-Warming通量向量分裂法,扩散项采用中心差分格式离散,时间导数项采用具有TVD(total variation diminishing)性质的二阶Runge-Kutta法离散。使用该湍流模型对德国航空航天研究试验院的L030-4叶栅亚声速和超声速流动进行了计算,在进行网格无关性研究基础上,获得了流场压力等值线云图、马赫数等值线云图,压力系数分布等结果。计算显示侧偏平均脉动速度在激波附近区域受到明显抑制,避免了激波附近区域出现过大的湍流黏性。数值计算结果与实验符合比较好,初步验证了PAFV湍流模型可以有效地模拟该类型的叶栅流动。

高负荷平面叶栅风洞流场品质改进措施数值研究

为研究不同改进措施对平面叶栅风洞流场品质的改进效果与机理,以高负荷扩压叶栅为研究对象,利用数值模拟方法研究了试验器进口段上下侧壁抽吸措施、出口导流尾板措施、抽吸与尾板组合措施对叶栅风洞流场品质的改进效果。研究结果表明:上下侧壁抽吸减小了进口段的静压梯度,减弱了上下侧壁附面层对进口流体流向的影响;导流尾板减弱了外界大气对上侧壁附近通道的影响;抽吸与尾板组合措施结合了侧壁抽吸与导流尾板的优点,使得进口静压分布均匀,通道间压差阻力差异减小,对叶栅试验段流场品质的改进效果优于其他方案。在当前马赫数工况下可使叶栅进口64%的通道范围马赫数误差<0.01,冲角误差<0.5°,有三个连续通道出口周期性较好,并且轴向密流比<1.15。

跨声速压气机转子的二次流旋涡结构

为了明确跨声速轴流压气机内部流场结构,数值模拟了NASA Rotor37转子,结合λ2准则分析流场参数,探索流动的规律和旋涡结构。研究发现,压气机转子的旋涡模型主要由马蹄涡、壁角涡、径向涡、脱落涡、泄漏涡、诱导涡和分离涡等7个旋涡组成。马蹄涡吸力面分支耗散,压力面分支向相邻的吸力面发展。壁角涡与脱落涡位于叶根角区,引起流动损失和角区失速。径向涡位于激波后吸力面的分离区内,它扩大吸力面分离、引起低能流体向叶顶堆积。激波与叶尖泄漏在叶顶通道中形成3涡:泄漏涡、诱导涡和分离涡,而叶栅通道出口存在分离涡和由泄漏涡与诱导涡合成的叶顶通道涡。泄漏涡与诱导涡破碎在流道中间产生的堵塞区,分离涡造成吸力面尾缘的低速区,共同触发跨声速压气机的失稳。