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.

Topological analysis of plasma flow control on corner separation in a highly loaded compressor cascade

In this paper, flow behavior and topology structure in a highly loaded compressor cascade with and without plasma aerodynamic actuation （PAA） are investigated. Streamline pattern, total pressure loss coefficient, outlet flow angle and topological analysis are considered to study the effect and mechanism of the plasma flow control on corner separation. Results presented include the boundary layer flow behavior, effects of three types of PAA on separated flows and performance parameters, topology structures and sequences of singular points with and without PAA. Two separation lines, reversed flow and backflow exist on the suction surface. The cross flow on the endwall is an important element for the comer separation. PAA can reduce the undertuming and overturning as well as the total pressure loss, leading to an overall increase of flow turning and enhancement of aerodynamic performance. PAA can change the topology structure, sequences of singular points and their corresponding separation lines. Types II and III PAA are much more efficient in controlling comer separation and enhancing aerodynamic performances than type I.

Effects of Plasma Aerodynamic Actuation on Corner Separation in a Highly Loaded Compressor Cascade

This paper reports experimental results on the effects of plasma aerodynamic actua- tion （PAA） on corner separation control in a highly loaded, low speed, linear compressor cascade. Total pressure loss coefficient distribution was adopted to evaluate the corner separation control effect in wind tunnel experiments. Results of pressure measurements and particle image velocime- try （PIV） show that the control effect of pitch-wise PAA on the endwall is much better than that of stream-wise PAA on the suction surface. When both the pitch-wise PAA on the endwall and stream-wise PAA on the suction surface are turned on simultaneously, the control effect is the best among all three PAA types. The mechanisms of nanosecond discharge and microsecond discharge PAA are different in corner separation control. The control effect of microsecond discharge PAA turns out better with the increase of discharge voltage and duty cycle. Compared with microsec- ond discharge PAA, nanosecond discharge PAA is more effective in preventing corner separation when the freestream velocity increases. Frequency is one of the most important parameters in plasma flow control. The optimum excitation frequency of microsecond discharge PAA is 500 Hz, which is different from the frequency corresponding to the case with a Strouhal number of unity.

Aerodynamic performance of high-turning curved compressor cascade with boundary layer suction

The impact of boundary layer suction on the aerodynamic performance of a high-turning compressor cascade was numerically simulated and discussed.The aerodynamic performance of a curved and a straight cascade with and without boundary layer suction were comparatively studied at several suction flow rates.The results showed that boundary layer suction dramatically improved the flow behavior within the flow passage.Moreover,higher loading over the whole blade height,lower total pressure loss,and higher passage throughflow were achieved with a relatively small amount of boundary layer removal.The integration of curved blade and boundary layer suction contributed to better aerodynamic performance than the cascades with only curved blade or boundary layer suction used,and the more favorable effect resulted from the weakening of the three dimensional effects of the boundary layer close to the endwalls.

Experimental and Numerical Investigations of Shock-Wave Boundary Layer Interactions in a Highly Loaded Transonic Compressor Cascade

Experimental and numerical investigations were conducted to investigate the variations of shock-wave boundary layer interaction(SBLI) phenomena in a highly loaded transonic compressor cascade with Mach numbers.The schlieren technique was used to observe the shock structure in the cascade and the pressure tap method to measure the pressure distribution on the blade surface.The unsteady pressure distribution on blade surface was measured with the fast-response pressure-sensitive paint(PSP) technique to obtain the unsteady pressure distribution on the whole blade surface and to capture the shock oscillation characteristics caused by SBLI.In addition,the Reynolds Averaged Navier Stokes simulations were used to compute the three-dimensional steady flow field in the transonic cascade.It was found that the shock wave patterns and behaviors are affected evidently with the increase in incoming Mach number at the design flow angle,especially with the presence of the separation bubble caused by SBLI.The time-averaged pressure distribution on the blade surface measured by PSP technique showed a symmetric pressure filed at Mach numbers of 0.85,while the pressure field on the blade surface was an asymmetric one at Mach numbers of 0.90 and 0.95.The oscillation of the shock wave was closely with the flow separation bubble on the blade surface and could transverse over nearly one interval of the pressure taps.The oscillation of the shock wave may smear the pressure jump phenomenon measured by the pressure taps.

In uence of Endwall Boundary Layer Suction on the Flow Fields of a Critically Loaded Di usion Cascade

Boundary layer suction is an e ective method used to delay separations in axial compressors. Most studies on bound?ary layer suction have focused on improving the performance of compressors,whereas few studies investigated the influence on details of the flow fields,especially vortexes in compressors. CFD method is validated with experi?mental data firstly. Three single?slot and one double?slot endwall boundary layer suction schemes are designed and investigated. In addition to the investigation of aerodynamic performance of the cascades with and without suction,variations in corner open separation,passage vortex,and concentration shedding vortex,which are rarely seen for the flow controlled blades in published literatures,are analyzed. Then,flow models,which are the ultimate aim,of both baseline and aspirated cascades are established. Results show that single?slot endwall suction scheme adjacent to the suction surface can e ectively remove the corner open separation. With suction mass flow rate of 0.85%,the overall loss coe cient and endwall loss coe cient of the cascade are reduced by 25.2% and 48.6%,respectively. Besides,this scheme increases the static pressure rise coe cient of the cascade by 3.2% and the flow turning angle of up to 3.3° at 90% span. The concentration shedding vortex decreases,whereas the passage vortex increases. For single?slot suction schemes near the middle pitchwise of the passage,the concentration shedding vortex increases and the passage vortex is divided into two smaller passage vortexes,which converge into a single?passage vortex near the trailing edge section of the cascade. For the double?slot suction scheme,triple?passage vortexes are presented in the blade passage. Some new vortex structures are discovered,and the novel flow models of aspirated compressor cascade are proposed,which are important to improve the design of multi?stage aspirated compressors.

亚声速可压缩流场叶片边界层热线测速方法研究

为了能够开发一种简单有效的亚声速可压缩流场叶片边界层速度测量方法,从而为叶型设计和相关数值研究工作提供支撑,本文围绕热线风速测量技术,针对可压缩流场密度变化对热线标定结果的影响,以及实际测量中速度、密度耦合而无法直接获取的问题,通过理论分析,提出了适用于边界层测量的恒定压力热线标定方法和引入叶表稳态静压进行速度解耦的方法,并对所提出方法的主要误差进行了分析评估。在此基础上进行了热线标定和边界层速度测量试验验证,明确了恒定压力热线标定数学模型系数随压力的线性变化规律,同时针对温度非线性影响提出了一种基于过热比调整的修正方法,该方法能够将约13℃的温度偏差对热线电压的影响降低到1%以内,进一步简化了恒定压力热线标定流程,结合基于叶表稳态静压的速度解耦方法,为亚声速可压缩流场叶片边界层瞬态速度测量提供了一种简单可行的高频响测速方法。

轮廓度误差对超声速压气机叶栅气动性能的影响

[目的]旨在评估轮廓度误差对压气机气动性能的影响,并为叶片鲁棒性设计提供参考。[方法]建立单峰值轮廓度误差分布数学模型,采用数值模拟方法,研究压力面和吸力面不同轮廓度组合误差对超声速压气机平面叶栅气动性能的影响。[结果]结果表明:吸力面轮廓度误差分布是影响叶栅总压损失的关键因素,随着吸力面轮廓度峰值误差位置向下游移动,总压损失系数逐渐降低;压力面和吸力面误差分布对气流折转角和静压升系数的影响趋势相反。对较低来流马赫数的叶栅,吸力面误差对气流折转角和静压升均起主导作用;对较高来流马赫数的叶栅,压力面误差对气流折转角和静压升影响明显。激波位置和激波强度、激波后扩张通道的流道型线综合决定了叶片表面和叶栅流道内的流动状态,使得近吸力面侧流动损失增大,近压力面侧流动损失减小,其综合效果决定了叶栅损失、气流折转角和静压升的变化。[结论]结果对指导跨声速压气机设计、加工和超差审理均具有重要意义。

基于马蹄涡前缘吸气控制的压气机叶栅流动机理

为探究从源头抑制角区分离的流动控制方法,采用数值模拟方法,以NACA65叶栅为研究对象,利用叶栅前缘端壁吸气技术控制马蹄涡,结合拓扑分析对叶栅通道内的三维流动结构进行精确重构,以研究前缘端壁吸气控制马蹄涡进而改善叶栅通道流场结构的机理。结果表明:叶栅前缘端壁吸气技术可以有效推迟马蹄涡形成并削弱其强度,同时在吸气狭缝末端形成一对反旋流向涡对,在向下游发展过程中与通道涡相互作用;前缘端壁吸气通过控制马蹄涡,降低端壁边界层厚度,叶栅前缘通道涡发展受限,由回流组成的叶表分离被抑制;前缘及压力面侧吸气(EPS)直接作用于马蹄涡压力面分支,通道涡强度进一步被削弱,角区分离模式由闭式分离转为不完全闭式分离。最后,对比最优吸气系数下不同方案出口总压损失,发现当吸气量为进口质量流量的0.2%时,EPS方案出口截面总压损失降低5.8%;并且通过调整吸气系数,可以获得较好的变工况控制性能。

可控扩散叶型吸力面峰值等熵马赫数位置对叶栅气动性能影响

通常亚音压气机叶型表面等熵马赫数分布符合可控扩散规律,并且吸力面峰值马赫数位置靠前叶栅气动性能较好。采用自动优化方法,设计出给定吸力面峰值等熵马赫数位置可控扩散叶型,分析此位置对叶栅气动性能的影响规律。研究结果表明:对于可控扩散转子和静子叶型,在设计工况下,当吸力面峰值等熵马赫位置位于0.20倍轴向弦长时,吸力面附面层沿流程快速发展,造成叶栅损失大幅增加;当吸力面峰值等熵马赫数位置为0.10~0.15倍轴向弦长时,设计进气角近似位于叶栅低损失进气角范围中,且低损失范围内损失较低。

载荷分布规律对开槽压气机叶型气动性能的影响

航空压气机叶片通道内流动呈强逆压梯度,为了减小流动损失、扩大稳定工作范围,针对压气机静子叶型提出一种新型开槽叶片,槽道由叶片前缘进气吸力面出气,使用来流速度冲量有效抑制吸力面附面层的发展。采用计算机数值模拟方法,研究不同吸力面峰值等熵马赫数位置的可控扩散叶型开槽对叶栅气动性能的影响。研究结果表明:在设计工况下,开槽可有效抑制吸力面附面层发展,降低叶栅损失,增加气流转角;吸力面峰值等熵马赫数位置越向尾缘,在整个进气角范围内,开槽降低损失程度越大,并且由于攻角越大吸力面附面层越厚,开槽降低损失程度越大。

压力敏感涂料技术在涡轮叶片和平面叶栅中的应用进展

由于传统压力测量技术难以实现涡轮叶片表面压力分布的连续监测,因此,开发新型测量技术变得尤为迫切。压力敏感涂料作为一种非接触式的光学表面压力测量技术,其以低成本、高空间分辨率、全域测量能力以及对气动流场影响小的显著优点,已经成为压力测量领域的一个研究热点。本文从实验装置设计、测量结果分析等多个维度,详细介绍了压力敏感涂料技术在涡轮叶片和平面叶栅压力测量中的应用现状,深入讨论了实验设备参数选择、测量布局优化等关键问题。同时,结合国内外的研究与应用案例,探讨了压力敏感涂料技术在叶片表面压力测量领域所遇到的挑战和发展前景。

小型涡轮平面叶栅试验段设计

Performance improvement of the high-load transonic turbine is the key method of improving the thrustto-weight ratio or the power density of gas turbine engines. In order to investigate the flow behaviors inside the high load turbine cascades, a linear turbine cascade test section is designed, which enables the Schlieren photography and static pressure measurement along the cascade profile can be conducted. Variable pitch is realized in the test section to achieve different Zweifel coefficients. Due to the capability limitation of the air supplier, the test section is designed to have only5 blade channels with shortened blade height to achieve high Mach number flow conditions. Numerical investigations were carried out to investigate the wall effect and its in fluences on the flow fields inside the test section. The result indicates that the shape of the connecting part of the test section has a significant influence on the flow similarity among different blade passages. With the proper design, a good repetition flow is achieved between neighbored blade passages.

开式风洞超声速压气机流场起动的数值研究

在开式风洞超声速平面叶栅试验中,从试验启动到叶栅建立超声速流动状态的过程,即超声速流场起动问题,已成为公认的难题。为建立可行的开式风洞超声速流场起动方法,奠定开式超声速风洞的使用基础,基于某超声速风洞,以超声速压气机平面叶栅为研究对象,开展三维数值仿真研究;分析试验条件下超声速流场起动失败的原因,制定三种流场起动方案。结果表明:起动失败的原因为叶栅前缘形成了一道强正激波;仅提高风洞进口总压无法建立叶栅超声速流动状态;仅增大下壁溢流缝宽度可起动超声速叶栅流场,但有效叶栅流道数量减少,壁面附面层增厚;保持上、下壁溢流缝宽度在1倍栅距以上,在栅前上、下壁设置超声速墙并进行抽吸,可有效起动超声速流场,相邻流道出口马赫数最大波动0.01,出口气流角最大波动0.09°,周期性满足试验需求。

Trailing-edge shock loss control with self-sustaining synthetic jet in a supersonic compressor cascade

To effectively reduce the loss of strong shock wave at the trailing edge of the supersonic cascade under high backpressure,a shock wave control method based on self-sustaining synthetic jet was proposed.The self-sustaining synthetic jet was applied on the pressure side of the blade with the blow slot and the bleed slot arranged upstream and downstream of the trailing-edge shock,respectively.The flow control mechanism and effects of parameters were investigated by numerical simulation.The results show that the self-sustaining synthetic jet forms an oblique shock wave in the cascade passage which slows down and pressurizes the airflow,and the expansion wave downstream of the blow slot weakens the shock strength which can effectively change the Mach reflection to regular reflection and thus weaken the shock loss.And the suction effect can reduce loss near blade surface.Compared with the baseline cascade,the self-sustaining jet actuator can reduce flow losses by 6.73%with proper location design and vibration of diaphragm.

Experimental Investigation of the Transonic Shock Oscillation Characteristics in a Heavy-Duty Gas Turbine Compressor Cascade

This paper presents an experimental study of the self-sustained transonic shock oscillating behaviors in a heavy-duty gas turbine compressor cascade under the inlet Mach number of 0.85,0.90 and 0.95.The transonic shock patterns and the surface flow structures are captured by schlieren imaging and oil flow visualization.The time-averaged and instantaneous transonic shock oscillating behaviors at the near choke point and the near stall point are investigated by the Anodized Aluminum Pressure-Sensitive Paint(AA-PSP)surface pressure measurement.The normal passage shock dominant pattern and the detached bow shock dominant pattern at the near choke point and the near stall point are experimental characterized,respectively.The passage shock oscillation behaviors at the near choke point have been observed to undergo periodic pressure perturbations of the shock shift between the upstreamλshock feet mode and the downstreamλshock feet mode.The detached bow shock oscillation behaviors at the near stall point have been observed to undergo the pressure perturbations of the shock cycle movement between the upstream detached bow shock mode and the downstream detached bow shock mode.The differences between the shock shift mode and the shock cycle movement mode lead to the different streamwise oscillation travel ranges and different shock intensity variations under the same inlet Mach number.

进口附面层对串列叶栅气动性能的影响研究

为探究进口附面层对串列叶栅气动性能的影响,以高亚声速压气机常规叶栅及其改型串列叶栅为研究对象,基于数值方法对比分析了不同进口附面层厚度对两叶栅的总体性能及三维角区分离的影响。结果表明:随着进口附面层厚度增加,原型叶栅三维角区分离范围沿展向逐渐增大,而串列叶栅三维角区分离范围的增加主要体现在周向。由于进口附面层的存在,串列叶栅前后叶片之间的缝隙射流与近端壁低能流体的相互作用所产生的通道涡、诱导涡及角涡是总压损失增大的主要原因。串列叶栅与原型叶栅相比能有效降低总压损失和提高静压升,然而附面层厚度的增加会减小这一优势;进口附面层厚度相对叶高的比值为0%,5%,12.5%时,串列叶栅的出口总压损失系数较原型叶栅分别降低11.1%,5.5%和4.1%,静压升系数分别提高7.4%,6.5%和6.4%。

叶栅吸力面吸气位置与角区分离的关联性分析

高压压气机后面级的叶片吸力面三维角区内易堆积大量的低能流体。为了指导吸气槽设计,达到量化分析叶栅吸力面吸气槽位置与角区分离的关联性的目的,以一高负荷压气机叶栅为研究对象,采用计算流体力学仿真软件讨论了全叶高吸气槽轴向位置变化对叶栅性能的影响,并以角区分离起始位置到叶栅前缘之间的距离(ZCS)作为量化标准。研究表明:在设计工况附近,总压损失随着吸气槽轴向位置从前向后移动表现出了先减小后增大的趋势,且叶片吸力面吸气槽在不同工况下存在最佳轴向位置。在0°迎角时,当吸气槽距离分离点间的轴向距离为0.33倍的ZCS时为最佳,可使总压损失系数降低10.9%。这种量化结果借助角区分离结构实现了吸气槽轴向位置设计的标准化指导。

轴流压气机叶型气动优化设计

以某轴流压气机动叶根部10%叶高典型截面叶型为研究对象,采用CFD软件计算分析不同气流角下叶型流场损失及气动特性。结果表明:随着入口气流角增加,出口总压损失先减小后增大,接近设计角度时总压损失最小。同时,运用基于自主开发的参数化叶型优化设计平台,采用三维N-S方程流场计算、网格自动生成、人工神经网络与遗传算法寻优相结合的方法对其进行了叶型型线优化设计,通过流场分析,优化效果显著。