發(fā)布時(shí)間：2018-08-28 07:18

【摘要】：端壁區(qū)內(nèi)復(fù)雜的粘性流動(dòng)嚴(yán)重降低多級軸流壓氣機(jī)的氣動(dòng)性能,現(xiàn)有研究表明,包括葉尖間隙流以及二次流在內(nèi)的端壁損失通常占壓氣機(jī)內(nèi)流動(dòng)總損失的50%~70%,因此深入認(rèn)識(shí)多級軸流壓氣機(jī)內(nèi)特別是端壁區(qū)的流場結(jié)構(gòu)及其主要物理機(jī)制對提升高壓壓氣機(jī)乃至航空發(fā)動(dòng)機(jī)的氣動(dòng)性能有著重要的現(xiàn)實(shí)意義。鑒于高壓壓氣機(jī)后面級尺寸小、轉(zhuǎn)速高,開展詳細(xì)流場測量技術(shù)難度大、成本高、風(fēng)險(xiǎn)大且周期長,低速模擬實(shí)驗(yàn)得到了一定的應(yīng)用,但迄今為止國內(nèi)在該領(lǐng)域開展的實(shí)驗(yàn)研究仍相當(dāng)匱乏,極大地限制了我國航空發(fā)動(dòng)機(jī)高壓壓氣機(jī)設(shè)計(jì)能力的提升。本文以豐富全面的實(shí)驗(yàn)測量為主,定常/非定常CFD計(jì)算為輔,針對兩臺(tái)用于某驗(yàn)證機(jī)高壓壓氣機(jī)低速模擬實(shí)驗(yàn)的四級壓氣機(jī)以及某葉尖臨界型轉(zhuǎn)子A和非葉尖臨界型轉(zhuǎn)子B,開展了壓氣機(jī)氣動(dòng)性能、內(nèi)部流動(dòng)結(jié)構(gòu)的研究,獲得了壓氣機(jī)內(nèi)部的主要流動(dòng)結(jié)構(gòu),包括:類重復(fù)級的流動(dòng)特征,轉(zhuǎn)子葉頂泄漏流及其與主流的交界面隨流量的減小而逐漸與前緣齊平的流動(dòng)規(guī)律,靜子通道內(nèi)葉片附面層沿程發(fā)展及在兩端吸力面角區(qū)形成的一對嚴(yán)重影響其性能的角渦,具有較大弓形的靜子隨流量的減小在吸力面葉中位置首先出現(xiàn)回流等,揭示了多級軸流壓氣機(jī)內(nèi)復(fù)雜定常/非定常流動(dòng)相互作用的物理機(jī)制及高流動(dòng)損失的來源,并基于對原型壓氣機(jī)內(nèi)部復(fù)雜流動(dòng)結(jié)構(gòu)的認(rèn)識(shí),采用先進(jìn)葉片設(shè)計(jì)技術(shù)改善了原型壓氣機(jī)輪轂區(qū)的流動(dòng),提高了壓氣機(jī)的氣動(dòng)性能,為我國當(dāng)代高性能航空發(fā)動(dòng)機(jī)高壓壓氣機(jī)的設(shè)計(jì)提供了必要的支持。本文主要包括以下三部分研究工作:第1部分基于“模擬準(zhǔn)則”參數(shù),開展了低速原型4級類重復(fù)級軸流壓氣機(jī)氣動(dòng)設(shè)計(jì)及其內(nèi)部詳細(xì)流動(dòng)結(jié)構(gòu)的實(shí)驗(yàn)研究。氣動(dòng)設(shè)計(jì)包括模擬目標(biāo)的建立、低速模擬級總體參數(shù)確定、S2氣動(dòng)設(shè)計(jì)、重復(fù)級及類重復(fù)級葉型設(shè)計(jì),將類重復(fù)級設(shè)計(jì)以及進(jìn)口堵塞納入了低速模擬壓氣機(jī)的設(shè)計(jì)中,完善了低速模擬設(shè)計(jì)體系,有效提高了低速壓氣機(jī)設(shè)計(jì)的可靠性和低速模擬的精度。分析了實(shí)驗(yàn)中各主要?dú)鈩?dòng)參數(shù)的不確定度及影響不確定度的主要因素,流量系數(shù)和總總壓升系數(shù)的不確定度分別為0.42%和0.4%。通過通道間、轉(zhuǎn)子和靜子葉片通道內(nèi)、靜子表面靜壓分布及轉(zhuǎn)子葉頂區(qū)流場測量完整地呈現(xiàn)了整個(gè)模擬級內(nèi)部的詳細(xì)流場結(jié)構(gòu),包括類重復(fù)級流動(dòng)特征,葉頂泄漏流及其與主流交界面隨流動(dòng)工況變化的規(guī)律,靜子葉片通道內(nèi)形成的端壁角區(qū)對渦等。研究表明,原型壓氣機(jī)第3級轉(zhuǎn)子輪轂區(qū)葉片負(fù)荷偏高且存在著一定的流動(dòng)分離,同時(shí)靜子輪轂也出現(xiàn)了較大的流動(dòng)堵塞、分離及更高的總壓損失。第2部分基于對原型壓氣機(jī)內(nèi)部復(fù)雜流動(dòng)結(jié)構(gòu)的認(rèn)識(shí),開展了原型壓氣機(jī)第3級葉片改進(jìn)設(shè)計(jì)及改進(jìn)設(shè)計(jì)壓氣機(jī)性能及詳細(xì)內(nèi)部流動(dòng)結(jié)構(gòu)的實(shí)驗(yàn)研究。針對模擬級轉(zhuǎn)、靜子開展了積疊規(guī)律的參數(shù)化研究,明確了端部彎角的選擇以減小輪轂區(qū)的流動(dòng)損失,改進(jìn)方案包含如下要素:轉(zhuǎn)子“J”型積疊,增大的靜子幾何進(jìn)口角,靜子“前加載”技術(shù)以及更大的弓形積疊。實(shí)驗(yàn)結(jié)果表明,四級壓氣機(jī)效率提升了約1個(gè)百分點(diǎn),總總壓升系數(shù)提高了約1.4%,第3級總總壓升系數(shù)提高了約10%,失穩(wěn)流量和原型基本一致。改進(jìn)設(shè)計(jì)實(shí)驗(yàn)的流場結(jié)構(gòu)反映出了新的三維葉片設(shè)計(jì)特征,結(jié)合CFD計(jì)算結(jié)果研究了新的葉片造型改善壓氣機(jī)內(nèi)流動(dòng)結(jié)構(gòu)及氣動(dòng)性能的主要物理機(jī)制。對級間的非定常測量結(jié)果進(jìn)行頻譜分析、系綜平均等處理后發(fā)現(xiàn),上游轉(zhuǎn)子尾跡脫落引起靜子尾跡區(qū)兩側(cè)葉片通過頻率(fBPF)具有更高的能量及總壓均方根值,靜子附面層與更前面級靜子附面層的相互作用使高階諧波能量增強(qiáng),且這種現(xiàn)象在靠近葉尖及壓氣機(jī)出口時(shí)更明顯,在第3級出口葉尖區(qū)域2fBPF的能量幅值達(dá)到fBPF的8倍,該現(xiàn)象對于壓氣機(jī)內(nèi)氣動(dòng)噪音和振動(dòng)的控制均具有重要的意義。第3部分針對某葉尖失速型轉(zhuǎn)子A和某非葉尖失速型轉(zhuǎn)子B葉頂區(qū)域開展了詳細(xì)的數(shù)值研究。研究揭示了葉尖泄漏渦的主要流動(dòng)結(jié)構(gòu),對于非葉尖臨界型轉(zhuǎn)子B,在間隙高度62.5%以下的葉尖區(qū)域內(nèi),從葉尖間隙前緣流出的流體會(huì)卷吸成葉尖泄漏渦,且隨著間隙高度的增加,其占據(jù)的葉尖泄漏渦的位置由內(nèi)而外,而62.5%間隙高度以上的流動(dòng)則并不卷吸入葉尖泄漏渦內(nèi),而主要表現(xiàn)為二次泄漏。在近失速工況下,葉尖臨界型轉(zhuǎn)子A葉頂會(huì)出現(xiàn)自激非定常性,這種現(xiàn)象的出現(xiàn)與否強(qiáng)烈依賴于葉尖間隙尺寸,而與上游附面層厚度關(guān)系不大;在葉頂區(qū)形成兩個(gè)堵塞區(qū),其中葉尖二次渦和破碎的葉尖泄漏渦的相互作用形成的大堵塞區(qū)會(huì)產(chǎn)生大的流動(dòng)損失,將直接決定轉(zhuǎn)子的氣動(dòng)穩(wěn)定性。通過有效的流動(dòng)控制方法消除或至少減弱該二次渦的強(qiáng)度將有助于拓寬該轉(zhuǎn)子的穩(wěn)定工作范圍。
[Abstract]:Complex viscous flow in the end-wall region seriously degrades the aerodynamic performance of a multistage axial compressor. Existing studies have shown that the loss of the end-wall, including tip clearance flow and secondary flow, usually accounts for 50%-70% of the total loss of flow in the compressor. Therefore, a thorough understanding of the flow field structure in the multistage axial compressor, especially in the end-wall region, as well as its main physical properties is given. The mechanism has important practical significance for improving the aerodynamic performance of high-pressure compressors and even aero-engines. In view of the small size and high speed of the rear stage of high-pressure compressors, it is difficult to carry out detailed flow field measurement technology, high cost, high risk and long cycle, low-speed simulation experiments have been applied to a certain extent, but so far domestic research in this field has been carried out. Experimental studies are still scarce, which greatly limits the improvement of design capability of high-pressure compressors for aero-engines in China. In this paper, based on abundant and comprehensive experimental measurements, steady/unsteady CFD calculations are supplemented by two four-stage compressors used in low-speed simulation experiments of high-pressure compressors for a certain validator and a blade-tip critical rotor A and A. A non-tip critical rotor B was developed to study the aerodynamic performance and internal flow structure of a compressor. The main flow structures in the compressor were obtained, including the flow characteristics of the quasi-repetitive stage, the tip leakage flow and its interface with the mainstream gradually flattened with the leading edge as the flow rate decreased, and the blade boundary in the stator passage. A pair of angular vortices formed along the layer and in the corner region of the suction surface at both ends seriously affect the performance of the compressor. The stator with a larger bow first appears to be recirculated in the suction surface blade with the decrease of flow rate. The physical mechanism of complex steady/unsteady flow interaction and the source of high flow loss in a multistage axial compressor are revealed. Based on the understanding of the complex flow structure inside the prototype compressor, advanced blade design technology is adopted to improve the flow in the hub region of the prototype compressor and improve the aerodynamic performance of the compressor, which provides the necessary support for the design of the high-pressure compressor of the contemporary high-performance aeroengine in China. Based on the "simulation criteria" parameters, the aerodynamic design and detailed flow structure of a low-speed prototype 4-stage repetitive axial compressor were studied experimentally. The aerodynamic design includes the establishment of simulation objectives, the determination of the overall parameters of the low-speed simulation stage, the S2 aerodynamic design, the repetitive stage and the quasi-repetitive stage blade design, and the quasi-repetitive stage design. The inlet blockage was incorporated into the design of the low-speed simulated compressor, and the low-speed simulated design system was improved. The reliability and accuracy of the design of the low-speed compressor were effectively improved. The uncertainty of the main aerodynamic parameters and the main factors affecting the uncertainty were analyzed. The static pressure distribution on the stator surface and the measurement of the rotor tip flow field in the rotor and stator blade passages show the detailed flow field structure in the whole simulation stage, including the characteristics of the quasi-repetitive stage flow, the variation of the tip leakage flow and its interface with the main flow conditions. The results show that the blade load in the hub region of the third stage rotor of a prototype compressor is high and there is a certain flow separation. At the same time, the stator hub also has a large flow blockage, separation and higher total pressure loss. Part 2 is based on the recognition of the complex flow structure in the prototype compressor. In order to reduce the flow loss in the hub region, a parameterized study of stacking rule was carried out on the stator for the simulated stage rotation. The selection of the end bending angle was made clear to reduce the flow loss in the hub region. The improvement scheme includes the following elements: rotor "J" The experimental results show that the efficiency of the four-stage compressor is increased by about one percentage point, the total pressure rise coefficient is increased by about 1.4%, the total pressure rise coefficient of the third stage is increased by about 10%, and the unstable flow rate is basically the same as that of the prototype. The structure reflects the new three-dimensional blade design characteristics, and the main physical mechanism of improving the flow structure and aerodynamic performance of the compressor by using the new blade shape is studied with CFD calculation results. The blade passage frequency (fBPF) has higher energy and root mean square of total pressure. The interaction between stator boundary layer and stator boundary layer in the preceding stage enhances the high-order harmonic energy. This phenomenon is more obvious near the blade tip and compressor outlet. The energy amplitude of 2fBPF in the blade tip region of the third stage is eight times that of fBPF. Aerodynamic noise and vibration control in a compressor are of great importance. Part 3 is a detailed numerical study of the tip region of a blade-tip stall rotor A and a non-tip stall rotor B. The study reveals the main flow structure of tip leakage vortices. For non-tip critical rotor B, the blade with a clearance height of 62.5% or less has been investigated. In the tip region, the fluid flowing from the leading edge of the tip clearance will entrap into the tip leakage vortex. With the increase of the tip clearance height, the tip leakage vortex occupies the position from inside to outside, while the flow above 62.5% clearance height does not entrap into the tip leakage vortex, but mainly presents secondary leakage. Self-excited unsteadiness occurs at the tip of rotor A blade, which is strongly dependent on tip clearance size and has little to do with the thickness of upper boundary layer. Two blockage zones are formed at the tip of rotor A blade, in which the large blockage zone formed by the interaction of tip secondary vortex and broken tip leakage vortex will produce a large flow loss, which will be directly related to the thickness of upper boundary layer. The aerodynamic stability of the rotor is determined. Eliminating or at least weakening the strength of the secondary vortex by effective flow control method will help to widen the stable operating range of the rotor.
【學(xué)位授予單位】：南京航空航天大學(xué)
【學(xué)位級別】：博士
【學(xué)位授予年份】：2015
【分類號(hào)】：V233

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