本文關(guān)鍵詞：超音壓氣機(jī)轉(zhuǎn)子的起動(dòng)特性及內(nèi)部流動(dòng)組織研究　出處：《中國(guó)科學(xué)院工程熱物理研究所》2017年博士論文　論文類(lèi)型：學(xué)位論文

  更多相關(guān)文章： 超音壓氣機(jī) 超音平面葉柵 流動(dòng)機(jī)理 轉(zhuǎn)子 葉型設(shè)計(jì)

【摘要】：隨著級(jí)負(fù)荷的不斷增加,超音轉(zhuǎn)子進(jìn)口相對(duì)馬赫數(shù)不斷提高,當(dāng)進(jìn)口相對(duì)馬赫數(shù)超過(guò)1.75左右時(shí),轉(zhuǎn)子的起動(dòng)問(wèn)題和高速來(lái)流的高效減速增壓?jiǎn)栴}成為了限制轉(zhuǎn)子性能的關(guān)鍵。本文從超音平面葉柵的葉型設(shè)計(jì)出發(fā),對(duì)進(jìn)口相對(duì)馬赫數(shù)較高的內(nèi)激波式轉(zhuǎn)子的起動(dòng)特性及內(nèi)部流動(dòng)組織問(wèn)題進(jìn)行了研究,涉及的主要內(nèi)容如下:1、圍繞超音平面葉柵的唯一攻角特性、起動(dòng)點(diǎn)的計(jì)算、流道內(nèi)波系結(jié)構(gòu)的優(yōu)化組織等問(wèn)題,對(duì)超音平面葉柵的流動(dòng)機(jī)理進(jìn)行了研究。首先,基于對(duì)超音平面葉柵極限特征線上游流場(chǎng)的分析,提出了超音平面葉柵的外伸波損失計(jì)算模型。其次,提出了一種簡(jiǎn)單的方法來(lái)模擬超音平面葉柵處于起動(dòng)點(diǎn)時(shí)喉口前的主要流場(chǎng)結(jié)構(gòu),利用該方法研究了不同類(lèi)型的葉柵處于起動(dòng)點(diǎn)時(shí)喉口前的總壓恢復(fù)系數(shù)隨來(lái)流馬赫數(shù)的變化,將相關(guān)結(jié)果與本文提出的理想化的起動(dòng)過(guò)程相結(jié)合,建立了葉柵起動(dòng)點(diǎn)的計(jì)算方法。再次,針對(duì)起動(dòng)工況下的前緣內(nèi)伸波,建立了模型來(lái)計(jì)算其形狀。最后,以激波損失最小為原則,研究了超音平面葉柵流道內(nèi)的波系優(yōu)化組織問(wèn)題,明確了僅在喉口前的斜激波系的損失為零的理想情況下,斜激波系和結(jié)尾激波組成的波系的損失取得最小值。2、依據(jù)超音平面葉柵流動(dòng)機(jī)理的研究成果,提出了超音平面葉柵葉型的參數(shù)化設(shè)計(jì)方法。利用該方法設(shè)計(jì)了相關(guān)葉型,與實(shí)驗(yàn)數(shù)據(jù)或數(shù)值模擬結(jié)果的對(duì)比表明,采用本文提出的設(shè)計(jì)方法能夠?qū)Τ羝矫嫒~柵的起動(dòng)點(diǎn)、唯一攻角特性以及流道內(nèi)的波系結(jié)構(gòu)進(jìn)行有效的控制。在此基礎(chǔ)上,研究了該設(shè)計(jì)方法中關(guān)鍵控制參數(shù)對(duì)超音平面葉柵氣動(dòng)性能的影響。3、針對(duì)超音轉(zhuǎn)子,本文以通流面積的軸向變化率不變?yōu)樵瓌t,提出了超音平面葉柵葉型向轉(zhuǎn)子回轉(zhuǎn)面葉型轉(zhuǎn)換的關(guān)系式,完善了超音轉(zhuǎn)子的氣動(dòng)設(shè)計(jì)方法。以采用這種設(shè)計(jì)方法設(shè)計(jì)的超音轉(zhuǎn)子為研究對(duì)象,通過(guò)數(shù)值模擬,對(duì)超音轉(zhuǎn)子的起動(dòng)特性進(jìn)行了研究,分析了各葉高葉型的起動(dòng)點(diǎn)和對(duì)應(yīng)的平面葉柵的起動(dòng)點(diǎn)不同的原因。同時(shí),轉(zhuǎn)子氣動(dòng)性能的研究結(jié)果表明,轉(zhuǎn)子流場(chǎng)中各激波的強(qiáng)度得到了有效控制,設(shè)計(jì)點(diǎn)的性能和設(shè)計(jì)目標(biāo)符合較好。4、開(kāi)展了超音轉(zhuǎn)子內(nèi)部流動(dòng)特性的研究,分析了葉表邊界層內(nèi)氣流的徑向輸運(yùn)、葉頂泄漏流、分離流對(duì)流場(chǎng)的影響。結(jié)果表明,在結(jié)尾激波的作用下,下游葉表邊界層內(nèi)氣流的徑向輸運(yùn)強(qiáng)度大幅增加,這雖然使葉根附近結(jié)尾激波與邊界層相互作用引起的流動(dòng)分離區(qū)顯著減小,但導(dǎo)致了機(jī)匣附近的邊界層增厚。厚度明顯增加的邊界層在和結(jié)尾激波作用后,發(fā)生了明顯的分離。隨后,在與分離渦和泄漏渦的共同作用下,靠近壓力面的葉尖角區(qū)的高熵區(qū)影響范圍不斷擴(kuò)大。
[Abstract]:With the increasing of the stage load, the relative Mach number of the inlet of the supersonic rotor increases continuously, when the relative Mach number of the inlet exceeds about 1.75. The starting problem of rotor and the problem of high efficiency deceleration and supercharging of high speed flow have become the key to limit the performance of rotor. In this paper, the blade shape design of supersound plane cascade is introduced. The starting characteristics and internal flow structure of the inner shock rotor with relatively high Mach number are studied. The main contents are as follows: 1, the unique angle of attack around the supersound plane cascade. The flow mechanism of the supersound plane cascade is studied by calculating the starting point and optimizing the structure of the internal wave system of the channel. Firstly, the upstream flow field of the limit characteristic line of the supersound plane cascade is analyzed. A model for calculating the outreaching wave loss of a supersound planar cascade is proposed. Secondly, a simple method is proposed to simulate the main flow field structure in front of the throat when the supersound planar cascade is at the starting point. By using this method, the change of total pressure recovery coefficient in front of the throat of different types of cascades at the starting point is studied, and the correlation results are combined with the idealized starting process proposed in this paper. The method of calculating the starting point of cascade is established. Thirdly, the model is established to calculate the shape of the leading edge wave in the starting condition. Finally, the principle of minimum shock loss is taken as the principle. In this paper, the optimal structure of the wave system in the channel of the supersound plane cascade is studied, and it is clear that the ideal condition is that the loss of the oblique shock system in front of the throat is zero. The loss of the oblique shock system and the wave system composed of the end shock wave is minimum. 2, according to the research results of the supersound plane cascade flow mechanism. A parameterized design method for supersound planar cascade blade is presented. The relative blade profile is designed by using this method, and the results are compared with experimental data or numerical simulation results. The design method presented in this paper can effectively control the starting point, the unique angle of attack and the structure of the wave system in the channel. The influence of key control parameters on the aerodynamic performance of supersound planar cascade is studied. The axial change rate of flow area is not changed to the principle of supersound rotor. The relationship between the supersound plane cascade blade and the rotor rotary blade shape is proposed, and the aerodynamic design method of the supersound rotor is improved. The supersound rotor designed by this design method is taken as the research object, and the numerical simulation is carried out. The starting characteristics of supersound rotor are studied, and the causes of different starting points of each blade type and corresponding plane cascade are analyzed. The results show that the aerodynamic performance of the rotor is different. The intensity of shock waves in rotor flow field is effectively controlled, and the performance and design target of design point are in good agreement with .4. the study on the internal flow characteristics of supersound rotor is carried out. The influence of radial air flow, tip leakage flow and separation flow on the flow field in the blade surface boundary layer is analyzed. The results show that the radial transport intensity of the downstream blade surface boundary layer increases greatly under the action of the end shock wave. Although the flow separation zone caused by the interaction between the end shock wave and the boundary layer near the leaf root decreases significantly, it leads to the thickening of the boundary layer near the casing and the obvious increase of the thickness of the boundary layer after the action of the end shock wave. Then the influence range of the high entropy region near the pressure surface in the tip region of the leaf is continuously expanded under the joint action of the separation vortex and the leakage vortex.
【學(xué)位授予單位】：中國(guó)科學(xué)院工程熱物理研究所
【學(xué)位級(jí)別】：博士
【學(xué)位授予年份】：2017
【分類(lèi)號(hào)】：V231

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