脈沖設(shè)備二維高超進(jìn)氣道壓縮面激波—邊界層干擾顯示技術(shù)及應(yīng)用研究
發(fā)布時間：2018-08-05 09:23
【摘要】：在吸氣式高超聲速飛行器研究中,激波-邊界層干擾問題是一個研究多年但遠(yuǎn)未解決的難題。其中,高超進(jìn)氣道外壓縮面上的激波-邊界層干擾通過影響波系結(jié)構(gòu)、’邊界層狀態(tài)而對進(jìn)氣道性能產(chǎn)生重要影響。然而,在眾多激波-邊界層干擾研究中,關(guān)注高超進(jìn)氣道外壓縮面上情況較少,特別是多級壓縮拐角和前緣半徑對激波-邊界層干擾特性及進(jìn)氣道性能的影響規(guī)律,尚缺乏系統(tǒng)的研究。在激波邊界層干擾(包括進(jìn)氣道外壓縮面)研究中,既要研究波系分布,也要關(guān)注邊界層行為.。脈沖型高超聲速試驗(yàn)設(shè)備因能夠模擬高超聲速總溫和總壓試驗(yàn)條件,是常規(guī)風(fēng)洞的重要補(bǔ)充,’兩者結(jié)合,可覆蓋更全面的馬赫數(shù)、雷諾數(shù)和壁溫比范圍,也是開展激波邊界層干擾研究的重要試驗(yàn)平臺。由于試驗(yàn)時間極短,脈沖型高超聲速試驗(yàn)設(shè)備上尚缺少成熟的、可顯示邊界層分離情況的油流試驗(yàn)技術(shù)；同時由于高超聲速流場中的高溫和高速測試環(huán)境,常規(guī)的邊界層厚度測試技術(shù)如皮托探針無法使用,測試高超聲速邊界層厚度存在巨大困難。針對上述問題,本文在脈沖型高超聲速風(fēng)洞上,發(fā)展了兩類高超進(jìn)氣道激波邊界層干擾研究測試技術(shù)。首先,在激波風(fēng)洞上發(fā)展了基于聚焦紋影技術(shù)的高超聲速邊界層厚度測量方法,結(jié)合理論和數(shù)值模擬結(jié)果,提供了從聚焦紋影照片中提取密度邊界層厚度數(shù)據(jù)的方法,證明了基于密度梯度的邊界層厚度近似密度邊界層厚度,密度邊界層與溫度邊界層厚度相同,.對于空氣速度邊界層厚度略小于密度邊界層厚度,在應(yīng)用中改進(jìn)了聚焦紋影系統(tǒng)光源和成像系統(tǒng)。其次,在脈沖燃燒風(fēng)洞上發(fā)展了油流顯示技術(shù),結(jié)合脈沖燃燒設(shè)備運(yùn)行特點(diǎn),剖析了脈沖類設(shè)備上發(fā)展油流顯示技術(shù)的難點(diǎn)所在；分析了油膜受力情況和運(yùn)動機(jī)理,給出了油膜運(yùn)動速度的影響因素,提出了瞬時脈沖型風(fēng)洞上膜式油流試驗(yàn)準(zhǔn)則；提供了合適的顯示油粘度范圍‘,發(fā)展了油膜噴涂方法,解決了高溫試驗(yàn)氣體自發(fā)光問題,設(shè)計(jì)了實(shí)時攝像系統(tǒng)；在極短的有效試驗(yàn)時間內(nèi),采集了油流圖譜序列,油膜的動態(tài)發(fā)展過程驗(yàn)證了上述關(guān)系準(zhǔn)則作為邊界層厚度測量方法和油流顯示技術(shù)的應(yīng)用,結(jié)合其它測試和數(shù)值模擬手段,初步研究了多級壓縮拐角和前緣半徑對進(jìn)氣道性能和激波邊界層干擾特性的影響規(guī)律。在脈沖燃燒風(fēng)洞上名義馬赫數(shù)6和單位雷諾數(shù)5.4×106/m條件下,得到的油流圖譜序列表明,兩級壓縮楔拐角處初始分離角度介于100～150之間,三級壓縮楔模型第二拐角處流動未分離。在激波風(fēng)洞上名義馬赫數(shù)6和單位雷諾數(shù)3.4×107/m條件下,研究了三級平面壓縮高超聲速進(jìn)氣道前緣半徑的影響,獲得了壁面靜壓、壁面熱流、激波幾何特征、邊界層密度厚度分布數(shù)據(jù)。壁面靜壓從拐點(diǎn)上游開始爬升,在拐點(diǎn)后一定距離達(dá)到壓強(qiáng)平臺；前緣半徑增加,壓強(qiáng)平臺值減小,達(dá)到壓強(qiáng)平臺需要的距離增加,’意味著拐角處激波-邊界層干擾區(qū)范圍擴(kuò)大,尖前緣情況下激波-邊界層干擾區(qū)范圍是上游邊界層厚度的5倍,3mm鈍前緣情況下達(dá)到8倍。壁面熱流從拐點(diǎn)上游開始逐漸爬升,在拐點(diǎn)后達(dá)到一個峰值,然后下降直到下一個壓縮拐角；隨著前緣半徑的增加,拐角熱流峰值顯著減小,熱流峰值位置向上游移動。邊界層厚度分布在拐角上游達(dá)到上游最大值,過干擾區(qū)后邊界層厚度比入流邊界層厚度顯著減小,在下一個壓縮面上重新增長；隨著前緣半徑的增加,邊界層厚度增加。使用FLUNET軟件選用不同湍流模型得到的壁面靜壓都與試驗(yàn)結(jié)果吻合較好,但壁面熱流分布、邊界層密度厚度分布差異較大。隨著前緣半徑的增加,第一級激波角逐漸增加,第二級和第三級激波角逐漸減小,拐角處第二、三道激波根部彎曲程度和影響區(qū)域增大。
[Abstract]:In the study of air breathing hypersonic vehicles, the shock boundary layer interference problem is a difficult problem for many years, but the shock boundary layer interference on the hyperactive outer compression surface has an important influence on the performance of the intake port by influencing the structure of the wave system and the state of the boundary layer. However, in a number of shock wave boundary layers. In the perturbation study, it is not necessary to study the influence of the high intake air compression surface on the outer compression surface, especially the influence of the multistage compression corner and the radius of the front edge to the shock boundary layer interference and the inlet performance. Boundary layer behavior, pulse type hypersonic test equipment is an important supplement to conventional wind tunnel because of its ability to simulate high hypersonic total pressure test conditions. The combination of the two can cover a more comprehensive Maher number, Reynolds number and the range of wall temperature ratio. It is also an important test platform for the study of shock wave boundary layer interference. An oil flow test technique which can display boundary layer separation conditions is still lacking on the flush hypersonic test equipment. At the same time, due to the high temperature and high speed testing environment in the hypersonic flow field, the conventional boundary layer thickness testing technology, such as the pitot probe, is difficult to test the thickness of the high supersonic boundary layer. In this paper, we develop two kinds of high supersonic wave boundary layer interference in high supersonic wind tunnel. Firstly, the method of measuring the thickness of hypersonic boundary layer based on the focused schlieren technique is developed on the shock wind tunnel, and the density is extracted from the focused schlieren photos with the theoretical and numerical simulation results. The method of boundary layer thickness data shows that the thickness of the boundary layer based on the density gradient is approximately the thickness of the density boundary layer, the density boundary layer is the same as the thickness of the temperature boundary layer. The thickness of the air velocity boundary layer is slightly less than the thickness of the density boundary layer, and the light source and imaging system of the focused schlieren system are improved in application. Secondly, the pulse combustion is used. The oil flow display technology is developed in the wind tunnel. The difficulties in the development of oil flow display technology on pulse equipment are analyzed in combination with the characteristics of the pulse combustion equipment. The stress situation and movement mechanism of the oil film are analyzed. The influence factors of the motion velocity of the oil film are given, and the test criterion of the film type oil flow on the instantaneous pulse blast tunnel is put forward. The oil film spraying method is properly displayed and the oil film spraying method is developed. The self luminescence problem of the high temperature test gas is solved. The real-time camera system is designed. The oil flow map sequence is collected during the very short effective test time. The dynamic development process of the oil film proves that the above relationship criterion is used as the boundary layer thickness measurement method and the oil flow display. The effect of the multistage compression corner and the radius of the front edge on the performance of the intake port and the interference characteristics of the shock wave boundary layer is preliminarily studied with the use of other testing and numerical simulation methods. Under the condition of the nominal Maher number 6 and the unit Reynolds number of 5.4 x 106/m on the pulse combustion wind tunnel, the obtained oil flow map sequence shows that the two stage compression wedge abduction is shown. The initial separation angle at the corner is between 100~150, and the flow of the three stage compression wedge is not separated at second corners. Under the condition of the nominal Maher number 6 and the unit Reynolds number 3.4 x 107/m on the shock wind tunnel, the influence of the three level plane compression on the radius of the front edge of the high supersonic inlet is studied, and the wall static pressure, the wall heat flow, the shock geometry feature and the edge are obtained. The boundary layer density thickness distribution data. The wall static pressure begins to climb up the inflection point and reaches the pressure platform at a certain distance after the inflection point; the radius of the front edge increases and the pressure platform value decreases, which can increase the distance needed by the pressure platform, 'means that the disturbance area of the shock boundary layer is enlarged at the corner and the shock boundary layer in the front edge is the interference area. The range is 5 times the thickness of the upstream boundary layer, and the 3mm blunt front edge reaches 8 times. The wall heat flow gradually rises from the upstream of the inflection point, reaches a peak after the turning point, and then drops to the next compression corner. With the increase of the radius of the front edge, the peak heat flow peak decreases significantly, the peak position of the heat flow moves upstream. Boundary layer thickness. The thickness of the upper boundary layer is significantly lower than the inflow boundary layer thickness, and the thickness of the boundary layer is increased again. With the increase of the radius of the front edge, the thickness of the boundary layer increases with the increase of the radius of the front edge. The wall static pressure obtained by using different turbulence models using FLUNET software is in good agreement with the test results, but the wall wall is in good agreement with the test results. The distribution of surface heat flow and the thickness distribution of the boundary layer density vary greatly. With the increase of the radius of the front edge, the first shock angle increases gradually, the second and third magnitude shock angles gradually decrease, and the bending degree and the affected area of the second, third shock waves at the corner are increased.
【學(xué)位授予單位】：中國空氣動力研究與發(fā)展中心
【學(xué)位級別】：博士
【學(xué)位授予年份】：2015
【分類號】：V211.74
本文編號：2165354
[Abstract]:In the study of air breathing hypersonic vehicles, the shock boundary layer interference problem is a difficult problem for many years, but the shock boundary layer interference on the hyperactive outer compression surface has an important influence on the performance of the intake port by influencing the structure of the wave system and the state of the boundary layer. However, in a number of shock wave boundary layers. In the perturbation study, it is not necessary to study the influence of the high intake air compression surface on the outer compression surface, especially the influence of the multistage compression corner and the radius of the front edge to the shock boundary layer interference and the inlet performance. Boundary layer behavior, pulse type hypersonic test equipment is an important supplement to conventional wind tunnel because of its ability to simulate high hypersonic total pressure test conditions. The combination of the two can cover a more comprehensive Maher number, Reynolds number and the range of wall temperature ratio. It is also an important test platform for the study of shock wave boundary layer interference. An oil flow test technique which can display boundary layer separation conditions is still lacking on the flush hypersonic test equipment. At the same time, due to the high temperature and high speed testing environment in the hypersonic flow field, the conventional boundary layer thickness testing technology, such as the pitot probe, is difficult to test the thickness of the high supersonic boundary layer. In this paper, we develop two kinds of high supersonic wave boundary layer interference in high supersonic wind tunnel. Firstly, the method of measuring the thickness of hypersonic boundary layer based on the focused schlieren technique is developed on the shock wind tunnel, and the density is extracted from the focused schlieren photos with the theoretical and numerical simulation results. The method of boundary layer thickness data shows that the thickness of the boundary layer based on the density gradient is approximately the thickness of the density boundary layer, the density boundary layer is the same as the thickness of the temperature boundary layer. The thickness of the air velocity boundary layer is slightly less than the thickness of the density boundary layer, and the light source and imaging system of the focused schlieren system are improved in application. Secondly, the pulse combustion is used. The oil flow display technology is developed in the wind tunnel. The difficulties in the development of oil flow display technology on pulse equipment are analyzed in combination with the characteristics of the pulse combustion equipment. The stress situation and movement mechanism of the oil film are analyzed. The influence factors of the motion velocity of the oil film are given, and the test criterion of the film type oil flow on the instantaneous pulse blast tunnel is put forward. The oil film spraying method is properly displayed and the oil film spraying method is developed. The self luminescence problem of the high temperature test gas is solved. The real-time camera system is designed. The oil flow map sequence is collected during the very short effective test time. The dynamic development process of the oil film proves that the above relationship criterion is used as the boundary layer thickness measurement method and the oil flow display. The effect of the multistage compression corner and the radius of the front edge on the performance of the intake port and the interference characteristics of the shock wave boundary layer is preliminarily studied with the use of other testing and numerical simulation methods. Under the condition of the nominal Maher number 6 and the unit Reynolds number of 5.4 x 106/m on the pulse combustion wind tunnel, the obtained oil flow map sequence shows that the two stage compression wedge abduction is shown. The initial separation angle at the corner is between 100~150, and the flow of the three stage compression wedge is not separated at second corners. Under the condition of the nominal Maher number 6 and the unit Reynolds number 3.4 x 107/m on the shock wind tunnel, the influence of the three level plane compression on the radius of the front edge of the high supersonic inlet is studied, and the wall static pressure, the wall heat flow, the shock geometry feature and the edge are obtained. The boundary layer density thickness distribution data. The wall static pressure begins to climb up the inflection point and reaches the pressure platform at a certain distance after the inflection point; the radius of the front edge increases and the pressure platform value decreases, which can increase the distance needed by the pressure platform, 'means that the disturbance area of the shock boundary layer is enlarged at the corner and the shock boundary layer in the front edge is the interference area. The range is 5 times the thickness of the upstream boundary layer, and the 3mm blunt front edge reaches 8 times. The wall heat flow gradually rises from the upstream of the inflection point, reaches a peak after the turning point, and then drops to the next compression corner. With the increase of the radius of the front edge, the peak heat flow peak decreases significantly, the peak position of the heat flow moves upstream. Boundary layer thickness. The thickness of the upper boundary layer is significantly lower than the inflow boundary layer thickness, and the thickness of the boundary layer is increased again. With the increase of the radius of the front edge, the thickness of the boundary layer increases with the increase of the radius of the front edge. The wall static pressure obtained by using different turbulence models using FLUNET software is in good agreement with the test results, but the wall wall is in good agreement with the test results. The distribution of surface heat flow and the thickness distribution of the boundary layer density vary greatly. With the increase of the radius of the front edge, the first shock angle increases gradually, the second and third magnitude shock angles gradually decrease, and the bending degree and the affected area of the second, third shock waves at the corner are increased.
【學(xué)位授予單位】：中國空氣動力研究與發(fā)展中心
【學(xué)位級別】：博士
【學(xué)位授予年份】：2015
【分類號】：V211.74
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