超声/高超声速非均匀来流下曲面压缩系统研究
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摘要
本文针对典型的超燃冲压发动机工作环境,采用理论分析、数值模拟和风洞试验相结合的方法,研究了一种新型的弯曲压缩系统及其所形成的弯曲激波。
首先,讨论了二维平面压缩、等熵压缩和锥波压缩的设计方法,提出了形成弯曲激波的二维弯曲压缩系统的设计思想。
其次,在无粘条件下分析了弯曲压缩系统的初始斜板长度、压缩面总长度、形成弯曲压缩系统的小折线偏转角度以及递增偏转角度等不同因素对弯曲压缩系统性能的影响。选取NASA经典的进气道试验作为参照依据进行算例考核,考查了本文研究中所采用的数值模拟方法及软件对进气道特性的预测能力,分析和比较了数值模拟和试验结果,根据不同湍流模型对进气道性能预测结果,确定了本文研究中所适用的湍流模型。
再次,分别设计了壁面压力按等压力梯度变化和壁面马赫数按等马赫数梯度变化的弯曲压缩系统,应用斜激波理论和等熵压缩波理论,给出了可以快速、便捷地获得其弯曲激波形状参数、压缩面上性能参数、激波后性能参数以及出口性能参数的近似计算方法,和预估弯曲压缩系统型面及其形成的弯曲激波的估算公式。
然后,选取一个设计点马赫数5.3、壁面压力按等压力梯度变化的弯曲压缩系统,经附面层修正后的型面加工成模型,在南航Ma 3.85/5.3小型高超风洞完成了均匀和非均匀来流试验。试验结果表明:在均匀来流条件下,设计点马赫数5.3,弯曲压缩系统壁面压力按等压力梯度分布;非设计点马赫数3.85,壁面压力也基本呈等压力梯度分布;在非均匀来流条件下,马赫数3.85比马赫数5.3时壁面压力更接近等压力梯度分布。同时,验证了所提出的二维等压力梯度弯曲压缩系统的设计方法。
最后,比较了等压力梯度曲面压缩非常规二维进气道同二楔压缩和三楔压缩的常规二维进气道性能。等压力梯度曲面压缩二维进气道与三楔压缩二维进气道压比相当时,其压缩面长度比三楔压缩的常规二维进气道缩短20.5%,大大地减轻了进气道的重量。在非设计点非均匀来流条件下,等压力梯度设计曲面压缩二维进气道较之楔压缩二维进气道,有较高的流量系数和总压恢复,吞入36%唇口高度来流附面层后,等压力梯度设计曲面压缩二维进气道流量系数比二楔压缩二维进气道高0.6%,总压恢复高6.8%。三楔压缩二维进气道在吞入较厚附面层后会导致不起动。同时,比较了均匀来流和非均匀来流条件下顶板平面压缩+侧板平面压缩、顶板平面压缩+侧板曲面压缩、顶板曲面压缩+侧板平面压缩和顶板曲面压缩+侧板曲面压缩四种侧压式进气道性能。在均匀来流条件下,侧板采用等压力梯度设计曲面压缩系统的侧压式进气道性能较好;在非均匀来流条件下,顶板+侧板均采用等压力梯度设计曲面压缩系统的侧压式进气道具有较高的流量系数和总压恢复,出口流场畸变较小。
通过本文研究,给出了一种比目前已经使用的或者在研的压缩系统的长度要短、压缩面的压力梯度可控、对非均匀超声速来流有更好承受能力的弯曲压缩系统,为今后超燃冲压发动机进气道设计建立新的发展基础。
A detailed study is performed in current dissertation on the newly curved compression system and its curved shock wave at the typical scramjet working environment by theoretical analysis, wind tunnel testing and numerical simulation methods.
First, design methods of the two-dimensional plane compression, isentropic compression and cone wave compression are discussed. Then, the design method of a two-dimensional curved compression system that can form curved shock is given.
Secondly, analysis the effect of the initial ramp length, total compression surface length, deflection angle of a small curved line forming compression system and increased yaw angle on the performance of different curved compression systems in the inviscid condition. The evaluation of software to predict characteristics of the inlet is assessing by selected the classic NASA inlet experiment example. Analysis and comparison of the numerical simulation and experimental results, different turbulence models forecast the performance of inlet obtained. The selected turbulence model is given.
Then, constant pressure gradient wall and constant Mach number gradient wall are designed, quick and convenient approximate calculation method access to curved shock wave shape, performance parameters on the compression surface, performance parameters after the shock and export performance parameters is given by applications of oblique shock wave theory and the theory of isentropic compression. The estimation formulas of estimating the shape of curved compression surface and its shock wave are given.
Next, a constant pressure gradient compression system modified by boundary layer with design point Mach number 5.3 is selected, tested in NUAA Mach number 3.85/5.3 wind tunnel in uniform and non-uniform inflow.The results showed that: at the design point Mach number 5.3, the wall pressure distribution of bending compression system is constant pressure gradient; so does at the off-design Mach number 3.85. In the non-uniform flow conditions, wall pressure distribution is closer to the constant pressure gradient at Mach number 3.85 than at Mach number 5.3. All results prove that the design method is feasible.
Detailed study on the performance of a two-dimensional constant pressure gradient non-conventional compression inlet and two-dimensional wedge compression inlets are performed. At design point, the pressure ratio of a two-dimensional non-conventional compression inlet is higher , its length of the compression surface can be reduced about 20.5 percent, greatly reducing the inlet weight. In Mach number 4 uniform inflow, a two-dimensional non-conventional compression inlet gets a higher mass capture ratio and total pressure recovery. Swallowed inflow boundary layer about 36 percent height of inlet height, mass capture ratio and total pressure recovery of the two-dimensional non-conventional compression inlet is increased 0.6 percent and 4.8 percent, respectively than two wedges compression inlet, but three wedges compression inlet is unstarted.
In uniform and non-uniform inflow, the performance of three-dimensional sidewall compression inlets with plane sidewall + plane topwall, curved sidewall + plane topwall, plane sidewall + curved topwall,and curved sidewall +curved topwall,respetively are presented. In uniform inflow conditions, three-dimensional sidewall compression inlet with plane sidewall + curved topwall is of good performance. In non-uniform inflow, three-dimensional sidewall compression inlet with curved sidewall + curved topwall is of higher mass capture ratio and total pressure recovery .
The study of a curved compression system with a shorter length, compression surface pressure gradient controlled, better ability to withstand non-uniform supersonic inflow is presented. The research results will be founded a new base for development of future scramjet research.
首先,讨论了二维平面压缩、等熵压缩和锥波压缩的设计方法,提出了形成弯曲激波的二维弯曲压缩系统的设计思想。
其次,在无粘条件下分析了弯曲压缩系统的初始斜板长度、压缩面总长度、形成弯曲压缩系统的小折线偏转角度以及递增偏转角度等不同因素对弯曲压缩系统性能的影响。选取NASA经典的进气道试验作为参照依据进行算例考核,考查了本文研究中所采用的数值模拟方法及软件对进气道特性的预测能力,分析和比较了数值模拟和试验结果,根据不同湍流模型对进气道性能预测结果,确定了本文研究中所适用的湍流模型。
再次,分别设计了壁面压力按等压力梯度变化和壁面马赫数按等马赫数梯度变化的弯曲压缩系统,应用斜激波理论和等熵压缩波理论,给出了可以快速、便捷地获得其弯曲激波形状参数、压缩面上性能参数、激波后性能参数以及出口性能参数的近似计算方法,和预估弯曲压缩系统型面及其形成的弯曲激波的估算公式。
然后,选取一个设计点马赫数5.3、壁面压力按等压力梯度变化的弯曲压缩系统,经附面层修正后的型面加工成模型,在南航Ma 3.85/5.3小型高超风洞完成了均匀和非均匀来流试验。试验结果表明:在均匀来流条件下,设计点马赫数5.3,弯曲压缩系统壁面压力按等压力梯度分布;非设计点马赫数3.85,壁面压力也基本呈等压力梯度分布;在非均匀来流条件下,马赫数3.85比马赫数5.3时壁面压力更接近等压力梯度分布。同时,验证了所提出的二维等压力梯度弯曲压缩系统的设计方法。
最后,比较了等压力梯度曲面压缩非常规二维进气道同二楔压缩和三楔压缩的常规二维进气道性能。等压力梯度曲面压缩二维进气道与三楔压缩二维进气道压比相当时,其压缩面长度比三楔压缩的常规二维进气道缩短20.5%,大大地减轻了进气道的重量。在非设计点非均匀来流条件下,等压力梯度设计曲面压缩二维进气道较之楔压缩二维进气道,有较高的流量系数和总压恢复,吞入36%唇口高度来流附面层后,等压力梯度设计曲面压缩二维进气道流量系数比二楔压缩二维进气道高0.6%,总压恢复高6.8%。三楔压缩二维进气道在吞入较厚附面层后会导致不起动。同时,比较了均匀来流和非均匀来流条件下顶板平面压缩+侧板平面压缩、顶板平面压缩+侧板曲面压缩、顶板曲面压缩+侧板平面压缩和顶板曲面压缩+侧板曲面压缩四种侧压式进气道性能。在均匀来流条件下,侧板采用等压力梯度设计曲面压缩系统的侧压式进气道性能较好;在非均匀来流条件下,顶板+侧板均采用等压力梯度设计曲面压缩系统的侧压式进气道具有较高的流量系数和总压恢复,出口流场畸变较小。
通过本文研究,给出了一种比目前已经使用的或者在研的压缩系统的长度要短、压缩面的压力梯度可控、对非均匀超声速来流有更好承受能力的弯曲压缩系统,为今后超燃冲压发动机进气道设计建立新的发展基础。
A detailed study is performed in current dissertation on the newly curved compression system and its curved shock wave at the typical scramjet working environment by theoretical analysis, wind tunnel testing and numerical simulation methods.
First, design methods of the two-dimensional plane compression, isentropic compression and cone wave compression are discussed. Then, the design method of a two-dimensional curved compression system that can form curved shock is given.
Secondly, analysis the effect of the initial ramp length, total compression surface length, deflection angle of a small curved line forming compression system and increased yaw angle on the performance of different curved compression systems in the inviscid condition. The evaluation of software to predict characteristics of the inlet is assessing by selected the classic NASA inlet experiment example. Analysis and comparison of the numerical simulation and experimental results, different turbulence models forecast the performance of inlet obtained. The selected turbulence model is given.
Then, constant pressure gradient wall and constant Mach number gradient wall are designed, quick and convenient approximate calculation method access to curved shock wave shape, performance parameters on the compression surface, performance parameters after the shock and export performance parameters is given by applications of oblique shock wave theory and the theory of isentropic compression. The estimation formulas of estimating the shape of curved compression surface and its shock wave are given.
Next, a constant pressure gradient compression system modified by boundary layer with design point Mach number 5.3 is selected, tested in NUAA Mach number 3.85/5.3 wind tunnel in uniform and non-uniform inflow.The results showed that: at the design point Mach number 5.3, the wall pressure distribution of bending compression system is constant pressure gradient; so does at the off-design Mach number 3.85. In the non-uniform flow conditions, wall pressure distribution is closer to the constant pressure gradient at Mach number 3.85 than at Mach number 5.3. All results prove that the design method is feasible.
Detailed study on the performance of a two-dimensional constant pressure gradient non-conventional compression inlet and two-dimensional wedge compression inlets are performed. At design point, the pressure ratio of a two-dimensional non-conventional compression inlet is higher , its length of the compression surface can be reduced about 20.5 percent, greatly reducing the inlet weight. In Mach number 4 uniform inflow, a two-dimensional non-conventional compression inlet gets a higher mass capture ratio and total pressure recovery. Swallowed inflow boundary layer about 36 percent height of inlet height, mass capture ratio and total pressure recovery of the two-dimensional non-conventional compression inlet is increased 0.6 percent and 4.8 percent, respectively than two wedges compression inlet, but three wedges compression inlet is unstarted.
In uniform and non-uniform inflow, the performance of three-dimensional sidewall compression inlets with plane sidewall + plane topwall, curved sidewall + plane topwall, plane sidewall + curved topwall,and curved sidewall +curved topwall,respetively are presented. In uniform inflow conditions, three-dimensional sidewall compression inlet with plane sidewall + curved topwall is of good performance. In non-uniform inflow, three-dimensional sidewall compression inlet with curved sidewall + curved topwall is of higher mass capture ratio and total pressure recovery .
The study of a curved compression system with a shorter length, compression surface pressure gradient controlled, better ability to withstand non-uniform supersonic inflow is presented. The research results will be founded a new base for development of future scramjet research.
引文
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