高超声速咽式进气道流场特性和设计方法研究
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摘要
高超声速进气道,作为超燃冲压发动机的重要组成部分,一方面要为发动机提供所需的空气,另一方面对来流进行足够的压缩,以满足燃烧的需要,其性能已成为决定发动机乃至整个吸气式高超声速飞行器性能的关键因素。美国空军实验室的Malo-Molina等人于2005年提出了一种新型三维内收缩高超进气道——“咽”式进气道,数值模拟结果表明:“咽”式进气道与常规进气道相比,具有更好的综合性能。但是目前国外公开文献并没有对其具体设计方法、内部流场特性进行详细的阐述,关于“咽”式进气道的许多问题还需要进一步的深入研究。本文以寻求适合作为吸气式高超声速飞行器进气道的,在宽马赫数范围内(Ma=5,6,7)具有较高性能的流线追踪“咽”式进气道为主要目标,利用理论分析、数值模拟、风洞实验等手段,对流线追踪“咽”式进气道进行了较为系统的研究。
对“咽”式进气道无粘流场、设计理论及方法进行了详细介绍,采用数值模拟的方法比较了具有相同基本流场的矩形进气道和“咽”式进气道的流动特性和性能,并分析了不同马赫数,不同攻角以及粘性对其流动特性和性能的影响,使本文所述的流线追踪设计方法得到了验证。另外可知,与矩形进气道相比,“咽”式进气道对设计状态下的攻角变化不太敏感,在非设计状态下具有较高的流量捕获和压缩能力,由于其浸湿面积小,进气道内附面层增长缓慢,激波与附面层干扰较弱,总压恢复较高。
通过数值模拟,选择了较为合适的进气道设计参数,并为解决由于粘性的存在可能引起的基于无粘流场设计方法的失效问题,利用Cebeci和Bradshaw提供的有限差分法求解二维可压缩附面层方程的程序,根据求得的进气道附面层位移厚度对8-7(即俯仰平面内的斜激波由和自由来流呈8o夹角的斜压缩面产生;偏航平面内的斜激波由和自由来流呈7o夹角的斜压缩面产生)“咽”式进气道进行了附面层修正,比较分析了修正前后粘性计算以及未修正无粘计算得到的进气道性能参数。
对进行附面层修正前后的“咽”式进气道开展了不同马赫数和攻角条件下的风洞实验和数值模拟,实验测量了进气道壁面中心线沿程静压分布,分析了附面层修正、来流马赫数及攻角的变化对“咽”式进气道流场和性能的影响。研究表明,壁面中心线沿程静压的数值模拟结果与风洞试验结果吻合较好,特别在进气道唇口前的压缩面上,壁面静压几乎完全相同,只是在进气道内收缩段的压力分布存在误差,但变化趋势是一致的,数值模拟结果基本能反映进气道内的压力变化,具有一定的可信度;另外,附面层修正能大幅度的提高“咽”式进气道模型的总压恢复系数,减小其内收缩比,提高其起动性能,从而扩大了其工作范围,对进气道设计有着积极的作用。
运用本文提出的高超声速飞行器一体化设计思想,对基于“咽”式进气道的飞行器外形设计进行了初步研究。通过数值模拟,对设计出的满足技术要求的高超声速飞行器的气动特性进行了分析,为今后“咽”式进气道的工程应用做了一定的尝试和铺垫。
As the important component of scramjet, the hypersonic inlet not only provides the necessary air for the engine, but also compresses enough air to meet the needs of combustion, so its performance is a key factor to scramjet and airbreathing hypersonic vehicle. In 2005, Malo-Molina et al from U.S. Air Force Laboratory propose Jaws inlet, a kind of innovative inward turning inlet. Numerical simulation results indicate that Jaws inlet has better property compared to routine inlet. But now the foreign open literature does not dictate a detailed exposition on specific design method, the internal flow field of the Jaws inlet, a lot of problems need more thorough investigation.
In present paper,the main objective is finding in a wide range of Mach number (Ma = 5,6,7), high-performance stream tracing Jaws inlet, which is suitable for air-breathing hypersonic vehicle inlet. A system investigation on streamline tracing Jaws inlet has been carried out using theoretical analysis, numerical simulation, wind tunnel experiments and other means.
Inviscid basic flow field and design method of Jaws inlet was described in details. Jaws inlet was explored numerically and compared with a rectangular cross-section inlet in the same basic flow field. The performance and flow-field characteristic of Jaws inlet on different Mach number and attack angle was compared, and the effect of viscosity were also analyzed. The design method of streamline traced has been verified. In addition, compared to the rectangular cross-section, inlet performance of the Jaws inlet is insensitive to the angles of attack on design point; inlet has high mass capture and compression capability on off-design point; boundary layer growth of the develops slowly and shock boundary layer interaction is weak because of the diminished wetted area, and total pressure recovery coefficient is much higher.
Because the presence of viscosity, a more suitable inlet design parameters are selected by numerical simulation, in order to solve the possible failure of design problems based on inviscid flow. Cebeci and Bradshaw’s finite difference program is used to solve the two-dimensional compressible boundary layer equation, according to the achieved displacement thickness of the boundary layer, the 8-7 (the ramp in pitch plane is inclined at 8o to the free stream and in yaw plane is inclined at 7o to the free stream, yielding planar shocks).inlet internal profile is then corrected, and the inlet performance after correction is compared with uncorrected flow by numerical simulation. Jaws inlet with and without boundary layer correction were investigated using numerical simulation and hypersonic wind tunnel tests. Static pressure distribution along the centerline of the walls was measured. The performance of Jaws inlet on different Mach number and attack angle was compared, and the effect of boundary layer correction was also studied. The tunnel experimental data agree well with the numerical simulation. Especially at the wall before the lip, the static pressures almost have the same value. Although there is small deviation existing at inward contraction section, the pressure changing trends of the wall are consistent. The experimental results prove that the numerical simulation can reflect the pressure distribution with good reliability. In addition, boundary layer correction improves the total pressure recovery coefficient of Jaws inlet, reduce internal contraction ratio, and enhance the start performance. So the working range is enlarged, it has positive effect on the design of inlet.
Using the proposed hypersonic propulsion/airframe integration design, based on Jaws inlet, shape design of an aircraft was studied. By numerical simulation, hypersonic vehicle's aerodynamic characteristics are analyzed; the design meets the technical requirements. Some try and bedding are carried out for the future engineering application of Jaws inlet.
对“咽”式进气道无粘流场、设计理论及方法进行了详细介绍,采用数值模拟的方法比较了具有相同基本流场的矩形进气道和“咽”式进气道的流动特性和性能,并分析了不同马赫数,不同攻角以及粘性对其流动特性和性能的影响,使本文所述的流线追踪设计方法得到了验证。另外可知,与矩形进气道相比,“咽”式进气道对设计状态下的攻角变化不太敏感,在非设计状态下具有较高的流量捕获和压缩能力,由于其浸湿面积小,进气道内附面层增长缓慢,激波与附面层干扰较弱,总压恢复较高。
通过数值模拟,选择了较为合适的进气道设计参数,并为解决由于粘性的存在可能引起的基于无粘流场设计方法的失效问题,利用Cebeci和Bradshaw提供的有限差分法求解二维可压缩附面层方程的程序,根据求得的进气道附面层位移厚度对8-7(即俯仰平面内的斜激波由和自由来流呈8o夹角的斜压缩面产生;偏航平面内的斜激波由和自由来流呈7o夹角的斜压缩面产生)“咽”式进气道进行了附面层修正,比较分析了修正前后粘性计算以及未修正无粘计算得到的进气道性能参数。
对进行附面层修正前后的“咽”式进气道开展了不同马赫数和攻角条件下的风洞实验和数值模拟,实验测量了进气道壁面中心线沿程静压分布,分析了附面层修正、来流马赫数及攻角的变化对“咽”式进气道流场和性能的影响。研究表明,壁面中心线沿程静压的数值模拟结果与风洞试验结果吻合较好,特别在进气道唇口前的压缩面上,壁面静压几乎完全相同,只是在进气道内收缩段的压力分布存在误差,但变化趋势是一致的,数值模拟结果基本能反映进气道内的压力变化,具有一定的可信度;另外,附面层修正能大幅度的提高“咽”式进气道模型的总压恢复系数,减小其内收缩比,提高其起动性能,从而扩大了其工作范围,对进气道设计有着积极的作用。
运用本文提出的高超声速飞行器一体化设计思想,对基于“咽”式进气道的飞行器外形设计进行了初步研究。通过数值模拟,对设计出的满足技术要求的高超声速飞行器的气动特性进行了分析,为今后“咽”式进气道的工程应用做了一定的尝试和铺垫。
As the important component of scramjet, the hypersonic inlet not only provides the necessary air for the engine, but also compresses enough air to meet the needs of combustion, so its performance is a key factor to scramjet and airbreathing hypersonic vehicle. In 2005, Malo-Molina et al from U.S. Air Force Laboratory propose Jaws inlet, a kind of innovative inward turning inlet. Numerical simulation results indicate that Jaws inlet has better property compared to routine inlet. But now the foreign open literature does not dictate a detailed exposition on specific design method, the internal flow field of the Jaws inlet, a lot of problems need more thorough investigation.
In present paper,the main objective is finding in a wide range of Mach number (Ma = 5,6,7), high-performance stream tracing Jaws inlet, which is suitable for air-breathing hypersonic vehicle inlet. A system investigation on streamline tracing Jaws inlet has been carried out using theoretical analysis, numerical simulation, wind tunnel experiments and other means.
Inviscid basic flow field and design method of Jaws inlet was described in details. Jaws inlet was explored numerically and compared with a rectangular cross-section inlet in the same basic flow field. The performance and flow-field characteristic of Jaws inlet on different Mach number and attack angle was compared, and the effect of viscosity were also analyzed. The design method of streamline traced has been verified. In addition, compared to the rectangular cross-section, inlet performance of the Jaws inlet is insensitive to the angles of attack on design point; inlet has high mass capture and compression capability on off-design point; boundary layer growth of the develops slowly and shock boundary layer interaction is weak because of the diminished wetted area, and total pressure recovery coefficient is much higher.
Because the presence of viscosity, a more suitable inlet design parameters are selected by numerical simulation, in order to solve the possible failure of design problems based on inviscid flow. Cebeci and Bradshaw’s finite difference program is used to solve the two-dimensional compressible boundary layer equation, according to the achieved displacement thickness of the boundary layer, the 8-7 (the ramp in pitch plane is inclined at 8o to the free stream and in yaw plane is inclined at 7o to the free stream, yielding planar shocks).inlet internal profile is then corrected, and the inlet performance after correction is compared with uncorrected flow by numerical simulation. Jaws inlet with and without boundary layer correction were investigated using numerical simulation and hypersonic wind tunnel tests. Static pressure distribution along the centerline of the walls was measured. The performance of Jaws inlet on different Mach number and attack angle was compared, and the effect of boundary layer correction was also studied. The tunnel experimental data agree well with the numerical simulation. Especially at the wall before the lip, the static pressures almost have the same value. Although there is small deviation existing at inward contraction section, the pressure changing trends of the wall are consistent. The experimental results prove that the numerical simulation can reflect the pressure distribution with good reliability. In addition, boundary layer correction improves the total pressure recovery coefficient of Jaws inlet, reduce internal contraction ratio, and enhance the start performance. So the working range is enlarged, it has positive effect on the design of inlet.
Using the proposed hypersonic propulsion/airframe integration design, based on Jaws inlet, shape design of an aircraft was studied. By numerical simulation, hypersonic vehicle's aerodynamic characteristics are analyzed; the design meets the technical requirements. Some try and bedding are carried out for the future engineering application of Jaws inlet.
引文
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