高超声速稀薄流中横向喷流干扰特性实验研究

喷流干扰是高超声速飞行高精度控制的一种有效手段,研究者们以往大部分都主要集中于连续流条件下喷流干扰效应的机理研究,并给出了喷流干扰流场的典型结构,而稀薄流条件下喷流干扰特性的实验数据还十分匮乏.本文利用JFX爆轰激波风洞产生高超声速稀薄自由流,基于平板模型开展不同喷流压力和自由来流参数对横向喷流干扰特性影响的实验研究,采用高速纹影成像及图像处理技术,获得稀薄流条件下喷流干扰流场演化过程及流场结构的变化规律.相比于无喷流条件形成的流场,横向喷流与稀薄自由流相互作用形成的流场结构更为复杂,喷流压力由于受到稀薄来流的扰动,斜激波会短暂穿透喷流干扰流场并延伸至楔形体上部.喷流干扰流场内桶状激波的影响范围随着喷流压力的升高而逐渐变宽,位于三波点上游的斜激波空间位置不会随喷流压力的变化而改变,而位于三波点下游的弓形激波则向上游移动,当喷流压力过低时,桶状激波不会与其他两种激波交汇形成三波点.高超声速稀薄来流压力的降低同样会使桶状激波的影响范围变宽,弓形激波同样也会向上游移动,但基本不会对斜激波空间位置产生任何影响.     

侧向多喷口干扰复杂流动数值模拟研究

采用具有高分辨率的NND格式,通过数值求解N-S方程对典型外形多喷口侧向喷流复杂干扰流动进行了数值模拟. 为了提高计算效率,采用了LU-SGS隐式算法. 采用分块对接网格技术,生成高质量的贴体计算网格,精确模拟喷口截面. 对比分析了不同计算格式、限制器形式、网格拓扑及流动形态(层流与湍流)对喷流干扰流场结构和压力分布特性的影响,研究和分析了喷口附近流场的涡系结构、波系结构和喷流干扰引起的气动力特性. 在上述研究的基础上,针对典型飞行器外形的侧向喷流干扰特性进行了详细的数值模拟,得到了喷口参数(喷口位置、数目等)及来流条件对喷流干扰流场结构、气动力特性的影响规律,并对其流动机理进行了相应的分析. 研究表明,发展的针对多喷口侧喷干扰的数值计算方法是成功的,可以应用于飞行器侧向喷流干扰的流场结构分析及气动力特性数值预测.      

导弹多喷口侧向喷流流动数值模拟

针对典型导弹外形,通过求解三维Navier-Stoke方程开展了多喷口喷流与超声速来流的复杂干扰流动数值模拟,研究了来流马赫数、来流攻角、喷口数量及位置、喷流压力比对流场结构和弹体表面流动的影响.结果表明,相邻喷口间会产生明显的干扰,来流马赫数、攻角、喷口数量及位置和喷流压力都会对喷流与主流之间的干扰以及导弹表面流动产生明显影响,且在沿流向布置的多喷口中,最上游喷口决定了喷口前的流场结构.     

带喷流激波针流动特性实验研究

采用动态测力、动态测压和纹影等风洞实验技术,对加装了带喷流激波针的钝头体的绕流特性、稳定和非稳模态的形成条件和机理进行了研究.结果表明:带喷流激波针流场存在稳态和非稳态两种模态,超声速喷流的压比大于临界压比时流动处于稳定模态,反之则为非稳模态;增大激波针长度可减小钝头体阻力,但达到一定长度后,进一步减阻的效果不再显著;增大喷流压比能够有效减弱再附激波强度,有利于缓解单独激波针的肩部热斑问题;非稳模态下波系自激振荡对再附激波在钝头体表面所围的区域影响剧烈,振荡是周期性的,且存在确定的主导频率,主导频率随喷流压力比增大而减小;自激振荡的产生是由于喷流出口周围的反压在喷流压比小于临界压比时无法获得持续的平衡而导致.      

超声速多喷流干扰流场特性研究

研究了旋成体上超声速来流与超声速横向多喷流相撞产生的层流干扰流场特性. 数 值方法针对三维可压缩Navier-Stokes方程按二阶精度Roe格式进行离散,采用基于多区对 接网格技术的有限体积法. 数值模拟结果描述了多喷流干扰流场的空间结构以及激波/边界层 干扰引起的分离范围,探讨了沿流向等间距排列的喷口个数对表面和空间流场结构以及压力 分布的影响规律. 结果表明,第一喷口对多喷流干扰流场主要结构和喷口上游表面分离范围 起主导作用. 其中三喷流流场数值模拟的对称面激波结构与实验纹影结果进行对比,符合较 好.     

高温燃气喷流/主流相互干扰实验研究

报道关于高温燃气自由喷流(热喷流)、燃气喷流/主流干扰流对气动热环境影响的实验研究结果. 其意义在于: 抽象出高超声速飞行器实际飞行时燃气喷流及其干扰流的物理模型, 为高超声速飞行器防热需求提供实验依据. 实验主流由脉冲风洞提供,燃气喷流用氢氧燃烧驱动路德维希管的方式产生. 利用脉冲风洞驱动段压力信号自动控制热气源的产生以保证风洞主流与燃气喷流同步, 利用氢气、氮气和氧气的不同比例实现燃气喷流的热力学相似. 实验技术上完成了高温燃气喷流系统的参数采集与系统状态标定; 实验内容上开展了压缩拐角平板模型的气动热实验研究, 通过实验比较了只有主流流场、只有热喷流流场和既有主流流场又有热喷流流场(即干扰流场)3种工况的热流分布. 实验研究发现,热喷流/主流相互干扰会对压缩拐角平板上某一范围内的气动热环境造成显著影响, 热流峰值较无喷流流场高出一个量级.      

矢量喷流对细长体大迎角非对称流动影响研究

采用测压方法研究了矢量喷流对细长旋成体大迎角非对称流动的影响特性.实验结果表明:矢量喷流对细长旋成体大迎角非对称侧向力有明显的抑制作用,该抑制作用是通过喷流诱导作用,改变其空间绕流涡系结构的分布来实现的,但是矢量喷流的存在并不能改变大迎角机身空间绕流涡系的本质结构;随着迎角的增大,矢量喷流对细长旋成体大迎角非对称流动的影响区域不断前移,甚至影响到头部;随着喷流落压比的增加,矢量喷流对细长旋成体大迎角非对称侧向力的抑制作用加强,但当喷流落压比达到临界落压比后(即喷管出口处达到设计马赫数时),喷流影响作用将不会随喷流落压比的增加而改变.     

底部喷流及后掠翼前缘涡破碎的流动显示研究

本文介绍了在中国科技大学立式水洞中进行的底部喷流流动显示及底压测量实验以及对x型翼体组合体弹翼上前缘涡破碎的观察,系统地比较了有、无喷流及有、无尾罩下各种底部流态,为工程应用提供了实验依据.     

在超声速流动中激波与边界层的干扰特性

激波与物面边界层的干扰涉及可压缩流动的稳定性、转捩、分离等问题,直接影响到飞行器的阻力、表面热防护和飞行性能等工程技术问题。首先总结了前人对于激波与边界层的干扰所做的工作,之后重点研究和对比分析了超声速与跨声速流动中,正激波、斜激波以及头部激波对于飞行器层流和湍流边界层的干扰影响。激波强度的不同对边界层干扰作用不同,在强干扰情况下将会引起边界层分离和翼型失速。     

绕水翼空化流动及振动特性的实验研究

空化是发生在水力机械内部的一种水动力现象,其发展具有显著的非定常特性.空化流动中空穴的脱落以及溃灭会诱发结构振动,对水力机械的效率、噪声、安全性等造成影响. 研究空化流动中结构的振动特性具有重要的工程意义. 采用实验的方法研究了绕NACA66 水翼空化流动的空穴形态和水翼振动特性. 实验在一闭式空化水洞中进行. 采用高速摄像技术观测不同空化阶段的空穴形态,应用多普勒激光测振仪测量水翼的振动速度,并通过一套同步系统实现了高速相机和多普勒激光测振仪的同步触发和测量. 采用小波分析方法对不同空化阶段下的空穴形态和水翼振动数据在时域和频域中的特性进行了分析.对云状空化阶段的同步测量结果进行了研究,分析了振动与空穴发展过程的联系. 结果表明,随着空化数的降低,流场经历了无空化、初生空化、片状空化和云状空化4个阶段,水翼的振动强度呈逐渐增大趋势. 在片状空化和云状空化阶段,空穴脱落导致水翼振动,诱发的振动频率与空穴脱落频率相同. 对于云状空化,在附着型空穴生长阶段水翼发生高频小幅度振动,在空穴脉动和断裂脱落期间水翼表现为低频大幅振动.      

Oscillatory behavior of supersonic impinging jet flows

Oscillatory flows of a choked underexpanded supersonic impinging jet issuing from a convergent nozzle have been computed using the axisymmetric unsteady Navier--Stokes system. This paper focuses on the oscillatory flow features associated with the variation of the nozzle-to-plate distance and nozzle pressure ratio. Frequencies of the surface pressure oscillation and flow structural changes from computational results have been analyzed. Staging behavior of the oscillation frequency has been observed for both cases of nozzle-to-plate distance variation and pressure ratio variation. However, the staging behavior for each case exhibits different features. These two distinct staging behaviors of the oscillation frequency are found to correlate well if the frequency and the distance are normalized by the length of the shock cell. It is further found that the staging behavior is strongly correlated with the change of the pressure wave pattern in the jet shear layer, but not with the shock cell structure. Communicated by K. Takayama PACS 02.60.Cb; 47.40.−x; 47.40.Nm; 47.35.+I; 47.15.−x     

DIAMOND PORT JET INTERACTION WITH SUPERSONIC FLOW

Interaction flow field of the sonic air jet through diamond shaped orifices at different incidence angles （10 degrees, 27.5 degrees, 45 degrees and 90 degrees） and total pressures （0.10 MPa and 0. 46 MPa） with a Mach 5.0 freestream was studied experimentally. A 90 degrees circular injector was examined for comparison. Crosssection Mach number contours were acquired by a Pitot-cone five-hole pressure probe. The results indicate that the low Mach semicircular region close to the wall is the wake region. The boundary layer thinning is in the areas adjacent to the wake. For the detached case, the interaction shock extends further into the freestream, and the shock shape has more curvature, also the low-Mach upwash region is larger. The vortices of the plume and the height of the jet interaction shock increase with increasing incidence angle and jet pressure. 90 degrees diamond and circular injector have stronger plume vorticity, and for the circular injector low-Mach region is smaller than that for the diamond injector. Tapered ramp increases the plume vorticity, and the double ramp reduces the level of vorticity. The three-dimensional interaction shock shape was modeled from the surface shock shape, the center plane shock shape, and crosssectional shock shape. The shock total pressure was estimated with the normal component of the Mach number using normal shock theory. The shock induced total pressure losses decrease with decreasing jet incidence angle and injection pressure, where the largest losses are incurred by the 90 degrees, circular injector.     

超声速钝体逆向喷流减阻的数值模拟研究

为研究逆向喷流技术对超声速钝体减阻的影响,采用标准k-ε湍流模型,通过求解二维Navier-Stokes方程对超声速球头体逆向冷喷流流场进行了数值模拟,并着重分析了喷口总压、喷口尺寸对流场模态和减阻效果的影响。计算结果显示:随着喷流总压的变化,流场可出现两种流动模态,即长射流穿透模态和短射流穿透模态;喷流能使球头体受到的阻力明显减小;存在最大减阻临界喷流总压值(在所研究参数范围内最大减阻可达51.1%);在其它喷流物理参数不变时,随着喷口尺寸的增大,同一流动模态下的减阻效果下降。本文的研究对超声速钝体减阻技术在工程上的应用具有一定的参考价值。     

Computational study of a supersonic free jet rotating perpendicular to the streamwise direction

We study the flow structure of supersonic jets rotating perpendicular to the streamwise direction using RANS simulations, and we assess the performance of different turbulence model rotation corrections. The Coriolis and centrifugal terms were added to the equations of motion to perform calculations in this non-inertial (rotating) frame of reference. An explicit, cell-centred, finite-volume numerical method, coupled to a k?ε turbulence model, was used for the computations. The turbulence model rotation corrections of Howard et al. (1980), Park and Chung (1999), and Cazalbou et al. (2005) were attempted. In the absence of experimental data for jets rotating perpendicular to the streamwise direction, the rotation corrections were examined against the available measurements of a swirling jet; the comparison of the numerical and experimental data indicates that the Cazalbou et al. (2005 Cazalbou, J.B. 2005. Two-equation modeling of turbulent rotating flows. Physics of Fluids, 17(5): 1–14. [Crossref], [Web of Science ®] , [Google Scholar]) and Park and Chung (1999 Park, J.Y. and Chung, M.K. 1999. A model for the decay of rotating homogeneous turbulence. Physics of Fluids, 11(6): 1544–1549. [Crossref], [Web of Science ®] , [Google Scholar]) corrections improve the performace of the turbulence model. Simulations were then run of a supersonic jet rotating perpendicular to the stream direction at 0, 50, 100 and 150 rad/s, using no turbulence model rotation correction, and using the three rotation corrections. The results indicate that the Cazalbou et al. (2005 Cazalbou, J.B. 2005. Two-equation modeling of turbulent rotating flows. Physics of Fluids, 17(5): 1–14. [Crossref], [Web of Science ®] , [Google Scholar]) correction is more physical than the other two, as it yields results that are qualitatively consistent with the known effects of rotation: that turbulence is enhanced and suppressed on the concave and convex sides of a rotating jet centreline, respectively, and that the effect of rotation saturates as the rotation rate increases. The findings are in qualitative agreement with the available literature.     

Effects of the jet-to-crossflow momentum ratio on a sonic jet into a supersonic crossflow

Numerical investigation of a transverse sonic jet injected into a supersonic crossflow was carried out using large-eddy simulation for a free-stream Mach number M = 1.6 and a Reynolds number Re = 1.38 × 10⁵ based on the jet diameter. Effects of the jet-to-crossflow momentum ratio on various fundamental mechanisms dictating the intricate flow phenomena, including flow structures, turbulent characters and frequency behaviors, have been studied. The complex flow structures and the relevant flow features are discussed to exhibit the evolution of shock structures, vortical structures and jet shear layers. The strength of the bow shock increases and the sizes of the barrel shock and Mach disk also increase with increasing momentum ratio. Turbulent characters are clarified to be closely related to the flow structures. The jet penetration increases with the increase of the momentum ratio. Moreover, the dominant frequencies of the flow structures are obtained using spectral analysis. The results obtained in this letter provide physical insight in understanding the mechanisms relevant to this complex flow.     

Numerical investigations of radial jet reattachment flows

A finite volume computational scheme to solve the Navier-Stokes equations for the laminar flow fields of partially enclosed axial and radial jets impinging on a flat plate has been devised and tested. This scheme is based on the SIMPLEC technique. However, because of the backflow at the ‘outflow’ boundary, the SIMPLEC pressure correction technique has to be modified. The need for this modification, necessitated by the convergence failure, showed the ‘hidden’ pressure boundary condition of SIMPLE-type techniques. Test computations with the present scheme for flows in a channel with a built-in cylinder show that the location of the exit boundary affects very slightly the separated flow behind the cylinder. Computed Squire jet flows compare quite well with the available analytical solution. Finally, impinging radial jets have been computed for different Reynolds numbers. The results show the critical Reynolds number below which a steady solution is obtained and above which periodic and eventually chaotic flows result.     

Flow structure formed by interaction of a single microjet with a supersonic jet flow

The flow structure at the initial section of a supersonic underexpanded jet in the presence of a stationary artificial disturbance in the form of a single microjet is studied experimentally. The influence of gas-dynamic and geometric parameters of the microjet on the structure of the main supersonic flow and a significant effect of the microjet on the changes in the Pitot pressure in the shear layer of the supersonic jets are identified. Interaction between the microjet and the main jet flow generates disturbances of two types propagating in the main jet flow: a disturbance induced by the wake flow behind the microjet and a weaker disturbance in the form of a low-intensity shock wave (Mach wave type). __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 50, No. 3, pp. 104–111, May–June, 2009.     

Numerical simulation of gap effect in supersonic flows

The gap effect is a key factor in the design of the heat sealing in supersonic vehicles subjected to an aerodynamic heat load. Built on S-A turbulence model and Roe discrete format, the aerodynamic environment around a gap on the surface of a supersonic aircraft was simulated by the finite volume method. As the presented results indicate, the gap effect depends not only on the attack angle, but also on the Mach number.     

Three-dimensional effects on regular reflection in steady supersonic flows

The reflection of shock waves between two symmetrical wedges is investigated for the case of three-dimensional flows. Oblique shadowgraphs at various optical angles of yaw and pitch were used to examine the nature of fully three-dimensional flows, with wedge aspect ratios as low as 0.25 being considered. These images were used to construct surface models of the overall flow field for various reflection patterns and aspect ratios, which provides a visual indication of the flow field shape. For a Mach number of 3.1, and suitable wedge angles, the flow field with regular reflection on the tunnel centreline and Mach reflection further out is examined. The point of transition from regular reflection to the peripheral Mach surfaces is identified for various wedge angles and aspect ratios. It is shown that the transition points move outwards from the central plane as the aspect ratio decreases. This shows that three-dimensional flows favor regular reflection, because of the increasing curvature of the incident shock as the wedge becomes narrower, causing a decrease in the local angle of incidence. The height of the Mach stem is shown to be highly dependent on the geometry of the test wedge models. The Mach stem height decreases with aspect ratio due to the three-dimensional relieving effect, where the increase in lateral flow relieves the pressure over the surfaces of the wedges. Experimental evidence of the existence of the strong oblique shock solution in steady flows is presented.Received: 7 July 2003, Revised: 20 October 2003, Accepted: 6 November 2003, Published online: 10 February 2004PACS: 47.40.Nm Correspondence to: B.W. Skews     

Combination interaction of Taylor–Goertler vortices in a curved shear layer of a supersonic jet

Problems of origination and evolution of streamwise vortex structures in an initial region of the shear layer of a supersonic jet are discussed. Streamwise vortices are generated with the use of artificial microroughnesses on the internal surface of polished nozzles. Results of Pitot pressure distributions measured in a supersonic nonisobaric jet both in the radial and azimuthal directions are presented. Streamline curvature in the initial region of supersonic nonisobaric jets has a significant effect on evolution of streamwise vortex structures. Azimuthal heterogeneity corresponding to streamwise vortices in the shear layer is analyzed with the use of both the Fourier analysis and wavelet analysis. PACS 47.40.Ki, 47.20.Ft, 02.30.Nw