Separation control by vortex generator devices in a transonic channel flow

An experimental study was conducted in a transonic channel to control by mechanical vortex generator devices the strong interaction between a shock wave and a separated turbulent boundary layer. Control devices—co-rotating and counter-rotating vane-type vortex generators—were implemented upstream of the shock foot region and tested both on a steady shock wave and on a forced shock oscillation configurations. The spanwise spacing of vortex generator devices along the channel appeared to be an important parameter to control the flow separation region. When the distance between each device is decreased, the vortices merging is more efficient to reduce the separation. Their placement upstream of the shock wave is determinant to ensure that vortices have mixed momentum all spanwise long before they reach the separation line, so as to avoid separation cells. Then, vortex generators slightly reduced the amplitude of the forced shock wave oscillation by delaying the upstream displacement of the leading shock.     

HLLE++格式在高马赫数空腔流动模拟中的应用

空腔流动存在剪切层运动、涡脱落与破裂,以及激波与激波、激波与剪切层、激波与膨胀波和激波/涡/剪切层相互干扰等现象,流动非常复杂,特别是高马赫数(M>2)时,剪切层和激波更强,激波与激波干扰更严重,对数值格式的要求更高,既需要格式耗散小,对分离涡等有很高的模拟精度,又需要格式在激波附近具有较大的耗散,可以很好地捕捉激波,防止非物理解的出现。Roe和HLLC等近似Riemann解格式在高马赫数强激波处可能会出现红玉现象,而HLLE++格式大大改善了这种缺陷,在捕捉高超声速激波时避免了红玉现象的发生,同时还保持在光滑区域的低数值耗散特性。本文在结构网格下HLLE++格式的基础上,通过改进激波探测的求解,建立了基于非结构混合网格的HLLE++计算方法,通过无粘斜坡算例,验证了HLLE++格式模拟高马赫数流动的能力,并应用于高马赫数空腔流动的数值模拟,开展了网格和湍流模型影响研究,验证了方法模拟高马赫数空腔流动的可靠性和有效性。     

变高度圆柱诱导的激波边界层干扰

采用分区方法及Ｒｏｅ三阶流通量差分分裂格式求解雷诺平均Ｎ－Ｓ方程,湍流附加黏性系数用Ｂａｌｄｗｉｎ－Ｌｏｍａｘ模型计算,数值模拟了高超声速条件下变高度圆柱诱导的激波边界层层干扰,其流场的主要特性均与实验结果一致或规律相同,结果清晰地展示了流场结构以及气动载荷分布随柱高度的变化特征,产说明激波碰撞和旋涡运动都可能导致飞行器表面局部气动载荷的增加。     

数值研究平板方舵激波-湍流边界层干扰

数值研究了平板方舵激波-湍流边界层干扰流场。模拟出了分离激波与弓型激波砬撞后形成的“λ”激波结构;消晰地显示了分离区中的旋涡结构,发现流场中会出现二次分离涡,并从理论上分析了流场对称面涡心形态与非定常的关系,得到了涡心为不稳定螺旋点或出现极限环是非定常流动特征的新结论。     

激波与转捩边界层干扰非定常特性数值分析

激波与边界层干扰的非定常问题是高速飞行器气动设计中基础研究内容之一.以往研究主要针对层流和湍流干扰,在分离激波低频振荡及其内在机理方面存在着上游机制和下游机制两类截然不同的理论解释.分析激波与转捩边界层干扰下非定常运动现象有助于进一步加深理解边界层状态以及分离泡结构对低频振荡特性的影响规律,为揭示其产生机理指出新的方向.采用直接数值模拟方法对来流马赫数2.9,24?压缩拐角内激波与转捩边界层干扰下激波的非定常运动特性进行了数值分析.通过在拐角上游平板特定的流向位置添加吹吸扰动激发流动转捩,使得进入拐角的边界层处于转捩初期阶段.在验证了计算程序可靠性的基础上,详细分析了转捩干扰下激波运动的间歇性和振荡特征,着重研究了分离泡展向三维结构对激波振荡特性的影响规律,最后还初步探索了转捩干扰下激波低频振荡产生的物理机制.研究结果表明:分离激波的非定常运动仍存在强间歇性和低频振荡特征,其时间尺度约为上游无干扰区内脉动信号特征尺度的10倍量级;分离泡展向三维结构不会对分离激波的低频振荡特征产生实质影响.依据瞬态脉动流场的低通滤波结果,转捩干扰下激波低频振荡的诱因来源于拐角干扰区下游,与流场中分离泡的收缩/膨胀运动存在一定的关联.     

Effect of air jet vortex generators on a shock wave boundary layer interaction

The effect of upstream injection by means of continuous air jet vortex generators (AJVGs) on a shock wave turbulent boundary layer interaction is experimentally investigated. The baseline interaction is of the impinging type, with a flow deflection angle of 9.5° and a Mach number M _e  = 2.3. Considered are the effects of the AJVGs on the upstream boundary layer flow topology and on the spatial and dynamical characteristics of the interaction. To this aim, Stereoscopic Particle Image Velocimetry has been employed, in addition to hot-wire anemometry (HWA) for the investigation of the unsteady characteristics of the reflected shock. The AJVGs cause a reduction of the separation bubble length and height. In addition, the energetic frequency range of the reflected shock is increased by approximately 50%, which is in qualitative agreement with the smaller separation bubble size.     

Characteristics of unsteady type IV shock/shock interaction

Characteristics of the unsteady type IV shock/shock interaction of hypersonic blunt body flows are investigated by solving the Navier–Stokes equations with high-order numerical methods. The intrinsic relations of flow structures to shear, compression, and heating processes are studied and the physical mechanisms of the unsteady flow evolution are revealed. It is found that the instantaneous surface-heating peak is caused by the fluid in the “hot spot” generated by an oscillating and deforming jet bow shock (JBS) just ahead of the body surface. The features of local shock/boundary layer interaction and vortex/boundary layer interaction are clarified. Based on the analysis of flow evolution, it is identified that the upstream-propagating compression waves are associated with the interaction of the JBS and the shear layers formed by a supersonic impinging jet, and then the interaction of the freestream bow shocks and the compression waves results in entropy and vortical waves propagating to the body surface. Further, the feedback mechanism of the inherent unsteadiness of the flow field is revealed to be related to the impinging jet. A feedback model is proposed to reliably predict the dominant frequency of flow evolution. The results obtained in this study provide physical insight into the understanding of the mechanisms relevant to this complex flow.     

Experimental investigation of three-dimensional supersonic flow of a gas in the interference region of intersecting surfaces

A contemporary high-speed aircraft represents a complex three-dimensional configuration, where supersonic gas flow is accompanied by numerous local flow interaction zones, in particular, near the intersection of different surfaces. Such a flow is characterized by three-dimensional systems of shock and expansion waves, and close to the surfaces one finds interaction of boundary layers and, above all, interaction of shock waves with the boundary layer. In general, the angular configurations are formed by intersection or contact of nonplanar surfaces with swept-back or blunted leading edges. This makes it practically impossible to obtain a rigorous theoretiical solution to the problem of gas flow over these surfaces, and presents considerable difficulty in an experimental investigation. It is therefore of interest to study the physical features of gas flow in corner configurations of very simple form [1–3]. The present paper examines the results of an experimental investigation of typical features of symmetric and asymmetric interaction of compressive, expansive, and mixed flows in the interference region of planar surfaces intersecting at an angle of less than 180?.     

The Multi-point Optimization of Shock Control Bump with Constant-Lift Constraint Enhanced with Suction and Blowing for a Supercritical Airfoil

Both shock control bump (SCB) and suction and blowing are flow control methods used to control the shock wave/boundary layer interaction (SWBLI) in order to reduce the resulting wave drag in transonic flows. A SCB uses a small local surface deformation to reduce the shock-wave strength, while suction decreases the boundary-layer thickness and blowing delays the flow separation. Here a multi-point optimization method under a constant-lift-coefficient constraint is used to find the optimum design of SCB and suction and blowing. These flow control methods are used separately or together on a RAE-2822 supercritical airfoil for a wide range of off-design transonic Mach numbers. The RANS flow equations are solved using Roe’s averages scheme and a gradient-based adjoint algorithm is used to find the optimum location and shape of all devices. It is shown that the simultaneous application of blowing and SCB (hybrid blowing/SCB) improves the average aerodynamic efficiency at off-design conditions by 18.2 % in comparison with the clean airfoil, while this increase is only 16.9 % for the hybrid suction/SCB. We have also studied the SWBLI and how the optimization algorithm makes the flow wave structure and interactions of the shock wave with the boundary layer favorable.     

Time-resolved stereo PIV measurements of shock–boundary layer interaction on a supercritical airfoil

Time-resolved stereo particle-image velocimetry (TR-SPIV) and unsteady pressure measurements are used to analyze the unsteady flow over a supercritical DRA-2303 airfoil in transonic flow. The dynamic shock wave–boundary layer interaction is one of the most essential features of this unsteady flow causing a distinct oscillation of the flow field. Results from wind-tunnel experiments with a variation of the freestream Mach number at Reynolds numbers ranging from 2.55 to 2.79 × 10⁶ are analyzed regarding the origin and nature of the unsteady shock–boundary layer interaction. Therefore, the TR-SPIV results are analyzed for three buffet flows. One flow exhibits a sinusoidal streamwise oscillation of the shock wave only due to an acoustic feedback loop formed by the shock wave and the trailing-edge noise. The other two buffet flows have been intentionally influenced by an artificial acoustic source installed downstream of the test section to investigate the behavior of the interaction to upstream-propagating disturbances generated by a defined source of noise. The results show that such upstream-propagating disturbances could be identified to be responsible for the upstream displacement of the shock wave and that the feedback loop is formed by a pulsating separation of the boundary layer dependent on the shock position and the sound pressure level at the shock position. Thereby, the pulsation of the separation could be determined to be a reaction to the shock motion and not vice versa.     

Vortex sound

Vortex motion is the only source of aerodynamic sound production in low Mach number flow: the unsteady part of the vorticity distribution contributes linearly to the sound field. The following fundamental model flows, which illustrate the vorticity as the predominant sound source in unsteady flows, are discussed: An initially planar elliptic vortex; two identical coaxial initially elliptic vortex rings, where a special case is the leap-frogging of two identical circular rings. For head-on collision of two identical circular vortex rings and for several cases of vortex-body interaction good agreement between theory and experiment exists. If the Mach number is not low, other mechanisms have also to be considered. Here the theory is not yet fully developed. Experimental results for a vortex-airfoil interaction in transonic flow show that local flow separation and boundary layer as well as compressibility effects play a basic role. However, if the motion of vorticity would be known in subsonic flow, essential parts of the sound field could be calculated by the theory. — In addition, it is shown that the general theory is well suited to provide a better understanding of the scattering of sound waves by vortex motion, at least for long wave lengths.     

Formation and evolution of vortex rings induced by interactions between shock waves and a low-density bubble

The deformation and instability of a low-density spherical bubble induced by an incident and its reflected shock waves are studied by using the large eddy simulation method. The computational model is firstly validated by experimental results from the literature and is further used to examine the effect of incident shock wave strength on the formations and three-dimensional evolutions of the vortex rings. For the weak shock wave case (Ma?=?1.24), the baroclinic effect induced by the reflected shock wave is the key mechanism for the formation of new vortex rings. The vortex rings not only move due to the self-induced effect and the flow field velocity, but also generate azimuthal instability due to the pressure disturbance. For the strong shock wave case (Ma?=?2.2), a boundary layer is formed adjacent to the end wall owing to the approach of vortex ring, and unsteady separation of the boundary layer near the wall results in the ejection and formation of new vortex rings. These vortex rings interact in the vicinity of the end wall and finally collapse to a complicated vortex structure via azimuthal instability. For both shock wave strength cases, the evolutions of vortex rings due to the instability lead to the formation of the complicated structure dominated by the small-scale streamwise vortices.     

Inviscid vortex structures in the shock layers of conical flows around V wings

The applicability of the criteria of existence of inviscid vortex structures (vortex Ferri singularities) is studied in the case in which a contact discontinuity of the corresponding intensity proceeds from the branching point of the λ shock wave configuration accompanying turbulent boundary layer separation under the action of an inner shock incident on the leeward wing panel. The calculated and experimental data are analyzed, in particular, those obtained using the special shadow technique developed for visualizing supersonic conical streams in nonsymmetric, Mach number 3 flow around a wing with zero sweep of the leading edges and the vee angle of 2π /3. The applicability of the criteria of existence of inviscid vortex structures is established for contact discontinuities generated by the λ shock wave configuration accompanying turbulent boundary layer separation realized under the action of a shock wave incident on the leeward wing panel. Thus, it is established that the formation of the vortex Ferri singularities in a shock layer is independent of the reason for the existence of the contact discontinuity and depends only on its intensity.     

Three-dimensional centrifugal instabilities development inside a parallelepipedic open cavity of various shape

This experimental study reports flow developments inside a parallelepipedic cavity of variable shape and dimensions. That flow is generated by the interaction between a laminar boundary layer and the cavity, which creates shear-layer oscillations. The aim is to understand the three-dimensional flow morphology varying the Reynolds number and the cavity shape. Flow visualizations are obtained in a plane situated inside the cavity in order to get the dynamical structures. Dimensional analysis of the cavity flow teaches that three dimensionless numbers are necessary for the flow reduction. This is confirmed by experimental results pointing thresholds of appearance of instabilities identified for some combinations of Reynolds number and geometric parameters. The key mechanisms for their existence are centrifugal effects induced by a vortex of spanwise axis with sufficient intensity, and viscous effects due to the wall confinement of the cavity. Their destruction is linked to flow transition to turbulence above a limiting convective velocity generated by the vortex of spanwise axis. These instabilities are generally present in a spanwise row of counter-rotating pairs of vortices, but for some cases, isolated pairs are also identified. Secondary modulations of primary instabilities are also present for particular parameters. Results permit to discriminate the relevant scales associated with the shear layer and the inner cavity flow.     

Method of mean-mass parameters for a three-dimensional boundary layer in a flow with vorticity

When a gas flows with hypersonic velocity over a slender blunt body, the bow shock induces large entropy gradients and vorticity near the wall in the disturbed flow region (in the high-entropy layer) [1]. The boundary layer on the body develops in an essentially inhomogeneous inviscid flow, so that it is necessary to take into account the difference between the values of the gas parameters on the outer edge of the boundary layer and their values on the wall in the inviscid flow. This vortex interaction is usually accompanied by a growth in the frictional stress and heat flux at the wall [2, 3]. In three-dimensional flows in which the spreading of the gas on the windward sections of the body causes the high-entropy layer to become narrower, the vortex interaction can be expected to be particularly important. The first investigations in this direction [4–6] studied the attachment lines of a three-dimensional boundary layer. The method proposed in the present paper for calculating the heat transfer generalizes the approach realized in [5] for the attachment lines and makes it possible to take into account this effect on the complete surface of a blunt body for three-dimensional laminar, transition, or turbulent flow regime in the boundary layer.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 1, pp. 80–87, January–February, 1981.     

Shock boundary layer interaction under the influence of a normal suction slot

The phenomenon of shock boundary layer interaction of a shock train under the influence of a normal suction slot is studied. In previous work, it was found that a normal, circumferential suction slot is sufficient to stabilize the primary shock of a shock train in as much as that the back pressure of the shock train can be increased until the shock train gradually changes into a single normal shock. Based on the experimental and numerical results, a flow model was derived which explains the transition of a shock train into a single shock under the influence of boundary layer suction. In this work, the normal shock boundary layer interaction model is validated against flow cases with different upstream Mach and Reynolds numbers. For that purpose three different nozzle flows are investigated at various total pressure levels. In a second step, the flow model is extended to the oblique shock case, correlating the suction mass flow with the total pressure distribution of the incoming boundary layer and the static pressure downstream of the oblique shock. Finally, the influence of the suction cavity pressure onto the shock boundary layer interaction is considered.     

Strong vortex/boundary layer interactions

This is the second of two papers on the interaction between a longitudinal vortex pair, produced by a delta-wing at angle of attack, and a turbulent boundary layer developing on a flat plate. In the first paper only the outer parts of the vortices entered the boundary layer whereas in this paper the vortices merge with it. In the resultant interaction, the boundary layer between the vortices is kept thin by lateral divergence and a three-dimensional separation line is formed outboard of each vortex. Turbulent, momentum-deficient fluid containing longitudinal vorticity is entrained from the boundary layer along these lines and wrapped around the vortices. As a consequence, the turbulent region of the vortices increases in size and the circulation slowly decreases. It is shown that the flow near the separation line and in the vortices is complicated, and this interaction is expected to be more difficult to calculate than the first. Detailed mean flow and turbulence measurements are reported.     

Stability and modal analysis of shock/boundary layer interactions

The dynamics of oblique shock wave/turbulent boundary layer interactions is analyzed by mining a large-eddy simulation (LES) database for various strengths of the incoming shock. The flow dynamics is first analyzed by means of dynamic mode decomposition (DMD), which highlights the simultaneous occurrence of two types of flow modes, namely a low-frequency type associated with breathing motion of the separation bubble, accompanied by flapping motion of the reflected shock, and a high-frequency type associated with the propagation of instability waves past the interaction zone. Global linear stability analysis performed on the mean LES flow fields yields a single unstable zero-frequency mode, plus a variety of marginally stable low-frequency modes whose stability margin decreases with the strength of the interaction. The least stable linear modes are grouped into two classes, one of which bears striking resemblance to the breathing mode recovered from DMD and another class associated with revolving motion within the separation bubble. The results of the modal and linear stability analysis support the notion that low-frequency dynamics is intrinsic to the interaction zone, but some continuous forcing from the upstream boundary layer may be required to keep the system near a limit cycle. This can be modeled as a weakly damped oscillator with forcing, as in the early empirical model by Plotkin (AIAA J 13:1036–1040, 1975).