Bifurcation of a transonic ideal-gas flow around a symmetric airfoil with a concavity

A transonic flow around a symmetric airfoil with a concavity in its central part has been numerically investigated. The dependences of the lift coefficient on the Mach number of the incident flow M_∞ and on the angle of attack α were determined. It is shown that, depending on M_∞, in the flow past the upper part of the indicated airfoil there can arise one or two supersonic-flow regions. It has been established that, at fairly large angles of attack, the coalescence and separation of supersonic-flow regions are realized in a discrete way. For these angles of attack, singular Mach numbers M_s, in the neighborhood of which the structure of the flow is transformed, were determined and the physical processes occurring in this case were analyzed. It was found that the flow being considered is characterized by a large hysteresis in M_∞. __________ Translated from Inzhenerno-Fizicheskii Zhurnal, Vol. 80, No. 4, pp. 63–68, July–August, 2007.     

Control of the aerodynamic characteristics of wing airfoils by nonstationary energy supply in transonic flow

The possibility of controlling the aerodynamic characteristics of wing airfoils in transonic regimes of flight using one-sided pulse-periodic energy supply has been studied. The flow structure near the symmetric airfoil at different angles of attack and its aerodynamic characteristics as functions of the value of energy in its nonsymmetric (about the airfoil) supply have been determined by numerical solution of two-dimensional nonstationary gasdynamic equations. A comparison of the obtained results and the data of calculation of flow past the airfoil at different angles of attack without energy supply has been made. It has been established that a prescribed lift can be obtained, using energy supply, with a much higher fineness ratio of the airfoil than that in the case of flow past it at an angle of attack. Translated from Inzhenerno-Fizicheskii Zhurnal, Vol. 82, No. 1, pp. 18–22, January–February, 2009.     

基于微秒脉冲激励的飞翼模型等离子体流动控制试验研究

为了改善飞翼布局的大迎角气动特性,采用飞翼全模和半模分别在低速和跨声速风洞中开展了微秒脉冲介质阻挡放电等离子体流动控制的试验研究。通过流动显示和测力的试验方法研究了等离子体流动控制的主要作用机制和激励频率与激励电压等对飞翼模型失速特性的影响规律,验证了微秒脉冲介质阻挡放电等离子体流动控制技术从低速到亚声速的有效性,有效的试验最高马赫数Ma达到0.6、雷诺数Re达到3.05×10⁶。试验研究表明：微秒脉冲介质阻挡放电等离子体通过非定常微尺度压缩波扰动的形式作用于翼面流场,通过频率耦合机制减弱模型的前缘分离涡、抑制翼面的流动分离;无量纲频率F⁺是影响等离子体流动控制效果的重要参数;在低速风洞试验风速V=30 m/s时,无量纲频率F⁺=0.35～1.06的控制效果较好,可将模型的最大升力系数提高25%以上、失速迎角推迟4°;在跨声速风洞试验马赫数Ma=0.6时,无量纲频率F⁺=0.22和F⁺=0.44的控制效果较好,可将模型的最大升力系数分别提高4.72%、4.77%,失速迎角分别推...     

翼型绕流分离的微楔控制

基于大涡模拟(LES, Large Eddy Simulation),结合3阶WENO(Weighted Essentially Non-Oscillatory)格式与沉浸边界法(IBM, Immersed Boundary Method),对NACA0012翼型在来流马赫数为Ma=0.5以及10°攻角下,有无微型三角楔涡流发生器作用的三维绕流进行了数值模拟,对比了两者的流场结构变化。数值结果表明,微楔能显著改变翼型绕流的尾涡结构,并能缩小其表面边界层的分离,提高升力,并对微型三角楔涡流发生器的流体分离抑制机理进行了讨论。     

NACA0012翼型抖振现象实验研究

通过在风洞侧壁安装动态压力传感器测量侧壁脉动压力的方法针对NACA0012翼型开展了零攻角状态下跨声速抖振特性的实验研究,并测量了翼型沿弦向与展向的压力分布。从时域和频域两个方面分析了脉动压力结果,得到了NACA0012翼型的跨声速抖振起始马赫数和抖振频率等抖振特性。研究结果表明:风洞二维性较好,侧壁区域的压力情况能够反映核心流压力变化情况;抖振发生在实验马赫数Ma_exp = 0.88~0.89范围内;随着马赫数的增大,抖振频率有所增加。     

振荡压气机叶栅叶片表面非定常响应以及气弹稳定性分析

发展并验证了一种适用于叶轮机内部非定常跨音流动诱导的叶片气弹问题的高效、准确的数值模拟方法。采用有限体积的多块结构化网格形式,多重网格方法加速收敛,隐式的双时间步时间推进,Spalart-Allmaras（S-A）湍流模型求解非定常雷诺平均Navier-Stokes方程。通过气动弹性标准算例10,叶片在高亚音和跨音流动下做弯曲振动,分析了流动状态、折合频率以及叶片间相位角对叶片表面非定常气动力响应以及叶栅气弹稳定性的影响。分析结果表明激波在此跨音振荡压气机叶栅中起失稳作用,叶片间相位角对气弹稳定性的影响在高折合频率下被加强。     

Effects of an airfoil and shock waves on vortex shedding process behind a square cylinder

Summary Effects of an airfoil and shock waves on vortex shedding process behind a square cylinder have been examined experimentally at a Mach number of about 0.91 and at a Reynolds number (based on the side lengthD of the square cylinder) of about 4.2×10⁵. The main experimental parameter is the spacing ratioL/D, and is varied from 1.125 to 5.5, whereL is the spacing between the square cylinder and the airfoil.It is found that similarly to the case at subcritical Mach numbers at the supercritical Mach number there exist three patterns of the flow around the square cylinder and airfoil arranged in tandem depending upon the spacing ratioL/D: In the first flow pattern with small spacing ratio, the downstream airfoil is enclosed completely in the vortex formation region of the square cylinder. In the second flow pattern, the shear layers separating from the square cylinder reattach to the airfoil. In the third flow pattern with large spacing the shear layers roll up upstream of the airfoil. The Strouhal number at the supercritical Mach number is higher than that at the subcritical Mach numbers. Shock waves hasten the vortex shedding behind the square cylinder by decreasing the area of asymmetrical part of the vortex formation region with respect to the wake axis, and let the streamwise length of the separating shear layers longer than otherwise.With 8 Figures     

弹性特征对跨声速抖振边界的影响研究

跨声速抖振直接影响飞机结构强度和疲劳寿命,易引发飞行事故。对于刚性静止翼型,其跨声速抖振特性仅与来流状态参数（马赫数、迎角和雷诺数）有关。该文通过释放翼型的俯仰自由度,采用非定常雷诺平均N-S方程建立了流场和结构的耦合求解方法,研究翼型结构弹性特征对跨声速抖振始发边界的影响。研究发现,结构固有频率对弹性支撑翼型的抖振边界影响较大,系统会在更小的来流迎角下失稳;而质量比对抖振边界影响较小。弹性支撑翼型的抖振始发迎角较刚性静止翼型的抖振起始迎角降低了近1°,因此翼型的弹性特征是跨声速抖振研究的重要因素。     

Accelerated and robust finite volume Navier-Stokes solver for all speeds

The paper describes the results obtained from a multigrid accelerated Navier-Stokes solver. The method is based on 2-D explicit cell-centred finite volume Reynolds averaged Navier-Stokes (N-S) flow solver for speeds from near-incompressible Mach numbers to high hypersonic Mach numbers including flows at high angles of attack. The time integration is done using a hybrid 5-stage Runge-Kutta local time stepping scheme. With the help of a simple technique, the capability of the Jameson-Schmidt-Turkel numerical dissipation scheme has been enhanced to include hypersonic flows. The iterative procedure is accelerated significantly by incorporating a multigrid technique which has been used in all computations up to about supersonic speeds. Systematic numerical experiments were conducted to evolve guidelines to generate airfoil grid that could offer reliable flow simulations. The computed results are in very good agreement with experimental data where available, especially from the point of view of predicting large suction peaks and shock locations where considerable departures are often seen in the literature. Further, the highly accelerated computations make this code a useful tool of practical interest in preliminary aerodynamic design.     

基于当地流活塞理论的气动弹性稳定性分析方法研究

基于改进的当地流活塞理论,推导了用模态坐标表示的弹性振动翼面的非定常广义气动力表达式。通过与非定常Euler方程的比较,证明该气动力模型具有较高的数值精度,放宽了原始活塞理论对翼型厚度、迎角、马赫数的限制。再耦合两自由度结构运动方程,在时域和频域内实现了超音速、高超音速气动弹性的稳定性分析。应用算例计算了一系列二元机翼和三维舵面的气动弹性特性。通过与耦合非定常Euler方法或实验结果的比较,证明该方法在保证精度的同时大大提高了超音速气动弹性的计算效率,在颤振分析中有较高的工程实用价值。     

飞机T型尾翼跨音速颤振特性研究

由于飞机T型尾翼的结构与气动布局特点,T型尾翼颤振计算不能套用常规尾翼的分析方法,而需要考虑平尾面内运动以及静升力等因素的影响。而跨音速空气压缩性效应和非定常气动力计算的不准确性,使得T型尾翼跨音速颤振计算更加困难,准确性较低。因此,需要采用试验为主计算为辅的方法来研究飞机T型尾翼跨音速颤振特性。针对某T型尾翼结构,用ZAERO软件等价片条势流跨音速颤振(ZTAIC)方法计算T型尾翼跨音速颤振特性,研究了马赫数、风洞气流密度和平尾迎角对T型尾翼颤振特性的影响。通过升力系数斜率空气压缩性修正计算方法和跨音速颤振模型风洞试验方法得到了飞机T型尾翼的跨音速颤振的凹坑曲线和空气压缩性特性,两种方法得到结果一致。     

基于动态近似边界条件的气动弹性数值模拟

发展了一种利用欧拉方程计算非定常气动力的数值方法,通过在固定物面边界上满足动态近似边界条件计算出非定常气动力,避免了在每个时间步重新生成网格或需用动网格技术进行网格变形处理过程,提高了计算效率。运用这种方法计算了一系列非定常气动力算例,并与非结构动网格准确边界条件下的欧拉方程解和实验数据进行了比较,进一步分析了翼型俯仰角和马赫数对非定常气动力相对误差的影响。将气动力解算器与结构方程耦合进行气动弹性数值模拟,计算了跨音速具有S型颤振边界的二元气动弹性标准算例-Isogaiwing。算例结果表明,利用动态近似边界条件的欧拉方程具有简便、高效的特点,并能在小振幅情况下得到与精确边界条件精度相当的非定常流场解,还可以用于气动弹性分析。     

Influence of the Curvature of an Airfoil on the Structure of Transonic Flow

A study is made of the dependence of the pattern of transonic flow about an airfoil on the change in its shape and in the parameters of the incoming air flow. Consideration is given to a family of airfoils in a channel modeling the working section of a wind tunnel. The formation and propagation of weak shock waves in local supersonic regions near the central part of the airfoil has been investigated numerically. The finite-difference method based on a substantially nonoscillatory scheme of second order of accuracy has been employed to solve the Euler equations describing nonviscous gas flow. It has been shown that the sensitivity of the flow pattern to a change in the Mach number of the free flow substantially increases with decrease in the curvature of the airfoil.     

Microtab对翼型跨音速抖振始发特性的影响

研究了microtab对翼型跨音速抖振始发特性的影响。采用SST湍流模型对RANS方程进行封闭,以NACA0012翼型为对象,用CFD方法计算了在其后缘附近安装microtab后,翼型的跨音速流场。对比分析了在翼型后缘附近的上表面、下表面单独安装和上下表面对称安装microtab后,对翼型跨音速抖振始发特性的影响,并研究了microtab的高度和弦向安装位置对跨音速抖振边界的控制作用。结果表明：在翼型后缘附近的上表面加装microtab能够提高抖振始发攻角,并且随着microtab高度的增加,抖振始发攻角有增大的趋势。在翼型后缘附近的下表面加装microtab能够提高抖振始发时的升力系数,且抖振始发时的升力系数随着microtab高度的增加而增大。microtab的弦向安装位置的变化对于抖振始发攻角影响较小,但是对抖振始发时的升力系数有明显的影响。         

某跨音速高压涡轮非定常流动数值模拟研究

通过求解三维非定常雷诺平均N-S方程模拟某跨音速高压涡轮非定常流场,研究涡轮内非定常流动特征。通过对静子尾迹及静子尾缘激波和转子叶排之间的相互干涉过程进行详细分析,发现定常/非定常模拟方法获得的涡轮总体性能参数基本一致但流场存在较大差异。静子尾迹是导致涡轮流场非定常性的重要因素之一:在转子叶栅通道中部和下部,静子尾迹和转子叶片附面层及下通道涡发生明显干涉,并导致通道中下部损失周期性波动幅度较大,此外尾迹和下通道涡间的干涉作用在转子尾缘处诱导出高频脱落涡。静子尾缘激波也是导致涡轮流场非定常性的原因之一,激波和转子叶片作用形成复杂的波系结构,对涡轮流场影响显著:一方面激波/附面层干涉导致转子和静子的吸力面产生周期性变化的高温区域;另一方面激波撞击叶片导致叶片表面的气流在激波后出现分离,对转子静压分布产生影响,使得转子叶片表面载荷出现明显的非定常性,进而导致涡轮输出功的周期性波动十分剧烈。     

非线性伺服气动弹性的时域数值模拟

运用非定常流场求解器 ,直接求解翼面作任意运动的时域非定常气动力。将拉氏域内的伺服传递函数转化成时域内的伺服状态方程。通过时域气动力、结构状态方程和伺服状态方程的耦合求解 ,实现伺服气动弹性的时域数值模拟。可以方便地模拟如跨音速、大迎角及高超音速等带有不同非线性的伺服气动弹性响应。通过数值模拟发现 ,伺服系统会降低原有气动弹性系统的稳定性 ,结构陷波器的引入可以减弱伺服系统和结构模态的耦合。应用算例给出了大迎角下翼面的极限环响应和 M=7时伺服颤振速度随迎角变化的非线性特性。     

Controlling aerodynamic characteristics of airfoils by means of periodic pulsed energy supply

The possibility of controlling the aerodynamic characteristics of lifting airfoils by means of local periodic pulsed energy supply in transonic flight regimes has been studied. Using a numerical solution of two-dimensional nonstationary equations of gasdynamics, a change in the flow structure near a symmetric airfoil is determined as dependent on the amount of energy supplied to the lower side of the airfoil in a range of attack angles.     

Electric discharge control of flow separation on oblique airfoil

The possibility of controlling flow separation on an oblique airfoil using dielectric-barrier discharge has been experimentally studied. The experiments were performed at subsonic flow velocities in a broad range of the angle of attack. The results of measurements of the velocity and surface pressure fields and an analysis of the flow patterns show that the application of electric discharge allows the interval of the angles of attack for separation-free flow past the airfoil to be significantly increased. Various discharge regimes have been studied, including those with continuous activation by single voltage pulses with a frequency of 0.5–5 kHz and by pulse trains at a repetition rate of 1–100 Hz. The efficiency of the flow separation control has been studied as dependent on the electrical parameters, frequency characteristics, and position of the discharge relative to the flow separation line.     

Numerical simulation of unsteady cavitating flows using a homogenous equilibrium model

 Numerical simulations of two-dimensional cavity flows around a flat plate normal to flow and flows through a 90^∘ bent duct are performed to clarify unsteady behavior under various cavitation conditions. A numerical method applying a TVD-MacCormack scheme with a cavitation model based on a homogenous equilibrium model of compressible gas-liquid two-phase media proposed by the present authors, is applied to solve the cavitating flow. This method permits the simple treatment of the whole gas-liquid two-phase flow field including wave propagation and large interface deformation. Numerical results including detailed observations of unsteady cavity flows and comparisons of predicted results with experimental data are provided. Received: 5 August 2002 / Accepted: 6 January 2003     

Numerical investigation of supersonic flow past blunt bodies of intricate shape at an angle of attack and slip angle

A supersonic flow of viscous homogeneous gas past blunt bodies of intricate shape at an angle of attack and slip angle is investigated numerically within the model of complete three-dimensional viscous shock layer using the time relaxation method. The main regularities are studied of the general structure of flow and of the distribution of pressure and heat flux along the surface. An analysis is performed of their dependence on the shape of the body, angle of attack and slip angle, Mach and Reynolds numbers, and on other determining parameters of the problem. The accuracy and range of validity of a number of approximate approaches to the solution of the problem are estimated.