The flow of the supersonic jets contains shock waves, vortices, turbulence and acoustic waves. The numerical simulation method and the mechanism of the shock associated noise of the supersonic jets have been topics of general interest. This paper contains two parts. Firstly, we briefly review the numerical studies on the fundamental problems of the shock associated noise of the supersonic jets. It includes the numerical methods for the shock associated noise, and the models of the supersonic jets, the axisymmetric and three dimensional supersonic jets. For the numerical method, we introduce the technique to reduce the non-physical oscillation and a criterion to design the smoothness indicator for high order shock capturing schemes. Secondly, we introduce our recent results based on the direct numerical simulations and experi-mental verification, including the localization of the axisymmetric modes, the development of trapped waves and the evolution of the flapping modes.
Inspired by the silent flight capability of owls, the serrated trailing edge design is considered as an effective method to reduce the turbulent boundary layer-trailing edge interference noise. In this study, the near-field flow and noise characteristics of a NACA 0012 airfoil with additional serrated trailing edges are investigated in detail using an implicit large eddy simulation approach with Reynolds number Re = $9.6 \times {10^4}$, far-field Mach number Ma = 0.1631, and angle of attack $\alpha = {4^ \circ }$. The simulation adopts unstructured grids with 70 million degrees of freedom. In this particular calculation, a small sawtooth-shaped rough strip is added to the airfoil surface to facilitate the fast transition to turbulence for both straight and serrated trailing edge cases. At an angle of attack of 4°, an increase in noise radiation is observed with respect to that at an angle of attack of 0°, with a deflection of the primary radiation direction and a noise reduction of about 2.5 dB in this direction. The flow analysis shows that the sawtooth induces the regularly distributed vortex pair structures at its sides, which facilitates noise reduction in the far-field. The analysis of the wall pressure fluctuation shows that the sawtooth mainly changes the space-time correlation properties near the trailing edges, and the space-time correlation properties of the pressure cannot be described by the existing velocity-based Taylor or elliptical correlation models. In addition, the sawtooth suppresses the noise radiation while causing some loss to the aerodynamic performance of the airfoil.
The spinning mode synthesizer is of great significance to the study of the propagation and radiation of aerodynamic noise in the cylindrical duct with flow as well as the evaluation of noise reduction with sound liners, working by controlling the array of loudspeakers and generating the specific distribution of the acoustical mode inside the duct. High-order circumferential modes, radial modes and their mixtures should be excited accurately with the manipulation of the multiple-rings array of loudspeakers. However, due to the systematic error of the loudspeaker, significant interference modes are generated simultaneously with the target mode, remarkably affecting the accuracy of mode excitation. A mode excitation method based on the least square and global calibration is proposed to motivate the target duct mode inside a flow duct and eliminate the influence of loudspeakers' system error on mode synthesizing. Through the global modeling of the system error of the loudspeaker sound source, the global calibration factor and flow field correction factor are introduced to convert the problem of solving the system error of each loudspeaker in the case of the flow field into the mode identification problem of a single loudspeaker in the case of no flow. The complex calibration factor of each loudspeaker is solved by matrix transformation and the least square method, after which the amplitude and phase excitation correction of the loudspeaker is realized. The presented method is applied to the spinning mode synthesizer in SJTU, whose experimental results indicate that the intensity of the interference modes is significantly suppressed, and the modal coefficient signal-to-noise ratio (SNRA) of the target excitation modes is greater than 15 dB within the operating frequency range.
Ffowcs Williams–Hawkings (FW–H) equation is the extension of the Lighthill’s acoustic analogy equation for sound prediction with moving boundaries. However, the spurious sound often arises from vortex structures crossing through permeable FW–H surfaces. This work aims to approximate the contribution of the vortex structures to far-field sound using the Lighthill stress tensor flux and subtract the resulting spurious sound. Based on the frequency-domain Lighthill stress tensor quadrupole correction model, a quadrupole correction model is proposed to account for the effect of a moving integral surface on the spurious sound. Based on the frozen flow assumption and far-field approximation of the FW–H equation’s Green’s function, the proposed model incorporates the FW–H surface’s velocity into the integrand of the quadru-pole correction model by solving an algebraic equation of the quadrupole volume integral term. The proposed model is validated by the far-field sound prediction of flows over a circular cylinder and a convecting vortex.
The acoustic liners produce vibroacoustic response under the excitation of sound waves at high sound pressure level, and the rigid structure assumption is no longer applied. Their structural vibrations have a certain impact on sound absorption performance. The work presented here is an experimental study on the influence of panel vibrations on sound absorption and vibroacoustic response, and the influence law of vibroacoustic response on acoustic impedance under different perforated plate geometric parameters is obtained through parametric research. The experimental results show that the vibration of the perforated plate causes resistance to the generation of peaks or dips at the structural resonance frequency, and the sound absorption coefficient generates extra absorption peaks or dips that cannot be understood assuming rigid acoustic liners. The increase of the perforation rate and sound pressure level suppresses the influence of vibration, and there is a critical perforation rate. Perforated plate parameters affect the characteristics of resistance changes caused by structural vibration at high sound pressure levels. The phase difference between the small holes and the panel near the structure resonance frequency changes abruptly, resulting in an increase in the relative velocity and a change in the sound absorption performance.
Taking the NACA65(12)–10 blade as the object, a linear microphone array based on the SODIX (SOurce DIrectivity modeling in the cross-spectral matriX) method is used to study the leading-edge (LE) noise directivity of the baseline and the effect of the wavy LE on the LE noise directivity. First, a SODIX data processing program was developed, and the program was validated by numerical simulation. The validation results show that the data processing program has a good accuracy with an error less than 0.26 dB. Then, a linear array with 31 microphones is designed to identify the LE noise directivity of the baseline and the wavy LE blade experimentally in a semi-anechoic chamber. Within the measured degree range of 40°–142°, the directivity of LE noise shows a characteristic of typical dipole sound sources with a peak occurs at 130°. Besides, the higher the frequency is, the more obvious of the ‘lobe’ distribution of the LE noise directivity is. Further analysis shows that the wavy LE with various amplitudes and wavelengths especially with larger amplitudes can reduce LE noise in measured angle ranges especially among 90°–120°. And the maximum value is 7.71 dB for A30W20.
粒子图像测速技术目前已经发展成为实验流体力学领域应用最广泛的非接触激光测试方法之一,为认知复杂流动机理提供直观的流场信息.本文基于超声速流场PIV技术研究实践,针对示踪粒子布撒器设计、粒子松弛特性模型构建、激波流场测试分析、超声速平板湍流边界层结构分析等方面具体问题的研究和认识,从理论、定量化的角度深入分析了应用于超声速流场PIV技术现阶段依然存在的问题.从应用于超声速流场PIV技术的原理出发,针对高速复杂流场的PIV测试现状,总结了应用于超声速流场PIV技术发展过程中的光学部件、示踪粒子及布撒系统所遇到的一系列挑战,以及国内外利用PIV技术在高速复杂流场研究中所取得的成就,针对PIV技术能否适用于高超声速流场的测量做了系统化地探索.并根据实践经验提出了应用于超声速流场PIV技术未来的发展方向:通用的精确的PIV方法不存在,必须从具体研究的流动机理角度改造相应的PIV测试手段.
高速列车进入隧道时,会产生压缩波,压缩波沿隧道内传播至隧道端口后形成向外辐射的微气压波。本文介绍了采用动模型实验平台在200~350km/h速度范围内对60m双向隧道模型的隧道壁面压力波和出口微气压波开展的实验研究。首先分析了实验数据的有效性;其次给出了初始压缩波最大值随时间的衰减变化规律和微气压幅值随实验速度的变化特性;最后研究了流线形头型长度对微气压波幅值的影响。实验结果表明:在实验速度范围内,隧道压力波和出口微气压波无量纲值保持一致,但隧道出口微气压波与流线型头型长度只能定性描述,定量关系难以确定。
在节段模型风洞试验中,两端设置端板可以有效减小端部效应对风压分布的影响,从而保证气流在模型周围的二维流动,其中端板尺寸是影响端板效果的主要参数。为了明确不同尺寸端板对矩形断面气动特性的影响,以桥梁节段模型中最常见的3种宽高比(B/H分别为1、5和10)的二维矩形断面为研究对象,通过刚性模型测压试验,研究了端板尺寸对各模型的气动力、风压分布和斯托罗哈数St的影响。研究结果表明:模型的端部效应不仅仅对端部附近的风压有影响,对中间位置处风压的影响也不容忽视,设置端板是获得准确试验结果的重要保证;随着断面宽高比(B/H)逐渐增大,端部效应影响的程度和范围逐渐减小;随着端板尺寸的增大,模型背风面风压绝对值逐渐增大并趋向一稳定值;抑制端部效应的最小端板尺寸与结构的风迎角有关,风迎角增大,所需的端板也相应增大;有无端板对斯托罗哈数St也有明显影响。
流体推力矢量技术不采用机械偏转,以流动控制方式实现推力转向,有望成为一种更加高效的推力矢量控制方法。目前实现流体推力矢量的主要方法有激波矢量法、双喉道方法、逆流控制方法和同向流方法等,对以上方法选择具有共性的计算与试验数据,对喷管的推力矢量效率、推力损失和流量系数进行了对比分析。结果表明激波矢量方法、双喉道方法和逆流方法能够在大落压比范围内(NPR=1.89~10)实现推力矢量控制,并且具有俯仰/偏航耦合甚至多轴控制的潜力。相比激波矢量法和逆流方法,双喉道和同向流方法在减少推力损失和提高矢量效率上占有优势,不足之处是双喉道方法对喉道进行控制限制了流量系数,而同向流方法的适用落压比范围受到严重限制。为寻求更加高效的矢量喷管技术,国内外相继发展了多种新概念流体推力矢量方法,对每种方法的控制原理、潜在优势和存在的问题挑战进行了探讨,新方法着眼于从喷流出口下游进行控制,对主流的干扰很小,值得深入研究,同时也为流体推力矢量的下一步研究方向提供了借鉴参考。
通过刚性模型测压风洞试验,研究了圆柱的气动阻力、气动升力系数和风压系数随雷诺数的变化规律,从流场分布的角度分析了气动力变化的原因,并研究了雷诺数影响下的流场在圆柱轴向的相关性。结果表明:在亚临界雷诺数区域,在时间平均上流场沿模型两侧呈对称分布,雷诺数对平均阻力系数和流场影响较小,平均升力系数基本为零。在临界雷诺数区域,随着特定区域大负压区的出现,流场不再对称,出现不容忽视的平均升力和脉动升力。在超临界雷诺数区域,随着对称侧大负压区的出现,流场恢复对称状态,平均升力基本消失。雷诺数对流场的轴向相关性有显著的影响。在雷诺数较低时(亚临界区域),卡门涡在轴向上的尺度相对较大,而随着雷诺数的提高,该尺度逐渐减小,各断面流场的相关性降低。
活塞式内燃发动机是现代工业中应用最为广泛的动力机械装置。由于其内部燃料喷射、蒸发、燃烧等复杂的工作过程会对发动机的结构可靠性、能量利用效率和污染物生成产生极大影响,研究内部过程的物理机理并确定控制策略对于发动机的设计和改进具有重要的科学意义和实用价值。近年来,为更加深入理解发动机内部工作过程,研究人员广泛采用光学诊断试验技术来测量发动机缸内流动和燃烧特性。本文首先介绍了各类用于模拟发动机工作过程的试验台架(如定容燃烧弹、快速压缩机、光学发动机等)。在此基础上,分析了各类光学诊断技术的基本原理及其在发动机研究中的应用。光学诊断技术分为两类进行讨论,分别是基于传统光学的传统诊断技术(如纹影法、双色法等)和基于激光的先进诊断技术(如粒子图像测速法、激光诱导荧光法等)。光学诊断技术可在多尺度下测量缸内温度、物质浓度、液滴粒径等参数,为准确评估发动机喷油、蒸发、燃烧过程提供试验依据。更重要的是,光学诊断技术为更加深入理解高温高压环境下流动、燃烧的物理/化学机理提供了可能性,为开发高功率、高能效、低排放的先进发动机提供可靠的试验手段,同时为研究人员未来开展基础试验研究、更加深入地理解发动机工作过程提供指导。
地面风洞试验和飞行试验是研究高超声速飞行器气动加热的主要手段。针对临近空间复杂气动外形高超声速飞行器气动热环境研究的需要,分析探讨了国内气动热试验及测量技术的发展情况。分析了临近空间高超声速飞行器外形特征以及飞行剖面、边界层转捩和气动热环境特性等,进而分析了气动热环境风洞试验模拟理论,介绍了适用于气动热研究的风洞试验设备及其模拟能力,重点讨论了适用于不同类型风洞的热流测量技术发展近况、存在的问题和发展趋势;在以长时间、高热流、高壁温为主要特征的高超声速飞行试验中,无法应用风洞环境下的热流测量技术,因而介绍了目前飞行试验中采用的气动热测量技术,讨论了根据结构温度反辨识表面热流存在的问题,以及热流传感器表面的"冷点效应"、表面催化特性等因素对飞行试验气动热测量的影响,提出了后续工作中应重点研究和解决的临近空间飞行器气动热环境测量技术问题。
被动式燃烧诊断技术是利用火焰自发射辐射信息进行燃烧诊断的一项技术,具有非接触、对环境要求不高、系统紧凑、易于实施等特点,在燃烧场在线测量及诊断中具有独特优势。首先,分析了各类燃烧诊断技术的优势及局限;其次,结合华中科技大学煤燃烧国家重点实验室开展的被动式燃烧测量诊断研究工作,从火焰发射光谱、火焰图像处理、热辐射成像技术三个方面介绍了自发辐射燃烧诊断技术的基本原理及研究现状,利用这三种技术,可实现燃烧状态定性分析以及燃烧流场中温度、组分体积分数等燃烧关键信息的定量计算;最后,指出了自发辐射燃烧诊断技术的发展趋势,即:获得更丰富的检测信号、更高的检测分辨率和精度以及更多的检测结果。
火箭冲压组合发动机包含多个工作模态,不同模态灵活组合的优势使其具有宽速域和广空域的工作特点,兼具加速和巡航的优点.火箭冲压组合发动机燃烧室中存在着亚声速、跨声速和超声速共存的流动结构,具有流动速度高、混合时间短、反应强度大、燃烧空间受限和波系结构复杂等特点.围绕火箭射流的强剪切性、燃烧模式的多样性和燃烧过程的动态性,分析了火箭冲压组合发动机的流动与燃烧特征,总结了面向发动机的高速湍流燃烧研究进展,研究了火箭冲压组合发动机中超声速反应混合层的生长特性、燃烧模式与空间释热分布和动态燃烧特性等问题.通过对碳氢燃料详细化学动力学机理的简化、校验,获得了分别适合于工程计算和细致燃烧机理研究的总包反应与框架机理.从火箭射流主导的反应混合层生长模型,宽范围、变来流工作中流动燃烧过程的不确定性和碳氢燃料动力学的简化与加速算法研究出发,提出了火箭冲压组合发动机基础研究中需要突破的问题,为认识发动机中多尺度燃烧机理、优化多模态燃烧组织提供参考.
高超声速边界层感受性是边界层转捩预测与控制的关键环节,其对高超声速飞行器研究至关重要。目前关于高超声速边界层感受性的实验研究仍然十分匮乏,为了更好地理解高超声速边界层感受性过程并指导该领域的实验研究,文章梳理了近20年来国际上高超声速边界层感受性问题的研究内容,包括对自由流扰动和壁面扰动的感受性,并主要介绍了Fedorov的前缘感受性理论和模态转化机制。最后总结了自由流扰动中感受性的不同发展路径。