S弯进气道优化对其内流场及性能影响研究

在数值研究大量附面层吸入对某半埋入式S弯进气遭内部流场及气动性能影响的基础上,以ISIGHT软件为平台对其进行优化,并详细对比优化前后进气道内部流场结构及性能变化,结果表明:因吸入大量附面层内低能流体,进气道内出现流动分离,周向总压畸变和旋流畸变相对均匀进气工况均显著增加;优化后,旋流畸变和周向总压畸变分别下降约44.46%和4.09%,中心线趋于前后缓急相当,扩压器前段截面面积缓慢递增,而在接近出口时急速增加,气流在此区间迅速扩压;不同厚度附面层吸入工况下,优化后进气道气动性能相比优化前均有所改善,但流动分离现象始终存在。     

满足斯贝发动机要求的进气道模型试验研究

一、引言 进气道的地面状态性能是进气道各种状态下的重要性能之一.大家知道,设计点取在高空高速状态下的进气道,当在地面状态工作时,进气道性能严重恶化.斯贝发动机对地面工作状态下规定的指标较高,它要求进气道的总压恢复系数σ达0.905,周向总压畸变指数DC_(60)为-0.25.为了使本进气道与斯贝发动机     

高外压缩比的高超声速内乘波进气道设计

为了提高内乘波式进气道的起动性能,扩大其稳定工作的马赫数范围,以得到一种高外压缩比例的高超声速进气道,在ICFC二维基本流场的基础上进行改进,提高流场的外压缩比例,减小流场长度,并基于改进后的基本流场生成内乘波式进气道。数值模拟结果表明:在来流马赫数6的设计状态,该进气道保持了内乘波式进气道的优点:流量捕获能力强(流量系数99.94%),显著减弱外流阻力等;较基于原ICFC流场得到的进气道,该高外压缩比的进气道长度更短,且性能略有提升(出口马赫数为2.78时总压恢复系数为0.459),能在马赫数3.8~6范围内起动工作,较原进气道具有更佳的低马赫数起动能力。     

进气畸变对压气机稳定性的影响

在低速大尺寸单级压气机实验器上,实验研究了动态进气畸变对压气机稳定边界的影响。并采取在静子叶片表面埋入微型压力传感器的方法,分析了压气机流场的动态压力特性。实验结果表明,动态总压畸变的旋转频率和旋转方向对压气机稳定性有很大的影响,当畸变相对转速n在0~+0．5之间时,压气机失速推迟;分析动态压力的频谱特性后知道,在压气机失速前一般存在与失速相关的压力扰动,动态总压畸变影响这种压力扰动,从而影响压气机的不稳定流动状态。     

基于Isight的二元进气道压缩楔射流控制参数优化

二元进气道常用于宽速域吸气式飞行器。宽速域飞行器的飞行速域较大,进气道要兼顾高低速条件下的飞行要求,这存在一定的困难。利用射流进行前体激波控制,在一定程度上可以改善流场,并提升进气道性能,但现有的射流激励方案仅是将激波推至唇口,不一定使得进气道达到最优性能或造成射流流量过多损失,因此射流控制参数的优化是一个重要问题。基于Isight软件搭建优化流程,采用Hooke-Jeeves优化方法,以射流角度、射流宽度以及射流位置作为优化变量,流量系数作为约束条件,总压恢复系数最大作为目标函数进行优化,探究了来流Mach数为6时不同射流参数对进气道性能的影响。结果表明,Hooke-Jeeves优化方法可以应用于进气道前体射流控制参数优化问题,优化后的进气道能够满足流量系数的要求,射流角度优化后的总压恢复系数相对于无射流方案提升18%,综合优化后的总压恢复系数相对于仅优化射流角度提升2.82%。     

基于内乘波概念的TBCC进气道过渡模态研究

TBCC(涡轮基组合循环)在过渡模态的工作,是这种新型高速动力系统的关键问题。本文针对基于内乘波概念的内并联式TBCC进气道进行了其过渡模态流场的研究。首先,应用本团队自主提出的内乘波式进气道设计概念,提出一种新型TBCC进气道设计方案。然后,验证了其在设计马赫数4.0下实现全流量捕获,在模态转换时,分流板的分流是有效的。针对其在过渡模态马赫数2.5的性能,通过基于数值模拟的设计参数分析,发现在内参考收缩比为3.11时进气道流通面积合适可实现自起动、且综合性能较优。研究还发现,分流板偏转角度范围越小,分流段流动越好、通流出口总压恢复系数也越高。     

风向对甲板直升机旋翼流场结构的影响

采用数值模拟方法对船体与直升机旋翼的复合流场进行求解,研究了不同风向时复合流场的流线形态与湍动能分布、旋翼桨叶表面静压分布与旋翼升阻特性.数值模拟的正确性由缩比船模风洞试验验证.研究结果表明:旋翼流场与船体流场间存在相互干扰;侧风会增大复合流场的湍流范围,加剧旋翼周围流场的湍流强度和旋翼桨叶的挥舞振动;右舷15°风向条件不利于直升机的甲板悬停.     

涡轮后机匣流场性能实验研究

对低压涡轮后机匣进行实验研究。对不同来流马赫数及不同来流气流角下的涡轮后机匣流场参数进行测量,分析其在不同来流条件下的整流及损失情况。研究结果表明:支板叶型为前加载且按最大厚度积叠的涡轮后机匣在设计点及非设计点状态都将气流从偏离轴向偏转到基本接近轴向,且总压损失都较小。来流马赫数的改变对后机匣出口气流角影响不大,但来流马赫数的增加会导致其总压损失增加;而来流气流角改变对其出口气流角及总压损失的影响相对较小,但出口气流角及总压损失沿径向的分布会随来流气流角的改变发生变化。     

S弯进气道出口畸变对风扇性能影响研究

为了研究某半埋入式S弯进气道出口畸变对其后风扇级性能的影响,分别将其原型及优化后模型与风扇级对接进行进气道加风扇级全流道数值研究。结果表明:全周畸变条件下,风扇级效率及增压比相比均匀进气工况明显下降,堵塞流量下降约1.20%,最高效率下降约3.77%;优化后风扇稳定工作范围有较大幅度提升,堵塞流量增加约0.19%,效率最大值增加约0.23%;受动叶旋转影响,进气道出口截面底部显著低能流体区在发展至动叶前缘的过程中逐渐减小,畸变流体与非畸变流体出现相互掺杂趋势,且这一现象在进气道优化前后始终存在。     

不同来流马赫数下激光能量形成高超声速进气道虚拟唇口的特性

用基于三维非定常可压缩雷诺平均Navier-Stokes方程的有限体积法计算了马赫数低于设计值6时一种高超声速进气道的性能参数,发现其性能存在明显下降。为提升进气道性能,将功率为15kW的激光能量注入进气道固体唇口前的流场中,形成虚拟唇口,马赫数为4.5,5.0和5.5时,计算得到来流捕获率分别提高了34%,20.6%和15.6%。绘制了不同马赫数下来流捕获率达到峰值时的流场压强云图,说明了虚拟唇口的特性及形成机制。结果表明:来流马赫数越低,来流捕获率越小,但相对于无能量注入时的来流捕获率的提升程度越明显;在不同来流马赫数条件下,通过改变激光能量引致的激波结构和位置,可达到最优状态,即激波与进气道前缘斜激波相交后的透射波打在进气道肩部位置的状态。     

Aeroelastic stability of complete rotors with application to a teetering rotor in forward flight

The derivation of a set of non-linear coupled flap-lag-torsion equations of motion for moderately large deflections of an elastic, two-bladed teetering helicopter rotor in forward flight is concisely outlined. The following degrees of freedom are included in the mathematical model: rigid body flapping, rigid body lead-lag, elastic bending in flap and lead-lag, blade root torsion, and shaft torsion. Quasi-steady aerodynamic loads are considered and the effects of reversed flow are included. The aeroelastic stability of the complete rotor is investigated by using a linearized system of equations of motion. The equilibrium position about which the equations are linearized is obtained by considering the trim state of the helicopter, in true or simulated forward flight conditions. The sensitivity of the aeroelastic stability boundaries to interblade structural and mechanical coupling is illustrated by comparing the complete rotor stability boundaries with those obtained from a single blade analysis for a number of hover and forward flight cases.     

流管实验装置中声传播计算的模态方法

流管实验装置是测量有流动情况下航空发动机消声短舱内声衬声阻抗的主要装置。本文发展了一种解析的模态匹配方法进行在平均流有声衬条件下矩形流管中声传播的计算。用同伦方法求解特征值问题,并与用环绕积分求解的结果进行比较。声场通过轴向阻抗间断面的声压和声质点速度积分相等计算。第一个算例是无流动、硬壁、有限长、考虑端口反射的情况,并与北航流管实验台测量数据进行了对比;第二个算例为有流动情况下有限长声衬管道不考虑端口反射的声场计算,它与文献中NASA流管实验结果和CAA计算结果符合得很好。     

Analysis of liner performance using the NASA Langley Research Center Curved Duct Test Rig

The NASA Langley Research Center Curved Duct Test Rig (CDTR) is designed to test aircraft engine nacelle liner samples in an environment approximating that of the engine on a scale that approaches the full scale dimensions of the aft bypass duct. The modal content of the sound in the duct can be determined and the modal content of the sound incident on the liner test section can be controlled. The effect of flow speed, up to Mach 0.5 in the test section, can be investigated. The results reported in this paper come from a study to evaluate the effect of duct configuration on the acoustic performance of single degree of freedom (SDOF) perforate-over-honeycomb liners. Variations of duct configuration include: asymmetric (liner on one side and hard wall opposite) and symmetric (liner on both sides) wall treatment; inlet and exhaust orientation, in which the sound propagates either against or with the flow; and straight and curved (outlet is offset from the inlet by one duct width) flow path. The effect that duct configuration has on the overall acoustic performance is quantified. The redistribution of incident mode content is shown, in particular the mode scatter effect that liner symmetry has on symmetric and asymmetric incident mode shapes. The Curved Duct Test Rig is shown to be a valuable tool for the evaluation of acoustic liner concepts.     

Comparison of the Noise Directivity Pattern of the Main Rotor of a Helicopter for Flight and Hover Modes

Acoustical Physics - The noise of the main rotor of a helicopter in hover and horizontal flight modes is numerically simulated. To calculate the acoustic characteristics caused by the nonstationary...     

变截面超音速汽液两相流升压过程的研究

本文针对超音速汽液两相流在变截面通道中的升压过程进行了实验研究,得到了变截面混合腔中超音速汽液两相流的压力变化规律,通过实验结果的分析得出了变截面混合腔中的渐缩部分的压力分布与出口压力无关、变截面通道的渐扩部分为一个单相流体扩压管、压力突变发生在变截面混合腔喉部和随着出口压力升高而激波强度增大、变截面超音速汽液两相流升压技术具有定流量特征等结论。     

基于特征值方法的旋翼尾迹稳定性分析

给出一个旋翼尾迹线性化稳定性分析的方法.在该方法中,尾迹涡线被离散为直线涡段,尾迹的扰动归结为涡元端点的扰动,考虑了桨尖涡的自诱导和涡线的互诱导以及桨尖涡与桨叶的实际干扰.使用该方法,分别以UH-1H和AH-1G模型旋翼为例,对悬停和前飞状态的旋翼尾迹的稳定性进行计算和分析.结果表明:旋翼尾迹运动存在大于0的特征值,是内在不稳定的,且最大发散率随波数变化呈现出一定规律性;前飞与悬停状态不同,其最大发散率减小,不稳定性减弱.     

Helicopter noise in hover: Computational modelling and experimental validation

The aeroacoustic characteristics of a helicopter rotor are calculated by a new method, to assess its applicability in assessing rotor performance in hovering. Direct solution of the Euler equations in a noninertial coordinate system is used to calculate the near-field flow around the spinning rotor. The far-field noise field is calculated by the Ffowcs Williams–Hawkings (FW–H) method using permeable control surfaces that include the blade. For a multiblade rotor, the signal obtained is duplicated and shifted in phase for each successive blade. By that means, the spectral characteristics of the far-field noise may be obtained. To determine the integral aerodynamic characteristics of the rotor, software is written to calculate the thrust and torque characteristics from the near-field flow solution. The results of numerical simulation are compared with experimental acoustic and aerodynamic data for a large-scale model of a helicopter main rotor in an open test facility. Two- and four-blade configurations of the rotor are considered, in different hover conditions. The proposed method satisfactorily predicts the aerodynamic characteristics of the blades in such conditions and gives good estimates for the first harmonics of the noise. That permits the practical use of the proposed method, not only for hovering but also for forward flight.     

Numerical simulation of the flow with a pseudo-shock in an axisymmetric expanding duct with a frontal inlet

The results of the numerical modeling of a flow with a pseudo-shock in an axisymmetric duct are presented. The duct included a frontal inlet with the initial funnel-shaped compression part and the cylindrical throat part as well as the subsequent expanding diffuser. To create a flow with a pseudo-shock, the duct was throttled with the use of the outlet converging insert. Numerical computations of the axisymmetric flow have been conducted on the basis of the solution of the Reynolds-averaged Navier?Stokes equations and with the use of the k-ω SST turbulence model. As a result of computations, such parameters of the flow were determined as the location of the beginning of the pseudo-shock, the length of its supersonic part, the velocity profiles in different cross sections of the pseudo-shock, the pressure distribution on the duct wall, the total pressure recovery factor, and others. The behavior of these parameters at the freestream Mach number М = 6 was analyzed versus the diffuser opening angle and different degree of the inlet duct throttling.     

Ventilation duct with concurrent acoustic feed-forward and decentralised structural feedback active control

This paper presents theoretical and experimental work on concurrent active noise and vibration control for a ventilation duct. The active noise control system is used to reduce the air-borne noise radiated via the duct outlet whereas the active vibration control system is used to both reduce the structure-borne noise radiated by the duct wall and to minimise the structural feed-through effect that reduces the effectiveness of the active noise control system. An elemental model based on structural mobility functions and acoustic impedance functions has been developed to investigate the principal effects and limitations of feed-forward active noise control and decentralised velocity feedback vibration control. The principal simulation results have been contrasted and validated with measurements taken on a laboratory duct set-up, equipped with an active noise control system and a decentralised vibration control system. Both simulations and experimental results show that the air-borne noise radiated from the duct outlet can be significantly attenuated using the feed-forward active noise control. In the presence of structure-borne noise the performance of the active noise control system is impaired by a structure-borne feed-through effect. Also the sound radiation from the duct wall is increased. In this case, if the active noise control is combined with a concurrent active vibration control system, the sound radiation by the duct outlet is further reduced and the sound radiation from the duct wall at low frequencies reduces noticeably.