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1.
The sensitivity of the flow along the nozzle and in the test section of high enthalpy wind tunnels to the thermochemical response
of the nozzle expansion process, as well as effects on the pressure and heat transfer distributions over the Electre blunt
cone standard test model, are examined in the framework of properly characterizing the test section flow field in such facilities.
Particularly sensitive to the thermochemical behaviour of the nozzle flow, in the facilities under consideration, are the
static pressure, static temperature and Mach number, whereas stagnation point (pitot) pressure and heat transfer data or freestream
velocity are inadequate for the characterization of the thermochemical state of the flow. The Electre and nozzle wall pressure
data in the F4 arc jet wind tunnel suggest, in contrast to nonequilibrium computations, that the flow in the F4 nozzle is
close to equilibrium. In the HEG and, to some extent, the T5 piston-driven shock tunnels, there are indications that significant
heat losses occur in the reservoir. Lastly, simple semi-empirical formulations for stagnation point heating are shown to perform
reasonably well in high enthalpy flow conditions. 相似文献
2.
高马赫数超燃冲压发动机技术研究进展 总被引:1,自引:0,他引:1
吸气式高超声速飞行在空间运输和国家空天安全领域具有极高价值,超燃冲压发动机是其核心动力装置.目前飞行马赫数4.0~7.0超燃冲压发动机技术日趋成熟,发展更高速的飞行动力技术成为今后临近空间竞争焦点之一.本文对飞行马赫数8.0~10.0的高马赫数超燃冲压发动机技术进行了分析和综述.首先论述其亟待解决的关键问题和技术,分别包括高焓离解与热化学非平衡效应、超高速气流燃料增混与燃烧强化技术、高超声速燃烧与进气压缩的匹配及工作模态、高焓低雷诺数边界层流动及其控制方法、高焓低密度流动/燃烧的热防护技术,以及高马赫数发动机的地面试验风洞技术.然后,进一步介绍了国内外高焓激波风洞与驱动技术以及国内外典型的地面和飞行试验进展.进而针对推进和热防护的总体性能评估、高马赫数发动机内凸显的高焓离解与热化学非平衡效应、超高速气流燃料增混和燃烧强化技术综述了相关研究进展及结论,讨论了高马赫数超燃冲压发动机的可行性以及各关键技术的特点.最后进行了总结并对后续研究提出了几点建议. 相似文献
3.
Tao Geng Fei Zheng Andrey V. Kuznetsov William L. Roberts Daniel E. Paxson 《Flow, Turbulence and Combustion》2010,84(4):653-667
Pulsed combustion is receiving renewed interest as a potential route to higher performance in air breathing propulsion and
ground based power generation systems. Pulsejets offer a simple experimental device with which to study unsteady combustion
phenomena and validate simulations. Previous computational fluid dynamics (CFD) simulations focused primarily on pulsejet
combustion and exhaust processes. This paper describes a new inlet sub-model which simulates the fluidic and mechanical operation
of a valved pulsejet head. The governing equations for this sub-model are described. Sub-model validation is provided through
comparisons of simulated and experimentally measured reed valve motion, and time averaged inlet mass flow rate. The updated
pulsejet simulation, with the inlet sub-model implemented, is validated through comparison with experimentally measured combustion
chamber pressure, inlet mass flow rate, operational frequency, and thrust. Additionally, the simulated pulsejet exhaust flowfield,
which is dominated by a starting vortex ring, is compared with particle imaging velocimetry (PIV) measurements on the bases
of velocity, vorticity, and vortex location. The results show good agreement between simulated and experimental data. The
inlet sub-model is shown to be critical for the successful modeling of pulsejet operation. This sub-model correctly predicts
both the inlet mass flow rate and its phase relationship with the combustion chamber pressure. As a result, the predicted
pulsejet thrust agrees very well with experimental data. 相似文献
4.
A combined experimental and numerical investigation of flow control actuation in a short, rectangular, diffusing S-shape inlet duct using a two-dimensional tangential control jet was conducted. Experimental and numerical techniques were used in conjunction as complementary techniques, which are utilized to better understand the complex flow field. The compact inlet had a length-to-hydraulic diameter ratio of 1.5 and was investigated at a free-stream Mach number of 0.44. In contrast to the baseline flow, where the flow field was fully separated, the two-dimensional control jet was able to eliminate flow separation at the mid-span portion of the duct and changed considerably the three-dimensional flow field, and ultimately, the inlet performance. A comparison between the baseline (no actuation) and forced flow fields showed that secondary flow structures dominated both flow fields, which is inevitably associated with total pressure loss. Contrary to the baseline case, the secondary flow structures in the forced case were established from the core flow stagnating on the lower surface of the duct close to the aerodynamic interface plane. High fidelity spectral analysis of the experimental results at the inlet’s exit plane showed that the baseline flow field was dominated by pressure fluctuations corresponding to a Strouhal number based on hydraulic diameter of 0.26. Not only did the two-dimensional tangential control jet improve the time-averaged pressure recovery at the inlet exit plane (13.3% at the lower half of the aerodynamic interface plane), it essentially eliminated the energy content of the distinct unsteady fluctuations which characterized the baseline flow field. This result has several implications for the design of a realistic engine inlet; furthermore, it depicts that a single non-intrusive static pressure measurement at the surface of the duct can detect flow separation. 相似文献
5.
6.
A combined experimental and numerical investigation of the flow field in a short, rectangular, diffusing S-shape inlet duct was conducted. The inlet duct had a length-to-hydraulic diameter ratio of 1.5 and an inflow Mach number of 0.44. The flow field was diagnosed utilizing stereoscopic particle image velocimetry, surface static pressure measurements, and high frequency total pressure measurements both on the lower surface and at the duct’s aerodynamic interface plane. To complement the experimental investigation and to aid in understanding the flow field associated with this complex geometry, a numerical flow simulation was undertaken. The flow field exhibited massive flow separations and shear layer formations at both turns of the compact inlet. Moreover, secondary flow structures along the duct’s lower surface and along the duct’s side walls were identified. It was shown that the two counter-rotating flow structures along the duct’s lower surface resulted in high levels of total pressure loss at the aerodynamic interface plane. A high fidelity spectral analysis of the pressure signals at the aerodynamic interface plane and along the lower surface was conducted and demonstrated that a high frequency surface static pressure sensor could identify flow separation in a non-intrusive fashion, allowing for future use in a closed-loop control scheme for active flow control. This work was part of a more comprehensive study which was to utilize active flow control to improve performance metrics of such compact inlets. 相似文献
7.
8.
To investigate the effect of different disturbances in the upstream, we present numerical simulation of transition for a hypersonic boundary layer on a 5-degree half-angle blunt cone in a freestream with Mach number 6 at 1-degree angle of attack. Evolution of small disturbances is simulated to compare with the linear stability theory (LST), indicating that LST can provide a good prediction on the growth rate of the disturbance. The effect of different disturbances on transition is investigated. Transition onset distributions along the azimuthal direction are obtained with two groups of disturbances of different frequencies. It shows that transition onset is relevant to frequencies and amplitudes of the disturbances at the inlet, and is decided by the amplitudes of most unstable waves at the inlet. According to the characteristics of environmental disturbances in most wind tunnels, we explain why transition occurs leeside-forward and windside-aft over a circular cone at an angle of attack. Moreover, the indentation phenomenon in the transition curve on the leeward is also revealed. 相似文献
9.
Aerodynamic experimentation with ducted models as applied to hypersonic air-breathing vehicles 总被引:1,自引:0,他引:1
Yu. P. Goon’ko 《Experiments in fluids》1999,27(3):219-234
A methodology of experimentation in high supersonic wind tunnels for studying aerodynamic characteristics of hypersonic flying
vehicles powered by air-breathing engines is discussed. Investigations of such total aerodynamic forces as drag, lift and
pitching moment at testing the models are implicit when the air flow through the model ducts is accomplished so that to provide
the simulation of the external flow around the airplane and flow over the inlets, but the operating engines and, hence, the
exhaust jets are not modeled. The methods used for testing such models are based on the measurement of duct stream parameters
alongside with the balance measurement of aerodynamic forces acting on the models. In the tests, aerometric tools are used
such as narrow metering nozzles (plugs), pitot and static pressure probes, stagnation temperature probes and pressure orifices
in walls of the model duct. The aerometric data serve to determine the flow rate and momentum of the stream at the duct exit.
The internal non-simulated forces of the model ducts are also determined using the conservation equations for energy, mass
flow and momentum, and these forces are eliminated from the aerodynamic test results. The techniques of the said model testing
have been well developed as applied to supersonic aircraft, however their application for hypersonic vehicles whose models
are tested at high supersonic speeds, Mach number M
∞>4, implies some specific features. In the present paper, the results of experimental and theoretical study of these features
are discussed. Some experimental data on aerodynamics of hypersonic aircraft models received in methodological tests are also
presented. The tunnel experiments have been carried out in the Mach number range M
∞=2–6.
Received: 25 July 1996/ Accepted: 14 December 1998 相似文献
10.
爆轰燃烧具有释热快、循环热效率高的特点. 斜爆轰发动机利用斜爆轰波进行燃烧组织, 在高超声速吸气式推进系统中具有重要地位. 以往研究主要关注斜爆轰波的起爆、驻定以及波系结构等, 缺少从整体层面出发对斜爆轰发动机开展推力性能分析. 本文将斜爆轰发动机内的流动和燃烧过程分解成进气压缩、燃料掺混、燃烧释热和排气膨胀4个基本模块并分别进行理论求解, 建立了斜爆轰发动机推力性能的理论分析模型. 在斜爆轰波系研究成果的基础上, 选取了过驱动斜爆轰、Chapman?Jouguet斜爆轰、过驱动正爆轰和斜激波诱导等容燃烧等4种燃烧模式来描述燃烧室内的燃烧释热过程, 并对比分析了不同燃烧模式对发动机比冲性能的影响. 此外, 还获得了不同来流参数、燃烧室参数和进排气参数等对发动机推力的影响规律, 发现来流马赫数和尾喷管的膨胀面积比是发动机理论燃料比冲的主要影响因素. 最后, 结合以往关于受限空间内斜爆轰波驻定特性等方面的研究成果, 提出了斜爆轰发动机燃烧室的设计方向. 相似文献
11.
On the influence of elevated surface temperatures on hypersonic shock wave/boundary layer interaction at a heated ramp model 总被引:2,自引:0,他引:2
Although important flow parameters as Mach number, Reynolds number and total enthalpy can be reproduced in most hypersonic experiments quite well, due to different surface temperature effects in wind tunnel and flight, scaling as well as specific flow properties of shock wave/boundary layer interactions are different. This especially holds for short-duration facilities like, e.g. shock tunnels where due to short running times the models remain more or less at ambient temperature. To overcome this shortcoming, an experimental study has been conducted using a preheatable ramp model with 15° ramp angle. This allowed us to adjust the surfaces to an arbitrary temperature just before the experiment started. Pressure and heat flux measurements clearly showed the effect of varying surface and free stream temperatures. These results are supported by schlieren pictures and infrared measurements. The comparison of the measurements with theoretical and numerical results shows a good agreement. Separation bubble scaling laws proposed by Katzer and Davis have been applied and partially confirmed using the local conditions of the boundary layer at separation. 相似文献
12.
为了研究乘波体几何外形参数和飞行参数对前体/进气道一体化设计的影响,采用理论分析和数值模拟相结合的方法,以马赫数Ma=6和攻角α=0为设计状态、进气道总压恢复系数和前体阻力系数为目标函数,对乘波体前体/进气道进行了优化设计,并在此基础上研究了攻角、马赫数、前缘半径、前体宽度对气动参数的影响。结果表明:该乘波体前体/进气道构型具有良好的攻角特性,总压恢复系数比基准构型提高17.79%,阻力系数比基准构型降低78.5%,符合高超声速飞行器高升力、低阻力的要求,且非常适合小攻角高超声速巡航飞行;为了得到较高升阻比的前体,在前缘半径R≤2mm的范围内进行流场反设计时,可以将设计马赫数的取值比预期低一些。 相似文献
13.
S. V. Guvernyuk A. F. Zubkov K. M. Pichkhadze V. S. Finchenko A. I. Shvets 《Moscow University Mechanics Bulletin》2009,64(5):123-125
Some results of experimental studies conducted in a wind tunnel at the Mach number M = 1.78 for a blunt body of small elongation are discussed. The effect of the attack angle on the drag and lift coefficients as well as on the static stability and the pressure center position is considered. 相似文献
14.
Experimental results on the shock structure of dual co-axial jets are presented. The effects of the geometric parameters of
the inner nozzle, jet static pressure ratio (ratio of the exit plane static pressures of the inner and outer nozzles) and
the ratio of outer to inner nozzle throat area on the shock structure were studied. A superimposed outer and inner jet structure
was observed in the schlieren photographs. The inner flow is compressed by the outer flow resulting in the formation of a
Mach disc and an exit shock. A parameter incorporating the effect of Mach number of the inner nozzle and jet static pressure
ratio was found to correlate the observations regarding the Mach disc location. 相似文献
15.
钝头体激波诱导振荡燃烧现象的数值模拟 总被引:3,自引:0,他引:3
采用一种改进的化学非平衡流解耦方法对轴对称Euler反应流方程解耦处理,对流项采用五阶WENO格式离散,化学反应源项的刚性采用简化的隐式方法处理,时间步进采用二阶精度的Runge-Kutta方法,对H2/Air预混气在来流Ma=4.48和Ma=4.79时的激波诱导振荡燃烧现象进行了数值研究. 对比分析了网格尺度的影响,发现计算结果对法向网格尺度比较敏感,流向网格密度的变化对结果影响不大;Ma=4.48时,采用敏感性分析方法对各反应模型进行了对比分析,J和B-W模型在实验所处的温度和压力范围内能够比较准确的预测诱导时间,所得的振荡频率与实验结果相符,所揭示的振荡机理与McVey和Toong振荡机理吻合,而JM模型预测的诱导时间偏长,其振荡频率低于实验观测值;Ma=4.79时,J模型所得的振荡频率与实验值吻合,而B-W模型释热时间短,且对流场波动比较敏感,流场扰动引起了瞬时局部爆震现象,破坏了振荡的规律性;五阶WENO空间离散格式的应用使计算结果精度较好. 相似文献
16.
J. Lepicovsky 《Experiments in fluids》2008,44(6):939-949
An extensive experimental study into the nature of the separated flows on the blade suction surface of modern transonic fans
is described in this paper. The study was a subtask of a larger experimental effort focused on blade flutter excited by flow
separation in the blade tip region. The tip sections of airfoils on transonic fan blades are designed for precompression and
consequently they differ from sections on the rest of the blade. The blade tip section was modeled by a low aspect ratio blade
and therefore most of the blade tested was exposed to the secondary flow effects. The aim of this work was to supply reliable
data on flow separation on transonic fan blades for validation of future analytical studies. The experimental study focused
on two visualization techniques: surface flow visualization using dye oils and schlieren (and shadowgraph) flow visualization.
The following key observations were made during the study. For subsonic inlet flow, the flow on the suction surface of the
blade was separated over a large portion of the blade, and the separated area increased with increasing inlet Mach number.
For the supersonic inlet flow condition, the flow was attached from the leading edge up to the point where a bow shock from
the upper neighboring blade imposed on the blade surface. Downstream, there was a separated flow region in which air flowed
in the direction opposite the inlet flow. Finally, past the separated flow region, the flow reattached to the blade surface.
For subsonic inlet flow, the low cascade solidity resulted in an increased area of separated flow. For supersonic flow conditions,
the low solidity resulted in an improvement in flow over the suction surface. 相似文献
17.
Computational fluid dynamics analysis of diffuser performance in gas-powered jet pumps 总被引:2,自引:0,他引:2
Jet pump diffuser performance is analyzed, both in terms of past experimental work dealing with the high inlet flow distortions involved and in the sense that this problem is amenable to predictive investigation by computational fluid dynamics techniques. In these highly nonuniform flow conditions, diffusers are seen to justify their inclusion in a jet pump design, for regaining static pressure downstream of the vacuum chamber, even though their performance in effectiveness terms is lowered by about two thirds at high inlet glow distortion levels. A satisfactory correlation has been found between outlet and inlet conditions and diffuser area ratio, extending well beyond past experimental published results for diffuser geometry and distorted inlet flows. 相似文献
18.
19.
Simulation of unsteady hypersonic combustion around projectiles in an expansion tube 总被引:1,自引:0,他引:1
The temporal evolution of combustion flowfields established by the interaction between wedge-shaped bodies and explosive
hydrogen-oxygen-nitrogen mixtures accelerated to hypersonic speeds in an expansion tube is investigated. The analysis is carried
out using a fully implicit, time-accurate, computational fluid dynamics code that we recently developed to solve the Navier-Stokes
equations for a chemically reacting gas mixture. The numerical results are compared with experimental data from the Stanford
University expansion tube for two different gas mixtures at Mach numbers of 4.2 and 5.2. The experimental work showed that
flow unstart occurred for both the Mach 4.2 cases. These results are reproduced by our numerical simulations and, more significantly,
the causes for unstart are explained. For the Mach 5.2 mixtures, the experiments and numerical simulations both produced stable
combustion. However, the computations indicate that in one case the experimental data were obtained during the transient phase
of the flow; that is, before steady state had been attained.
Received 7 February 2000/ Accepted 20 February 2001 相似文献