Use of shock tunnels for hypersonic propulsion testing
Use of shock tunnels for hypersonic propulsion testing
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DOI：
10.2514/6.1999-2447
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发表时间：
1999-06
期刊：
影响因子：
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通讯作者：
H. Olivier;M. Habermann;M. Bleilebens
中科院分区：
文献类型：
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作者：
H. Olivier;M. Habermann;M. Bleilebens
激波风洞是模拟高超声速飞行和再入过程中高超声速流动条件的有力工具。对于推进试验，已经证明它们不仅能够测量复杂推进系统的外部和进气流场，而且能够测量发动机内部的重要现象，如燃料混合和超音速燃烧。在亚琛激波风洞中，用简单和复杂的坡道几何形状进行了TH2实验，以研究典型的进气流动现象。为此，采用了高马赫数、高雷诺数的流动条件。但对于高超声速飞行器来说，要在风洞中进行尽可能精确和逼真的模拟，就需要尽可能接近地模拟飞行过程中的壁温。因此，在第一次尝试中，通过电阻加热元件从内侧加热24度坡道模型。到目前为止，达到的最高温度达到690K。为了增加流动的总焓，已经开发了一种爆轰驱动器，它将在不久的将来取代激波风洞TH2的传统氦驱动器。到目前为止，已经在激波管模式下进行了6m短驱动段的爆轰驱动实验。用该装置(THD)获得的结果表明，爆轰驱动概念在各种流动条件下应用于激波风洞是可行的。亚琛激波风洞TH2是亚琛激波风洞的激波管，内径140毫米，壁厚80毫米。驱动段和从动段的长度分别为6m和15.4m。该建筑由美国航空航天学会授予*教授、成员、研究工程师Coypright 01999。版权所有。激波隧道是专门为使用这样的隧道而建造的。一道800 mm厚的钢筋混凝土墙将驾驶室和驱动段隔开，作为保护墙，同时也用于支撑隧道的后坐吸收系统。驱动部分和从动段由一个双隔膜室隔开，该隔膜室在最大压力下使用两个10 mm厚的不锈钢板作为隔膜，由铣刀以十字的形式刻印。另一个黄铜或铜板隔膜位于从动段和喷嘴入口处之间。整个管子的最大工作(稳定)压力为1500巴。驱动器可以电加热到最大T4为600K。有两个锥形喷嘴可用，一个半开角度为5.8度，出口直径为586 mm，另一个为10.5度，出口直径为572 mm。对于最后一个，另外两个截锥允许喷嘴出口直径分别为1m和2m。喷嘴喉部直径和试验段马赫数也可以通过插入不同的喉部段来改变。标称出口马赫数为7的异型喷嘴是可用的。图1显示了带有5.8度锥形喷嘴的激波风洞的侧视图。在过去的几年中，隧道主要在表1所列的测试条件下运行。氦用作驱动气体，合成空气用作测试气体。这些条件的主要目的是研究真实气体效应对高超声速空气动力学的影响。因此，停滞温度在1500K到4700K之间变化，这涵盖了下部的完美气体行为和上限的真实气体行为，并在弓激波之后发生了显著的氧解离。当然，不仅可以产生这些试验条件，还可以产生激波风洞运行特性内的各种其他试验段流动。特别是为了研究高超声速发动机的进气流动，标定了单位雷诺数为1650万/m的试验条件X。
Shock tunnels represent a powerful tool for simulating hypersonic flow conditions as they occur during hypersonic flight and reentry. For propulsion testing it has been demonstrated that they are capable not only to allow the outer and intake flow field of complex propulsion systems but also important phenomena inside of the engines like fuel mixing and supersonic combustion. In the Aachen shock tunnel TH2 experiments have been performed with simple and complex ramp geometries to study typical intake flow phenomena. For this a high Mach number, high Reynolds number flow condition is used. But for hypersonic flight vehicles a simulation in wind tunnels as accurate and realistic as possible requires to simulate the high wall temperatures occuring in flight as close as possible. Therefore, in a first attempt a 24 degree ramp model was heated from the inner side by electrical resistance heating elements. The maximum temperature achieved so far amounts to 690 K. In order to increase the total enthalpy of the flow, a detonation driver has been developed which in near future will replace the conventional helium driver of the shock tunnel TH2. Up to now experiments have been performed with the detonation driver in the shock tube mode with a short 6 m long driven section. Results achieved with this facility (THD) show the feasibility of the detonation driver concept for shock tunnel applications over a wide range of flow conditions. THE AACHEN SHOCK TUNNEL TH2 The shock tube of the Aachen shock tunnel has an inner diameter of 140 mm with a wall thickness of 80 mm. The lengths of the driver and driven section are 6 m and 15.4 m, respectively. The building which *Professor, Member AIAA tResearch Engineer Coypright 01999 by the American Institute of Aeronautics and Astronautics Inc. All rights reserved. houses the shock tunnel was especially built for the use of such tunnels. A 800 mm steel-enforced concrete wall which separates the rooms for driver and driven section serves as a protecting wall but is also used for supporting the recoil absorbing system of the tunnel. Driver and driven section are separated by a double-diaphragm chamber which at maximum pressure utilizes two 10 mm thick stainless steel plates as diaphragms scored in the form of a cross by a milling cutter. Another diaphragm of brass or copper sheet is located between the driven section and the nozzle entrance. The maximum operating (steady) pressure of the complete tube is 1500 bar. The driver can electrically be heated to a maximum T4 of 600 K. There are two conical nozzles available, one with a half opening angle of 5.8 degree with an exit diameter of 586 mm and one with 10.5 degree and an exit diameter of 572 mm. For the last one two other truncated cones allow nozzle exit diameters of 1 m and 2 m. The nozzle throat diameter and therefore, the test section Mach number can also be changed by inserting different throat pieces. A contoured nozzle is available for a nominal exit Mach number of 7. In Fig. 1 a side view to scale is shown of the shock tunnel with the 5.8 degree conical nozzle. During the last years the tunnel has mainly been operated under test conditions which are listed in Table 1. Helium is used as driver gas and synthetic air as test gas. The main purpose of these conditions is to study the influence of real gas effects on hypersonic aerodynamics. Therefore, the stagnation temperature is varied between 1500 K and 4700 K which covers perfect gas behaviour at the lower and real gas behaviour at the upper limit with significant oxygen dissociation behind the bow shock. Of course, not only these test conditions can be generated but also a variety of other test section flows which are within the operating characteristics of the shock tunnel. Especially for studying the intake flow in a hypersonic propulsion engine, test condition X with an unit Reynolds number of 16.5 million per meter was calibrated.