高超声速进气道起动特性机理研究
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摘要
以超声速燃烧冲压发动机为动力的吸气式高超声速飞行器为高速洲际旅行以及从大气层向外层空间运输物资提供了一种可行的廉价方案。国际上吸气式高超声速飞行器的研究已经进入到实质性试飞验证阶段。在不远的将来,超燃冲压发动机技术可望率先在高超声速巡航导弹上获得成功应用,对国防建设产生重要影响。进气道是超燃冲压发动机首当其冲的气动部件,发动机稳定、高效的工作以进气道起动为前提。由于高超声速进气道流动的复杂性,目前还没有较好的理论对进气道起动性能进行准确的全面预报。国际上已经开展的许多飞行试验也因不同程度的遭遇到进气道不起动而未能完全达到试验目标。进气道起动能力的预报方法、实验检测方法以及进气道不起动、自起动、起动/起动迟滞、激波振荡等现象是目前研究的热点和难点。开展高超声速进气道起动/起动以及激波振荡特性研究,对十预报进气道不起动、揭示进气道激波振荡机理、探索控制进气道不起动以及激波振荡的有效方法,达到优化进气道设计、提高进气道性能、拓宽进气道稳定工作范围的目的,具有重要意义。本文以脉冲型设备激波风洞为实验平台,发展了高超声速进气道多种流动特征的快速实验检测方法,对进气道起动/起动以及激波振荡等流动现象进行观测,并结合适当的数值模拟,分析了流动的内在机理。
本文首先考察了激波风洞流场的非定常建立过程。为了保证激波风洞喷管的迅速起动,真空罐的初始压强通常很低,风洞流场建立过程中存在非定常高马赫数气流占主导的现象。然后开展了激波风洞与进气道流场建立过程相耦合的研究,考察了真空罐初始压强变化对风洞喷管起动波系以及进气道起动过程的影响,为进一步拓展激波风洞的研究范围,提供了依据。通过对喷管起动过程非定常波系结构的完整描述,弥补了以往进气道脉冲起动研究中使用简化波系的不足,指出风洞流场建立过程的非定常高马赫数气流对进气道起动过程有辅助作用。真空罐的初始压强越低,风洞流场建立越快,这种辅助作用越明显,也是进气道起动的直接原因。
在充分认识激波风洞运行特征的基础上,对激波风洞辅助进气道起动的效应加以利用和抑制,可以为进气道流动研究快速地提供丰富多彩的实验数据。本文在不影响激波风洞喷管迅速起动的前提下,发展了一种在进气道内预先设置轻质堵块的方法,抑制激波风洞运行初期辅助进气道起动的效应。该方法能够迫使进气道在风洞运行初期出现不起动,待轻质堵块被吹出后,流道恢复畅通,进而考察进气道是否具有自起动能力。采用高速纹影拍摄同步壁面压强测量的手段,对二元高超声速进气道的起动特性进行了研究。通过实验结果与数值模拟以及进气道起动能力经验关系式的对比分析,考核了所发展的进气道自起动能力检测方法的可靠性。在激波风洞中进一步研究了内收缩比变化对进气道起动特性的影响。对于较大内收缩比的进气道,利用激波风洞辅助进气道起动的效应,进气道能够起动并维持在起动状态；而在预先设置轻质堵块之后,在相同的来流条件下,同一构型的进气道在轻质堵块被吹出后,进气道最终流态是不起动的。这一特点表明,在激波风洞中开展进气道起动/起动双解区实验,不需要额外的变几何机构辅助,具有一定的潜力和优势。在尖前缘进气道前缘安装钝化部件后,对钝前缘进气道的实验表明,前缘钝化后进气道的起动能力下降,其中外压缩面前缘钝化对进气道自起动能力的影响较大。外压缩面前缘钝化后,外压缩面上的边界层厚度增加,进气道内收缩段入口截面的流场畸变程度增大、平均马赫数下降,这些因素不利于进气道起动。
进一步拓展激波风洞的潜力,延长试验时间,本文开展了进气道激波振荡研究。通过在隔离段出口事先设置固定的堵塞节流楔块模拟反压升高,观测了二元高超声速进气道在不同堵塞度下的流动特征。结果表明,在小堵塞度下,进气道能够“忍受”节流导致的反压并保持起动状态。堵塞度超过临界值,进气道不起动并出现激波振荡现象。在激波振荡过程中,隔离段后部形成向上游传播的激波,并最终导致进气道不起动。逆流而上的激波的成因与此前流道上游产生的压缩波/波在节流段的反射相关。随着隔离段出口堵塞度的增大,激波振荡的频率有所升高,主要是由于逆流而上的激波的平均速度增大。保持内流道的长度基本不变,改变进气道的内收缩比,对不同构型进气道进行的实验表明,激波振荡的流态和频率特性与进气道的几何构型相关。小内收缩比大喉道的进气道在激波振荡过程中,分离激波能够进入隔离段,振荡频率相对较高。大内收缩比小喉道的进气道(不能自起动)在激波振荡过程中,分离激波始终位于外压缩面上,唇口上方存在溢流,振荡幅度较低而振荡频率较高。
本文的研究表明,通过采取适当的实验方法,可以进一步发挥激波风洞运行方式灵活、建造和运行费用较为低廉的特点,为高超声速进气道研究提供一种灵活快捷的途径。
The scramjet engine powered air breathing hypersonic vehicle enable a low-cost way for both high-speed air travel and space access. The status of related researches is now approaching to a phase of flight tests. The scramjet technologies may firstly achieve a success in hypersonic missile application, and make changes in defense sys-tem in the future. The hypersonic inlet is an important and first-step component of the scramjet engine. For efficient operation, the scramjet engine requires the inlet to operate in a starting mode. However, the capability to accurately predict the inlet start-ing characteristic is still lacking, due to the complexity of inlet flows. In some hy-personic air breathing propulsion flight tasks, the test vehicles encountered inlet un-starting and the flight ended prematurely. The prediction and test method of hyper-sonic inlet starting characteristic, such as starting, self-starting, hysteresis in starting, and shock wave oscillations, are still hot and difficult issues. The understanding in hypersonic inlet starting/un-starting and shock wave oscillations are important for the prediction/detection and suppression of inlet un-starting and shock wave oscillations, the improvement of inlet design and performance, and extension of inlet stable oper-ation range. Therefore, additional thorough and detailed wind tunnel experiments are necessary. In the present study, a simple way to test various behaviors of hypersonic inlets was developed in a shock tunnel (pulse facility). Then, the flow patterns of inlet starting/un-starting and shock wave oscillations were observed and discussed with the combination of CFD.
The flow establishment in a shock tunnel was examined at first. To shorten the starting process of the tunnel nozzle, a very low initial pressure is usually used in the dump tank, which results in an unsteady flow with a high Mach number dominating the nozzle starting process. The coupling of a shock tunnel starting process and an inlet starting process were studied, to cover the shortage of some earlier investigations in literature. The effects of initial pressure in the dump tank are shown. It appears that the high Mach number unsteady flow has a capability of aiding the inlet to start. It is also found that the lower the initial pressure is, the faster the nozzle starting process is resulted, and the stronger the aid-starting capability is achieved.
Based on the previous coupling study, a new test method was developed to suppress the aid-starting capability of the shock tunnel without affecting much of the test time. It adopts a light obstacle as a flow plug, that can be rapidly blown out of the inlet, to choke the inlet flow for un-starting at the initial stage of the shock tunnel operation. Simultaneous high speed Schlieren imaging and surface pressure measurements were applied to determine whether the inlet could be restarted after the obstacle action. The initiation of un-start and restart process were observed. A generic hypersonic inlet that was able to self-start was archived in the shock tunnel. With the help of the newly developed test method, the flow characteristics in starting/un-starting, known as dual-solution area, were also observed in an inlet with a large internal concentration ratio. The internal contraction ratio limits for inlets self-starting and un-starting were obtained and compared with the literature. The effects of leading edge bluntness on the inlet starting characteristic were also examined. It appears that the leading edge bluntness play a quite important role in the inlet starting process.
The usable test time of the shock tunnel was extended, and a wedge shaped flow plug was pre-mounted near the isolator exit. A set of two-dimensional hypersonic inlets with different internal contraction ratios was tested with various exit throttling ratios. The results indicate that the backpressure generated by the throttling device can be toler-ated and the inlet can maintain the starting mode at low throttling ratios, whereas un-start flows are initiated from the near-choke throttling ratios and a shock wave oscillation ap-pears. The Schlieren movie demonstrates that the upstream-propagating shocks in the duct play important roles during the oscillation cycles and that the formations of the upstream-propagating shocks are related to the downstream-propagating compression waves/shock waves that encounter the throttling section. The frequency of shock wave oscillation increases with increasing exit throttling ratios, primarily because of the ac-celeration of the upstream propagation. The critical throttling ratio at which the shock wave oscillations occur varies with the inlet internal concentration ratios, and the shock wave oscillation flow patterns and frequencies also relate to the inlet configurations.
Based on these experiments, shock tunnels with sufficiently long running times can be a useful tool to investigate various flow characteristics in hypersonic inlets.
本文首先考察了激波风洞流场的非定常建立过程。为了保证激波风洞喷管的迅速起动,真空罐的初始压强通常很低,风洞流场建立过程中存在非定常高马赫数气流占主导的现象。然后开展了激波风洞与进气道流场建立过程相耦合的研究,考察了真空罐初始压强变化对风洞喷管起动波系以及进气道起动过程的影响,为进一步拓展激波风洞的研究范围,提供了依据。通过对喷管起动过程非定常波系结构的完整描述,弥补了以往进气道脉冲起动研究中使用简化波系的不足,指出风洞流场建立过程的非定常高马赫数气流对进气道起动过程有辅助作用。真空罐的初始压强越低,风洞流场建立越快,这种辅助作用越明显,也是进气道起动的直接原因。
在充分认识激波风洞运行特征的基础上,对激波风洞辅助进气道起动的效应加以利用和抑制,可以为进气道流动研究快速地提供丰富多彩的实验数据。本文在不影响激波风洞喷管迅速起动的前提下,发展了一种在进气道内预先设置轻质堵块的方法,抑制激波风洞运行初期辅助进气道起动的效应。该方法能够迫使进气道在风洞运行初期出现不起动,待轻质堵块被吹出后,流道恢复畅通,进而考察进气道是否具有自起动能力。采用高速纹影拍摄同步壁面压强测量的手段,对二元高超声速进气道的起动特性进行了研究。通过实验结果与数值模拟以及进气道起动能力经验关系式的对比分析,考核了所发展的进气道自起动能力检测方法的可靠性。在激波风洞中进一步研究了内收缩比变化对进气道起动特性的影响。对于较大内收缩比的进气道,利用激波风洞辅助进气道起动的效应,进气道能够起动并维持在起动状态；而在预先设置轻质堵块之后,在相同的来流条件下,同一构型的进气道在轻质堵块被吹出后,进气道最终流态是不起动的。这一特点表明,在激波风洞中开展进气道起动/起动双解区实验,不需要额外的变几何机构辅助,具有一定的潜力和优势。在尖前缘进气道前缘安装钝化部件后,对钝前缘进气道的实验表明,前缘钝化后进气道的起动能力下降,其中外压缩面前缘钝化对进气道自起动能力的影响较大。外压缩面前缘钝化后,外压缩面上的边界层厚度增加,进气道内收缩段入口截面的流场畸变程度增大、平均马赫数下降,这些因素不利于进气道起动。
进一步拓展激波风洞的潜力,延长试验时间,本文开展了进气道激波振荡研究。通过在隔离段出口事先设置固定的堵塞节流楔块模拟反压升高,观测了二元高超声速进气道在不同堵塞度下的流动特征。结果表明,在小堵塞度下,进气道能够“忍受”节流导致的反压并保持起动状态。堵塞度超过临界值,进气道不起动并出现激波振荡现象。在激波振荡过程中,隔离段后部形成向上游传播的激波,并最终导致进气道不起动。逆流而上的激波的成因与此前流道上游产生的压缩波/波在节流段的反射相关。随着隔离段出口堵塞度的增大,激波振荡的频率有所升高,主要是由于逆流而上的激波的平均速度增大。保持内流道的长度基本不变,改变进气道的内收缩比,对不同构型进气道进行的实验表明,激波振荡的流态和频率特性与进气道的几何构型相关。小内收缩比大喉道的进气道在激波振荡过程中,分离激波能够进入隔离段,振荡频率相对较高。大内收缩比小喉道的进气道(不能自起动)在激波振荡过程中,分离激波始终位于外压缩面上,唇口上方存在溢流,振荡幅度较低而振荡频率较高。
本文的研究表明,通过采取适当的实验方法,可以进一步发挥激波风洞运行方式灵活、建造和运行费用较为低廉的特点,为高超声速进气道研究提供一种灵活快捷的途径。
The scramjet engine powered air breathing hypersonic vehicle enable a low-cost way for both high-speed air travel and space access. The status of related researches is now approaching to a phase of flight tests. The scramjet technologies may firstly achieve a success in hypersonic missile application, and make changes in defense sys-tem in the future. The hypersonic inlet is an important and first-step component of the scramjet engine. For efficient operation, the scramjet engine requires the inlet to operate in a starting mode. However, the capability to accurately predict the inlet start-ing characteristic is still lacking, due to the complexity of inlet flows. In some hy-personic air breathing propulsion flight tasks, the test vehicles encountered inlet un-starting and the flight ended prematurely. The prediction and test method of hyper-sonic inlet starting characteristic, such as starting, self-starting, hysteresis in starting, and shock wave oscillations, are still hot and difficult issues. The understanding in hypersonic inlet starting/un-starting and shock wave oscillations are important for the prediction/detection and suppression of inlet un-starting and shock wave oscillations, the improvement of inlet design and performance, and extension of inlet stable oper-ation range. Therefore, additional thorough and detailed wind tunnel experiments are necessary. In the present study, a simple way to test various behaviors of hypersonic inlets was developed in a shock tunnel (pulse facility). Then, the flow patterns of inlet starting/un-starting and shock wave oscillations were observed and discussed with the combination of CFD.
The flow establishment in a shock tunnel was examined at first. To shorten the starting process of the tunnel nozzle, a very low initial pressure is usually used in the dump tank, which results in an unsteady flow with a high Mach number dominating the nozzle starting process. The coupling of a shock tunnel starting process and an inlet starting process were studied, to cover the shortage of some earlier investigations in literature. The effects of initial pressure in the dump tank are shown. It appears that the high Mach number unsteady flow has a capability of aiding the inlet to start. It is also found that the lower the initial pressure is, the faster the nozzle starting process is resulted, and the stronger the aid-starting capability is achieved.
Based on the previous coupling study, a new test method was developed to suppress the aid-starting capability of the shock tunnel without affecting much of the test time. It adopts a light obstacle as a flow plug, that can be rapidly blown out of the inlet, to choke the inlet flow for un-starting at the initial stage of the shock tunnel operation. Simultaneous high speed Schlieren imaging and surface pressure measurements were applied to determine whether the inlet could be restarted after the obstacle action. The initiation of un-start and restart process were observed. A generic hypersonic inlet that was able to self-start was archived in the shock tunnel. With the help of the newly developed test method, the flow characteristics in starting/un-starting, known as dual-solution area, were also observed in an inlet with a large internal concentration ratio. The internal contraction ratio limits for inlets self-starting and un-starting were obtained and compared with the literature. The effects of leading edge bluntness on the inlet starting characteristic were also examined. It appears that the leading edge bluntness play a quite important role in the inlet starting process.
The usable test time of the shock tunnel was extended, and a wedge shaped flow plug was pre-mounted near the isolator exit. A set of two-dimensional hypersonic inlets with different internal contraction ratios was tested with various exit throttling ratios. The results indicate that the backpressure generated by the throttling device can be toler-ated and the inlet can maintain the starting mode at low throttling ratios, whereas un-start flows are initiated from the near-choke throttling ratios and a shock wave oscillation ap-pears. The Schlieren movie demonstrates that the upstream-propagating shocks in the duct play important roles during the oscillation cycles and that the formations of the upstream-propagating shocks are related to the downstream-propagating compression waves/shock waves that encounter the throttling section. The frequency of shock wave oscillation increases with increasing exit throttling ratios, primarily because of the ac-celeration of the upstream propagation. The critical throttling ratio at which the shock wave oscillations occur varies with the inlet internal concentration ratios, and the shock wave oscillation flow patterns and frequencies also relate to the inlet configurations.
Based on these experiments, shock tunnels with sufficiently long running times can be a useful tool to investigate various flow characteristics in hypersonic inlets.
引文
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