Investigation of dual ignition for a detonation-driven shock tunnel in forward driving mode

A detonation-driven shock tunnel is useful as a ground test facility for hypersonic flow research.The forward detonation driving mode is usually used to achieve high-enthalpy flows due to its strong driving capability.Unfortunately,the strong detonation wave front results in diaphragm fragments that disturb the test flow and scratch the nozzle or test models.In this study,a dual ignition system was developed to burst a metal diaphragm without fragmentation in the forward driving mode.A series of experiments were conducted to validate the proposed technique.The influences of the delay time setting on the test conditions were investigated in detail.Numerical simulations were also conducted to obtain a better understanding of the wave processes in the shock tube.The results showed that the dual ignition system solved the diaphragm issues in the forward driving mode.The test time was shortened due to the additional ignition close to the primary diaphragm;the smaller the delay time,the shorter the effective test time.However,a small amount of time loss is considered worthwhile because the severe diaphragm problems have been solved.

The Feature of Weak Shock Wave Propagated in a Overlong Tunnel

Many experiment researches have been developed before. But most of them were carried out with the condition that the tunnel’s ratio of length and diameter (x/D) is under 1000. Recently, the process that compression wave convents into shock wave in the overlong tunnel has also been paid attention. In this paper, features of shock wave as it propagates through a overlong tunnel is investigated, rupturing thin films at the entrance to obtain three kinds of shock wave with different intensities. Then study their features respectively during they propagates through a overlong tunnel with x/D over 6000 at most. Comprehend shock wave more deeply by comparing the results of experiments.

TWO DIMENSIONAL SUPERSONIC SHOCK TUNNEL AND ITS APPLICATION

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THE INTERACTION OF A SHOCK WAVE WITH THE BOUNDARY LAYER IN A REFLECTED SHOCK TUNNEL

The influence of a nontotal reflection on the interaction of a reflected shock wave with the boundary layer in a reflected shock tunnel has been investigated. The calculating method of the velocity, the temperature and the Mach number profiles in the boundary layer in reflected shock fixed coordinates has been obtained. To account for equilibrium real gas effects of nitrogen, the numerical results show that the minimum Mach number in the boundary layer has been moved from the wall into the boundary layer with the increasing of the incident shock Mach number. The minimum Mach number, the shock angle in the bifurcated foot and the jet velocity along the wall to the end plate are reduced owing to the Increasing of the area of nozzle throat. The numerical results are in good agreement with measurements.

Study on the Propagation Law of Shock Wave Pressure in Tunnels with Different Materials

The propagation of shock wave pressure in the tunnel is greatly affected by the tunnel structure,shape,material and other factors,and there are great differences in the propagation law of shock wave pressure in different kinds of tunnels.In order to study the propagation law of shock wave pressure in tunnels with different materials,taking the long straight tunnel with the square section as an example,the AUTODYN software is used to simulate the explosion of TNT in the concrete,steel and granite tunnel,and study on the variation law of shock wave pressure in tunnels with different materials.By using dimensional analysis and combined with the results of numerical simulation,a mathematical model of the propagation law of shock wave pressure in the tunnel is established,and the effectiveness of the mathematical model is verified by making the explosion test of the warhead in the reinforce concrete tunnel.The results show that the same mass of TNT explodes in the tunnel with different materials,and the shock wave overpressure peak at the same measuring point is approximate in the near field.However,there is a significant difference in the middle-far fields from the explosion center,the shock wave overpressure peak in the steel tunnel is 20.76%and 34.82%higher than that of the concrete and the granite tunnel respectively,and the shock wave overpressure peak in the concrete tunnel is 24.91%higher than that in the granite tunnel.Through the experimental verification,getting the result that the maximum relative deviation between the measured value and the calculated value of the shock wave overpressure peak is 11.85%.Therefore,it is proved that the mathematical model can be used to predict the shock wave overpressure peak in the tunnel with different materials,and it can provide some reference for the power evaluation of warhead explosion in the tunnel.

长直坑道内爆炸冲击波相似规律研究

通过长直坑道内炸药爆炸后冲击波的超压峰值变化研究坑道内爆炸冲击波的相似规律。使用AUTODYN软件建立长直坑道仿真模型,采用映射方法将3.25 kg当量的球形TNT炸药在空气中的爆炸状态映射到长直坑道中,采用经典试验结果对模型进行验证。根据几何相似原理,通过改变坑道尺寸、坑道截面形状、炸点位置和炸药当量建立多组不同的模型,模拟计算得到坑道内爆炸冲击波的超压峰值,分析超压峰值的相似规律。结果表明,保持坑道截面积不变、改变截面形状及保持爆心距与炸药当量的比值不变、改变炸药当量,各模型计算得到的冲击波超压峰值具有相似规律。采用本文方法可对其他满足几何相似原理的坑道内爆炸冲击波参数进行预测,为地下空间设施的安全设计提供参考。

EFFECTS OF INCOMING FLOW ASYMMETRY ON SHOCK TRAIN STRUCTURES IN CONSTANT-AREA ISOLATORS

To simulate the actual flowfield at the exit of the supersonic/hypersonic inlet, a wind tunnel is designed to study the flow in the scramjet isolator under the asymmetric incoming flow. And compression fields in the isolator are investigated using wall static and pitot pressure measurements. Three incoming Mach numbers are considered as 1.5, 1.8 and 2. Results show that the increase of the asymmetry of the flow at the isolator entrance leads to the increase of the shock train length in the isolator for a given pressure ratio. Based on the analysis of the flow asymmetry effect at the isolator entrance on the shock train length, a modified correlation is proposed to calculate the length of the shock train. Predicted results of the proposed correlation are in good agreement with the experimental data.

强震区隧道洞口段新型墙-板抗减震措施的效果分析

为了进一步提高强震区铁路隧道的安全以及抗震性能,依托包银铁路甘德尔山隧道,运用数值模拟软件FLAC 3D,针对强震区隧道洞口段墙-板和墙-板-减震层的抗减震作用效果进行研究。研究结果表明,与不采用抗减震措施相比较,采取墙-板抗震措施后,横向位移减小了5.95%,竖向位移减小了45.87%,最小安全系数增大了40.91%~49.25%,最大主应力的抗震作用效果达到了63.56%,最小主应力的抗震作用效果为4.04%,最大剪应力抗震作用效果为8.35%;采取墙-板-减震层减震措施后,横向位移减小了13.11%,竖向位移减小了43.88%,最小安全系数增大了65.82%~74.50%,最大主应力的减震作用效果达到了75.11%,最小主应力的减震作用效果为10.21%,最大剪应力减震作用效果为12.80%。采用墙-板-减震层的抗减震措施优于墙-板结构。推荐采用墙-板-减震层措施对甘德尔山隧道进行抗减震加固设计。

OXYHYDROGEN COMBUSTION AND DETONATION DRIVEN SHOCK TUBE

The performance of combustion driver ignited by multi-spark plugs distributed along axial direction has been analysed and tested. An improved ignition method with three circumferential equidistributed ignitors at main diaphragm has been presented, by which the produced incident shock waves have higher repeatability, and better steadiness in the pressure, temperature and velocity fields of flow behind the incidence shock, and thus meets the requirements of aerodynamic experiment. The attachment of a damping section at the end of the driver can eliminate the high reflection pressure produced by detonation wave, and the backward detonation driver can be employed to generate high enthalpy and high density test flow. The incident shock wave produced by this method is well repeated and with weak attenuation. The reflection wave caused by the contracted section at the main diaphragm will weaken the unfavorable effect of rarefaction wave behind the detonation wave, which indicates that the forward detonation driver can be applied in the practice. For incident shock wave of identical strength, the initial pressure of the forward detonation driver is about 1 order of magnitude lower than that of backward detonation.

Simplified design of rock cavern concrete lining to resist shock loading

A simplified method was proposed for the design of concrete lining in underground rock cavern/tunnel against shock loading. The loading may result from the detonation of explosives on ground surface or ground penetration projectiles exploding adjacent to the cavern/tunnel. The resulting problem necessitates the solution of the dynamics of a beam loaded by a transient pressure uniformly distributed over the span. According to mechanical characteristics of the system with rock bolt and shotcrete, a dynamic support design method based on equivalent single degree of freedom （SDOF） was put forward. The SDOF method was applied to obtaining the maximum displacement at the mid-span of the beam, which is often the controlling factor in the blast-resistant design. In the formulation of the problem, the proposed method combines the phenomena of spalling and structural dynamics theory. An example is provided to demonstrate the applicability of this simplified method.

Air Plasma Mitigation of Shock Wave

Shock wave is a detriment in the development of supersonic aircrafts;it increases flow drag as well as surface heating from additional friction;it also initiates sonic boom on the ground which precludes supersonic jetliner to fly overland. A shock wave mitigation technique is demonstrated by experiments conducted in a Mach 2.5 wind tunnel. Non-thermal air plasma generated symmetrically in front of a wind tunnel model and upstream of the shock, by on-board 60 Hz periodic electric arc discharge, works as a plasma deflector, it deflects incoming flow to transform the shock from a well-defined attached shock into a highly curved shock structure. In a sequence with increasing discharge intensity, the transformed curve shock increases shock angle and moves upstream to become detached with increasing standoff distance from the model. It becomes diffusive and disappears near the peak of the discharge. The flow deflection increases the equivalent cone angle of the model, which in essence, reduces the equivalent Mach number of the incoming flow, manifesting the reduction of the shock wave drag on the cone. When this equivalent cone angle exceeds a critical angle, the shock becomes detached and fades away. This shock wave mitigation technique helps drag reduction as well as eliminates sonic boom.

Shock Wave Mitigation by Air Plasma Deflector

When the spacecraft flies much faster than the sound speed (~1200 km/h), the airflow disturbances deflected forward from the spacecraft cannot get away from the spacecraft and form a shock wave in front of it. Shock waves have been a detriment for the development of supersonic aircrafts, which have to overcome high wave drag and surface heating from additional friction. Shock wave also produces sonic booms. The noise issue raises environmental concerns, which have precluded routine supersonic flight over land. Therefore, mitigation of shock wave is essential to advance the development of supersonic aircrafts. A plasma mitigation technique is studied. A theory is presented to show that shock wave structure can be modified via flow deflection. Symmetrical deflection evades the need of exchanging the transverse momentum between the flow and the deflector. The analysis shows that the plasma generated in front of the model can effectively deflect the incoming flow. A non-thermal air plasma, generated by on-board 60 Hz periodic electric arc discharge in front of a wind tunnel model, was applied as a plasma deflector for shock wave mitigation technique. The experiment was conducted in a Mach 2.5 wind tunnel. The results show that the air plasma was generated symmetrically in front of the wind tunnel model. With increasing discharge intensity, the plasma deflector transforms the shock from a welldefined attached shock into a highly curved shock structure with increasing standoff distance from the model;this curved shock has increased shock angle and also appears in increasingly diffused form. In the decay of the discharge intensity, the shock front is first transformed back to a well-defined curve shock, which moves downstream to become a perturbed oblique shock;the baseline shock front then reappears as the discharge is reduced to low level again. The experimental observations confirm the theory. The steady of the incoming flow during the discharge cycle is manifested by the repeat of the baseline shock front.

马赫数10条件下冲压发动机内氢气燃烧特性试验

针对高马赫数超燃冲压发动机面临的高效燃烧组织挑战,本文提出了一种凹腔后缘激波强化的高马赫数超声速燃烧组织技术,并设计了一套燃烧室采用双侧对称布置凹腔结构的三维发动机试验模型。采用OH*基化学发光光谱诊断与壁面测压相结合的试验手段,在自由活塞驱动激波风洞的名义Ma=10流场中,对凹腔上游横向氢气射流的燃烧特性进行了研究,并讨论了模拟流场的重复性,给出了氢气燃烧演化特征、火焰稳定结构及释热特性。不同车次的总压等流场参数表明试验流场具有较高的重复性,可保障氢气燃烧特性的可复现性。通过观察试验过程中OH*基动态发光特征发现,在高焓激波风洞发动机试验中采用燃料提前喷注的方法使发动机流道在空气主流到来之前充盈大量的氢气,进而在主流到达发动机内的瞬间形成所谓“激波管流动-燃烧”效应,使来流空气与氢气接触面发生自点火与剧烈燃烧,产生显著不同于发动机正常工作情况下的点火与燃烧机制,但随着主流趋于平稳,“激波管流动-燃烧”效应消失,在高总温气流的自点火效应与凹腔后缘的X型激波耦合作用下,火焰稳定在凹腔上游近壁面区的氢气射流尾迹区和凹腔后缘附近的全流场中。通过分析壁面压力分布特征发现,凹腔后缘的X型激波实现了燃烧的强化与火焰的稳定,并获得了最显著的释热压升。这些结果表明高马赫数冲压发动机可利用凹腔后缘X型激波强化燃烧和稳定火焰。

海拔高度对长直坑道内爆炸冲击波传播的影响

为有效表征不同海拔坑道内爆炸冲击波的传播特征,利用非线性显式动力学有限元软件AUTODYN,研究了海拔高度对长直坑道内爆炸冲击波传播的影响规律,探讨了高海拔环境对坑道内冲击波传播的影响,基于量纲分析,建立了适用于不同海拔高度典型坑道内冲击波峰值超压的计算模型,并通过数值计算进行了验证。结果表明:随着海拔高度升高,坑道内爆炸冲击波波阵面传播速度与径向的冲击波参数偏差增大,平面波形成距离增加,冲击波峰值超压降低;在0~4000 m范围内,海拔高度每升高1000 m,冲击波冲量降低约0.91%。结合Sachs无量纲修正方法和量纲分析,推导出不同海拔高度冲击波峰值超压的理论分析模型,模型计算结果与数值计算结果的相对偏差不大于10%,能够为高海拔环境下坑道内爆炸冲击波的传播提供理论依据。

薄膜热流计与原子层热电堆热流传感器的激波风洞试验对比

激波风洞试验的主要测试量是热流密度,且多采用薄膜热流计进行测试,但薄膜热流计热流测试结果缺乏直接验证手段,存在测热结果不确定度偏大等问题。原子层热电堆(ALTP)热流传感器响应时间短且线性度好,结合高精度可溯源的热流传感器标定实验,可保证ALTP热流传感器测热结果的准确性和可靠性。在多个流场条件下的激波风洞试验中开展了薄膜热流计和ALTP热流传感器测热结果交叉对比验证。对比试验显示:ALTP热流传感器和薄膜热流计在不同流场参数下测热结果相对稳定,中高热流下两者测热结果相差在8%以内。结合对比标定溯源链以及激波风洞试验测热结果的讨论,展现了利用ALTP热流传感器在激波风洞试验中在线标定薄膜热流计的可行性。

风洞来流参数影响的气动热数据不确定度评估

激波风洞是开展气动热预测的重要手段,气动热数据的准确预示对降低高超声速飞行器热防护系统设计裕度、增加飞行器有效载荷和射程有重要意义,因此对激波风洞气动热试验数据开展不确定度研究很有必要。对激波风洞重复性运行导致的热流不确定度进行研究,首先采用蒙特卡洛方法,对重复运行58车次的激波风洞测量数据进行分析,获得了风洞来流参数随机分布不确定性。结果表明来流速度的相对不确定度较小,为1.02%;而来流压力的相对不确定度较大,为3.54%。然后选取来流温度、来流密度、来流速度、壁面温度4个不确定性输入变量,采用拉丁超立方抽样法生成样本,通过气动热数值计算获得钝锥模型壁面热流数据,采用非嵌入式多项式混沌方法进行不确定度量化和敏感性分析。结果表明在钝锥头部和锥体大部分区域,风洞重复性运行导致的热流不确定度约为4%~5%;敏感性分析结果显示壁面热流对来流速度最敏感。因此如果需要降低热流数据不确定度,就需要降低风洞来流速度不确定度,这需要通过将激波管激波马赫数控制在更小的偏差范围内来实现。

隧道内障碍物对爆破冲击波传播特性影响研究

隧道掌子面开挖爆破产生的冲击波对施工安全构成极大威胁,为研究隧道内障碍物对掌子面开挖爆破产生的冲击波传播特性影响,依托宁攀高速公路胜利隧道,采用动力分析软件ANSYS/LS-DYNA模拟爆破冲击波在隧道内障碍物段的传播,并采用现场测试方法对数值模拟结果进行验证。研究表明:正常衰减的冲击波传播至障碍物面上方2 m处超压迅速增长4.1%,通过障碍物后由于绕流作用超压增长2.2%,障碍物上方冲击波总体呈现增强-衰减-再增强-再减小的特性;障碍物中心轴线前5 m处反射冲击波超压开始高于初始冲击波,反射冲击波传播至障碍物面时超压峰值增长至初始冲击波的2.4倍;障碍物对正后方0.5 m范围内超压峰值有一定的降低作用,0.5~2 m范围内冲击波超压有所增长,随后恢复为正常超压衰减。现场测试结果表明,障碍物前2.5 m处形成2个超压峰值,经障碍物表面反射第2个冲击波超压峰值为初始峰值的2.95倍。

大尺寸自由活塞激波风洞自由流参数诊断方法

高焓激波风洞自由流具有一定程度的热化学非平衡,使得流场参数的确定存在困难。针对大尺寸自由活塞高焓激波风洞的总温2 700~4 700 K、总压5.6~20.3 MPa条件,采用非接触吸收光谱技术、接触测量技术等测试技术,结合多温度多组分数值模拟方法,共同诊断喷管名义马赫数10出口自由流的参数。结果显示,温度测量值与计算值最大偏差小于7.5%,速度测量值与计算值最大偏差小于5%,NO浓度测量值与计算中最大偏差小于12%。

基于立方型气体状态方程的激波风洞准一维流动数值研究

采用基于立方型气体状态方程的准一维流动数值模拟方法研究了反射式高焓激波风洞的真实气体流动,重点关注了高压真实气体效应对风洞全场流动时空结构和驻室区气流参数的影响,并以理论分析揭示了高压真实气体效应对激波管内流动的作用机理。研究表明:对于以冷高压气体驱动的激波风洞,使用考虑分子体积和分子间作用力的真实气体状态方程能够更准确地描述气体的状态和风洞内的流动状况。高压真实气体效应主要在冷驱动气体中发生作用,其作用效果主要是使当地声速增大,从而使得入射稀疏波和反射稀疏波的传播速度加快;另一方面,高压气体效应在高温气体效应较显著的被驱动气体中作用微弱,且对激波管产生激波的强度和激波后的流动状态影响甚微。稀疏波的加快传播改变了激波管波系的相干时空关系。提前抵达的稀疏波可在一定情况下侵蚀激波风洞的有效试验时间。对于所测试的激波风洞构型,在150 MPa氢气驱动110 kPa氮气的工况下,高压效应导致的有效试验时间缩短约38%。适当加长驱动段长度和采用高温气体驱动均可有效减弱高压真实气体效应的影响。

真实气体效应和热化学反应对激波风洞流场特征参数的影响研究

真实气体效应会改变气体热物性,热化学反应则会影响激波波后状态,两者将共同影响激波风洞流场特征参数。利用真实气体状态方程和热化学平衡假设,建立了一套真实气体热物性和波后参数快速计算方法,研究了真实气体热物性参数的变化及热化学反应对波后参数的影响。通过考虑真实气体和热化学非平衡效应的准一维计算方法,结合气动理论分析,研究了真实气体效应和热化学反应对风洞流场时空结构和驻室参数的影响。结果表明,真实气体的声速、焓以及熵等参数与理想气体存在显著差异,且偏差随着压强的增加而增大。热化学反应使得激波波后气体温度和压强分别低于和高于定比热比假设的结果,且波后状态接近热化学平衡。真实气体效应则使得膨胀波波头传播速度增加及入射激波马赫数减小,影响风洞流场的时空结构,并可能导致有效实验时间减小。