Freejet experiments were conducted in the hypersonic flight-duplicated shock tunnel (JF-12) to understand the detonation combustion mode in the combustor of a standing oblique detonation ramjet (Sodramjet) engine prototype. Besides the traditional schlieren imaging to capture the oblique shock positions in the combustor, we implemented laser absorption spectroscopy of hydroxyl (OH) radical with three laser beams around the shock front. The Sodramjet engine combustor provides an ideal condition to measure OH because its molar fraction within the detonation front is estimated to exceed 0.03, which is much higher than other combustors. However, the harsh environment of freejet shock tunnel test poses a great challenge to the reliability of measurement systems. In this work, near-infrared tunable laser diode sensors centered at 1528 nm were chosen to measure the line-of-sight average of OH partial pressure. Continuous absorption measurements were conducted near the shock front, indicating a thin OH region after shock compression that could only be achieved by detonation combustion. The time-varied OH partial pressure was well predicted by the computational fluid dynamics (CFD) prediction. Our measurement results verified the detonation combustion of Sodramjet engine prototype and showed that in situ, real-time detonation diagnostics based on laser absorption spectroscopy is possible in freejet experiments of shock tunnel.

The first free-jet experiment of a kerosene-fueled oblique detonation engine, conducted in a large-scale hypersonic shock tunnel, is reported in this paper. A novel initiation-control technique using a small on-wedge trip is proposed to overcome the initiation issue of oblique detonation waves (ODWs) encountered when liquid hydrocarbon fuel is used. The results show that the kerosene-fueled ODW fails to initiate within the length-limited combustor without the trip but is successfully initiated when the trip is used, which demonstrates the effectiveness of the proposed initiation-control method. The feasibility of kerosene-fueled oblique detonation propulsion technology is also demonstrated. Novelty and significance statement The feasibility of liquid-hydrocarbon-fueled oblique detonation propulsion technology was demonstrated by performing free-jet shock tunnel experiments of a kerosenefueled oblique detonation engine. To address the initiation issue of oblique detonation waves encountered when liquid hydrocarbon fuel such as kerosene is used, a novel but simple initiation-control technique was proposed and experimentally validated. Stabilized oblique detonation waves were implemented in the combustor, and the detailed flow structures of different oblique detonation combustion modes were clarified. The proposed technique and the results provide significant reference to the future development of oblique detonation engines. (c) 2023 The Combustion Institute. Published by Elsevier Inc. All rights reserved.

In this study, a theoretical method is proposed to solve shock relations coupled with chemical equilibrium. Not only shock waves in dissociated flows but also detonation waves in combustive mixtures can be solved. The global iterative solving process is specially designed to mimic the physical and chemical process in reactive shock waves to ensure good stability and fast convergence in the proposed method. Within each global step, the single-variable equations of normal and oblique shock relations are derived and solved with the Newton iteration method to reduce the complexity of the problems, and the minimization of free energy method of NASA (National Aeronautics and Space Administration) is adopted to solve equilibrium compositions. It is demonstrated that the convergent process is stable and very close to the real chemical-kinetic process, and high accuracy is achieved in the solutions of normal and oblique reactive shock waves. Moreover, the proposed theoretical method has also been applied to many problems associated with reactive shocks, including the stability of oblique detonation wave, bow detonation over a sphere, and shock reflection in dissociated air. The great importance of using chemical equilibrium to theoretically predict the theoretical range of the wedge angle for a standing oblique detonation wave (the standing window of the oblique detonation wave), the stand-off distance of bow detonation wave and the transition criterion of shock reflection in dissociated air with high accuracy have been addressed.

斜爆轰发动机和激波诱导燃烧冲压发动机在高马赫数吸气式发动机中具有重要应用前景,但是斜爆轰发动机是否具有足够大的净推力,还是一个未知的问题,因此需要对高马赫数冲压发动机的推进性能以及提高推力的方法进行理论研究.本文主要分为3部分.第1部分理论研究了超燃冲压发动机中的爆燃波和爆轰波的传播特性.保证发动机稳定燃烧是提高推力的前提.通过对爆燃波和爆轰波传播特性研究,得到了影响发动机燃烧稳定性的关键参数和物理规律.第2部分研究了发动机处于热壅塞临界状态下的燃烧规律和推力特性.在临界状态下,燃烧室入口气流速度正好等于爆轰波传播速度,二者处于平衡状态,这是发动机推进性能的理论上限.第3部分研究了提高高马赫数超燃冲压发动机推力的理论方法.对于高马赫数冲压发动机,燃烧室入口气流速度远远大于爆轰波的传播速度,这部分速度差就是提高推力的理论空间.对于马赫数Ma≥12的超燃冲压发动机,理论上燃烧产生的爆燃波或激波不会引起发动机不起动,因此可以通过进一步添加燃料和氧化剂的方法来提高其推力.理论分析结果表明,对于高马赫数超燃冲压发动机,不但燃烧流场是容易稳定的,而且可以有很多方法来进一步提高推力.

为实现低马赫数条件下斜爆震波的快速起爆,以及提高斜爆震波对来流扰动的抗干扰能力,提出了通过在诱导区内设置一个小鼓包来控制斜爆震波起爆的方案,并通过数值求解二维多组分欧拉方程研究了楔面和鼓包诱导的斜爆震波.结果表明:通过在诱导区内设置鼓包可实现斜爆震波的提前起爆;且在较宽马赫数范围内,其起爆位置相对固定.鼓包诱导的斜爆震波主要有强耦合和弱耦合两种起爆形式,起爆形式与鼓包和斜激波的相对位置有关.此外,来流扰动会使楔面诱导的斜爆震波剧烈振荡,不利于斜爆震发动机的稳定工作;鼓包的存在对来流扰动所引发的斜爆震波振荡具有一定的抑制作用.

The problem of engine unstart of scramjets has not been resolved. In this paper, the mechanism of engine unstart is discussed from the point of view of shock/shock interaction and deflagration-to-detonation transition. The shock/shock interaction leads to the nonlinear, transient and discontinuous process of the supersonic combustion flow field. This process is similar to the deflagration-to-detonation transition process. If the velocity of pre-combustion shock wave is faster than the velocity in the isolator, it will propagate upstream and cause the engine unstart. The C-J detonation velocity is defined as the stable operation boundary of scramjets, which is the maximum shock wave produced by combustion theoretically. The scramjets will work stable if the velocity in the isolator is faster than the corresponding C-J detonation velocity. The combustion characteristics and propulsive performance of scramjets is theoretically analyzed by using C-J detonation theory. For high Mach number scramjets, the velocity in the isolator is much faster than the C-J detonation velocity. Therefore, extra fuel and oxygen can be injected into the combustor to increase the thrust as long as the shock wave velocity driven by the combustion products is slower than the air velocity in the isolator. The theoretical results agree well with the existing experimental results, which can be used as a baseline for the development of scramjets.

The problem of engine unstart of scramjets has not been resolved. In this paper, the mechanism of engine unstart is discussed from the point of view of shock/shock interaction and deflagration-to-detonation transition. The shock/shock interaction leads to the nonlinear, transient and discontinuous process of the supersonic combustion flow field. This process is similar to the deflagration-to-detonation transition process. If the velocity of pre-combustion shock wave is faster than the velocity in the isolator, it will propagate upstream and cause the engine unstart. The C-J detonation velocity is defined as the stable operation boundary of scramjets, which is the maximum shock wave produced by combustion theoretically. The scramjets will work stable if the velocity in the isolator is faster than the corresponding C-J detonation velocity. The combustion characteristics and propulsive performance of scramjets is theoretically analyzed by using C-J detonation theory. For high Mach number scramjets, the velocity in the isolator is much faster than the C-J detonation velocity. Therefore, extra fuel and oxygen can be injected into the combustor to increase the thrust as long as the shock wave velocity driven by the combustion products is slower than the air velocity in the isolator. The theoretical results agree well with the existing experimental results, which can be used as a baseline for the development of scramjets.

To validate the concept of free oblique detonation waves (ODWs) induced by non-intrusive energy deposition applied in ODW engines (ODWEs), numerical simulations are performed by modeling the laser spark as a high-temperature and high-pressure zone embedding in the high-speed incoming flow and by solving the non-dimensional reactive Euler equations with one-step chemistry. The initiation process, flow structures, and self-adapted characteristics of such kinds of free ODWs in an open space are discussed with varying dimensionless source energy and inflow Mach number. The results show that free ODWs can be initiated successfully by local non-intrusive energy deposition. In particular, four ignition regimes, namely, the subcritical regime, the critical I regime, the critical II regime, and the supercritical regime, which are analogous to the one-dimensional direct initiation of detonation, are distinguished by varying the dimensionless source energy. However, the typical energy required by each ignition regime is higher than that in the one-dimensional case due to multi-dimensional energy transmission and instabilities. Moreover, the dimensionless source energy should be moderate to obtain a stable free ODW flow field because extremely low source energy leads to ignition failure, while extremely high source energy causes premature ignition. When the inflow Mach number varies, it is found that free ODWs can adjust automatically to a nearly Chapman–Jouguet state, indicating the self-sustaining nature of free detonation waves. Due to this feature, ODWE performance will be relatively stable despite the change in inflow Mach numbers if free ODWs are utilized in the combustor.

The propagation theory of one-dimensional detonation is complete and accurate, while prediction of two-dimensional oblique detonation propagation with large scale and high accuracy is still difficult primarily due to the treatment of wedge wall during the calculation and the modelling of viscosity. In this paper, the space-time correlation between two-dimensional steady oblique detonation induced by finite wedge and one-dimensional unsteady detonation supported by piston is investigated by numerical simulations of multi-species Euler equations with H2-Air detailed chemical kinetics. The initiation and propagation process of detonation wave, together with its interaction with rarefaction waves are numerically analyzed from the perspective of both space and time. The results show that under the same overdriven degree, the wave structure, wall parameters and profile variation calculated from one-dimensional case fit well both qualitatively and quantitatively with two-dimensional case after a certain space-time transformation, which validates the space-time correlation between one- and two-dimensional detonation waves. The difference mainly lies in the transition process between different stages of detonation development, such as the transition from over-driven oblique detonation wave to near Chapman-Jouguet (CJ) oblique detonation wave under the effect of rarefaction waves. Since most features of oblique detonation waves over finite wedge can be obtained efficiently by one-dimensional numerical calculation of piston-driven detonation and space-time transformation, the present work provides a feasible way to understand the spatial structure of oblique detonations waves, including detonation initiation, formation of over-driven oblique detonation and cellular structure downstream. Besides, this paper also provides a novel method to distinguish the effect of wall compression and boundary layer by comparing one- and two-dimensional numerical results. Moreover, the results imply that two-dimensional flow field over wedges of different shapes can be obtained with satisfying accuracy by altering the velocity of piston, which greatly reduces the time, cost and complexity of numerical simulation during the design of combustion chamber in an oblique detonation wave engine.

为了研究斜爆轰发动机的稳定燃烧机理,开展了飞行马赫数9的斜爆轰发动机的数值模拟研究和试验研究.设计了全尺度斜爆轰发动机模型,发动机的总长度为2.8m.采用两级进气道压缩,每级压缩角度均为15°.利用三个小支板在进气道前缘主流核心区中进行氢气的喷射和混合.采用带化学反应的雷诺平均N-S方程、SST k-ω模型以及9组分19步反应的基元反应模型,对氢气混合过程和发动机燃烧过程进行了数值模拟研究.结果表明,氢气在进气道内混合得比较均匀,在燃烧室内获得了稳定的斜爆轰流场和正爆轰流场.在激波风洞中开展了马赫数9状态下的斜爆轰发动机稳定燃烧机理试验研究,在50ms的风洞有效试验时间内获得了持续稳定的斜爆轰流场,试验结果与数值模拟结果吻合较好,表明在试验中形成了斜爆轰波.研究结果证明了斜爆轰发动机的技术可行性.

Initiation and stabilization of oblique detonation waves (ODWs) are important to the successful application of oblique detonation engines (ODEs), which, however, have been rarely studied under realistic combustor conditions. In this study, the flow structures, stabilization characteristics and potential thrust performance (under different combustor's geometries with different ODW reflection locations) in a typical hydrogen-fueled ODE combustor are numerically studied by solving the two-dimensional multi-species Reynolds-averaged conservation equations with a detailed hydrogen combustion mechanism. Results suggest that all the detonation waves/shock waves can be stabilized in the space-confined combustor, and the boundary layer separation induced by the ODW-boundary layer interaction is found crucial to determining the types of combustion mode in the combustor. Except for the expected ODW-induced combustion, fast combustion induced by a stabilized overdriven normal detonation wave (NDW) may exist in the combustor simultaneously (even up to a large extent, >73.7%). It is demonstrated that the stabilization of the overdriven NDW in the combustor can be attributed to the formation of an effective aerodynamic convergent-divergent nozzle that quickly accelerates the subsonic flow behind the NDW to supersonic, preventing downstream disturbances from propagating upstream. Benefiting from the chemical equilibrium shift caused by the expansion effect of the flow, more heat is released to compensate for the compression loss and the simulated thrust performance is shown not deteriorate significantly even with a large percentage of NDW-induced combustion existing in the ODE combustor. This work would be beneficial to the future developments of the ODEs.

The deflagration-to-detonation transition (DDT) process is of great importance to both combustion theory and industry safety. In this study, the propagating mechanism of Chapman-Jouguet (C-J) deflagration is studied. Firstly, three models are put forth to decouple the C-J detonation front. These three models are (a) to introduce an expansion parameter into the one-dimensional energy equation, (b) to increase the activation energy of the chemical reaction model and (c) to decouple the shock wave from the flame front by artificial method. The C-J deflagration is obtained after the C-J detonation is decoupled by one-dimensional numerical simulations with different models, chemical reaction kinetics and numerical schemes. Secondly, the propagating mechanism of C-J deflagration is discussed. For the C-J deflagration with a propagating velocity of about 1/2 C-J detonation, the static temperature behind the leading shock wave is too low to ignite the combustion. But, the total temperature of the flow induced by the leading shock wave is high enough to ignite the mixture. The induced flow is slowed down by the rarefaction waves form the wall and its static temperature increases. The flame and the leading shock wave propagate with almost the same velocity and the double-discontinuity structure of the flow field keeps stable. The propagating velocity equals to the sound speed of the combustion products, which is about 1/2 C-J detonation velocity.

In this study, the performances of a Mach 9 oblique detonation engine fueled by hydrogen are numerically investigated by solving the multi-species reactive Reynolds-averaged Navier-Stokes (RANS) equations with a detailed combustion mechanism. The fuel is perpendicularly pre-injected into the core airflow in the engine inlet by three parallel strut-injectors. It is demonstrated that mixing can be enhanced by the baroclinic effect of oblique shock waves, the incipient expansion of fuel jets and the intensive momentum exchange of vertical jets into the crossflow, resulting in a well-mixed fuel-air core flow before entering the combustor. Analyses of the two most dangerous zones that bear potential of pre-ignition suggest that no pre-combustion occurs in the inlet. Benefiting from the floor bleed structure, the upstream movement of the shock waves stops at the combustor's entrance and they remain stabilized in the combustor thereafter. Finally, the combustor is proved to work under the stable detonation mode of combustion, and the supersonic fuel-air mixture is fast burnt through the steady detonation waves generated in the combustor. The concept of the pre-injection oblique detonation engine has been numerically demonstrated, which provides a significant reference to further experimental studies and future engineering applications.