This study focuses on the stationary characteristics of oblique detonation across various reaction rate distributions by using two-dimensional Euler equations coupled with a two-step kinetic model, focusing on the effects of the reduced activation energy and the exothermic reaction rate constant. The results show that increasing the reduced activation energy while decreasing the rate constant of the exothermic reaction postpones the initiation of oblique detonation and stabilizes the initiation structure. This paper innovatively uses these two parameters to regulate the exothermic reaction rate distribution while ensuring that the lengths of the induction and exothermic reaction zones remain constant. It is found that the exothermic reaction rate distribution significantly affects the stationary characteristics of oblique detonation waves. For a constant exothermic zone length, as the activation energy increases, the initiation position of the oblique detonation shifts downstream until it reaches a specific critical threshold. Beyond this point, the initiation structure of the oblique detonation undergoes a rapid transition to instability, evolving into an unstable oblique detonation configuration. This critical behavior occurs because cases with larger activation energy exhibit heightened sensitivity to overdrivenness.

Hypersonic stage separation is considered a crucial and challenging problem that involves complex aerodynamic interference and unsteady aerodynamics, which directly determines the success or failure of the two-stage-to-orbiter (TSTO) vehicle. The longitudinal stage separation (LSS) scheme was proposed to improve the safety of the parallel-staged TSTO stage separation, wherein the orbiter separates along the booster’s upper surface with tiny or even no gaps resulting in weak interference. Therefore, the hypersonic flow past the TSTO vehicle at Mach 7 during the LSS is examined by the free-flight experimental test in the JF-12 shock tunnel and laminar CFD simulation. The high-speed pneumatic ejection to launch vehicle model system incorporated into the shock tunnel is developed to conduct the LSS test within the short test duration. In addition, high-speed visualization and image-processing techniques were used to investigate the separation behavior of TSTO. The flow patterns, unsteady measured wall pressure, and the orbiter’s separating trajectory are studied and compared between experiment and simulation. The results between the experiment and CFD show a good agreement with each other. The LSS flow pattern is governed by the type I and VI weak shock wave-shock wave interaction, and short-term shock reflection occurs between stages when the orbiter’s afterbody separates from the booster’s leading edge. Moreover, the aerodynamic interference of LSS for parallel-staged TSTO is weak due to the tiny interstage gap, and the orbiter separates from the booster successfully.

The distribution of exothermic reaction rates is jointly influenced by reduced activation energy and reaction rate constant. This study focuses on the effect of distribution of exothermic reaction rates on detonation wave propagation instability, specifically under conditions where the length of the induction and exothermic reaction remains constant. It is found that the distribution variation of exothermic reaction rates significantly influences the detonation wave propagation characteristics. Specifically, under conditions of high activation energy, the exothermic reaction rate profile exhibits a smoother distribution but becomes more prone to perturbations. This heightened sensitivity, coupled with the augmented overdriven degree associated with pulsating detonation and cellular detonation wave propagation, further exacerbates the instability characteristics of detonation waves. Especially to the two-dimensional detonation waves with high activation energies, the distribution of exothermic reaction rates becomes more sensitive to these displacements, reinforcing the transverse shock wave and leading to a transformation of the wavefront and cellular structure towards more unstable configurations. This research delves into the intricate interactions between the distribution of exothermic reaction rates and detonation wave instability, aiming to provide an explanatory of detonation instability.

This study delves into the effect of stratified fluid on the stationary characteristics of oblique detonation within a confined combustor, with the objective of providing a reference for the advancement of oblique detonation engines. Oblique detonation engines possess remarkable advantages such as high thermal cycle efficiency and are highly suitable for high-Mach-number flight. However, ensuring the stationary of detonation waves remains a formidable challenge. In this research, two-dimensional Euler equations augmented by a two-step kinetic model are employed to conduct numerical analyses. The study reveals that in both unconfined and confined combustors, the oblique detonation wave exhibits distinct basic structures. Stratified gas with varying widths exerts a profound impact on the stationary of the oblique detonation wave, giving rise to diverse reflection structures, including the regular reflection mode, Mach reflection mode, and the novel mode where the Mach stem is connected to a normal detonation wave. The formation mechanism of these complex structures is investigated, with particular emphasis on the crucial role of the triple-wave point. The presence of an inert gas layer within the stratified fluid configuration not only alters the Mach number behind the shock/detonation wave but also modifies the angles of the reflected shock wave in both the inert gas layer and the combustible gas layer. These changes lead to a series of unique interactions that ultimately govern the behavior of the oblique detonation wave.

The present study delves into the examination of the stereoscopic cells and wavefront structures characterizing the propagation of three-dimensional detonation waves within square ducts. Leveraging numerical solutions derived from three-dimensional reactive Euler equations, incorporating an induction-exothermic reaction kinetic model, this work reveals the distinct classification of three modes of detonation waves based on the direction of propagation and the phase characteristics of transverse shock waves on the wavefront. This paper delineates the presence of two different types of phenomena: duct wall slapping waves due to shock-wall collisions and internal slapping waves resulting from shock interactions. Furthermore, this investigation exposes the existence of two distinct types of triple-wave lines on the wavefront: the first comprising a strong Mach disk, a weak Mach disk, and a transverse shock wave; the second characterized by a weak Mach disk, an incident shock wave, and a transverse shock wave. Notably, the pressure behind the first type of triple-wave line is observed to be the highest. It elucidates the transition from two- to three-dimensional detonation waves, revealing that the prevalence of transverse shock waves on the wavefront in the rectangular and diagonal modes is twofold and quadruple, respectively, when compared to their two-dimensional counterparts within identical ducts/channels.

No AccessTechnical NotesExperimental Study on Aerodynamic Heating of Hypersonic Boundary-Layer BlowingZongxian Li, Meikuan Liu, Guilai Han, Dagao Wang and Zonglin JiangZongxian Li https://orcid.org/0000-0002-7372-2552Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Meikuan Liu https://orcid.org/0000-0002-2900-8861Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Guilai Han https://orcid.org/0000-0001-6442-2577Chinese Academy of Sciences, 100190 Beijing, People's Republic of China, Dagao WangChinese Academy of Sciences, 100190 Beijing, People's Republic of China and Zonglin JiangChinese Academy of Sciences, 100190 Beijing, People's Republic of ChinaPublished Online:3 Sep 2024https://doi.org/10.2514/1.J064454SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookXLinked InRedditEmail About References [1] Camillo G. 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All requests for copying and permission to reprint should be submitted to CCC at www.copyright.com; employ the eISSN 1533-385X to initiate your request. See also AIAA Rights and Permissions www.aiaa.org/randp. TopicsAerodynamicsAeronauticsAerospace SciencesAerothermodynamicsBoundary LayersFluid DynamicsHeat FluxThermodynamic PropertiesThermodynamicsThermophysics and Heat TransferVortex DynamicsWind Tunnels KeywordsHeat Flux DistributionAdverse Pressure GradientAerodynamic HeatingHypersonic Boundary LayerWind Tunnel TestsBoundary Layer ControlAcknowledgmentsThis work was supported by the National Natural Science Foundation of China (12132017), the National Key Research and Development Program of China (2022YFB3207000), the Youth Cross Team CAS (JCTD-2022-02), and the Youth Innovation Promotion Association CAS (2020019).Digital Received29 May 2024Accepted31 July 2024Published online3 September 2024

Hypersonic stage separation is a significant process, probably involving complex aerodynamic interaction, which determines the survival of two-stage-to-orbit (TSTO) vehicles. The longitudinal stage separation (LSS) scheme is proposed to improve the safety of the parallel-staged TSTO vehicle, where a small interstage gap may result in weak interference. Therefore, an experimental and numerical study of LSS for the parallel-staged TSTO vehicle at Mach 7 with different angles of attack (AoA) is carried out. The dynamic interaction, including variations in the shock structure, wall pressure distribution, and unsteady aerodynamics, is investigated by testing and numerical simulation. The LSS experiments for the TSTO vehicle were performed using a high-speed pneumatic ejection launch system in the JF-12 shock tunnel, and the method was developed using high-speed visualization and image processing techniques to capture the separating trajectory. The numerical simulations were carried out using the overset grid method and solving the Navier–Stokes equations coupled with the rigid body dynamics equations to obtain the laminar flows over the TSTO vehicle during LSS. The qualitative and quantitative comparison of the test and numerical results showed good agreement in terms of aerodynamic performance, flowfield pattern, wall pressure, and separation trajectory. They show that the small interstage gap of the LSS leads to weak type I and VI shock–shock interactions, with short-duration weak shock reflection at a higher AoA. Furthermore, no shock reflection or interstage gap is observed at lower AoA. Moreover, no stage recontact is observed, and the safety and feasibility of LSS for parallel-staged TSTO vehicles are demonstrated.

The incident shock attenuation phenomenon in shock tube has received widespread interest because of its inevitable influence on experimental gas properties. However, few studies have investigated the cascading effects of the resulting nonuniformity on shock reflection in shock tunnels before the nozzle. This paper describes a numerical study on the unsteady reflection of a decelerating incident shock driven by a decelerating piston, as a simplification of the complex nonideal factors. The initial decay and uniform parameter distributions are generated based on the non-inertial frame. The results indicate that the existence of the nonuniform area forces the reflected shock wave region to undergo a transition process of attenuation and then stability. The final values of the gas parameters in zone 5 will, therefore, deviate from those given by the traditional relation for an ideal shock tube, which are only dependent on the terminal shock Mach number. This affects the determination of the total temperature T5, which is difficult to measure directly. We discuss the reconstruction of the nonuniform region with the attenuation trajectory of the incident shock wave and find that there is no one-to-one correspondence between this trajectory and the resulting nonuniformity, which introduces additional uncertainties to the predictions. Thus, an analogy method is developed through the multilevel block building algorithm, allowing the total temperature T5 to be determined using the total pressure p5 considering the effect of nonuniformity. Further assessments verify the applicability of this method in cases with multidimensional and viscous interactions.

When a force test is conducted in a shock tunnel,vibration of the Force Measurement System(FMS)is excited under the strong flow impact,and it cannot be attenuated rapidly within the extremely short test duration of milliseconds order.The output signal of the force balance is coupled with the aerodynamic force and the inertial vibration.This interference can result in inac-curate force measurements,which can negatively impact the accuracy of the test results.To elimi-nate inertial vibration interference from the output signal,proposed here is a dynamic calibration modeling method for an FMS based on deep learning.The signal is processed using an intelligent Recurrent Neural Network(RNN)model in the time domain and an intelligent Convolutional Neural Network(CNN)model in the frequency domain.Results processed with the intelligent mod-els show that the inertial vibration characteristics of the FMS can be identified efficiently and its main frequency is about 380 Hz.After processed by the intelligent models,the inertial vibration is mostly eliminated from the output signal.Also,the data processing results are subjected to error analysis.The relative error of each component is about 1％,which verifies that the modeling method based on deep learning has considerable engineering application value in data processing for pulse-type strain-gauge balances.Overall,the proposed dynamic calibration modeling method has the potential to improve the accuracy and reliability of force measurements in shock tunnel tests,which could have significant implications for the field of aerospace engineering.

Due to the complex aerodynamic interaction, the safe separation of two stages is one of the challenges for the successful launch of a two-stage-to-orbit (TSTO) vehicle. The unsteady hypersonic flow past the parallel-staged TSTO model during stage separation at Ma = 6.7 and Re = 8.86 x 10(5) m(-1) is numerically studied using laminar flow simulation. The TSTO model consisted of a waverider and a spaceplane as booster and orbiter, respectively. The effect of the center of gravity (CoG) of the orbiter on the unsteady aerodynamic interference during stage separation of TSTO is analyzed in detail with 0.65 = l(CoG)/l(o) = 0.80. In addition, the aerodynamic characteristics, dynamic behaviors, and unsteady wall pressure variation are compared in different cases. The results show that the CoG regime is limited to 5% of the orbiter length for absolutely safe separation, i.e., 0.70 < l(CoG)/l(o) < 0.75. As for the unsuccessful separation, the orbiter tends to fly nose-down if l(CoG)/l(o) = 0.70 while tending to pitch or somersault when l(CoG)/l(o) = 0.80. Furthermore, the pitching moment of the orbiter, which is influenced by the interstage shock wave-boundary layer interaction and shock-shock interaction, dominates the separation safety, and the specific flow mechanisms concerning the separation behavior associated with aerodynamic interference in different cases are analyzed in detail.

两级入轨(TSTO)飞行器或将成为下一代天地运输往返系统,其具有低成本、高效率和多用途等优点,但是两级分离成功与否将直接决定入轨任务的成败.目前的并联式TSTO飞行器多采用横向级间分离,该方法会在两级间产生复杂强气动干扰而直接增加了分离风险,所以探索一种可以避免或减弱两级强气动干扰的新分离方式是十分必要的.提出并着重分析了一种并联式TSTO纵向级间分离(LSS)方案,即轨道级在助推级背面沿着飞行方向分离,对其进行了动态分离过程的数值研究.针对新分离方案,设计了一种由宽速域乘波体和可重复使用空天飞机分别作为助推级和轨道级的TSTO组合飞行器,在高超声速条件下,采用重叠动网格技术分析了不同来流攻角(AOA)下的纵向分离流动机理、非定常壁面压力分布及气动特性变化规律.结果表明:TSTO纵向分离过程中仅存在VI型激波干扰和激波汇聚等简单的弱干扰类型,两级间无明显的激波反射或激波边界层干扰;非定常压力分布特性表明助推级前缘激波是轨道级受力变化的主要影响因素;纵向分离过程中,助推级受到的气动干扰力载荷小于轨道级.此外,不同来流攻角下,两级气动干扰流场结构具有相似性,并给出了实现安全纵向分离的攻角条件.

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.

高精准度气动力测量是激波风洞试验中的关键技术.在开展测力试验时,测力系统在风洞流场起动瞬间的冲击激励下产生振动,但振动信号无法在较短的有效试验时间内快速衰减,导致天平输出信号中耦合了惯性干扰.基于深度学习技术,对激波风洞天平信号在频域内开展数据处理,并针对动态信号的频域特征进行卷积神经网络建模分析,旨在消除测力信号中的惯性干扰.在频域模型训练样本和验证样本的结果分析中,天平信号的大幅惯性振动干扰被消除,达到预期的结果,验证频域建模分析方法的有效性和可靠性.此外,对处理结果进行误差分析,进一步验证该方法在激波风洞天平数据处理中具有较好的工程应用价值.

Standing oblique detonation is a unique pressure-gain combustion phenomenon for hypersonic ramjet propulsion, and its research has been related with supersonic combustion in scramjet engines since its births, for example, absent treatment in its early stage and re-consideration in recent decades. Standing oblique detonations and supersonic combustion share the same features of supersonic chemically-reacting flows, and can be considered as different flow development stages. Combustion instability in a chemically-reacting flow is reviewed first to identify its fundamental mechanisms, and the upstream-propagating shock wave is identified as one of intrinsic characteristics and taken as the key problem for developing hypersonic ramjet propulsion. Critical conditions for the standing oblique detonation are summarized as a theoretical base for standing oblique detonation ramjet engines. Three key parameters are included, that is, the maximum heat that can drive local flow states from supersonic to sonic after combustion, the critical inflow Mach number of combustors, at which supersonic combustion becomes stable, and the critical wedge angle at which a standing oblique detonation can be initiated. The evolution of the standing oblique detonation is reviewed by placing emphasis on its complex wave structure that was found to develop via three stages, that is, shock-induced initiation, the decaying stage and the fully-developed stage. Finally, progress in experimental research is reviewed with detailed discussions on stabilization of the standing oblique detonation, experimental methods and development of adequate test facilities. In conclusion, the stable operation of hypersonic ramjet propulsion is a critical issue to approach its engineering application, and the standing oblique detonation ramjet engine is recommended as a promising candidate, deserving more attention in the future.

A longitudinal stage separation (LSS) scheme for a parallel-arrangement two-stage-to-orbit (TSTO) vehicle is proposed and analyzed in detail, in which the orbiter moves along the upper surface of the booster. A TSTO concept comprising a waverider and a spaceplane was designed to numerically investigate the dynamic characteristics of the LSS at Mach 7. The influence of spike models assembled at the nose of the orbiter on the LSS at different angles of attack (AoA) (i.e., spike with half cone, spike with half cone-disk, and hemispheric spike) was explored. Moreover, the aerodynamic interference and characteristics were analyzed and compared for different spiked configurations. The aerodynamic interference during LSS is simple and weak, which is only associated with type VI shock/shock interaction, with rapid increases in the axial force when the shock waves of both stages converge. Furthermore, the model with a half cone spike has the best performance in drag reduction by 7%, whereas the model with a half cone-disk spike has the worst performance that increases axial force. The TSTO model with a half cone spike at [Formula: see text] and 5 deg is advantageous in LSS because of the high drag reduction and weak aerodynamic interference.

Air-blowing is one of the techniques for active flow control and thermal protection system of hypersonic vehicles. Introducing air into the hypersonic boundary layer alters the cross-sectional profile of the boundary layer, thereby influencing the boundary-layer transition. This study investigates the active air-blowing control effects on the hypersonic flat-plate boundary layer under various blowing mass flow rates and incoming Mach numbers by solving the Reynolds-averaged Navier–Stokes equations with the Langtry–Menter four-equation transitional shear stress transport model. The study examined alterations in the blowing boundary-layer profiles under two conditions: natural and bypass transition, induced by different blowing flow rates. Blowing significantly alters the sonic line and boundary-layer profile characteristics, triggering blowing oblique shock and causing alterations in the instability mechanisms of the two transition states. A higher Mach number intensifies compressibility effects, stabilizing the boundary layer and leading to an increase in the thickness of the blowing boundary layer and air film.

Learn to design and improve state-of-the-art aerodynamic ground testing facilities in this comprehensive reference book, with particular focus on high-enthalpy shock tunnels. Including the latest advances in detonation-driven high-enthalpy shock tunnels, readers will discover how to extend test time with brand new concepts and duplicate real hypersonic flight test conditions. Through a systematic approach, the book describes technologies for a variety of different drivers in hypersonic and high-enthalpy shock tunnels. The fundamental theories for hypersonic and high-enthalpy shock tunnels are described step-by-step, with examples throughout, providing an accessible introduction. Built on years of real-world experience, this book examines in detail the advantages and challenges of improving test flow qualities, including increasing total pressure and enthalpy, model scale amplification and test-time extending for different types of shock tunnel drivers. This is an ideal companion handbook for aerospace engineers as well as graduate students.

The interactions between oblique and bow shock waves are significant problems related to the aerodynamic force and thermal performance of hypersonic vehicles, but few studies have considered the dynamic effect of the body's motion on the phenomena. In this work, a numerical study on the oblique and bow shock waves ahead of an elliptic cylinder rotating with a forced-oscillation approach was conducted at Mach 5 by solving the unsteady, two-dimensional Navier–Stokes equations in a non-inertial coordinate system. The hysteresis loops of aerodynamic coefficients were analyzed first, and it was found that the moment is sensitive to rotation. Then, two different hysteresis forms were found at positive and negative angles of rotation (AOR), corresponding to cases with the interference point above or below the wall, respectively. When AOR is positive, the rate-dependent transition hysteresis among various shock interaction types causes the movement of strong flow structures (reflected shock wave, Mach stem, and jet) to always lag behind the body's motion. When AOR is negative, besides the evolution hysteresis of flow structures, two unusual patterns between Edney Types III and VI were observed on different transition paths, which led to very different peak pressures. Also discussed are the driving mechanisms associated with the effect of the subsonic region and the downstream boundary of the interaction zone, as well as the modulating action of the formed virtual Laval flow channel. Additional simulations were performed to study the effect of rotation speed on the transition boundary and the transition structures between Types III and VI.

As the most important measuring device in aerodynamic testing, the wind tunnel balance is used to measure the magnitude, direction, and point of the aerodynamic loads (forces and moments) acting on the test model. The accuracy of the measurement is directly related to the static calibration of the wind tunnel balance, which establishes the mapping relationship between the balance output signals and aerodynamic loads on the calibration equipment. This paper explores the possibility of improving the calibration performance of the strain-gauge balance in the calibration system AiBCS, developed by Institute of Mechanics of Chinese Academy of Science, using the convolutional neural network (CNN). The applicable conditions, validity, and reliability of CNN in the balance calibration are discussed and evaluated. Results obtained by the CNN-based calibration method and the traditional polynomial fitting method are analyzed and compared. It turns out that the CNN-based calibration method can effectively reduce the load interference between various balance components, yielding a significantly improved performance. Consequently, the deep-learning technology shows great application potential in calibrating wind tunnel balance.

Hypersonic liquid film cooling technology is to press out the cooling medium through a series of slits or holes, creat a low-temperature cooling film in the boundary layer of the surface of the aircraft to prevent the aerodynamic heating of the aircraft by hypersonic airflow. As an active cooling method, it has great application potential in surface thermal protection of hypersonic vehicle.In this paper, numerical methods and VOF model are used to study the spreading of liquid film at 25km flight altitude and Ma5 airflow. The evolution process and cooling mechanism of liquid film on a flat plate are discussed through the incident velocity, Angle, surface tension and viscosity coefficient of different cooling medium. The results show that under the action of air flow, the liquid film develops downstream to the wall surface, the existence of the liquid film leads to the boundary layer separation, and the continuous liquid film will be broken into liquid blocks at a certain position,and then further broken into droplets. The change of incident conditions and liquid properties will affect the development of the liquid film along the flow direction, which is manifested in the position of the fracture point and the thickness of the continuous liquid film. Within the computational domain set in this paper, the wall heat flow is reduced by 80～95 percent, and the cooling efficiency of the liquid film on the wall varies with the the change of the liquid film morphology.