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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.

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.

Pressure-gain combustion has gained attention for airbreathing ramjet engine applications owing to its better thermodynamic efficiency and fuel consumption rate. In contrast with traditional detonation induced by a single wedge, the present study considers oblique shock interactions attached to double wedges in a hypersonic combustible flow. The temperature/pressure increases sharply across the interaction zone that initiates an exothermic reaction, finally resulting in an Oblique Detonation Wave (ODW). Compared with the case for a single-wedge ODW, the double-wedge geometry has great potential to control the initiation of the ODW. As a tentative study, two-dimensional compressible Euler equations with a two-step induction-reaction kinetic model are used to solve the detonation dynamics triggered by a double wedge. The effects of the wedge angles and wedge corner locations on the initiation structures are investigated numerically. The results show an ODW complex comprising three Oblique Shock Waves (OSWs), an induction zone, a curved detonation front, and an unburned/low-temperature gas belt close to the surface of the second wedge. Both the increasing wedge angle and downstream wedge corner location lead to an abrupt OSW-ODW transition type, whereas the former corresponds to the shock–shock interaction and the later has a greater effect on the exothermic chemical process. Analysis of the shock polar and flow scale confirms that the OSW-ODW initiation structure mainly depends on the coupling of shocks and heat release in a confined initiation zone.

This study investigates the flowfield patterns and distributions of surface heat flux due to interactions among three-dimensional shock waves at the junction of the body and wing of an aircraft by ...

Hypersonic high-enthalpy wind tunnels have been a challenge to ground tests in aerospace research area for decades and its test flow uniformity is one of the most important parameters for evaluating wind tunnel performances. Regarding to the performance requirement, theories and methods for designing hypersonic flow nozzles at high enthalpy conditions are quite difficult, but very interesting topics, especially when air molecule dissociations take place in wind tunnel test gas reservoirs. In this paper, fundamental theories and important methods for nozzle designs are briefly reviewed with the emphasis on two-dimensional axisymmetric nozzles for hypersonic high-enthalpy wind tunnels, including the Method of Characteristics (MOC), the graphic design method, the Sivells method, the theory for boundary correction, and the CFD-based design optimization methods. These theories and methods had been proposed based on several physical issues, respectively, which play important roles in nozzle flow expansion processes. These issues cover the expansion wave generation and reflection, the boundary layer development, the real gas effect of hypersonic high-enthalpy flows. Difficulties arising from applications of these methods in hypersonic high-enthalpy nozzle design are discussed in detail and the state of the art of the nozzle design technologies that have reached for decades is summarized with some brief comments. Finally, the prospect for the hypersonic nozzle design methods, and its numerical and experimental verifications are provided with from authors’ viewpoint for readers’ reference.

A novel α-cyanostilbene-containing tetraphenylethylene-based discotic LC with 3,4,5-tridecyloxy benzene units at the terminal was synthesized by Ti-catalyzed McMurry coupling, Pd-catalyzed Suzuki c...

Hypersonic airbreathing propulsion is one of the top techniques for future aerospace flight, but there are still no practical engines after seventy years' development. Two critical issues are identified to be the barriers for the ramjet-based engine that has been taken as the most potential concept of the hypersonic propulsion for decades. One issue is the upstream-traveling shock wave that develops from spontaneous waves resulting from continuous heat releases in combustors and can induce unsteady combustion that may lead to engine surging during scramjet engine operation. The other is the scramjet combustion mode that cannot satisfy thrust needs of hypersonic vehicles since its thermos-efficiency decreases as the flight Mach number increases. The two criteria are proposed for the ramjet-based hypersonic propulsion to identify combustion modes and avoid thermal choking. A standing oblique detonation ramjet (Sodramjet) engine concept is proposed based on the criteria by replacing diffusive combustion with an oblique detonation that is a unique pressure-gain phenomenon in nature. The Sodramjet engine model is developed with several flow control techniques, and tested successfully with the hypersonic flight-duplicated shock tunnel. The experimental data show that the Sodramjet engine model works steadily, and an oblique detonation can be made stationary in the engine combustor and is controllable. This research demonstrates the Sodramjet engine is a promising concept and can be operated stably with high thermal efficiency at hypersonic flow conditions.

Hypersonic air-breathing propulsion is one of the key techniques for future aviation and the ground aerodynamic testing for full scale test models with sufficient test time at flight conditions is of fundamental importance for verifying hypersonic air-breathing engines. Based on the backward detonation-driven concept, the hypersonic flight-duplicated shock tunnel (or JF-12 shock tunnel) has been successfully constructed and calibrated. This facility is capable of reproducing airflow for Mach numbers ranging from 5 to 9 at an altitude of 25–50 km, with a test duration of more than 100 ms. To quantify the performance of the shock tunnel, experiments were conducted to investigate the aerodynamic characteristics of the test flows and the effects of several critical techniques that play important roles in the operation of the shock tunnel. The stagnation pressure was constant within ±5 $$\%$$ and the average stagnation pressure varied by less than 0.048 $$\%/$$ ms. The variation of the stagnation pressure in repeated experiments is less than 2.0 $$\%$$ , indicating the good repeatability of the wind tunnel. The non-uniformity of the Mach number in the core flow field at the nozzle exit was within ±2.5 $$\%$$ . Additional, a uniform flow field is established upstream of the nozzle exit. The axial gradients of the flow field are small since the Mach number varies less than 1.7 $$\%/$$ m. Findings regarding the ignition technology, diaphragm ruptures, detonation driver capacity, incident shock-wave decay, and tunnel operation mode are also presented. The findings of this study are not only helpful for operating the shock tunnel, but can also assist the future development of hypersonic wind tunnels.

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 strong background radiation in high enthalpy hypersonic shock tunnels has posed severe challenges for measurement using luminescent coatings. We proposed a solution for reducing background radiation from time-resolved temperature-sensitive paint (TSP) measurement in a hypersonic flow with Ma = 6.5 and T0 = 3525 K. The TSP was applied on an inlet ramp model, and the images were taken by a high-speed camera at 2 kHz under a modulated excitation. The strong background radiation led to a low signal-to-noise ratio and significant errors for the first half of the 130-ms test duration. Accordingly, three noise reduction methods were developed and evaluated based on temporal reconstruction, spatial reconstruction and robust principal component analysis (RPCA), respectively. The RPCA method showed the best performance that successfully recovered high-quality TSP data for a majority of test duration (t ≥ 40 ms).

Hypersonic and high-enthalpy wind tunnels and their measurement techniques are the cornerstone of the hypersonic flight era that is a dream for human beings to fly faster, higher and further. The great progress has been achieved during the recent years and their critical technologies are still in an urgent need for further development. There are at least four kinds of hypersonic and high-enthalpy wind tunnels that are widely applied over the world and can be classified according to their operation modes. These wind tunnels are named as air-directly-heated hypersonic wind tunnel, light-gas-heated shock tunnel, free-piston-driven shock tunnel and detonation-driven shock tunnel, respectively. The critical technologies for developing the wind tunnels are introduced in this paper, and their merits and weakness are discussed based on wind tunnel performance evaluation. Measurement techniques especially developed for high-enthalpy flows are a part of the hypersonic wind tunnel technology because the flow is a chemically reacting gas motion and its diagnosis needs specially designed instruments. Three kinds of the measurement techniques considered to be of primary importance are introduced here, including the heat flux sensor, the aerodynamic balance, and optical diagnosis techniques. The techniques are developed usually for conventional wind tunnels, but further improved for hypersonic and high-enthalpy tunnels. The hypersonic ground test facilities have provided us with most of valuable experimental data on high-enthalpy flows and will play a more important role in hypersonic research area in the future. Therefore, several prospects for developing hypersonic and high-enthalpy wind tunnels are presented from our point of view.

Due to the involvement of powder materials in dust explosion hazards and detonation experiments, it is imperative to analyze the heterogeneous detonation in a polydisperse suspension with a continu...

When aircraft enter/reenter the atmosphere at near and super-orbital speeds, such aircraft need to endure challenges posed by a hypervelocity flow environment. High enthalpy gas generated by strong shock wave compression causes violent aerodynamic heating to such aircraft. At this time, accurately predicting the aerodynamic thermal load becomes extremely difficult. Expansion tubes/tunnels are one of the few types of qualified ground facilities that can simulate hypervelocity flow. The JF-16 expansion tunnel at the Institute of Mechanics, Chinese Academy of Sciences, has successfully generated hypervelocity flows with speeds of more than 10 km/s. Accordingly, experiments have been conducted measuring the hypervelocity aerothermal and wall catalytic activity. The stagnation heat flux of a spherical model at high temperature and strong scour test flow with near and super-orbital velocities was successfully obtained at microsecond resolutions. It was found that the experimental stagnation heat flux is higher than the predicted levels according to the empirical formula and CFD simulations when the total enthalpy of the test gas is greater than 45 MJ/kg. In addition, the experimental stagnation heat flux based on a fully catalytic wall (with a Cu coating) is 53.93% higher than that based on a non-catalytic wall (with an Al2O3 coating). The numerical analysis indicated that 22.5% O and 37.3% N atoms remain at the stagnation point of the model under the non-catalytic condition. Meanwhile, under the fully catalytic condition, the atoms are all combined and cause a significant increase in the heat flux at the stagnation point. This analysis confirmed the experimental results.

In this study, oblique-shock/bow-shock interference is theoretically and numerically studied with two incident shock waves. The transition criteria between the two modes of multiple shock–shock interference, i.e., the concomitant-jet (CJ) and dual-jet (DJ) modes, are given. The oblique shock relationship and shock polar analysis are utilized to obtain the analytical solution of the transition condition. The theoretical results indicate the existence of a dual solution interval (DSI) that widens with increasing Mach number and narrows with increasing deflection angle induced by the first incident shock wave. The DSI obtained by numerical simulation is considerably narrower than that theoretically predicted due to the advanced CJ→DJ and DJ→CJ transitions. The analysis reveals that the transitions are advanced due to the downstream disturbance and secondary waves in the flow field.

The stationary characteristics of the oblique detonation wave (ODW) induced by the double wedge with an expansion corner are investigated using two-dimensional Navier–Stokes equations along with a two-step induction-exothermic kinetic model. The results show that the detached ODW can be reattached by expansion waves induced by the double wedge so that the standing window of ODW can be expanded. The re-standing position of ODW depends on the location and strength of the expansion waves, which are governed by the first wedge length L and the corner angle between the first and second wedge surface θC. There is a critical angle reattachment that determines whether the ODW can be reattached by expansion waves, and this critical angle increases as wedge length increases. However, the detached ODW cannot be reattached when the wedge length is increased to a critical value regardless of the wedge corner. The re-standing position moves downstream with the increment of θC until the last Mach wave tangent to the subsonic zone behind the strong overdriven ODW because no more Mach waves interact with the initiation zone. Moreover, the comparison of viscous and inviscid fields demonstrates that a shorter wedge length is necessary for the viscous field to reattach the ODW because the recirculation zone forms a gas wedge that extends the first wedge surface.

In this Letter, the dynamic effect of rotation on unsteady shock interaction between an oblique shock wave and bow shock ahead of an elliptic cylinder is numerically investigated. First, a non-deforming grid method based on a non-inertial reference frame is applied for supersonic/hypersonic flow around an arbitrarily moving body. A simulation considering a rotating elliptic cylinder is then conducted at Ma5 with forced-oscillation approach. The results show that the shock interaction structure during rotation has obvious hysteresis compared with the static condition, and transition among different types of shock interaction is also affected by rotation direction and speed. The mechanisms are briefly addressed.

Symbolic regression (SR), as a special machine learning method, can produce mathematical models with explicit expressions. It has received increasing attention in recent years. However, finding a concise, accurate expression is still challenging because of its huge search space. In this work, a divide and conquer (D&C) scheme is proposed. It tries to divide the search space into a number of orthogonal sub-spaces based on the separability feature inferred from the sample data (dividing process). For each sub-space, a sub-function is learned (conquering process). The target model function is then reconstructed with the sub-functions according to their separability patterns. To this end, a separability pattern detecting technique, bi-correlation test (Bi-CT), is also proposed. Note that the sub-functions could be determined by any of the existing SR methods, which makes D&C easy to use. The D&C powered SR has been tested on many symbolic regression problems, and the study shows that D&C can help SR to get the target function more quickly and reliably.