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The forward shock wave induced by the detonation wave propagates in the inlet and restrains the air from being injected into the rotating detonation engine, and the stable continuous propagation of the detonation wave is affected. Numerical simulation and theoretical derivation of air flow velocity and detonation wave height were carried out in a rotating detonation combustor. The influence of the forward shock wave on air flow velocity and air flow velocity on detonation wave height was analyzed. A non-premixed air-breathing rotating detonation combustor was numerically simulated with the two-dimensional Navier–Stokes equations and a kerosene/air reaction model based on the RYrhoCentralFoam solver. Analytical equations were derived and validated to calculate the air flow velocity across the forward shock wave and the height of the detonation wave. The results show that the air flow velocity across the forward shock wave is inversely proportional to the detonation wave velocity, incoming flow temperature, and pressure ratio at the forward shock wave. In addition, the angle of the forward shock wave is inversely proportional to the velocity of the detonation wave and incoming flow and is proportional to the incoming flow temperature and the pressure ratio. The forward shock wave leads to a blocked zone of injection in the combustor, as the inflow velocity decreases or even reverses across the shock wave. As a result, the forward shock wave induced by the detonation wave restrains the recovery of the reactant fill zone, which has an influence on the height of the detonation wave. [ABSTRACT FROM AUTHOR]

When a helicopter rotor undergoes flow separation, the drag of the rotor increases substantially, as does the power demand, which seriously affects the aerodynamic performance and flight safety of the helicopter. Therefore, it is crucial to research how to suppress the flow separation of rotor blades. An active control technique based on a coflow jet (CFJ) at the rotor blade tip was employed in this study to suppress the flow separation over the rotor. The mechanisms and behavior were investigated. The rotor flow field was numerically simulated by solving the Reynolds-averaged Navier–Stokes equations with the finite volume method. The turbulence model was k − ω SST, and the rotor motion was simulated using the overset mesh technique. After applying CFJ control, the airflow from the injection slot at the leading edge of the rotor increased the energy of the mainstream in the near-wall area, which enhanced its ability to resist the adverse pressure gradient. A flow separation was effectively suppressed, both on the advancing and retreating sides, which improved the aerodynamic performance of the rotor. During the whole rotation period, the thrust coefficient of the rotor increased by up to 5.6%, the moment coefficient decreased by as much as 26.8%, and the equivalent lift-to-drag ratio increased by up to 44.0%. Moreover, the effects of the CFJ parameters on the flow separation suppression of the rotor are researched. These results may provide a foundation for the development of aerodynamic performance improvement for helicopter rotor based on a CFJ. [ABSTRACT FROM AUTHOR]

The oblique detonation induced by a two-dimensional semi-infinite wedge is simulated numerically with the Navier–Stokes equations and a detailed H2/air reaction model based on the open-source program-Adaptive Mesh Refinement in Object-oriented C++. A spatially seventh-order-accurate weighted essentially non-oscillatory scheme is adopted for the convective flux discretization. The formation and evolution of the oblique detonation induced by wedges at different angles and inflow conditions are investigated, and a prediction model for the oblique detonation flow field is proposed. The results show that the formation of the oblique detonation flow field can be divided into two processes. The first process is similar to the oblique shock flow field with unreactive inflow. When the inflow passes through the wedge, the oblique shock wave starts to form at the tip, followed by the unstable curved shock surface and triple point. In this process, a thin reaction layer is formed on the wedge front, but the thickness of the reaction layer is almost constant. The second process is similar to the process of deflagration to detonation. As the reaction rate increases, the deflagration front is fixed on the wedge, the reaction layer thickens, and the deflagration front gradually approaches the oblique shock wave. When the deflagration front is coupled with the oblique shock wave, the oblique detonation is formed. Moreover, a theoretical prediction model for the triple point location is proposed. Compared with the numerical simulation results, the theoretical model prediction for the position of the transition point of the oblique shock wave–oblique detonation wave is relatively acceptable. [ABSTRACT FROM AUTHOR]

In this work, a discrete gas-kinetic scheme (DGKS) based on the arbitrary Lagrangian–Eulerian (ALE) method is proposed for the simulation of moving boundary problems. The governing equations are the ALE-based Navier–Stokes equations, which are discretized using the finite volume method. Starting from a circular function-based Boltzmann equation, a grid motion term is introduced to obtain the Boltzmann equation in ALE form. Based on the moment relations and Chapman–Enskog analysis, the moment of particle velocity and distribution function are summed to obtain the fluxes. The DGKS expression in the ALE framework can then be derived. In this method, the flux at the cell interface can be calculated from the local solution of the Boltzmann equation, which is physically realistic and makes the algorithm more stable. As DGKS is based on a multidimensional particle velocity model, it is not necessary to use approximate values for the reconstruction process. In addition, DGKS can simultaneously handle inviscid and viscous fluxes when simulating viscous flow problems, resulting in a higher degree of consistency. Finally, several moving boundary examples are simulated to validate the ALE-DGKS method. The results show the algorithm was observed to achieve second-order accuracy and can solve moving boundary problems effectively. [ABSTRACT FROM AUTHOR]

Abstract: The low momentum flux ratio jet in the HyShot II scramjet combustor is studied by DES (Detached Eddy Simulation) and RANS (Reynolds-Averaged Navier–Stokes) methods. The flow structure near the injector, shock pattern in the symmetry plane as well as the instantaneous coherent structures are presented and explained. Further insight into the flow physics is obtained by visualizing instantaneous coherent structures. The formation of Ω-shaped vortices, which was previously observed in experiments but never well-studied numerically, is discussed in detail. A new schematic of flow physics is proposed to enhance the understanding of the low momentum flux ratio jet. Compared to the DES result, the RANS method is unable to capture the dynamics of turbulent structures. The DES method provides much detailed information about mixing patterns and a more reliable mixing efficiency than the RANS result. The RANS method over-predicts the eddy-viscosity during turbulence modeling and suppresses unsteady turbulent fluctuations by time averaging, which results in a 25% over-estimation of the mixing efficiency. [Copyright &y& Elsevier]

A linear lattice Boltzmann flux solver (LLBFS) is developed based on the high-precision finite-volume method (FVM) and is used to simulate problems involving acoustic propagation. The least-squares-based finite difference (LSFD) is used to construct the high-precision FVM, which discretizes the linear Navier–Stokes equations (LNSE). In terms of the flux calculation for the interface of the control cell, the LLBFS is derived by comparing the macroscopic and microscopic variables through the Chapman–Enskog (C-E) multi-scale analysis, which establishes the relationship between the linear lattice Boltzmann equation (LBE) and the LNSE. The algorithm has the following advantages. First, the proposed LLBFS inherits the advantages of the lattice Boltzmann flux solver (LBFS), i.e., the local solution of the lattice Boltzmann equation is used to calculate the fluxes; handling the inviscid and viscous fluxes simultaneously makes for a more consistent algorithm and it can be expanded to the multi-dimensional scheme without approximation. Second, solving the LNSE using the high-precision FVM on unstructured grids, which enables problems involving acoustic waves interacting with complex geometries can be simulated conveniently. To validate the accuracy and robustness of the proposed method, several cases are simulated, including acoustic waves propagating in an inviscid and viscous mean flow and interacting with complex geometries. [ABSTRACT FROM AUTHOR]

Hypersonic vehicles with Mach numbers greater than five exhibit complex shock wave interactions, accompanied by locally enhanced heat flux near the shock-affected surface. To investigate the unclear interstage thermal environment in a Two-Stage-To-Orbit system, a model composed of a wide-speed-range vehicle and a reusable rocket is employed in an aerothermodynamic study of the Two-Stage-To-Orbit system considering the interstage interactions performed at freestream Mach number = 6. The innovations of this paper are the complex model, which is highly similar to the real situation, and the study of the thermal environment between stages near different reflection positions, whereby the shock wave interference that causes different forms of heat flux is divided into two categories. The thermal design of the Two-Stage-To-Orbit system should especially consider the head of the wide-speed-range vehicle, interstage distance, and the swept angle of canted fins and wings, as well as their width. This paper adopts the second-order AUSMDV (a variant of the Advection Upstream Splitting Method) scheme for spatial discretization and Lower–Upper Symmetric Gauss–Seidel method for time discretization in formulating the Reynolds-averaged Navier–Stokes equations, which are then solved by the finite volume method and Menter's shear stress transport k- ω turbulence model. Numerical analysis sheds light on the complex shock structure and the reasons for the considerable increase in wall temperature and aerothermal loads at critical parts of various reflected shock positions. The incident shock wave generated by the head apex of the rocket constantly reflects and gradually weakens between the two stages, which brings about shock wave/boundary-layer interaction and gives rise to elevated heat flux. The moderate incident shock and the weak reflected shock are primarily responsible for the lack of boundary-layer separation, although the expansion caused by the convex curvature of the head of the wide-speed-range vehicle also contributes. A non-reflected part of the incident shock wave interacts with the windward surface of the canted fins and wings, producing shock wave/boundary-layer interaction and a V-shaped heat flux distribution that extends to the leading edges of the wings. As the angle of attack decreases, the intensity of the incident shock and the reflected shock gradually weakens, leading to reduced heat flux on the windward surface of the wide-speed-range vehicle. The range over which the incident shock impinges the walls also narrows, corresponding to the smaller heat flux distribution. These aerothermal studies comprehensively determine the law of thermal changes, which is significant for the design of thermal protection in Two-Stage-To-Orbit systems. • The interstage thermal environment of real parallel vehicle is studied. • The shock wave causes different forms of heat flux at various reflection positions. • The shock wave interaction is divided into two categories. • The incident shock wave reflects between stages, leading to elevated heat flux. • A non-reflected part of the incident shock causes a V-shaped heat flux distribution. [ABSTRACT FROM AUTHOR]

Numerical simulation and analysis are carried out on interactions between a 2D/3D conical shock wave and an axisymmetric boundary layer with reference to the experiment by Kussoy et al. [AIAA J, 1980, 18(12): 1477–1487], in which the shock was generated by a 15° half-angle cone in a tube at 15° angle of attack (AOA). Based on the Reynolds-averaged Navier–Stokes equations and Menter's SST turbulence model, the present study uses the newly developed WENO 3 − PRM 1 , 1 2 scheme and the PHengLEI CFD platform for the computations. First, computations are performed for the 3D interaction corresponding to the conditions of the experiment by Kussoy et al., and these are then extended to cases with AOA = 10° and 5° For comparison, 2D axisymmetric counterparts of the 3D interactions are investigated for cones coaxial with the tube and having half-cone angles of 27.35°, 24.81°, and 20.96° The shock wave structure, vortex structure, variable distributions, and wall separation topology of the interaction are computed. The results show that in 2D/3D interactions, a new Mach reflection-like event occurs and a Mach stem-like structure is generated above the front of the separation bubble, which differs from the model of Delery for 2D planar shock wave/boundary layer interaction. A new interaction model is established to describe this behavior. The relationship between the length of the circumferentially unseparated region in the tube and the AOA of the cone indicates the existence of a critical AOA at which the length is zero, and a prediction of this angle is obtained using an empirical fit, which is verified by computation. The occurrence of side overflow in the windward meridional plane is analyzed, and a quantitative knowledge is obtained. To further elucidate the characteristics of the 3D interaction, the scale and structure of the vortex and the pressure and friction force distributions are presented and compared with those of the 2D interaction. [ABSTRACT FROM AUTHOR]

The performance of rotating detonation engines is strongly affected by the multiple detonation waves that occur, although the reason for their formation has not yet been fully explained. In this paper, two-dimensional unfolded rotating detonation waves are simulated to investigate the effects of the equivalence ratios on the number of detonation waves. The compressible reactive Navier–Stokes equations with a detailed CH 4 /O 2 chemistry model are solved using third-order hybrid weighted essentially non-oscillatory/centered-difference numerical methods on a structured adaptive mesh. Compared with the typical two-dimensional rotating detonation wave structures in the homogeneous environment, the flow field is more unstable in the more realistic inhomogeneous environment. As the detonation propagates, coupling, decoupling, and further coupling occur due to the discontinuity of the detonation wave. The fuel mixing is enhanced by the flow field behind the wave, and the accumulation of fuel and the generation of local hot spots lead to the formation of multiple detonation propagation modes. As the equivalence ratio increases in the inhomogeneous environment, the number of detonation waves gradually increases. The intensity of the detonation waves, height of reactants, and heat release are particularly high when the equivalence ratio is close to 1.0. Moreover, equivalence ratios in the range 0.8–1.4 have little effect on the average thrust, but multiple detonation waves can be used to stabilize the thrust in the combustion chamber. [ABSTRACT FROM AUTHOR]