Search results

The purpose of this paper is to study the effects of boundary‐layers bleeding on performance parameters of hypersonic inlets.

The inner flowfield of a hypersonic inlet at different bleeding rates is simulated with a Reynolds‐averaged Navier‐Stokes solver using a renormalization group k‐ε turbulence model.

In contrast with no bleeding, the performance parameter of hypersonic inlets without backpressure is reduced slightly, but the flow uniformity is improved. The interaction between boundary layers and shocks is weakened at the action of the bleeding, which leads to that the boundary‐layers separations at the entrance of the isolator caused by the high‐backpressure occur later, and it can improve the maximum backpressure ratio of hypersonic inlets. With the bleeding rate increasing, the maximum backpressure ratio of hypersonic inlets is added, while the total‐pressure recovery coefficient and mass‐captured coefficient are reduced.

This paper is a useful reference to the design and performance improvement of hypersonic inlets and propulsion systems.

PARALLEL with the quest for higher speeds in aircraft of the traditional type, there has been a remarkable advance in rocket and missile design in recent years. The German legacy of the V2 gave both East and West a long start in this development, and full advantage of this has been taken in the period of intensive research of the past ten years. Lately the possibility of space travel has made a very profound appeal to people in all spheres and this has caused the pace of research in this field to be stepped up.

This paper describes the numerical solutions of type‐IV shock‐on‐shock interactions in hypersonic thermochemical nonequilibrium air flows around blunt bodies. The Navier‐Stokes equation solver for a chemically reacting and vibrationally relaxing gas mixture was applied to the present problem, where the concepts of the Advection Upstream Splitting Method (AUSM) and the Lower‐Upper Symmetric Gauss‐Seidel (LU‐SGS) method were basically employed along with the two‐temperature thermochemical model of Park. The aerodynamic heating with or without the shock‐on‐shock interaction to a sphere and circular cylinders are simulated for a hypersonic nonequilibrium flow. The numerical results show that typical type‐IV shock‐on‐shock interaction pattern with a supersonic jet structure is also formed in a high‐enthalpy thermochemical nonequilibrium flow, and that the contribution of convective heat flux in the translational/rotational mode to the total heat flux is dominant. Furthermore, the inherent unsteadiness of nonequilibrium type‐IV shock‐on‐shock interaction flowfield is discussed briefly.

The current interest in hypersonic flows is leading to significant effort both to develop CFD methods and also to provide experimental data for their evaluation. In our research we attempt to integrateCFD and experiments as closely as possible so much so that most of our experimental model designs are based upon preliminary flow field computations in order to identify likely regions of importance and distribute instrumentation as efficiently as possible. The experiments must also have the CFD requirements clearly in mind. In particular we consider it important to separate evaluation on the numerics (essentially the algorithm) from modelling of the physics (which includes the uncertainties of turbulence modelling) to this end our experiments include laminar studies, for both attached and separated flows, for which the physical equations are known exactly, as well as turbulent flow studies. This paper concentrates mainly on our CFD efforts and presents details of a high resolution solver for viscous flows together with their predictions for a range of problems which are the subject of our current and planned experiments.

As to different angles of attack and nonlinear problems caused by high temperatures in coexisting hypersonic aircraft, people mainly rely on fluid software for research but lack analysis of flow mechanisms. Owing to computational difficulties, few people use numerical algorithms to combine them for discussion. Hence, this study aims to make a deep inquiry into the laminar flow and heat transfer of compressible Newtonian fluid in hypersonic aircraft with small attack angles.

In this paper, on the basis of mass, momentum and energy conservation laws, the governing equations of the hypersonic boundary layer are established. Viscosity, specific heat capacity and thermal conductivity are considered nonlinear functions concerning temperature. In virtue of the MacCormack finite difference method, the stationary numerical solutions are solved directly, and the validity of the algorithm is verified.

The results demonstrate that at Mach number 5, compared to the 0° attack angle, the maximum temperature near-wall at the 3° attack angle increases by about 25%. An enjoyable phenomenon is discovered, where the position corresponding to the maximum wall shear force shifts back as the attack angle and Mach number increase. The relationship between the near-wall maximum temperature versus attack angle and Mach number is fitted through numerical calculation results.

Empirical formulas can be used to estimate heat transfer characteristics at small attack angles, which will guide the design of aircraft thermal protection systems.

This study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.

Two geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.

The array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.

This paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

The purpose of this paper is to perform the simulation to explore the gap flow field under a hypersonic air flow. Thermal protection systems of hypersonic vehicles generally consist of thermal insulation tiles, and gaps between these tiles probably cause a severe local aerodynamic thermal effect.

The discretizations of convection flux term and temporal term in the governing equation with chemical equilibrium, respectively, take AUSM+-up flux-vector splitting scheme and the implicit lower-upper symmetric Gauss–Seidel method. Based on these, the flow field in a deep gap is simulated by means of the computer codes that the authors have written.

The numerical results show that the heat flux distribution in a gap has a good agreement with experimental results. Importantly, the distribution of heat flux is “U” shaped and the maximum of the heat flux occurs at the windward corner of a gap.

To explore the gap flow field under a hypersonic air flow, which is a chemically reacting, all speed and viscous flow, a novel model with an equivalent ratio of specific heats is presented. The investigation in this paper has a guide for the design of the thermal protection system in hypersonic vehicles.

A finite difference method for solving the quasi one‐dimensional non‐equilibrium hypersonic flow equations in a diverging nozzle is presented and discussed. In chemically reacting flows the system of equations to be solved is very stiff. Some reactions may be several orders of magnitude faster than others and generally, they are much faster than the convective process except for very high Ma numbers. For this reason the development of a numerical scheme whose stability is independent of the chemical reaction rates is of importance. The main advantage of this scheme is the conservation of each chemical component, the positivity of densities and vibrational energies, as well as its relative simplicity, which results in a fast computer code.

To develop a high‐order compact finite‐difference method for solving flow problems containing shock waves.

A numerical algorithm based on high‐order compact finite‐difference schemes is developed for solving Navier‐Stokes equations in two‐dimensional space. The convective flux terms are discretized by using advection upstream splitting method (AUSM). The developed method is then used to compute some example laminar flow problems. The problems considered have a range of Mach number that corresponds to subsonic incompressible flow to hypersonic compressible flows that contain shock waves and shock/boundary‐layer interaction.

The paper shows that the AUSM flux splitting and high‐order compact finite‐difference methods can be used accurately and robustly in resolving shear layers and capturing shock waves. The highly diffusive nature of conventional flux splitting especially on coarse grids makes them inaccurate for boundary layers even with high‐order discretization.

This paper presents a high‐order numerical method that can accurately and robustly capture shock waves without deteriorating oscillations and resolve boundary layers and shock/boundary layer interaction.

With the continuous increase of space activities and insight into space exploration, the velocity of near space hypersonic vehicles becomes higher and higher, which leads to the aerothermodynamic phenomenon getting worse around a vehicle; therefore, the exploration of numerical scheme applicability is essential for hypersonic flow simulations.

The implicit finite volume schemes are derived from axisymmetric Navier–Stokes equations for chemical equilibrium flow and programmed in Fortran. Taking the atmosphere at 30 km as an example, the performance of spatial discretization schemes such as AUSMPW and AUSMPW+ are analyzed in a range of Mach numbers from 17 to 32.

The AUSMPW scheme appears pressure jump near the stagnation if the Mach number is over 18, but AUSMPW+ scheme shows better performance in comparison.

This study will help the aerothermodynamic design in near space hypersonic vehicles.