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A comparison of the accuracy of the central discretization scheme with artificial dissipation and the upwind flux‐difference TVD scheme has been made for the compressible Navier‐Stokes equations for high Reynolds number flows. First, a comparison is made on two one‐dimensional model problems. Then the schemes are compared on flat plate boundary layer flow. It is shown that a central scheme basically has poor accuracy due to the isotropic nature of the artificial dissipation. An upwind scheme decomposes the flow into different components and adapts the dissipation to the velocity of the components. The associated ansitropic dissipation results in a good accuracy. It is further discussed how a central discretization scheme with artificial dissipation can be improved at the expense of the same complexity of an upwind scheme.

The steady compressible Navier—Stokes equations coupled to the k—ε turbulence equations are discretized within a vertex‐centered finite volume formulation. The convective fluxes are obtained by the polynomial flux‐difference splitting upwind method. The first order accurate part results directly from the splitting. The second order part is obtained by the flux‐extrapolation technique using the minmod limiter. The diffusive fluxes are discretized in the central way and are split into a normal and a tangential contribution. The first order accurate part of the convective fluxes together with the normal contribution of the diffusive fluxes form a positive system which allows solution by classical relaxation methods. The source terms in the low‐Reynolds k‐ε equations are grouped into positive and negative terms. The linearized negative source terms are added to the positive system to increase the diagonal dominance. The resulting positive system forms the left hand side of the equations. The remaining terms are put in the right hand side. A multigrid method based on successive relaxation, full weighting, bilinear interpolation and W‐cycle is used. The multigrid method itself acts on the left hand side of the equations. The right hand side is updated in a defect correction cycle.

This paper aims to propose a precise turbulence model for vehicle aerodynamics, especially for vehicle window buffeting noise.

Aiming at the fact that commonly used turbulence models cannot precisely predict laminar-turbulent transition, a transition-code-based improvement is introduced. This improvement includes the introduction of total stress limitation (TSL) and separation-sensitive model. They are integrated into low Reynolds number (LRN) k-ε model to concern transport properties of total stress and precisely capture boundary layer separations. As a result, the ability of LRN k-ε model to predict the transition is improved. Combined with the constructing scheme of constrained large-eddy simulation (CLES) model, a modified LRN CLES model is achieved. Several typical flows and relevant experimental results are introduced to validate this model. Finally, the modified LRN CLES model is used to acquire detailed flow structures and noise signature of a simplified vehicle window. Then, experimental validations are conducted.

Current results indicate that the modified LRN CLES model is capable of achieving acceptable accuracy in prediction of various types of transition at various Reynolds numbers. And, the ability of this model to simulate the vehicle window buffeting noise is greater than commonly used models.

Based on the TSL idea and separation-sensitive model, a modified LRN CLES model concerning the laminar-turbulent transition for the vehicle window buffeting noise is first proposed.

This paper aims to explore the potential of transition prediction methods for modelling transitional shock wave/boundary layer interactions. The study is fuelled by the strong interest of researchers and airframe manufacturers in reducing the drag of vehicles flying at transonic speeds. The principle of drag reduction via flow laminarity is valid, provided there is no need for the flow to sustain large pressure gradients or shocks. This is true, as laminar boundary layers are less resistant to flow separation.

It is, therefore, worthwhile to assess the performance of CFD methods in modelling laminar boundary layers that can be tripped to turbulent just before an interaction with a shock. In this work, the CFD solver of Liverpool University is used. The method is strongly implicit, and, for this reason, the implementation of intermittency-based models requires special attention. The Navier–Stokes equations, the transport equations of the kinetic energy of turbulence and the turbulent frequency are inverted at the same time as the transport equations for the flow intermittency and the momentum thickness Reynolds number.

The result is stable and robust convergence even for complex three-dimensional flow cases. The method is demonstrated for the flow around the V2C section of the TFAST EU, F7 project. The results suggest that the intermittency-based model captures the fundamental physics of the interaction, but verification and validation are needed to ensure that accurate results can be obtained. For this reason, comparisons with the TFAST experiments is put forward as a means of establishing confidence in the transition prediction tools used for shock/boundary layer interaction simulation.

At the moment, experimental data for transonic transitional buffet are not yet available, although this will change in the near future.

The required CPU time is neither insignificant not prohibitive for routine computations.

Reducing aircraft drag without compromising on stall characteristics will result in lower fuel consumption and contribute to a greener and more economic flight for passengers.

To the authors’ knowledge, this is the first time that transitional buffet has been addressed.

This paper aims to provide a validation of a state‐of‐the‐art methodology for computing three‐dimensional transitional flows in turbomachinery.

The Reynolds‐averaged Navier‐Stokes equations for compressible flows are solved. Turbulence is modeled using an explicit algebraic stress model and k−ω turbulence closure. A numerical method has been developed, based on a cell‐centered finite volume approach with Roe's approximate Riemann solver and formally second‐order‐accurate MUSCL extrapolation. The method is validated versus two severe test cases, namely, the subsonic flow through a turbine cascade with separated‐flow transition; and the transonic flow through a compressor cascade with transitional boundary layers, shock‐induced separation and corner stall. For the first test case, the transition model of Mayle for separated flow has been employed, whereas, for the second one, the transition has been modeled employing the Abu‐Ghannam and Shaw correlation.

The comparison of numerical results with the experimental data available in the literature shows that, for such complex flow configurations, an improved numerical solution could be achieved by employing transition models. Unfortunately, the available models are case‐dependent, each of them being suitable for specific applications.

A state‐of‐the‐art numerical methodology has been developed and applied to compute very complex flows in turbomachinery. Through an original analysis of the results, the merits and limits of the considered approach have been assessed. The paper points up the fundamental role of transition modeling for turbomachinery flow simulations.

The purpose of this study is to provide the description of a computational methodology to model the combined propulsive systems of hydrogen propelled air-breathing scramjet vehicles and to evaluate the pollutant and climate-changing emissions.

Emissions indexes of nitrogen oxide (EINO) and water vapour released by the air turbo rocket (ATR) and dual mode ramjet (DMR) engines of the STRATOFLY air-breathing, hypersonic scramjet vehicle, propelled by hydrogen/air were evaluated. ATR engine operation was assessed for several cruise conditions in both subsonic and supersonic flight regimes in Ecosimpro software, which is an object-oriented thermodynamic design and simulation platform. ATR combustor inlet flow conditions play a key role in the computation of species mass fractions, and these conditions are highly dependent on turbomachinery performance and engine flight regime. A propulsive operational database was created by varying mass flow rates of fuel and flight conditions such as cruise speed and altitude to investigate possible engine operations. The all-inlet conditions in this map are provided to the Cantera-Python chemical/combustion chemistry solver implementing a specially designed and formulated 0D kinetic-thermodynamic methodology successfully used to model and simulate the electric spark ignition required to activate the combustion process of the reacting mixture in the ATR combustion chambers, whereas the coupled aero-thermodynamic/aero-propulsive 0D/1D code i.e. Scramjet PREliminary Aerothermodynamic Design (SPREAD), designed and developed by the Italian Aerospace Research Centre (CIRA) was used for DMR calculations. Results show low emissions of NO according to the optimized design of the ATR; on the other hand, a tuning of operational conditions is needed for DMR, with its complete re-design to be more conclusive. Analogously, the released amount of water vapour is in good agreement with the required combustion efficiency and the expected propulsive performance.

Results show low emissions of NO according to the optimized design of the ATRs; on the other hand, a tuning of operational conditions is needed for DMR, with its complete re-design to be more conclusive. Analogously, the released amount of water vapour is in good agreement with the required combustion efficiency and the expected propulsive performance.

Applications of innovative 0D/1D chemical kinetic methodology and in-house codes.

The capability to predict and evaluate the motor pressure during each phase by means of a numerical analysis can significantly increase the efficiency of the preliminary design process with a reduction of both the motor development and operational costs. This paper aims to perform numerical simulation to analyze the ignition transient in solid rocket motor by solving Euler equation coupled with some semi-empirical correlations. These relations take into account the main phenomena affecting the ignition transient. Coupling relationships include the heat transfer of the gas to the propellant and erosive burning rate relationship.

The current research effort divides motor into series of control volumes along the port axis, and the variation of port area, burning surface and burning rate along the port are taken into account. A set of governing equations are then solved using explicit, time-dependent, predictor-corrector finite difference method. The numerical model helps to capture and embed shock wave associated with igniter flow within the solution. Second-order artificial viscosity dampens out the numerical oscillations due to sharp gradient within the flow field. The developed computer code predicts the start-up characteristics of motor. The study also provides comparison of simulation results with in-house experimental motor.

Simulations are performed with and without erosive burning to demonstrate that the flow model is a good physical approximation of motor. Numerical results calculated by this model without erosive burning are not in good agreement with experimental results. This minor discrepancy has motivated the inclusion of erosive burning in numerical model. The simulated results are then compared with the experimental data for head-end and rear-end pressure. The agreement between simulation and experiment is remarkable. In summary, major finding of this study is that unsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively.

Unsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively. However, in systems where two- and three-dimensional effects are pre-dominant, one would require to develop a more elaborate, multi-dimensional model which will allow for further understanding of the flow behavior and eventually lead to modeling of rocket motors with more complex geometries.

The close agreement between experimental and simulation results can be considered as forced to some degree, because the general mathematical model of erosive burning contains a free variable erosive burning exponent. However, in future, this variable can be established a priori by erosive burning tests.

The solid propellant ignition process consists of series of rapid events and must be completed in a fraction of a second. An understanding of the dynamics of ignition has become increasingly vital with the development of larger and more sophisticated solid propellant rocket motors. This research effort provides the simulation framework to predict and evaluate the motor pressure during each phase by means of a numerical analysis, thus significantly increasing the efficiency of the preliminary design process with a reduction of both the motor development and operational costs.

Steady‐state two‐dimensional solutions to the full compressible Navier‐Stokes equations are computed for laminar convective motion of a gas in a square cavity with large horizontal temperature differences. No Boussinesq or low‐Mach number approximations of the Navier‐Stokes equations are used. Results for air are presented. The ideal‐gas law is used and viscosity is given by Sutherland’s law. An accurate low‐Mach number solver is developed. Here an explicit third‐order discretization for the convective part and a line‐implicit central discretization for the acoustic part and for the diffusive part are used. The semi‐implicit line method is formulated in multistage form. Multigrid is used as the acceleration technique. Owing to the implicit treatment of the acoustic and the diffusive terms, the stiffness otherwise caused by high aspect ratio cells is removed. Low Mach number stiffness is treated by a preconditioning technique. By a combination of the preconditioning technique, the semi‐implicit discretization and the multigrid formulation a convergence behaviour is obtained which is independent of grid size, grid aspect ratio, Mach number and Rayleigh number. Grid converged results are shown for a variety of Rayleigh numbers.

A numerical scheme that has already proved to be efficient and accurate for laminar heat transfer is extended for turbulent, axisymmetric heat transfer calculations. The extended scheme is applied to the steady‐state heat transfer of axisymmetric turbulent jets, impinging onto a flat plate. Firstly, the low‐Reynolds version of the standard k‐ε model is employed. As is well known, the classical k‐ε turbulence model fails to predict the heat transfer of impinging jets adequately. A non‐linear k‐ε model, with improved ε‐equation, yields much better results. The numerical treatment of the higher order terms in this model is described. The effect on the heat transfer predictions of a variable turbulent Prandtl number is shown to be small. It is also verified that the energy equation can be simplified, without affecting the results. Results are presented for the flow field and the local Nusselt number profiles on the plate for impinging jets with different distances between the pipe exit and the flat plate.

The purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed.

The LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile.

It was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth.

The proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before.

This paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.