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In this study, 2D density-based SST K-turbulence model with compressibility effect is used to observe the flow separation and shock wave interactions of the flow. The wall static pressure and Mach number differences are also evaluated. This study aims to discuss the aforementioned objectives

This study outlines the evaluation of the performance of a 2D convergent–divergent nozzle with various triangular jet tab configurations that can be used for effective thrust vectoring of aerial vehicles.

From the study, it is seen that the shadow effect induced by the tab with a height of 30% produces higher oblique wave deflection and higher thrust deflection at the exit nozzle. The numerical calculation concluded that thrust vector efficiency of 30% jet tab is, 0.46%. In the case of 10% jet tab height the thrust vector efficiency is higher, i.e. 1.647%.

2D study.

The optimization will open up a new focus in TVC that can be implemented for effective attitude control in aircrafts.

Used in future aircrafts.

The influence of shadowing ratio with different tab heights at different Mach numbers has not been reported in the previous studies. Few of the studies on jet tab are focused on the acoustic studies and not pertaining to the aerodynamic aspects. The multi jet configuration, the combination of location, shapes and other parametric analysis have not been covered in the previous studied.

An exposition of the need of supersonic aircraft to have a variable geometry intake and a fully variable convergent‐divergent nozzle for optimum performance. POWERPLANTS for supersonic aircraft bear only a superficial resemblance to the propulsive units of their subsonic counterparts. A typical subsonic turbojet powerplant consists of a high compression engine with a short fixed intake, and, possibly, a variable convergent nozzle. The supersonic aircraft on the other hand requires a powerplant with a sophisticated variable geometry intake having its own automatic control system and a fully variable convergent‐divergent nozzle in order to extract the full performance throughout the speed range from the slightly lower pressure ratio engine.

This paper aims to numerically investigate the influence of shock wave and freestream turbulence interaction on the parabolic and spherically blunted nose cones at supersonic speed.

Using density-based solver, the three-dimensional steady-state simulation is carried out. The working fluid is calorically perfect that obeys ideal gas law and the no-slip boundary conditionis given to the surface of the nose cone. Pressure far-field boundary condition is imposed at the boundary of the computational domain by giving freestream Mach number, freestream static pressure and temperature.

The growth rate of the boundary layer is faster on the spherically blunted nose cone, hence, the overall drag force is higher than the parabolic nose cone. Temperature at the edge of the boundary layer is increased due to the early ampli-fication of instabilities by the upstream disturbance. In this sense, the effects of freestream turbulence depend on its level, freestream conditions, strength and type of shock wave and zone of influence.

Simulations are carried out for the flow Mach number 2.0 at zero angles of attack for the freestream conditions of the flow at an altitude of 10,000 m.

The phenomenon of shock wave–turbulence interaction occurs in flow regimes from transonic to hypersonic speeds and finds a wide range of applications, especially in the design of aircraft and missiles configurations.

The phenomenon of compression wave and freestream turbulence interaction around the commonly used nose cones in the case of aircraft, missiles, etc., is investigated. The performance characteristics such as aerodynamic drag, boundary layer dynamics and the nature of flow around the different nose cones at zero angle of attack are illustrated.

A simple convection algorithm for simulation of time‐dependent supersonic and hypersonic flows of a perfect but viscous gas is described. The algorithm is based on conservation and convection of mass, momentum and energy in a grid of rectangular cells. Examples are given for starting flow in a shock‐tube and oblique shocks generated by a wedge, at Mach numbers up to 30.4. Good comparisons are achieved with well‐known perfect gas flows.

This work aims to describe the physical and numerical modeling of a CFD solver for hypersonic flows in thermo-chemical non-equilibrium. This paper is the second of a two-part series that concerns the application of the solver introduced in Part I to adaptive unstructured meshes.

The governing equations are discretized with an edge-based stabilized finite element method (FEM). Chemical non-equilibrium is simulated using a laminar finite-rate kinetics, while a two-temperature model is used to account for thermodynamic non-equilibrium. The equations for total quantities, species and vibrational-electronic energy conservation are loosely coupled to provide flexibility and ease of implementation. To accurately perform simulations on unstructured meshes, the non-equilibrium flow solver is coupled with an edge-based anisotropic mesh optimizer driven by the solution Hessian to carry out mesh refinement, coarsening, edge swapping and node movement.

The paper shows, through comparisons with experimental and other numerical results, how FEM + anisotropic mesh optimization are the natural choice to accurately simulate hypersonic non-equilibrium flows on unstructured meshes. Three-dimensional test cases demonstrate how, for high-speed flows, shocks resolution, and not necessarily boundary layers resolution, is the main driver of solution accuracy at walls. Equally distributing the error among all elements in a suitably defined Riemannian space yields highly anisotropic grids that feature well-resolved shock waves. The resulting high level of accuracy in the computation of the enthalpy jump translates into accurate wall heat flux predictions. At the opposite end, in all cases examined, high-quality but isotropic unstructured meshes gave very poor solutions with severely inadequate heat flux distributions not even featuring expected symmetries. The paper unequivocally demonstrates that unstructured anisotropically adapted meshes are the best, and may be the only, way for accurate and cost-effective hypersonic flow solutions.

Although many hypersonic flow solvers are developed for unstructured meshes, few numerical simulations on unstructured meshes are presented in the literature. This work demonstrates that the proposed approach can be used successfully for hypersonic flows on unstructured meshes.

This study aims to present the numerical study on supersonic jet mixing characteristics of the co-flow jet by varying lip thickness (LT). The LT chosen for the study is 2 mm, 7.75 mm and 15 mm.

The primary nozzle is designed for delivering Mach 2.0 jet, whereas the secondary nozzle is designed for delivering Mach 1.6 jet. The Nozzle pressure ratio chosen for the study is 3 and 5. To study the mixing characteristics of the co-flow jet, total pressure and Mach number measurements were taken along and normal to the jet axis. To validate the numerical results, the numerical total pressure values were also compared with the experimental result and it is proven to have a good agreement.

The results exhibit that, the 2 mm lip is shear dominant. The 7.75 mm and 15 mm lip is wake dominant. The jet interaction along the jet axis was also studied using the contours of total pressure, Mach number, turbulent kinetic energy and density gradient. The radial Mach number contours at the various axial location of the jet was also studied.

The effect of varying LT in exhaust nozzle plays a vital role in supersonic turbofan aircraft.

Supersonic co-flowing jet mixing effectiveness by varying the LT between the primary supersonic nozzle and the secondary supersonic nozzle has not been analyzed in the past.

WHEN talking about airbreathing engines it is now generally understood that they are either turbine engines when the maximum flight Mach number is subsonic or moderately supersonic, or ramjets when the Mach number is definitely high. When trying to meet the propulsion requirements from take‐off to a high enough speed the joint use of both engine types has to be considered. In such case most people would think of the ramjet as taking over the propulsion task from the turbine engine when reaching a certain value of the flight Mach number, or more precisely of the air stagnation temperature, above which the turbine engine is no longer able to operate. The most elementary view is that of presenting it as a limitation in the engine structure, with improvements calling for the use of better materials. Bringing thermo‐dynamics into the picture shows that increased air stagnation temperature results in a deterioration in the cycle efficiency of the turbine engine proper and this may result in the specific fuel consumption of the turbojet becoming higher than that of the ramjet. Such a performance limitation can be shifted to higher Mach numbers while using increased turbine intake temperatures. The consideration of the aerodynamics of the internal flow brings out another type of limitation due to the difficulty of keeping the operating line of the turbojet over the flight profile far enough from the surge limit though within the range of good compresser efficiency. Variable geometry in the compressor and turbine stators may produce some improvement in this respect.

The purpose of this paper is to analyze the effect of pressure fluctuations on the combustion efficiency of the hydrogen fuel injected into the supersonic oxidizing cross flow. The pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave.

The pressure fluctuations are imposed on inlet air flow and also on the fuel flow stream. Two different situations are considered: the combustion chamber once without and again with the inlet standing oblique shock wave. The unsteady turbulent reacting flow solver is developed to simulate the supersonic flow field in the combustion chamber with detail chemical kinetics, to predict the time-variation of the combustion efficiency due to the imposed pressure fluctuations.

The results show that the response of the reacting flow field depends on both the frequency of fluctuations and the existence of the inlet shock wave. In addition, the inlet standing shock wave has some attenuating role, but the reacting flow shows an amplifying role on imposed oscillations which is also augmented by imposing anti-phase fluctuations on both inlet and fuel flow streams.

This study is performed to analyze the instabilities in the supersonic combustion which has not been considered before in this manner.

Partially‐decoupled upwind‐based total‐variation‐diminishing (TVD) finite‐difference schemes for the solution of the conservation laws governing two‐dimensional non‐equilibrium vibrationally relaxing and chemically reacting flows of thermally‐perfect gaseous mixtures are presented. In these methods, a novel partially‐decoupled flux‐difference splitting approach is adopted. The fluid conservation laws and species concentration and vibrational energy equations are decoupled by means of a frozen flow approximation. The resulting partially‐decoupled gas‐dynamic and thermodynamic subsystems are then solved alternately in a lagged manner within a time marching procedure, thereby providing explicit coupling between the two equation sets. Both time‐split semi‐implicit and factored implicit flux‐limited TVD upwind schemes are described. The semi‐implicit formulation is more appropriate for unsteady applications whereas the factored implicit form is useful for obtaining steady‐state solutions. Extensions of Roe's approximate Riemann solvers, giving the eigenvalues and eigenvectors of the fully coupled systems, are used to evaluate the numerical flux functions. Additional modifications to the Riemann solutions are also described which ensure that the approximate solutions are not aphysical. The proposed partially‐decoupled methods are shown to have several computational advantages over chemistry‐split and fully coupled techniques. Furthermore, numerical results for single, complex, and double Mach reflection flows, as well as corner‐expansion and blunt‐body flows, using a five‐species four‐temperature model for air demonstrate the capabilities of the methods.