Inlet performance is an important field in aerodynamic design of aerial vehicle engines. This study focuses on numerical investigation of Mach number effects on a supersonic axisymmetric mixed compression inlet performance. For this purpose, a density based finite volume CFD code has been developed. A structured multi-block grid and an explicit time discretization of Reynolds averaged Navier-Stokes (RANS) equations have been used. Furthermore, Roe’s approximated Riemann solver has been utilized for computing inviscid flux vectors. Also, the monotone upstream centered schemes for conservation laws (MUSCL) extrapolation with Van Albada limiter have been used to obtain second order accuracy. In addition, Spalart-Allmaras one-equation turbulence model has been used to close the governing equations. The code is validated in three test cases by comparing numerical results against experimental data. Finally, the code has been utilized for numerical simulation of a specific supersonic mixed compression inlet. The effects of free stream Mach number on performance parameters, including mass flow ratio (MFR), drag coefficient, total pressure recovery (TPR), and flow distortion (FD) have been discussed and investigated. Results show that increase in Mach number, leads to decrease in TPR and drag coefficient; however, MFR and FD increase. Also, FD variations with respect to other performance parameters are significant, such that increase in Mach number from 1.8 to 2.2 leads to more than 100% FD increment while increase in MFR is less than 10%. By using this code it will be possible to design, performance parametric study, and geometrical optimization of axisymmetric supersonic inlet.

The flow quality inside a supersonic axisymmetric mixed compression air intake designed for the freestream Mach number of 2. 0 has been investigated experimentally and numerically in this study. The numerical study was used to analyze the shock configurations inside the intake. The flow in a supersonic intake is always irreversible due to the shock waves and boundary layers. A useful tool for studying flow quality entering the engine is the investigation of entropy generation due to various factors. In this study, the accuracy of the numerical results is evaluated by the experimental data at first and then the entropy generation inside intake is studied for different back pressures. Results indicated that reduction of the pseudo-shock length results in the significant decrease of entropy generation. Furthermore, role of the pressure fluctuations in the entropy generation was examined and it is observed that pressure fluctuations could have a significant effect on the irreversibility of the flow. According to the results, by increasing the exit blockage ratio from 55% to 62. 5%, the rate of entropy generation will be reduced by 33% due to the reduction of peuso-shock length, reduction in the flow separation at the end of diffuser and reduction of pressure fluctuations.

This study examines forced oscillations under supercritical conditions in a supersonic air inlet designed for a Mach number of 2. The preparation of inlet geometry, fluid flow simulations, and result postprocessing are performed using Ansys software and some in-house Fortran and Matlab numerical codes. Turbulence modeling was performed using the kω−SST turbulent model which has been approved in the past. The disturbances at the inlet’s exit are due to flow fluctuations in the combustion chamber. These are modeled by applying a sinusoidal excitation function in this study. The excitation function’s key parameters are amplitude and frequency. Despite the complex flow in the field, one of the main findings of this study is the observation of a frequency similar to the excitation frequency. Furthermore, the study investigates the effects of different excitation parameters on the upstream movement of fluctuations, and as the most important achievement the physical phenomena of supersonic flow are fully described.

Sudden expansion of flow in supersonic flow regime has gained relevance in the recent pasts for a wide run of applications. A number of kinematic as well as geometric parameters have been significantly found to impact the base pressure created within the suddenly expanded stream. The current research intends to create a predictive model for base pressure that is established in the abruptly extended stream. The artificial neural network (ANN) approach is being utilized for this purpose. The database utilized for training the network was assembled utilizing computational fluid dynamics (CFD). This was done by the design of experiments based L27 Orthogonal array. The three input parameters were Mach number (M), nozzle pressure ratio (NPR) and area ratio (AR) and base pressure was the output parameter. The CFD numerical demonstrate was approved by an experimental test rig that developed results for base pressure, and used a nozzle and sudden extended axissymmetric duct to do so. The ANN architecture comprised of three layers with eight neurons in the hidden layer. The algorithm for optimization was Levenberg-Marquardt. The ANN was able to successfully predict the base pressure with a regression coefficient R2 of less than 0. 99 and RMSE=0. 0032. The importance of input parameters influencing base pressure was estimated by using the ANN weight coefficients. Mach number obtained a relative importance of 47. 16% claiming to be the most dominating factor.

A numerical analysis of the inlet-combustor interaction and flow structure through a scramjet engine at a flight Mach number M=6 with parallel injection (Strut with circular inlet) is presented in the present research article. Three different angles of attack (a=-4o, a =0o a =4o) have been studied for parallel injection.The scramjet configuration used here is a modified version of DLR scramjet model. Fuel is injected at supersonic speed (M=2) through a parallel strut injector. For parallel injection, the shape of the strut is chosen in a way to produce strong stream wise vorticity and thus to enhance the hydrogen/air mixing inside the combustor. These numerical simulations are aimed to study the flow structure, supersonic mixing, and combustion phenomena for the three different types of geometries along with circular shaped strut configuration.

In this research, a lattice Boltzmann model is proposed to simulate free convection flow through pressure filtration. High compressibility effect needs to be considered near the critical point. A Poiseuille flow has been used as the first example to examine the effects of boundary condition model used in this study. It has been shown that the encountered error is of second order, which is considered to be desirable. The effect of the present boundary condition on the stability of solution to a Rayleigh-Benard problem has also been demonstrated. Finally, the filtered pressure equations have been implemented to model flow of a supercritical fluid in a cavity. The results are in good agreements with available data in the literature.

A series of experiments were conducted to study impacts of the free-stream Mach number, back pressure, and bleed on the stability of a supersonic intake. The ow stability is related to the buzz phenomenon, i. e. the oscillation of all shock waves of the intake, which may also occur when the intake mass ow rate is decreasing. In this study, the intake was axisymmetric with Mach number of 2. 0. The results showed that stability margin of the intake decreased when the freestream Mach number increased for both bleedo  and bleed-on cases. In the con guration without bleed, the frequency of buzz oscillation increased when the freestream Mach number decreased or when back pressure increased. By applying bleed and, consequently, preventing separation of the ow, the intake became more stable and the shocks oscillated with a smaller amplitude during the buzz phenomenon. Also, when the bleed was applied, the buzz triggering mechanism varied from the Dailey criterion to the Ferri one, which considerably changed stability characteristics of the intake.

In the recent years, achievement of a more accurate numerical method appropriate for different flow regimes to capture discontinuities with less oscillation and numerical errors has been of interest by many researchers. The specific comment in this paper is the comparison of the performance of artificial dissipation and upwind methods in solving the Euler equations for internal compressible flows in a wide range of inlet Mach numbers. In this work, we examine the ability of the AUSM+ upwind method, and the Scalar and Cusp artificial dissipation methods for flows with very low Mach number up to ultrasound and non-viscous flows in a convergent-divergent nozzle. The ability of the AUSM+ and Scalar methods in a 2D inviscid transonic flow between the turbine stator blades at both the supersonic and subsonic outlets is also studied. An excellent performance was observed for the AUSM+ method with more convergence speed and low numerical error in all flow regimes at a converging-diverging nozzle. Further, for the second case, the AUSM+ method coincides with the experimental results very well with lower numerical errors, and satisfies the mass conservation better than Scalar. It should be mentioned that the AUSM+ method is highly recommended for higher Mach numbers.

The flow field around a Sharp cone model configuration has been investigated by means of Schlieren facility in hypersonic shock tunnel. The time dependent evolution of flow around a cone of angle 11. 38° with base radius of 150mm has been visualized for a flow Mach number M = 6. 5. Experiments have been carried out with Helium as driver gas and air as test gas to visualize the hypersonic flow field. The flow establishment, steady state, and termination process of the hypersonic flow have been visualized for two different angles of attack, namely 0° &5° . Experimentally measured shock angle compares well with the theoretical and the computational study. The measured shock layer thickness compares well with the numerical simulation for both angles of attack.