ЧИСЛЕННОЕ МОДЕЛИРОВАНИЕ ОБТЕКАНИЯ СПУСКАЕМОГО АППАРАТА ПРИ СВЕРХЗВУКОВЫХ И ГИПЕРЗВУКОВЫХ СКОРОСТЯХ. СРАВНЕНИЕ ДВУХ СПУСКАЕМЫХ АППАРАТОВ. ЧАСТЬ 1

Investigation on wall shear stress measurement in supersonic flows with shock waves using shear-sensitive liquid crystal coating

A similitude has been obtained for an oscillating two-dimensional wedge in pitch with attached shock at high angle of incidence in hypersonic flow. A strip theory, in which flow at infinite span wise location is two dimensional and independent of each other is being used. Further this unites the similitude with piston theory to give one dimensional piston theory with large flow deflection. Closed form solution is obtained for stiffness and damping derivatives in pitch. The present theory is valid when the shock wave is attached with the nose of the wedge. With the increase in semi vertex angle of the wedge the stiffness and the damping derivatives are found to increase progressively, and with the increase in the Mach number both stiffness and damping derivatives are found to decrease then with increase in Mach number it becomes independent of Mach number. From the amalgamation of theory developed, we can evaluate the stiffness and damping derivatives for in pitch for a wide range of Mach numbers, for various pivot positions, and angle of attack. Significantly the same results are being validated with the analytical results of Liu and Hui [4] and Lighthill’s [3] theory for some cases.

Frontal cavity consists of complex unsteadiness which causes bow shock instabilities when the body is moving at supersonic or hypersonic speed through a fluid. These complex unsteady flow fields are an important practical concern in the aerospace applications. Understanding this flow field around the frontal cavity will help us to know more in depth about these instabilities to create effective control methods to avoid structural damages. The main objective of the present research is to analyze the bow shock instabilities at different Mach number around the frontal cavity at supersonic and hypersonic speeds by using the numerical concept of Detached Eddy Simulations (DES) in two dimensional axisymmetric domain. Three different freestream Mach numbers (2, 4 and 6) were considered in the present investigation in order to study the effect of Mach number on flow unsteadiness. It was observed that an increase in Mach number, the frequency of most dominant mode (bow shock pulsation) increases significantly. In addition to that, additional low amplitude sub-dominant modes were observed for all the cases of freestream Mach numbers studied in the present investigation.

A calculating method of shock wave oscillating frequency due to turbulent shear layer fluctuations in supersonic flow

Grid fin is widely used in the field of aerospace because of its unique advantages in the aerodynamic and structural performance. Meanwhile, the aerodynamic and structural characteristics of the grid fin are quite different from those of the conventional rudder. Deep understanding of the aeroelastic characteristics and mechanism of the grid fin is of great reference value for the fine design of grid fin. In this paper, an efficient reduction-order-method-based aeroelastic analysis method is established, and the method is verified by the flutter case of AGARD 445.6 wing. Finally, based on this method, the flowfield characteristics and the variation of flutter characteristics with Mach number are studied in detail. The results show that in the subsonic region, the transonic flutter “dip” of the grid fin occurs earlier than that of the conventional wing due to the acceleration effect of the grid fin on the airflow. In the transonic region, when the flowfield is characterized by a bow shock wave, the critical flutter speed increases slightly with the increase of Mach number. However, in the supersonic region, when the flowfield is characterized by an oblique shock wave, the flutter critical velocity increases substantially with the increase of Mach number. It is found that the variation law of the flutter boundary with Mach number is very similar to that of the normal force and the pitching moment with Mach number, and there is a close relationship between them. The ratio of flutter speed of the grid fin to the freestream velocity will reach the minimum value near . At the same time, it is discovered that the flutter instability of the grid fin is mainly caused by the coupling of the bending mode and torsional mode of the grid fin, and the instability mode is the torsional mode. Meanwhile, it should be noted that the flutter coupling characteristics of the grid fin differ greatly from those of the conventional rudder surface due to the great difference in structural characteristics and unsteady aerodynamic characteristics.

Various gaps in the surface of the supersonic aircraft have a significant effect on airflows. In order to predict the effects of attack angle, Mach number and width-to-depth ratio of gap on the local aerodynamic heating environment of supersonic flow, two-dimensional compressible Navier-Stokes equations are solved by the finite volume method, where convective flux of space term adopts the Roe format, and discretization of time term is achieved by 5-step Runge-Kutta algorithm. The numerical results reveal that the heat flux ratio is U-shaped distribution on the gap wall and maximum at the windward corner of the gap. The heat flux ratio decreases as the gap depth and Mach number increase, however, it increases as the attack angle increases. In addition, it is important to find that chamfer in the windward corner can effectively reduce gap effect coefficient. The study will be helpful for the design of the thermal protection system in reentry vehicles.

Automatic detection of oblique shocks and simple waves in schlieren images of two-dimensional supersonic steady flows

In the present paper a theory for 2D slender bodies at high angle of attack in hypersonic flow has been developed to determine the aerodynamic derivatives in pitch for different Mach numbers. The present theory has been applied to a sharp thick wedge with attached shock case. Using the theory, a relation for a piston moving in a cylinder at any velocity and relations for stiffness and damping derivatives are obtained for zero incidence of the wedge and it is found to be dependent on flight Mach number and wedge semi vertex angle. The present method includes the thin wedge case as well, which was covered by Lighthill piston theory, and applies only for small amplitude and low reduced frequency case. Effect of viscosity and secondary wave reflections have been neglected. The results are obtained for wedges of different semi vertex angles and Mach numbers. It is observed that the stiffness and Damping derivatives increase with increase in the semi vertex angle and decrease with an increase in Mach number. At very high Mach number the Mach number independence principle holds good.

Direct numerical simulation of compressible turbulent channel flows using the discontinuous Galerkin method

Effect of Mach number and plate thickness on the flow field and heat transfer characteristics of supersonic turbulent flow over a flat plate at different thermal boundary conditions

Spark ignition and stability limits of spray kerosene flames under subatmospheric pressure conditions

In most of the application occasions of supersonic and hypersonic flows, the flow characters are extremely complicated, involving a wide range of interacting scales strongly affected by compressibility, shock wave, instabilities, turbulence, and confinement of sidewalls. It has confined our scientific understanding of the referred flow mechanisms.

Descending a space vehicle in Mars’s atmosphere is challenging. The main challenge is the unusual heating while the descending spacecraft moves at hypersonic speeds. This paper presents a computational model of a hypersonic flow around a spacecraft descending into Mars’s atmosphere. The work focuses on aerodynamic heating and the thermal field around the vehicle at hypersonic speeds. The mathematical model considers a three-dimensional, compressible hypersonic flow with bow shock waves. OpenFOAM, an open-source CFD code, solves the governing equations of mass, momentum, and energy using a finite-volume approach. The solver is a density-based, shock-capturing approach based on the central-upwind schemes of Kurganov and Tadmor. In this investigation, the turbulence model is Spalart-Allmaras. A clustered mesh resolves the super-fine boundary layer near the walls where large gradients dominate. This study focuses on the range of hypersonic Mach numbers from 5 to 27. A grid refinement study and comparison with the analytical solution verifies the solution approach and validates the results. This work aims to present the temperature field on and about the vehicle. The paper presents regions of subsonic and supersonic flows over the vehicle, surface temperature, and the contribution of aerodynamic heating.

Results of a numerical study of the flow structure in the rectangular duct of a model air inlet with the interaction of the shock wave generated by the inlet cowl with the turbulent boundary layer on the inner duct surfaces are reported. A configuration involving a convergent entrance section and a constant-cross-section throat was analyzed. The numerical simulation of the three-dimensional flow with free-stream Mach number M = 4 was performed on the basis of the Reynolds-averaged Navier-Stokes equations and the k-ω SST model of turbulence. As a result of the calculations, the limit streamlines and the distributions of static pressure on the duct surfaces, as well as the distributions of total pressure and the fields of Mach number in duct cross sections were obtained. The structure of the flow is analyzed and the possibility of occurrence of longitudinal vortex structures in the corner region of the intake-duct throat is shown.

Chapter 15 - Compressible Flow