Hemisphere-cylinder in transonic flow, free-stream Mach number from 0.7 to 1.0

AIAA Journal ◽  
1975 ◽  
Vol 13 (10) ◽  
pp. 1411-1413 ◽  
Author(s):  
Tsuying Hsieh
Keyword(s):  
Mach Number ◽  
Transonic Flow ◽  
Free Stream ◽  
Free Stream Mach Number

scholarly journals Numerical solutions for transonic flow

1949 ◽  
Vol 196 (1044) ◽  
pp. 32-36 ◽  
Keyword(s):  
Mach Number ◽  
Exact Solutions ◽  
Compressible Fluid ◽  
Transonic Flow ◽  
Numerical Solutions ◽  
Free Stream ◽  
Two Dimensional ◽  
Transverse Axis ◽  
Two Dimensional Flow ◽  
Free Stream Mach Number

In a previous paper (Cherry 1947), the author has established a family of exact solutions for steady two-dimensional flow of a compressible fluid past a cylinder; the final formulae are given in theorem 6, equations (5.17) to (5.21). These formulae have now been evaluated (taking γ = 1.405) for the value T 1 = 0.05, corresponding to a free-stream Mach number of 0.510, and the streamlines are shown in figure 1. The cylindrical obstacle has a thickness ratio 0.93, but is markedly different from an ellipse, being almost exactly circular over its up- and downstream quadrants. The Mach number a t the ends of its transverse axis is 1.39. The flow is everywhere regular, but a small increase in the free-stream Mach number would be critical; a shock-line would begin to appear near the points on the surface where the tangent is inclined at about 25 or 30° to the direction of the free-stream.

Direct numerical simulation of flow past a transversely rotating sphere up to a Reynolds number of 300 in compressible flow

2018 ◽  
Vol 857 ◽  
pp. 878-906 ◽  
Author(s):  
T. Nagata ◽  
T. Nonomura ◽  
S. Takahashi ◽  
Y. Mizuno ◽  
K. Fukuda
Keyword(s):  
Numerical Simulation ◽  
Reynolds Number ◽  
Mach Number ◽  
Direct Numerical Simulation ◽  
Drag Coefficient ◽  
Rotation Rate ◽  
Free Stream ◽  
Rotating Sphere ◽  
Free Stream Mach Number ◽  
Low Reynolds

In this study, direct numerical simulation of the flow around a rotating sphere at high Mach and low Reynolds numbers is conducted to investigate the effects of rotation rate and Mach number upon aerodynamic force coefficients and wake structures. The simulation is carried out by solving the three-dimensional compressible Navier–Stokes equations. A free-stream Reynolds number (based on the free-stream velocity, density and viscosity coefficient and the diameter of the sphere) is set to be between 100 and 300, the free-stream Mach number is set to be between 0.2 and 2.0, and the dimensionless rotation rate defined by the ratio of the free-stream and surface velocities above the equator is set between 0.0 and 1.0. Thus, we have clarified the following points: (1) as free-stream Mach number increased, the increment of the lift coefficient due to rotation was reduced; (2) under subsonic conditions, the drag coefficient increased with increase of the rotation rate, whereas under supersonic conditions, the increment of the drag coefficient was reduced with increasing Mach number; and (3) the mode of the wake structure becomes low-Reynolds-number-like as the Mach number is increased.

scholarly journals Звуковой удар от тонкого тела и локальных областей нагрева сверхзвукового набегающего потока

2021 ◽  
Vol 91 (4) ◽  
pp. 558
Author(s):  
А.В. Потапкин ◽  
Д.Ю. Москвичев
Keyword(s):  
Mach Number ◽  
Supersonic Flow ◽  
Free Stream ◽  
Local Heating ◽  
Air Flow ◽  
Sonic Boom ◽  
Incoming Flow ◽  
Combined Method ◽  
Cold Flow ◽  
Free Stream Mach Number

The problem of a sonic boom generated by a slender body and local regions of supersonic flow heating is solved numerically. The free-stream Mach number of the air flow is 2. The calculations are performed by a combined method of phantom bodies. The results show that local heating of the incoming flow can ensure sonic boom mitigation. The sonic boom level depends on the number of local regions of incoming flow heating. One region of flow heating can reduce the sonic boom by 20% as compared to the sonic boom level in the cold flow. Moreover, consecutive heating of the incoming flow in two regions provides sonic boom reduction by more than 30%.

Calculation of External and Internal Transonic Flow Field of a Three-Dimensional S-Shaped Inlet

1985 ◽  
Author(s):  
Shen Huili ◽  
Luo Shijun ◽  
Ji Minggang ◽  
Xing Zongwen ◽  
Zhu Xin ◽  
...  
Keyword(s):  
Finite Difference ◽  
Flow Field ◽  
Transonic Flow ◽  
Present Method ◽  
Velocity Potential ◽  
Free Stream ◽  
Three Dimensional ◽  
Finite Difference Schemes ◽  
Finite Difference Equations ◽  
Free Stream Mach Number

A mixed finite difference method for calculating the external and internal transonic flow field around an s-shaped inlet is presented. Starting from the velocity potential equation and using Cartesian mesh and mixed finite difference schemes, the authors have obtained a system of finite difference equations and solved them with the aid of alternating line relaxations along two directions. Computations have been made for an s-shaped inlet with free stream Mach number M=0.8 at different angles of attack. Computed results are compared with those computed by perturbation method and with experimental results. Such a comparison shows that the present method is promising.

Probe Blockage Effects in Free Jets and Closed Tunnels

1975 ◽  
Vol 97 (4) ◽  
pp. 509-514 ◽  
Author(s):  
J. S. Wyler
Keyword(s):  
Mach Number ◽  
Flow Regime ◽  
Analytical Solutions ◽  
Transonic Flow ◽  
Free Stream ◽  
Pressure Measurements ◽  
Free Jets ◽  
Blockage Effect

The effect of probe blockage on the free-stream pressure and Mach number has been studied for the two limiting cases of open free jets and closed tunnels. Cylinder probe calibrations carried out in subsonic free jets show that the blockage effect is much greater than previous analytical solutions predict. The blockage effect in free jets was found to be of approximately the same magnitude as in closed tunnels. Generalized blockage corrections are derived which indicate the importance of minimizing blockage effects both when calibrating a probe and when using it to make pressure measurements, especially in the transonic flow regime.

Subsonic Flow Over Open Cavities: Part 1 — Analytical Models and Numerical Investigations

2016 ◽  
Author(s):  
Weidong Shao ◽  
Jun Li
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Shear Layer ◽  
Layer Thickness ◽  
Boundary Layer Thickness ◽  
Feedback Loop ◽  
Subsonic Flow ◽  
Free Stream ◽  
Open Cavities ◽  
Free Stream Mach Number

The aeroacoustical oscillation and acoustic field generated by subsonic flow grazing over open cavities has been investigated analytically and numerically. The tone generation mechanism is elucidated with an analytical model based on the coupling between shear layer instabilities and acoustic feedback loop. The near field turbulent flow is obtained using two-dimensional Large Eddy Simulation (LES). A special mesh is used to absorb propagating disturbances and prevent spurious numerical reflections. Comparisons with available experimental data demonstrate good agreement in both the frequency and amplitude of the aeroacoustical oscillation. The physical phenomenon of the noise generated by the feedback loop is discussed. The correlation analysis of primitive variables is also made to clarify the characteristics of wave propagation in space and time. The effects of free-stream Mach number and boundary layer thickness on pressure fluctuations within the cavity and the nature of the noise radiated to the far field are examined in detail. As free-stream Mach number increases velocity fluctuations and mass flux into the cavity increase, but the resonant Strouhal numbers slightly decrease. Both the resonant Strouhal numbers and sound pressure levels decrease with the increase of boundary layer thickness. Results indicate that the instability of the shear layer dominates both the frequency and amplitude of the aeroacoustical oscillation.

Second-order boundary-layer effects in hypersonic flow past axisymmetric blunt bodies

1964 ◽  
Vol 20 (4) ◽  
pp. 593-623 ◽  
Author(s):  
R. T. Davis ◽  
I. Flügge-Lotz
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Free Stream ◽  
Practical Interest ◽  
Second Order ◽  
The Body ◽  
Specific Flow ◽  
Flow Past ◽  
Free Stream Mach Number ◽  
Stability And Accuracy

First- and second-order boundary-layer theory are examined in detail for some specific flow cases of practical interest. These cases are for flows over blunt axisymmetric bodies in hypersonic high-altitude (or low density) flow where second-order boundary-layer quantities may become important. These cases consist of flow over a hyperboloid and a paraboloid both with free-stream Mach number infinity and flow over a sphere at free-stream Mach number 10. The method employed in finding the solutions is an implicit finite-difference scheme. It is found to exhibit both stability and accuracy in the examples computed. The method consists of starting near the stagnation-point of a blunt body and marching downstream along the body surface. Several interesting properties of the boundary layer are pointed out, such as the nature of some second-order boundary-layer quantities far downstream in the flow past a sphere and the effect of strong vorticity interaction on the second-order boundary layer in the flow past a hyperboloid. In several of the flow cases, results are compared with other theories and experiments.

Performance study of an inlet in supersonic flow

2012 ◽  
Vol 227 (1) ◽  
pp. 159-174 ◽  
Author(s):  
Mohammad R Soltani ◽  
Mohammad Farahani
Keyword(s):  
Mach Number ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Free Stream ◽  
Characteristic Curve ◽  
Pressure Recovery ◽  
Internal Boundary ◽  
Performance Study ◽  
Free Stream Mach Number

The performance characteristics of an axisymmetric inlet at its design and off-design operational conditions are experimentally investigated. The model is tested for wide ranges of free stream Mach numbers, M∞ = 1.5–2.5, and mass flow rates. For each test, the pressure recovery, the mass flow passing through the inlet and the pressure distribution over the spike and the cowl are measured. In addition, the shock wave formed in front of the inlet is visualized. The characteristic curve of the inlet is then obtained for each free stream Mach number. As the Mach number is increased, the pressure recovery is reduced, but the maximum value of the mass flow rate grows up. Variations of the mass flow affect the surface pressure over both the front portion of the cowl and the entire surface of the spike. Further, it has changed both pressure and Mach number at the end of the diffuser, which would consequently affect the performance of the propulsion system. In addition, contrary to the internal boundary layer, the external one far from the cowl lip has been found to be almost independent of the inlet mass flow rate for a constant free stream Mach number.

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