Parameters Affecting the Postshock Wave Relaxation Zone in an Oxygen Carbon Particle Suspension

1986 ◽  
Vol 108 (3) ◽  
pp. 360-365 ◽  
Author(s):  
O. Igra ◽  
G. Ben-Dor ◽  
I. Elperin
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Equilibrium State ◽  
Carbon Concentration ◽  
Carbon Particle ◽  
Particle Diameter ◽  
Relaxation Zone ◽  
Initial Carbon ◽  
State Changes ◽  
Shock Wave Mach Number

A parametric study was conducted regarding the effects of the initial carbon concentration, the initial carbon particle diameter, the shock wave Mach number and the preshock suspension pressure on the suspension behavior in the relaxation zone. The suspension was composed of oxygen gas seeded with small carbon particles. It was found that changing either the initial carbon concentration or the shock wave Mach number has a marked effect on the suspension behavior in the relaxation zone and on the eventually reached postshock equilibrium state. Changes in the initial carbon particle diameter (at a constant shock wave Mach number and carbon concentration) has no effect on the postshock equilibrium state, but it does affect the way in which the suspension reaches this equilibrium state. Changes in the preshock suspension pressure has only minor effects on the suspension behavior in the relaxation zone.

Experimental investigation of diaphragm material combinations in a small-scale shock tube

2020 ◽  
Vol ahead-of-print (ahead-of-print) ◽  
Author(s):  
Anugya Singh ◽  
Aravind Satheesh Kumar ◽  
Kannan B.T.
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Tube ◽  
Production Capacity ◽  
Incident Shock ◽  
Small Scale ◽  
Content Type ◽  
Shock Mach Number ◽  
Reinforcing Material ◽  
Shock Wave Mach Number

Purpose The purpose of this study is to experimentally investigate the trends in shock wave Mach number that were observed when different diaphragm material combinations were used in the small-scale shock tube. Design/methodology/approach A small-scale shock tube was designed and fabricated having a maximum Mach number production capacity to be 1.5 (theoretically). Two microphones attached in the driven section were used to calculate the shock wave Mach number. Preliminary tests were conducted on several materials to obtain the respective bursting pressures to decide the final set of materials along with the layered combinations. Findings According to the results obtained, 95 GSM tracing paper was seen to be the strongest reinforcing material, followed by 75 GSM royal executive bond paper and regular 70 GSM paper for aluminium foil diaphragms. The quadrupled layered diaphragms revealed a variation in shock Mach number based on the position of the reinforcing material. In quintuple layered combinations, the accuracy of obtaining a specific Mach number was seen to be increasing. Optimization of the combinations based on the production of the shock wave Mach number was carried out. Research limitations/implications The shock tube was designed taking maximum incident shock Mach number as 1.5, the experiments conducted were found to achieve a maximum Mach number of 1.437. Thus, an extension to further experiments was avoided considering the factor of safety. Originality/value The paper presents a detailed study on the effect of change in the material and its position in the layered diaphragm combinations, which could lead to variation in Mach numbers that are produced. This could be used to obtain a specific Mach number for a required study accurately, with a low-cost setup.

scholarly journals Scaling in the shock–bubble interaction

2003 ◽  
Vol 21 (3) ◽  
pp. 335-339 ◽  
Author(s):  
K. LEVY ◽  
O. SADOT ◽  
A. RIKANATI ◽  
D. KARTOON ◽  
Y. SREBRO ◽  
...  
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Flow Field ◽  
Vortex Ring ◽  
Speed Of Sound ◽  
Bubble Radius ◽  
Spherical Bubble ◽  
Bubble Interaction ◽  
Shock Wave Mach Number ◽  
Good Agreement

The passage of a shock wave through a spherical bubble results in the formation of a vortex ring. In the present study, simple dimensional analysis is used to show that the circulation is linearly dependent on the surrounding material speed of soundcsand the initial bubble radiusR. In addition, it is shown that the velocities characterizing the flow field are linearly dependent on the speed of sound, and are independent of the initial bubble radius. The dependence of the circulation on the shock wave Mach numberMis derived by Samtaney and Zabusky (1994) as (1 + 1/M+ 2/M2) (M− 1). Experiments were performed for slow/fast (air-helium) and fast/slow (air-SF6) interactions. Full numerical simulations were conducted resulting in good agreement. From the results, it is seen that in both cases, according to the proposed scaling, the vortex ring velocity is bubble radius independent. The numerical results for the slow/fast interaction show that the proposed Mach scaling is valid forM< 2. AboveM≅ 2, the topology of the bubble changes due to a competition between the upstream surface of the bubble and the undisturbed shock wave.

Finite Mach number spherical shock wave, application to shock ignition

2013 ◽  
Vol 20 (8) ◽  
pp. 082702 ◽  
Author(s):  
A. Vallet ◽  
X. Ribeyre ◽  
V. Tikhonchuk
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Spherical Shock Wave ◽  
Spherical Shock ◽  
Shock Ignition

Shock Wave - Film Cooling Interactions in Transonic Flows

2001 ◽  
Author(s):  
P. M. Ligrani ◽  
C. Saumweber ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Cross Flow ◽  
Injection Velocity ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Density Ratios ◽  
Film Cooling Holes

Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.

scholarly journals FLOW SIMULATION OVER REENTRY CAPSULE AT SUPERSONIC AND HYPERSONIC SPEEDS. THE COMPARE BETWEEN TWO REENTRY CAPSULES. PART 1

2020 ◽  
pp. 45-51
Author(s):  
Pavel Timofeev ◽  
◽  
Vladimir Panchenko ◽  
Sergey Kharchyk ◽  
◽  
...  
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Flow Direction ◽  
Numerical Algorithms ◽  
Flow Simulation ◽  
Ansys Fluent ◽  
Forward Shock ◽  
Hypersonic Speeds ◽  
Mach Numbers ◽  
Cfd Method

This study presents flow simulation over the reentry capsule at supersonic and hypersonic speeds. Numerical algorithms solve for the CFD method, which is produced using help ANSYS Fluent 19.2. The using GPU core to get a solution faster. The main purpose – flow simulation and numerical analysis reentry capsule; understand the behavior of supersonic and hypersonic flow and its effect on the reentry capsule; compare temperature results for the range Mach numbers equals 2–6. This study showed results on velocity counters, on temperature counters and vector of velocity for range Mach numbers equals 2–6. This study demonstrates the importance of understanding the effects of shock waves and illustrates how the shock wave changes as the Mach number increases. For every solves, the mesh had adapted for pressure gradient and velocity gradient to get the exact solution. As a result of the obtained solution, it is found that a curved shock wave appears in front of the reentry capsule. The central part of which is a forward shock. An angular expansion process is observed, which is a modified picture of the Prandtl- Mayer flow that occurs in a supersonic flow near the sharp edge of the expanding region. It is revealed that with an increase in the Mach number, the shock wave approaches the bottom of the reentry capsule, and there is also a slope of the shock to the flow direction, with an increase in the Mach number. The relevance and significance of this problem for the design of new and modernization of old reentry capsules.

The Influence of Initial Velocity and Temperature Lag on the Relaxation Zone behind a Shock Wave in a Gas-Solid Flow

1975 ◽  
Vol 189 (1) ◽  
pp. 367-373
Author(s):  
B. N. Cole ◽  
M. R. Baum ◽  
F. R. Mobbs
Keyword(s):  
Shock Wave ◽  
Energy Transfer ◽  
Initial Velocity ◽  
Solid Particles ◽  
Flow Properties ◽  
Equilibrium Conditions ◽  
Solid Flow ◽  
Relaxation Zone ◽  
Nozzle Flows ◽  
Kinetic Energy Transfer

Thermal and velocity lag prior to the gas discontinuity are shown to have a significant effect on the properties of the relaxation zone of a gas-solid flow shock wave. The results obtained by previous workers assuming equilibrium conditions are not therefore applicable to shock waves formed in gas-solid nozzle flows where considerable disequilibrium already exists. The variations of the relevant flow properties in the relaxation zone are interpreted physically in terms of the magnitude and sign of the heat, momentum and kinetic energy transfer between the solid particles and the gas.

Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition

2020 ◽  
Author(s):  
Hoshio Tsujita ◽  
Masanao Kaneko
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Secondary Flow ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Flow Behavior ◽  
Turbine Cascade ◽  
Blade Loading ◽  
Passage Vortex ◽  
Exit Mach Number

Abstract Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.

scholarly journals Mathematical Modeling of Carbon Flux Parameters for Low-Pressure Vacuum Carburizing with Medium-High Alloy Steel

Coatings ◽  
2020 ◽  
Vol 10 (11) ◽  
pp. 1075
Author(s):  
Haojie Wang ◽  
Jing Liu ◽  
Yong Tian ◽  
Zhaodong Wang ◽  
Xiaoxue An
Keyword(s):  
Carbon Concentration ◽  
Carbon Flux ◽  
Carbon Fluxes ◽  
Low Pressure ◽  
Process Conditions ◽  
Flux Method ◽  
Vacuum Carburizing ◽  
Initial Carbon ◽  
Vacuum Carburization ◽  
And Diffusion

Low-pressure vacuum carburizing adopts a pulse process mode to improve the carburizing efficiency and reduces gas and energy consumption. Carbon flux is the key to accurately control the time of strong infiltration and diffusion in each pulse. In order to obtain the carbon fluxes with various materials under diffident carburizing process conditions, an evenly segmented carbon flux method is proposed. A systematic study with each model using different materials (12Cr2Ni4A, 16Cr3NiWMoVNbE, and 18Cr2Ni4WA represent different initial carbon concentrations and different alloy compositions), carburizing temperatures, and carburizing pressures to determine the effect of these conditions on carbon flux is conducted. Compared with traditional segmented carbon flux method, an evenly segmented carbon flux method can predict the actual carbon flux more precisely and effectively in order to finely control the pulse carburization process. The paper also indicates that carbon fluxes increase with the increase of pressure. The optimal carburization pressure for low-pressure vacuum carburization is 300 Pa. Raising the carburization temperature to 980 °C instead of 920 °C can increase effective carbon flux by more than 30%. Among the material compositions, alloy content has the biggest impact over the carbon, initial carbon concentration the second, and saturated carbon concentration the third biggest impact.

A 3-D Thin Layer Navier-Stokes Code for Supersonic Laminar and Turbulent Flows

2002 ◽  
Author(s):  
Omid Abouali ◽  
Mohammad M. Alishahi ◽  
Homayoon Emdad ◽  
Goodarz Ahmadi
Keyword(s):  
Experimental Data ◽  
Shock Wave ◽  
Mach Number ◽  
Turbulent Flows ◽  
Turbulence Modeling ◽  
Cross Flow ◽  
Navier Stokes ◽  
Cross Sectional ◽  
Laminar And Turbulent Flows ◽  
Circumferential Pressure

A 3-D Thin Layer Navier-Stokes (TLNS) code for solving viscous supersonic flows is developed. The new code uses several numerical algorithms for space and time discretization together with appropriate turbulence modeling. Roe’s method is used for discretizing the convective terms and the central differencing scheme is employed for the viscous terms. An explicit time marching technique and a finite volume space discretization are used. The developed computational model can handle both laminar and turbulent flows. The Baldwin-Lomax model and Degani-Schiff modifications are used for turbulence modeling. The computational model is applied to a hypersonic laminar flow at Mach 7.95 around a cone at different incidence angles. The circumferential pressure distribution is compared with the experimental data. The cross-sectional Mach number contours are also presented. It is shown that in addition to the outer shock, a cross-flow shock wave is also present in the flow field. The cases of supersonic turbulent flows with Mach number 3 around a tangent-ogive with incidence angles of 6° and a secant-ogive with incidence angles of 10° are also studied. The circumferential pressure distributions are compared with the experimental data and the Euler code results and good agreement is obtained. The cross-sectional Mach number contours are also presented. It is shown that in this case also in addition to the outer shock, a cross-flow shock wave is also present at the incidence angle of 10°.

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