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.

scholarly journals FLOW SIMULATION OVER REENTRY CAPSULE AT SUPERSONIC AND HYPERSONIC SPEEDS. THE COMPARE BETWEEN TWO REENTRY CAPSULES. PART 2

2020 ◽  
pp. 52-60
Author(s):  
Pavel Timofeev ◽  
Keyword(s):  
Mach Number ◽  
Drag Coefficient ◽  
Numerical Algorithms ◽  
Flow Simulation ◽  
Ansys Fluent ◽  
Hypersonic Speeds ◽  
Higher Temperature ◽  
Mach Numbers ◽  
Lower Temperature ◽  
Cfd Method

This study is a continuation of the first part “Flow simulation over reentry capsule at supersonic and hypersonic speeds. The comparison between two reentry capsules. Part 1”. This study presents a comparison between two reentry capsules, which are made in Russia. 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 two tips reentry capsule understand the behavior of supersonic and hypersonic flow and its effect on the reentry capsule; compare temperature results and compare drag coefficient for the range Mach numbers equals 2–6. This study showed results on velocity counters, on temperature counters and drag coefficient 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 the “Prototype 2” reentry vehicle has a lower temperature a lateral surface, than the “Prototype 1”, but a higher temperature over the frontal segment at Mach number M = 6. A general trend has been observed, which is that the resistance drag coefficient decreases with increasing Mach number, which is associated with high height and low density and pressure values. The relevance and significance of this problem for the design of new and modernization of old reentry capsules.

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 Critical Mach Numbers of Flow around Two-Dimensional and Axisymmetric Bodies

Aerodynamics ◽  
2021 ◽  
Author(s):  
Vladimir Frolov
Keyword(s):  
Mach Number ◽  
Cross Section ◽  
Flow Simulation ◽  
The Body ◽  
Two Dimensional ◽  
Critical Mach Number ◽  
Method Of Integral Relations ◽  
Viscosity Factor ◽  
Mach Numbers ◽  
Axisymmetric Bodies

The paper presents the calculated results obtained by the author for critical Mach numbers of the flow around two-dimensional and axisymmetric bodies. Although the previously proposed method was applied by the author for two media, air and water, this chapter is devoted only to air. The main goal of the work is to show the high accuracy of the method. For this purpose, the work presents numerous comparisons with the data of other authors. This method showed acceptable accuracy in comparison with the Dorodnitsyn method of integral relations and other methods. In the method under consideration, the parameters of the compressible flow are calculated from the parameters of the flow of an incompressible fluid up to the Mach number of the incoming flow equal to the critical Mach number. This method does not depend on the means determination parameters of the incompressible flow. The calculation in software Flow Simulation was shown that the viscosity factor does not affect the value critical Mach number. It was found that with an increase in the relative thickness of the body, the value of the critical Mach number decreases. It was also found that the value of the critical Mach number for the two-dimensional case is always less than for the axisymmetric case for bodies with the same cross-section.

Optimization of Aerodynamic Parameters of Cropped Delta Wing with Fence at Sonic Mach Number

2019 ◽  
Vol 16 (2) ◽  
pp. 403-409
Author(s):  
M. P. Arun ◽  
M. Satheesh ◽  
Edwin Raja J. Dhas
Keyword(s):  
Shock Wave ◽  
Boundary Condition ◽  
Mach Number ◽  
Sound Speed ◽  
Delta Wing ◽  
Leading Edge ◽  
Supersonic Speed ◽  
Sweep Angle ◽  
Ansys Fluent ◽  
Aerodynamic Parameters

Manufacturing and maintaining different aircraft fleet leads to various purposes, which consumes more money as well as man power. Solution to this, nations that are leading in the field of aeronautics are performing much research and development works on new aircraft designs that could do the operations those were done by varied aircrafts. The foremost benefit of this delta wing is, along the huge rearward sweep angle, the wing’s leading edge would not contact the boundary of shock wave. Further, the boundary is produced at the fuselage nose due to the speed of aircraft approaches and also goes beyond the transonic to supersonic speed. Further, rearward sweep angle greatly worse the airspeed: wings under normal condition to leading edge, so permits the aircraft to fly at great transonic, subsonic, or supersonic speed, whereas the over wing speed is kept to minimal range than that of the sound speed. The cropped delta wing with fence has analysed in three cases: Fences at 3/4th distance from the centre, with fences at half distance from the centre and with fences at the centre. Further, the delta wing that cropped is exported to ANSYS FLUENT V14.0 software and analysed by making the boundary condition settings like sonic Mach number of flow over wing along with the angle of attack.

Three-Dimensional Passive and Active Control Methods of Shock Wave Train Physics in a Duct

2016 ◽  
Vol 08 (04) ◽  
pp. 1650047 ◽  
Author(s):  
Reza Kamali ◽  
Seyed Mahmood Mousavi ◽  
Danial Khojasteh
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
High Performance ◽  
Wave Train ◽  
Three Dimensional ◽  
Maximum Flow ◽  
Control Methods ◽  
Shock Train ◽  
Ansys Fluent ◽  
Starting Point

In the present work, the physics of a three-dimensional shock train in a convergent-divergent nozzle is numerically investigated. In this regards, the Ansys-Fluent Software with Algebraic Wall-Modeled Large-Eddy Simulation (WMLES) is used. To estimate precision and errors accumulation we used the Smirinov’s method; fine flow structures are obtained via Laplacian of density called shadowgraph and the shock parameter is defined as multiplication of flow Mach number by the normalized pressure gradient, in which shock wave structures are visible distinctly. The results are compared with the experimental data of Weiss et al. [Experiments in Fluids 49(2) (2010) 355–365], in the same conditions including geometry, boundary conditions, etc. The results show that there is good agreement with experimental trends concerning wall pressure and centerline Mach number profiles. Therefore, the focus of the present study is an assessment of various flow control methods to change the shock structures. Consequently, we investigated the effects of passive (bump and cavity) and active (suction and blowing) control methods on the starting point of shock, shock strength, minimum pressure, maximum flow Mach number, etc. All CFD investigations are carried out by High Performance Computing Center (HPCC).

Shock Wave–Film Cooling Interactions in Transonic Flows

2001 ◽  
Vol 123 (4) ◽  
pp. 788-797 ◽  
Author(s):  
P. M. Ligrani ◽  
C. Saumweber ◽  
A. Schulz ◽  
S. Wittig
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Film Cooling ◽  
Free Stream ◽  
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 deg. Free-stream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to free-stream 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 crossflow 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 free-stream 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 crossflow Mach number) is utilized. Effectiveness values measured with a supersonic approaching free-stream and shock waves then decrease as the injection crossflow 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.

The interaction of an isotropic field of acoustic waves with a shock wave

1995 ◽  
Vol 300 ◽  
pp. 383-407 ◽  
Author(s):  
Krishnan Mahesh ◽  
Sangsan Lee ◽  
Sanjiva K. Lele ◽  
Parviz Moin
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Kinetic Energy ◽  
Acoustic Waves ◽  
Sound Pressure ◽  
Linear Analysis ◽  
Sound Level ◽  
Pressure Level ◽  
Acoustic Fluctuations ◽  
Mach Numbers

Moore's (1954) inviscid linear analysis of the interaction of a shock wave with a plane acoustic wave is evaluated by comparison to computation. The analysis is then extended to study the interaction of an isotropic field of acoustic waves with a normal shock wave. The evolution of fluctuating kinetic energy, sound level and thermodynamic fluctuations across the shock wave are examined in detail.The interaction of acoustic fluctuations with the shock is notably different from that of vortical fluctuations. The kinetic energy of the acoustic fluctuationsdecreasesacross the shock wave for Mach numbers between 1.25 and 1.8. For Mach numbers exceeding 3, the kinetic energy amplifies by levels that significantly exceed those found in the interaction of vortical fluctuations with the shock. Upon interacting with the shock wave, the acoustic waves generate vortical fluctuations whose contribution to the far-field kinetic energy increases with increasing Mach number. The level of sound increases across the shock wave. The rise in the sound pressure level across the shock varies from 5 to 20 dB for Mach number varying from 1.5 to 5. The fluctuations behind the shock wave are nearly isentropic for Mach number less than 1.5, beyond which the generation of entropy fluctuations becomes significant.

The shape of a diffracting shock wave

1967 ◽  
Vol 29 (2) ◽  
pp. 297-304 ◽  
Author(s):  
B. W. Skews
Keyword(s):  
Shock Wave ◽  
Experimental Study ◽  
Mach Number ◽  
Sound Wave ◽  
Diffraction Theory ◽  
Incident Shock ◽  
Linearized Theory ◽  
Shock Mach Number ◽  
Mach Numbers ◽  
Good For

This paper describes an experimental study of the shape of a shock diffracting around a corner made up of two plane walls, for corner angles from 15 to 165° (in 15° steps) and shock Mach numbers from M0 = 1·0 to 4·0. The results are compared with profiles determined from the diffraction theory of Whitham (1957, 1959). The agreement is shown to be good for an incident shock Mach number of 3·0, and fair in other cases. The behaviour is found to follow the trends established by Lighthill (1949) in a linearized theory. Results for the Mach number of the wall shock are also presented. The shock does not degenerate to a sound wave even for large corner angles and low Mach numbers.

Design and Numerical Simulation of Convergent Divergent Nozzle

2016 ◽  
Vol 852 ◽  
pp. 617-624
Author(s):  
V. Ramji ◽  
Raju Mukesh ◽  
Inamul Hasan
Keyword(s):  
Mach Number ◽  
Computational Analysis ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Minimum Length ◽  
Ansys Fluent ◽  
Exit Pressure ◽  
Constant Height ◽  
Mach Numbers ◽  
Exit Mach Number

This works centers on the design of a De Laval (convergent - Divergent) nozzle to accelerate the flow to supersonic or hypersonic speeds and computational analysis of the same. An initial design of the nozzle is made from the method of characteristics. The coding was done in Matlab to obtain the contour of the divergent section for seven different exit Mach numbers viz. 2.5,3,3.5,4,4.5,5 and 5.5.To quantify variation in the minimum length of the nozzle divergent section with respect to the exit mach number, a throat of constant height (0.005m) and width (0.05m) was chosen for all the design. The area exit required for each mach no varying from 1 to 5.5 was plotted using isentropic relations and was also used to verify the exit area of the nozzle for each of those mach numbers. An estimate of the exit pressure ratio is obtained by using isentropic and normal shock relations. With this exit pressure ratio, a more refined verification is done by computational analysis using ANSYS Fluent software for a contour nozzle with exit Mach number 5.5. The spalart Allmaras and k-epsilon model were used for turbulence modeling.

scholarly journals Simulation of Back Pressure Effect on Behaviour of Convergent Divergent Nozzle

2013 ◽  
Vol 6 (1) ◽  
pp. 105-120
Author(s):  
Nazar Muneam Mahmood
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Gas Flow ◽  
Static Pressure ◽  
Flow Region ◽  
Back Pressure ◽  
Cross Sectional ◽  
Ansys Fluent ◽  
One Dimensional ◽  
Velocity Magnitude

In this research a simulation of steady flow of a gas through a convergent divergent nozzle which has a varying cross sectional area will be considered. The nature of the flow can be explained by considering how the flow and its characteristics in the nozzle changes as nthe back pressure Pb is decreased.The characteristics of gas flow i.e.(Mach number, static pressure, density, velocity magnitude and static temperature) distributions for the convergent divergent nozzle are implemented by using the ANSYS Fluent 12.1 software to solve the quasi-one dimensional nozzle flow.The reductions in the back pressure cannot affect conditions upstream of the throat. The nozzle is, therefore, choked. The shock wave increases the pressure, density and temperature and reduces the velocity and Mach number to a subsonic value, and as back pressure is further reduced to a certain value, the extent of the supersonic flow region increases, the shock wave moving further down the divergent portion of the nozzle towards the exit plane.

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