scholarly journals Unsteady shock wave dynamics in accelerating and decelerating flight

2021 ◽  
pp. 1-23
Author(s):  
I. Mahomed ◽  
H. Roohani ◽  
B.W. Skews ◽  
I.M.A. Gledhill
Keyword(s):  
Shock Wave ◽  
Steady State ◽  
Mach Number ◽  
Cone Angle ◽  
Shock Interactions ◽  
Acceleration Technique ◽  
Axial Acceleration ◽  
Wave Development ◽  
History Effect ◽  
Commercial Solver

Abstract Increasingly agile manoeuvre is an advantage in the flight of aircraft, missiles and aerial vehicles, but the principles of accelerating aerodynamics in the transonic regime are only now being fully investigated. This study contributes to the understanding of shock and separation effects on drag during axial acceleration, using a simple geometric configuration. Unsteady shock wave behavior was numerically investigated for an axisymmetric cone-cylinder using a commercial solver and the Moving Reference Frame acceleration technique. This acceleration technique was validated using unsteady numerical and experimental methods. The cone-cylinder was accelerated from Mach number 0.6 to Mach number 1.2 at 100g constant and deceleration was from Mach number 1.2 until Mach number 0.6 at –100g constant. Three cone angles were tested for the cone-cylinder with uniform cylinder diameter. Acceleration through the transonic Mach regime was characterised by a delayed and gradual shock wave development when compared to steady state, demonstrating a clear flow history effect. Deceleration through the transonic Mach regime was characterised by shock wave propagation from the base to the nose. New flow structures appeared during deceleration that do not have counterparts in the steady state, including shock interactions and propagating expansion-compression features. Gross changes in the unsteady drag coefficient curves for each cone-angle are explained with reference to unsteady shock wave behaviour for accelerating and decelerating motion.

scholarly journals Effect of wavy wall strut fuel injector on shock wave development and mixing enhancement of fuel and air for a scramjet combustor

2020 ◽  
Author(s):  
Obula Reddy Kummitha ◽  
K M Pandey
Keyword(s):  
Shock Wave ◽  
Mach Number ◽  
Shock Waves ◽  
Shear Layer ◽  
Wave Generation ◽  
Wavy Wall ◽  
Fuel Injector ◽  
Shock Wave Generation ◽  
Shear Mixing ◽  
Wave Development

Abstract The shear mixing and streamline vortices are the notable parameters to influence the air–fuel mixing in hypersonic flows. The shock wave development and Mach number significantly influence the shear mixing phenomenon. Hence, this research introduced an unconventional strut and tested its performance for the generation of shock waves at different flow conditions (M = 2,4,6). The Reynolds-averaged Navier–Stokes equations are solved to evaluate the performance of the new strut. Both the DLR scramjet strut injector and wavy wall strut injector are assessed for the shear mixing development. Turbulence for the association of shock waves, mixing layer, and the boundary layer has been modeled with the SST k-ω model. The variation in shock development and its interactions are investigated further with an increase in Mach number. The scramjet flow structure differentiation found the increased number of oblique shock waves with the wavy wall strut fuel injector. It increases the turbulence level with increased streamline vortices, turbulent intensity, and turbulent kinetic energy. The shock wave generation analysis at different Mach numbers (M = 2,4,6) found fewer interactions between the shock wave and shear layer with increased Mach number. From the examination of shock wave generation and its interaction with the shear layer and analysis of turbulent parameters, it is found that the wavy wall strut has an appreciable effect on shock-induced blend augmentation of fuel and air.

A numerical study on the single pulsed energy addition based unsteady wave drag reduction at varied hypersonic flow regimes

2020 ◽  
pp. 095441002097313
Author(s):  
Dathi SNV Rajasekhar Rao ◽  
Bibin John
Keyword(s):  
Steady State ◽  
Mach Number ◽  
Drag Reduction ◽  
Shock Layer ◽  
Blunt Body ◽  
Wave Drag ◽  
Flow Features ◽  
Energy Pulse ◽  
Energy Addition ◽  
Wave Drag Reduction

In this study, unsteady wave drag reduction in hypersonic flowfield using pulsed energy addition is numerically investigated. A single energy pulse is considered to analyze the time-averaged drag reduction/pulse. The blast wave creation, translation and its interaction with shock layer are studied. As the wave drag depends only on the inviscid aspects of the flowfield, Euler part of a well-established compressible flow Navier-Stokes solver USHAS (Unstructured Solver for Hypersonic Aerothermodynamics) is employed for the present study. To explore the feasibility of pulsed energy addition in reducing the wave drag at different flight conditions, flight Mach numbers of 5.75, 6.9 and 8.0 are chosen for the study. An [Formula: see text] apex angle blunt cone model is considered to be placed in such hypersonic streams, and steady-state drag and unsteady drag reductions are computed. The simulation results indicate that drag of the blunt-body can be reduced below the steady-state drag for a significant period of energy bubble-shock layer interaction, and the corresponding propulsive energy savings can be up to 9%. For energy pulse of magnitude 100mJ deposited to a spherical region of 2 mm radius, located 50 mm upstream of the blunt-body offered a maximum percentage of wave drag reduction in the case of Mach 8.0 flowfield. Two different flow features are found to be responsible for the drag reduction, one is the low-density core of the blast wave and the second one is the baroclinic vortex created due to the plasma energy bubble-shock layer interaction. For the same freestream stagnation conditions, these two flow features are noted to be very predominant in the case of high Mach number flow in comparison to Mach 5.75 and 6.9 cases. However, the ratio of energy saved to the energy consumed is noted as a maximum for the lower Mach number case.

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

Foreshock wave transmission into the magnetosheath and magnetosphere: results from global hybrid-Vlasov simulations

2021 ◽  
Author(s):  
Lucile Turc ◽  
Markus Battarbee ◽  
Urs Ganse ◽  
Andreas Johlander ◽  
Yann Pfau-Kempf ◽  
...  
Keyword(s):  
Solar Wind ◽  
Mach Number ◽  
Cone Angle ◽  
Low Frequency ◽  
Compressional Wave ◽  
Wave Transmission ◽  
Wave Power ◽  
Solar Wind Flow ◽  
Magnetic Pulsations ◽  
Wave Properties

<p>The foreshock, extending upstream of the quasi-parallel shock and populated with shock-reflected particles, is home to intense wave activity in the ultra-low frequency range.<em> </em>The most commonly observed of these waves are the “30 s” waves, fast magnetosonic waves propagating sunward in the plasma rest frame, but carried earthward by the faster solar wind flow. These waves are thought to be the main source of Pc3 magnetic pulsations (10 – 45 s) in the dayside magnetosphere. A handful of case studies with suitable spacecraft conjunctions have allowed simultaneous investigations of the wave properties in different geophysical regions, but the global picture of the wave transmission from the foreshock through the magnetosheath into the magnetosphere is still not known. In this work, we use global simulations performed with the hybrid-Vlasov model Vlasiator to study the Pc3 wave properties in the foreshock, magnetosheath and magnetosphere for different solar wind conditions. We find that in all three regions the wave power peaks at higher frequencies when the interplanetary magnetic field strength is larger, consistent with previous studies. While the transverse wave power decreases with decreasing Alfvén Mach number in the foreshock, the compressional wave power shows little variation. In contrast, in the magnetosheath and the magnetosphere, the compressional wave power decreases with decreasing Mach number. Inside the magnetosphere, the distribution of wave power varies with the IMF cone angle. We discuss the implications of these results for the propagation of foreshock waves across the different geophysical regions, and in particular their transmission through the bow shock.</p>

Simulation of a Steady-State Electron Shock Wave in a Submicron Semiconductor Device Using High-Order Upwind Methods

1991 ◽  
pp. 27-32 ◽  
Author(s):  
Emad Fatemi ◽  
Carl L. Gardner ◽  
Joseph W. Jerome ◽  
Stanley Osher ◽  
Donald J. Rose
Keyword(s):  
Shock Wave ◽  
Steady State ◽  
Semiconductor Device ◽  
High Order ◽  
State Electron ◽  
Upwind Methods

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.

scholarly journals A study of the interaction of a travelling shock wave with a solid-propellant rocket motor head end

2021 ◽  
Author(s):  
Robert Elian Feteanu
Keyword(s):  
Shock Wave ◽  
Rocket Motor ◽  
Combustion Chambers ◽  
Cfd Simulations ◽  
Radial Wave ◽  
Flow Models ◽  
Numerical Studies ◽  
Solid Propellant Rocket ◽  
Wave Development ◽  
Solid Rocket

Experimental and numerical studies have been undertaken to examine various aspects pertaining to the interaction of an incident travelling shock wave with a solid rocket motor's head end (forward section), in order to identify any potential gasdynamic mechanism of wave reinforcement pertinent to combustion instability behaviour in these motors. A cold-flow experiment, based on a shock tube scheme tailored to the present application, has proved to be useful in providing information surrounding the interaction process. Both experimental and numerical results (CFD simulations) confirm the existence of substantial transient radial wave development superimposed on the base reflected axial shock wave. These results illustrate the potential weakness of one-dimensional flow models for certain engineering applications, where important multidimensional phenomena, such as those observed in this work, may not be captured. By analogy to actual propulsion system combustion chambers, the transverse wave activity is potentially a factor in supporting an augmentation of the local combustion rate in the head-end region of a rocket motor combustor.

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.

Hypersonic Flow Separation in Shock Wave Boundary Layer Interactions

1992 ◽  
Author(s):  
A. Hamed ◽  
Ajay Kumar
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Flow Separation ◽  
Characteristic Length ◽  
Separated Flow ◽  
Surface Cooling ◽  
Flat Plates ◽  
Shock Interactions ◽  
Wave Boundary ◽  
Incipient Separation

This work presents an assessment of the experimental data on separated flow in shock wave turbulent boundary layer interactions at hypersonic and supersonic speeds. The data base consist of selected configurations where the only characteristic length in the interation is the incoming boundary layer thickness. It consists of two dimensional and axisymmetric interactions in compression corners or cylinder-flares, and externally generated oblique shock interactions with boundary layers over flat plates or cylindrical surfaces. The conditions leading to flow separation and the empirical correlations for incipient separation are reviewed. The effects of Mach number, Reynolds number, surface cooling and the methods of detecting separation are discussed.

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