Shock Waves最新文献

Pre-ignition is an undesired combustion event known to restrict kinetic modeling validation. Previous methane oxidation studies reported premature ignition as part of ignition delay time measurements in shock tubes. In this context, the effect on the pre-ignition propensity and auto-ignition behavior of stoichiometric methane mixtures at different dilution levels of (hbox {N}_2), Ar, He, and (hbox {CO}_2) was studied at 10 bar and 25 bar and temperatures between 1080 K and 1350 K. In addition to conventional sidewall pressure and endwall light emission measurements, a high-speed imaging setup was utilized to visualize the ignition process. Relevant physicochemical parameters to describe and predict the pre-ignition phenomenon were used. The results suggest that dilution levels up to (80%) of bath gas are not successful in mitigating early ignition occurrence and its effects at moderate pressures. Replacing (hbox {N}_2) by He was found to suppress early ignition at 10 bar, attributed to an enhanced dissipation of temperature inhomogeneities in the test gas section. The present findings demonstrate that (hbox {CO}_2) has potential for pre-ignition heat release mitigation, while Ar was confirmed to promote premature ignition. To the best of our knowledge, we present the first detailed study on pre-ignition mitigation for methane mixtures in shock tubes, where further insights into its ignition non-idealities are given.

A new reflected shock tunnel capable of generating hypersonic environments at realistic flight enthalpies has been commissioned at Sandia. The tunnel uses an existing free-piston driver and shock tube coupled to a conical nozzle to accelerate the flow to approximately Mach 9. The facility design process is outlined and compared to other ground test facilities. A representative flight-enthalpy condition is designed using an in-house state-to-state solver and piston dynamics model and evaluated using quasi-1D modeling with the University of Queensland L1d code. This condition is demonstrated using canonical models and a calibration rake. A 25-cm core flow with 4.6-MJ/kg total enthalpy is achieved over an approximately 1-ms test time. The condition was refined using analysis and a heavier piston, leading to an increase in test time. A novel high-speed molecular tagging velocimetry method is applied using in situ nitric oxide to measure the freestream velocity of approximately 3016 m/s. Companion simulation data show good agreement in exit velocity, pitot pressure, and core flow size.

When departure from the ideal gas equation of state is considered, the Noble–Abel stiffened gas model is an appealing and versatile candidate due to its simple form. The linear interaction approximation formalism is extended to consider non-ideal gas effects introduced by this equation of state. Kovásznay decomposition and adequate definition of the energy of disturbances are provided in the context of this equation of state. Changes with respect to ideal gas are investigated on transfer functions, critical angle, and compression factor. Those differences yield concrete effects on the damping and transfer of fluctuations across shock waves. Those changes are further illustrated by considering the interaction of an entropy spot with a Mach 3 stationary shock wave.

Detonation propagation in a stratified layer of combustible gas over an inert gas was investigated experimentally. The layer formed in a 12.7-mm-wide channel by opening a sliding door that initially separated a nitrogen-diluted stoichiometric hydrogen–oxygen mixture from argon, or nitrogen. As the lighter combustible gas layer spreads axially down the channel, diffusion across the interface produces a composition gradient across the layer height. A steady detonation wave, generated by deflagration-to-detonation transition in the driver section before the door location, was transmitted into the combustible layer. The axial distance the layer spreads and the amount of mass diffusion across the layer were controlled by the flame ignition delay time after the door opens. Schlieren video and soot foils were used to measure the extent of detonation propagation through the layer. It was shown that detonation propagation through the layer is self-limiting due to over-mixing at the layer leading edge. Three-dimensional numerical simulations, including viscous and multicomponent mass diffusion effects, predicted the composition distribution within the layer. The cell size distribution, calculated based on the theoretical ZND induction zone length, corresponding to the simulation composition distribution showed that a cell size gradient-based failure criterion successfully predicted the extent of propagation in the layer.

Separation shock unsteadiness in an incident-shock-induced interaction with and without control is evaluated in a Mach 2.05 flow using a (14^{circ }) shock generator. An array of mechanical vortex generators (MVGs) in the form of rectangular vanes (MVG1), ramp vanes (MVG2), and a delta ramp (MVG3) is placed (14delta ) upstream of the interaction region ((delta =5.2,{textrm{mm}}) being the local boundary layer thickness at the interaction). Among all the devices tested, MVG1 shows a maximum reduction of the separation length (about 28% relative to the no-control case). The spectra at separation also show a shift in the dominant frequency from 220 Hz without control to 539 Hz with MVG1. Interestingly, the peak rms (root mean square) value is seen to occur with control at much larger intermittency values ((upgamma _{upsigma ,{textrm{max}}}=0.92) for MVG1) in contrast to the no-control case in which it occurs at (upgamma _{upsigma ,{textrm{max}}}=0.5) as reported so far. The auto-correlation at the separation and reattachment shock locations indicates the presence of relatively small-scale structures with control as compared to the case without control. Out of all the control cases tested, MVG1 exhibits better separation control with a relatively lower unsteadiness level.

Normal detonation reflection generates a shock wave that exhibits complicated dynamics as it propagates through the incident detonation and post-detonation flow. Ideal models have historically neglected the influence of a finite detonation thickness on the reflected shock due to its small size relative to laboratory scales. However, one-dimensional numerical simulations show that the reflected shock accelerates to a large shock speed not predicted by ideal theory as it propagates through the incident detonation. Analysis with a derived shock-change equation identifies the principal role of the highly nonuniform upstream flow on producing the large shock acceleration. Simulations of detonation reflection show how a finite detonation thickness affects the entire trajectory of the reflected shock.

Computational fluid dynamics shows that a shock wave can detach from the sharp leading edge of a curved wedge at a wedge angle smaller than the classical maximum flow deflection as well as the sonic wedge angle. This is attributed to the inability of the sonic flow, at the wedge trailing edge, to pass as much mass flow as is being admitted through the shock wave attached at the leading edge. At this condition, the flow is unsteady, causing both the sonic surface and the shock to make adjustments in their shapes and positions to achieve a steady state with mass-flow balance. As a result, the shock wave becomes detached. Time-accurate CFD calculations show the gasdynamic details of the adjustment where the flow and the detached shock assume a steady state as the mass-flow imbalance gradually decreases to zero. This mechanism of shock detachment, occurring near the leading edge, is called local choking to distinguish it from shock detachment due to global choking that occurs because of flow choking at the exit of a convergent duct and to distinguish it as well from detachment due to an excessive leading-edge deflection. The local choking mechanism has been postulated to be a cause of shock detachment from doubly curved wedges. An analysis, based on curved shock theory and confirmed by CFD, shows that local choking and shock detachment from a doubly curved leading edge are dependent on Mach number, wedge angle, wedge curvature (both streamwise and cross-stream), and wedge length.

The Hugoniot shock wave velocity ((U_{textrm{S}}))–particle velocity ((u_{textrm{p}})) curve of polymethyl methacrylate (PMMA) was measured in an experiment using only 2.5 g of a high explosive. The thickness of the plate was varied to accurately determine (U_{textrm{S}}) at an arbitrary position in the PMMA. Image analysis was conducted to obtain the x–t diagram of shock wave propagating in PMMA along the axis of the explosive, and its derivative was used to obtain the on-axis (U_{textrm{S}}) at an arbitrary location. Using the pressure measurement results and (U_{textrm{S}}) values, the Hugoniot (U_{textrm{S}})–(u_{textrm{p}}) curve of PMMA was obtained by calculating (u_{textrm{p}}) from the momentum conservation law. The results are in very good agreement with the reported values for flat-plate impact experiments conducted using an impact gun. It was found that the Hugoniot (U_{textrm{S}})–(u_{textrm{p}}) curve of PMMA on the low-pressure side ((u_{textrm{p}} < 0.5, {textrm{km}}/{textrm{s}})) can be evaluated with high accuracy using a simple measurement method that does not use plane waves.

Turbulence modeling has the potential to revolutionize high-speed vehicle design by serving as a co-equal partner to costly and challenging ground and flight testing. However, the fundamental assumptions that make turbulence modeling such an appealing alternative to its scale-resolved counterparts also degrade its accuracy for practical high-speed configurations, especially when fully 3D flows are considered. The current investigation develops a methodology to improve the performance of turbulence modeling for a complex Mach 8.3, 3D shock boundary layer interaction (SBLI) in a double fin geometry.A representative two-equation model, with low-Reynolds-number terms, is used as a test bed. Deficiencies in the baseline model are first elucidated using benchmark test cases involving a Mach 11.1 zero pressure gradient boundary layer and a Mach 6.17 flow over an axisymmetric compression corner. From among different possibilities, two coefficients are introduced to inhibit the non-physical over-amplification of (i) turbulence production and (ii) turbulence length-scale downstream of a shock wave. The coefficients rely on terms already present in the original model, which simplifies implementation and maintains computational costs. The values of the coefficients are predicated on the distribution of turbulence quantities upstream of the shock; this ensures that the modifications do not degrade the model predictions in simpler situations such as attached boundary layers, where they are unnecessary. The effects of the modifications are shown to result in significant improvements in surface pressure and wall heat flux for the 3D SBLI test case, which contains numerous features not observed in 2D situations, such as 3D separation, skewed boundary layers, and centerline vortices. Considerations on the inflow values of turbulence variables and mesh resolution are provided.

Full-field direct simulation of sonic boom has only been applied to the analysis of axisymmetric geometries. In this work, a more realistic analysis of complex geometries over buildings is achieved by employing a combination of the following four numerical approaches: (i) a hierarchical structured adaptive mesh refinement method, (ii) a ghost fluid method for incorporating the immersed boundary conditions on the solid–fluid interfaces, (iii) a well-balanced finite volume method to allow stable stratification of the atmosphere, and (iv) a segmentation method of the computational domain to increase the efficiency of the computations. The three-dimensional Euler equations with a gravitational source term are solved over a stratified atmosphere. The simulation is split into two stages. Firstly, the entire flow field that involves a delta wing body is solved without buildings. Thereafter, the flow behaviors near the ground are recomputed considering rectangular and L-type buildings. Computational results show that the near- and far-fields waveforms are comparable to those from the wind tunnel experiment and the waveform parameter method, respectively. The waveform shape behind the shock waves is spiked due to the diffracted waves around buildings, with the spiking effect in L-type buildings being stronger than that in rectangular buildings. The pressure rises for rectangular and L-type buildings are significantly amplified due to double and triple reflections, respectively, each with an amplification factor comparable to the theoretical value. These results indicate that full-field simulation is promising for analyzing three-dimensional characteristics of sonic boom emanating from complex geometries passing over buildings.