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形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2021-11-04, PA2220039000

Current design constraints have encouraged the studies of aeroacoustics fields around compressible jet flows. The present work addresses the numerical study of unsteady turbulent jet flows for aeroacoustic analyses of main engine rocket plumes. A novel large eddy simulation (LES) tool is developed in order to reproduce high fidelity results of compressible jet flows which could be used for aeroacoustic studies with the Ffowcs Williams and Hawkings approach. The numerical solver is an upgrade of an existing Reynolds-averaged Navier-Stokes solver previously developed in the group. The original framework is rewritten in a modern fashion and intensive parallel computation capabilities have been added to the code. The LES formulation is written using the finite difference approach. The energy equation is carefully discretized in order to model the energy equation of the filtered Navier-Stokes formulation. The classical Smagorinsky model is the chosen subgrid scale closure for the present work. Numerical simulations of perfectly expanded jets are performed and compared with the literature in order to validate the new solver. Moreover, speedup and the computational performance of the code are evaluated and discussed. Flow results are used for an initial evaluation of the noise radiated from the rocket plume., 25 pages conference paper

Volcanic jet flows in explosive eruptions emit radio frequency signatures, indicative of their fluid dynamic and electrostatic conditions. The emissions originate from sparks supported by an electric field built up by the ejected charged volcanic particles. When shock-defined, low-pressure regions confine the sparks, the signatures may be limited to high-frequency content corresponding to the early components of the avalanche-streamer-leader hierarchy. Here, we image sparks and a standing shock together in a transient supersonic jet of micro-diamonds entrained in argon. Fluid dynamic and kinetic simulations of the experiment demonstrate that the observed sparks originate upstream of the standing shock. The sparks are initiated in the rarefaction region, and cut off at the shock, which would limit their radio frequency emissions to a tell-tale high-frequency regime. We show that sparks transmit an impression of the explosive flow, and open the way for novel instrumentation to diagnose currently inaccessible explosive phenomena., Comment: 9 pages, 6 figures

Sonic boom reduction has been an urgent need for the development of future supersonic transport, because of the heavy damage of noise pollution. This paper provides a novel concept for supersonic aircraft to reduce the sonic boom and drag coefficient, wherein a suction slot near the leading edge and an injection slot near the trailing edge on the airfoil suction surface are opened. To make sure of a zero net mass flux flow control, the mass flow sucked in near the leading edge is equal to the mass flow injected near the trailing edge. The diamond and NACA0008 airfoils are adopted as the baseline airfoil to verify the capability of the proposed design method. The effects of the suction and injection location, the suction and injection slot size, the mass flow rate and the attack angle on the ground boom signature and drag coefficient are studied in detail. The results show that the optimized airfoils with the suction and injection have benefits in both sonic boom reduction and wave drag reduction. And the reduction of the sonic boom intensity is more sensitive to the injection near the trailing edge than the suction near the leading edge. From the viewpoint of aerodynamics, opening the suction and injection slots will have no adverse effect on the aerodynamic performances of the supersonic aircraft and even increase the lift-drag ratio under some circumstances. For energy saving, the suction and injection slots can be selectively opened, which are opened when the supersonic aircraft flies over the city but are closed when the aircraft flies over the sea.

We have developed a new type of simple lattice Boltzmann (LB) model for the compressible Euler equations based on the collisionless kinetic-equation approach proposed by Sone. The model uses the collisionless kinetic equation in the streaming process, and modifies the distribution function to its Chapman-Enskog type at each time step. Compared with the current LB models which solve the kinetic equation of the BGK type, the proposed model is superior in the following two points: (i) Inviscid flows can be computed stably while there is no such model in the current LBM; (ii) The velocity distribution function does not need to be memorized in computation. We calculate various inviscid compressible flows described by the compressible Euler equations using our new one-, two-, and three-dimensional models. Numerical results show the capability of our scheme to simulate high-speed supersonic flows with shock waves.

This paper presents a modified analytical model to evaluate the trajectories of various lift separation sabot configurations. The aerodynamic forces acting on the sabot surfaces during a supersonic flight are modeled in the present analytical model by incorporating the pressures on the windward side of the sabot due to the detached/attached shock and its reflections and then integrated using the 3-DoF dynamical equations. The trajectory and the aerodynamic coefficients were obtained for these configurations at a projectile Mach number of 3. The sabot configurations, which include two new designs, are compared with each other and with the conventional free flight trajectory data of the conventional sabots. The mechanical interaction between the sabot and projectile is also addressed in the present work. The comparison shows that the new designs with the aerodynamic surfaces close to the center of gravity, lift-off from the projectile with minimal mechanical interaction compared to a conventional sabot.

Since proposed, the self-similarity variables based genuinely multidimensional Riemann solver is attracting more attentions due to its high resolution in multidimensional complex flows. However, it needs numerous logical operations in supersonic cases, which limit the method’s applicability in engineering problems greatly. In order to overcome this defect, a hybrid multidimensional Riemann solver, called HMTHS (Hybrid of MulTv and multidimensional HLL scheme based on Self-similar structures), is proposed. It simulates the strongly interacting zone by adopting the MHLLES (Multidimensional Harten-Lax-van Leer-Eifeldt scheme based on Self-similar structures) scheme at subsonic speeds, which is with a high resolution by considering the second moment in the similarity variables. Also, it adopts the MULTV (Multidimensional Toro and Vasquez) scheme, which is with a high resolution in capturing discontinuities, to simulate the flux at supersonic speeds. Systematic numerical experiments, including both one-dimensional cases and two-dimensional cases, are conducted. One-dimensional moving contact discontinuity case and sod shock tube case suggest that HMTHS can accurately capture one-dimensional expansion waves, shock waves, and linear contact discontinuities. Two-dimensional cases, such as the double Mach reflection case, the supersonic shock / boundary layer interaction case, the hypersonic flow over the cylinder case, and the hypersonic viscous flow over the double-ellipsoid case, indicate that the HMTHS scheme is with a high resolution in simulating multidimensional complex flows. Therefore, it is promising to be widely applied in both scholar and engineering areas.

The role of N2 in the upper atmosphere on the atomic oxygen (AO)-induced erosion of polyimide in low Earth orbit (LEO) and sub-LEO is investigated through ground-based experiments and flight data. The experiment is performed by adding an Ar beam at the same collision energy as an undecomposed O2 component in the AO beam formed by laser detonation to simulate the physical effect of simultaneous N2 collision in sub-LEO. The Ar beam is added by the dual-pulsed supersonic valve-equipped laser-detonation system developed at Kobe University. The experimental results indicate that the erosion of polyimide in the laser-detonation system is promoted by the presence of O2 and Ar in the beam, corresponding to N2 in the sub-LEO. On-ground experimental results are compared with in-orbit AO measurements. Previous space shuttle, international space station-based exposure experiments, as well as the world’s first real-time sub-LEO material erosion data aboard a super low altitude test satellite (SLATS) orbiting at an altitude of 216.8 km are presented. The SLATS data suggests the presence of an acceleration effect by N2 collision on AO-induced polyimide erosion, as predicted by ground-based experiments.

In this study, a single-channel supersonic cascade model is investigated experimentally at a freestream Mach number of 2.4 to obtain a better understanding of the flow field evolution during the throttling process. A flap is placed at the channel exit to choke the flow linearly. Measurements include 1-kHz schlieren imaging and 10-kHz simultaneous fast-response wall pressure. Three stages, namely attached flow, separated flow, and oscillatory flow, are identified in the throttling process. The joint time–frequency analysis and wall pressure spectrum contour exhibit the time evolution and spatial distribution of the pressure fluctuation. With the increase in backpressure, the pressure fluctuation in the low-frequency shock oscillation range of 40–400 Hz on the suction surface located in the separated flow gradually enhances. The power spectral, coherence, and phase analyses of the schlieren images describe the dominant oscillation structure and its relationship with other regions. During the separated flow, the pressure change in the subsonic separated region first lead to a change in the state of the separated shear layer, after which the shock waves in the shock train, move. The oscillatory flow is a process wherein the upstream shock wave oscillates, causing the entire downstream channel to fluctuate.

The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions (SWTBLIs) in the hypersonic flow was investigated using a scaling analysis, in which the interaction length normalized by the displacement thickness of boundary layer was correlated with a corrected non-dimensional separation criterion across the interaction after accounting for the wall temperature effects. A large number of hypersonic SWTBLIs were compiled to examine the scaling analysis over a wide range of Mach numbers, Reynolds numbers, and wall temperatures. The results indicate that the hypersonic SWTBLIs with low Reynolds numbers collapse on the supersonic SWTBLIs, while the hypersonic cases with high Reynolds numbers show a more rapid growth of the interaction length than that with low Reynolds numbers. Thus, two scaling relationships are identified according to different Reynolds numbers for the hypersonic SWTBLIs. The scaling analysis provides valuable guidelines for engineering prediction of the interaction length, and thus, enriches the knowledge of hypersonic SWTBLIs.

In order to understand the physical phenomenon of the reflected shock/turbulent boundary layer interaction, the Large Eddy Simulation (LES) is conducted to investigate shock wave and turbulent boundary layer interaction in a 12 ° compression ramp with inlet high Mach number of 2.9. Rescaling/recycling method is used as inflow turbulence generation technique and validated on a supersonic flat plate turbulent boundary layer. The flow field of recycling plane in the plate computation domain is obtained to give the inlet boundary condition for the LES computation. This paper focuses on the reflected shock/turbulent boundary layer interaction region, where the fine flow structure and instantaneous flow field are analyzed in detail. It is found that the unsteady motion of the shock wave leads to the increase of wall pressure fluctuation.

With the explosive development of aerospace science, the design of the new generation airliner at higher speeds is attracting more attentions. To achieve this goal, it is necessary to achieve accurate prediction of the aerodynamic heating / force loads and successful reduction of drag and heat flux. As a remedy for the existing studies which are based upon the CFD and wind tunnel tests, this study presents a flight test for the drag and heat reduction spike technology. The principal goals of this flight test were to provide reference for verifying the accuracy of the prediction technology on ground and promote the development of the drag and heat reduction technology. By adopting the OS-X rocket, the TT-0 test vehicle designed by Shenyang Aircraft Design & Research Institute reached a maximum Mach number of 5.8 and a maximum altitude of 38 km. Hypersonic and supersonic pressure data by pressure scanning valves and heat fluxes by gauges at different locations were obtained successfully. Also, heat fluxes obtained by in-house CFD code are illustrated in comparison with the flight data. The results indicate that the numerical errors are large in most cases. More technologies, such as more CFD codes and more numerical procedures, should be adopted to conduct studies on this issue in the future.

The combustion modes in two different scramjet combustors with the mass flow rates of 1.8 kg/s and 3.6 kg/s are experimentally investigated to explore the scaling effects on supersonic combustion with a Mach number 2.0 inflow. It is found that the scramjet combustor with a larger scale can broaden the flame rich blowout limit. As the Equivalence Ratio (ER) increases, the combustion in the small-scale combustor maintains in the cavity-stabilized mode, and the flamebase moves downstream along the cavity shear layer; however, the combustion in the large-scale combustor gradually transfers from the cavity-stabilized mode to the jet-wake-stabilized mode. The differences in the cavity residence time, the ignition delay time and the Damkohler number caused by different scales of the scramjet combustor are likely to account for the scaling effects on the combustion modes.

Impact flash occurs when objects collide at supersonic speeds and can be used for real-time damage assessment when weapons rely on kinetic energy to destroy targets. However, the mechanism of impact flash remains unclear. A series of impact flash experiments of flat-head long-rod projectiles impacting thin target plates were performed with a two-stage light gas gun. The impact flash spectra for 6061 aluminum at 1.3–3.2 km/s collision speeds were recorded with a high-speed camera, a photoelectric sensor, and a time-resolved spectrometer. The intensity of the impact flash exhibited a pulse characteristic with time. The intensity (I) increased with impact velocity (V0) according to I∝V0n , where n = 4.41 for V0 > 2 km/s. However, for V0 2 km/s, a continuous spectrum (thermal radiation background) was observed and increased in intensity with V0. However, for V0

Numerical aero-acoustic analysis was conducted on a weapon bay model with doors, incorporating radar cross section reduction features. The effect of angle of attack on the aero-acoustic response of the cavity was analysed at a transonic Mach number of 0.85, and at a supersonic Mach number of 1.20. It was found that incidence had influence on both mean-flow features and acoustic response. Further, linear and angular accelerations induced by the flow on doors revealed potential adverse fluid-structure coupling when results were compared with modal analysis. Again, angle of attack did influence the aeroacoustic effects on the cavity door structure.

A rotating supersonic combustion engine (RSCE) is tested with various initiator tube positions along the combustion chamber to determine its impact on the ignition process. The type of fuel used is methane with oxygen as the oxidizer. The initiation of RSCE is assessed for nearstoichiometric methane-oxygen mixture to slightly rich mixture. Successful initiation of rotating supersonic combustion in the RSCE was obtained when the position of the initiator tube is located higher end of the combustion chamber that indicates a proper mixing of reactants. The mixing of reactants improves further up the annulus chamber of the RSCE.

In this paper effects of material constructions on natural frequencies and critical aerodynamic pressures are investigated. It is assumed that the rectangular plate is made of a polymeric matrix reinforced with graphene nanoplatelets or carbon nanotubes. A general closed analytical method of solution is presented. It is demonstrated that three parameters define entirely the location of the critical flutter pressure. The influence of material properties and transverse shear effects is characterized by a set of multipliers. They can be easily adopted in design procedures.

This paper proposes using a Nonlinear Energy Sink (NES) to suppress the nonlinear aeroelastic response of laminated composite panels in supersonic airflows. Relevant aeroelastic equations are established using Hamilton’s principle and a finite element approach, drawing upon Von Karman’s large deflection theory and first order piston theory. The idea of the NES suppression region is proposed and the effects of NES parameters on the NES suppression region are studied in detail. The results show that the nonlinear aeroelastic responses of the panel can be completely suppressed by the Transient Resonance Capture (TRC); the appropriate NES parameter values can increase the critical dynamic pressure for flutter and suppress the nonlinear aeroelastic response effectively. Increasing the mass ratio of the NES can improve the NES suppression region; the nonlinear stiffness coefficient and damping of the NES within a specific range can suppress the nonlinear aeroelastic response. The most effective installation position for a NES is in a specific region behind the center-line of the panel in the direction of the airflow.

El patrón de flujo en toberas supersónicas lo definen los perfiles aerodinámicos de la geometría de las paredes internas, entre otros parámetros, donde la garganta es una sección crítica. En el presente trabajo, el objetivo es analizar el comportamiento del flujo en la sección recta de la garganta de una tobera cónica experimental de un motor de cohete sonda de combustible sólido. El flujo sobreexpandido se simuló con el código ANSYS-Fluent en un dominio computacional 2D, empleando el modelo RANS y el modelo de turbulencia de Menter, y la ecuación de Sutherland para la viscosidad en función de la temperatura. Se llevaron a cabo cinco casos de estudio para la longitud de garganta en el rango de 1-10 mm. Se obtuvo fluctuaciones de número de Mach, presión y temperatura, ondas de choque oblicuas en la sección de la garganta para la longitud de 10 mm; para longitudes menores disminuyó la intensidad de la magnitud del choque. Se concluye que para la longitud de garganta de 1 mm el flujo es transónico sin la presencia de choques oblicuos. En la sección divergente, las ondas de choque varían en su intensidad y cambian de posición.

Supersonic shear layer development affects mixing between the primary and secondary fluids in a supersonic ejector substantially. The splitter plate plays a key role in shear layer development in the near-trailing-edge region. In this paper, supersonic shear layers with various splitter plate thicknesses are calculated and analyzed. Two types of instability modes are found when various splitter plate thicknesses are used. The shock wave strength and absolute unstable region determine the instability modes. Shock waves can limit instability development but an absolute unstable region promotes instability development. Velocity thickness growth rates are found to be the result of the merging process between two or more adjacent vortices in the transition region. The number of vortices that merge depends on the ratio of the vortex shedding and nature frequencies. The velocity thickness growth rates are larger under the local instability mode than under the global instability mode since three vortices merge under the former mode but two vortices merge under the latter. This study may provide a helpful reference for the design of supersonic ejectors.

Проблематика. Найкращі плавці мають обтічну форму, яка забезпечує течію без відриву примежового шару і затримує ламінарно-турбулентний перехід. Сама їх форма може бути причиною малого опору та високої швидкості руху. Найшвидші риби, наприклад вітрильник, риба-меч, чорний марлін тощо, мають ще одну особливість своєї форми – дуже гострий ніс – рострум, призначення якого залишається незрозумілим. Поширена думка, що ці хижаки використовують рострум для проколювання здобичі, часто піддається сумніву. Мета. У цьому дослідженні ми аналізуємо гідродинамічні аспекти наявності роструму та можливе використання подібних корпусів для суперкавітуючих підводних апаратів і швидкого проникнення у воду. Ми ілюструємо, що форми з дуже гострим носом можуть бути корисними для гіперзвукових рухів, щоб виключити перегрів фюзеляжу транспортного засобу. Методика реалізації. Для моделювання розподілу тиску на тілах обертання з гострим носом використовуються відомі точні розв’язки рівнянь Ейлера для нестисливої рідини. Для моделювання суперкавітації та осесиметричної течії повітря за великих чисел Маха використовується теорія тонкого тіла. Результати. Розраховано тіла обертання з рострумом, подібні до тулубів найшвидших риб (вітрильника, риби-меч, чорного марліна) та відповідні коефіцієнти тиску і температури. Запропоновані форми забезпечують відсутність точoк гальмування потоку та високого тиску і температури на носику на до- та надзвукових швидкостях як у воді, так і в повітрі. Було оцінено опір таких тіл обертання для безвідривних, суперкавітаційних і надзвукових схем обтікання. Висновки. Запропоновано метод розрахунку осесиметричних тіл без точок гальмування потоку на їхній поверхні. Ця особливість форми дає змогу зменшити максимальний тиск і температуру на носику без значного збільшення опору. Такі форми з гострим увігнутим носиком можна рекомендувати для високошвидкісних безвідривних і суперкавітаційних тіл обертання та для гіперзвукового руху. Background. The best swimmers have a streamlined shape that provides a flow pattern without boundary layer separation and delays the laminar-to-turbulent transition. Their shape itself could be the reason of small drag and high locomotion velocity. The fastest fish, e.g., sailfish, swordfish, black marlin, etc. have another feature of their shape – a very sharp nose – rostrum, the purpose of which remains unclear. Popular belief that the rostrum is used by these predators to pierce their prey is often disputed. Objective. In this study, we analyze the hydrodynamic aspects of the rostrum presence and the possible use of similar hulls for supercavitating underwater vehicles and fast penetration into water. We illustrate that shapes with the very sharp nose could be useful for hypersonic motion in order to eliminate overheating of the vehicle fuselage. Methods. We use the known exact solutions of the Euler equations for the incompressible fluid to simulate the pressure distribution on the bodies of revolution with a sharp nose. The slender body theory is used to simulate the supercavitation and the axisymmetric air flows at high Mach numbers. Results. Bodies of revolution with a rostrum similar to trunks of the fastest fish (sailfish, swordfish, black marlin) and corresponding pressure ant temperature coefficients were calculated. The proposed shapes ensure no stagnation points and no high pressures and temperatures on their noses at sub- and supersonic speeds both in water and air. The drag on such bodies of revolution was estimated for attached, supercavitating and supersonic flow patterns. Conclusions. A method of calculation of axisymmetric bodies without stagnation points on their surface was proposed. This peculiarity of the shape allows diminishing the maximum pressure and temperature on the nose without a significant increase in drag. Such shapes with the sharp concave nose could be recommended for high-speed attached and supercavitating bodies of revolution and for the hypersonic motion. Проблематика. Лучшие пловцы имеют обтекаемую форму, которая обеспечивает течение без отрыва пограничного слоя и задерживает ламинарно-турбулентный переход. Сама их форма может быть причиной малого сопротивления и высокой скорости передвижения. Самые быстрые рыбы, например парусник, рыба-меч, черный марлин и т.д., имеют еще одну особенность формы – очень острый нос – рострум, назначение которого остается неясным. Распространенное мнение о том, что эти хищники используют рострум для прокалывания своей добычи, часто оспаривается. Цель. В этом исследовании мы анализируем гидродинамические аспекты наличия рострума и возможное использование подобных корпусов для суперкавитирующих подводных аппаратов и быстрого проникновения в воду. Мы иллюстрируем, что формы с очень острым носом могут быть полезны для гиперзвукового движения, чтобы исключить перегрев фюзеляжа транспортного средства. Методика реализации. Для моделирования распределения давления на телах вращения с острым носом ми используем известные точные решения уравнений Эйлера для несжимаемой жидкости. Для моделирования суперкавитации и осесимметричных течений воздуха при высоких числах Маха используется теория тонкого тела. Результаты. Рассчитаны тела вращения с рострумом, подобные туловищам самых быстрых рыб (парусника, рыбы-меч, черного марлина), и соответствующие коэффициенты давления и температуры. Предложенные формы гарантируют отсутствие точки торможения потока и высокого давления и температуры на носике на до- и сверхзвуковых скоростях как в воде, так и в воздухе. Проведены оценки сопротивления таких тел вращения для безотрывных, суперкавитационных и сверхзвуковых схем обтекания. Выводы. Предложена методика расчета осесимметричных тел без точек торможения на их поверхности. Эта особенность формы позволяет снизить максимальное давление и температуру на носике без значительного увеличения лобового сопротивления. Такие формы с острым вогнутым носиком могут быть рекомендованы для высокоскоростных безотрывных и суперкавитационных тел вращения, а также для гиперзвукового движения.

Thermodynamic models constitute one of the essential tools to properly design supersonic ejectors. However, by their simplistic nature, most of said models remain unable to properly integrate the adequate physics that takes place within the device. Most notably, the Fabri-choking theory constitutes the building block of the large majority of those models. However, it has recently been shown that the so-called compound-choking theory may be better suited to predict the behavior of a double choked ejector. In the present study, a new state-of-the-art thermodynamic model based on the compound-choking theory is presented. First, the algorithm of the on- and off-design model is laid out. Then, the link between Fabri- and compound-choking is clarified by comparing the model with its Fabri-choking counterpart. Characteristic curves are calibrated onto air and R134a experimental data. Finally, an analytical study is performed to show that imposing the compound-choking is actually equivalent to maximizing the mass flow rate within the ejector. © 2021 Elsevier Ltd

The article analyzes the peculiarities of gas turbine operation, as the main element of the gas turbine in conditions of high operating temperatures and pressure, examines the statistics of characteristic failures and malfunctions of gas turbines, which reduce the reliability of engines and proposed alternative ways to overcome them. Ways of raising the temperature of gases in front of the turbine of modern gas turbine and some systems of cooling of nozzle apparatus and working blades of turbines are considered. The directions of increasing the parameters of the workflow of gas turbines of high-temperature engines with the purpose of ensuring trouble-free operation of aviation equipment during operation are analyzed. Increased responsibility of the engine to ensure the safety, complexity of design and increase its cost, strict requirements for the combat readiness and survivability, expanding the combat use of aircraft is connected with the problem of creating the engine for the aircraft of the 5th generation, which would ensure the ability to operate at supersonic speeds with a maximum operating mode. The main trend in the development of engine designis a continuous increase in temperature in front of the turbine, in order to increase their thrust. Based on the results of analysis and statistical data of the gas turbine operation, ways to solve this problem are proposed. The main ways to increase the temperature of gases before the turbine include: production of blanks based on rapid prototyping, single-crystal manufacturing of blades, protection of surfaces from gas corrosion, oxidation and interaction of high operating temperatures through the use of modern methods of applying protective coatings, appropriate selection of the turbine cooling system, the use of turbo-injector engines. The main directions of technology for manufacturing gas turbine blades in the work make it possible to conduct a comparative analysis of aircraft cooling systems of different generations of leading countries in the world for the purpose of rational choice and efficiency of their application. Therefore, the need to ensure trouble-free operation of the turbine provides for a consistent finding of a reasonable compromise at all stages of the life cycle.

This paper presents an experimental study into dynamics of chamber pressure and heat release rate during self-excited spinning and standing azimuthal mode in NTO/MMH (nitrogen tetroxide/monomethylhydrazine) impinging combustion chambers. Nine cases including two combustion chamber configurations were conducted. The operating conditions of all unstable cases were located in the instability region according to Hewitt empirical correlation. The results show that chamber pressure oscillations keep pace with the corresponding OH* chemiluminescence intensity over the whole combustion region in the spinning and standing modes. It is indicated that the Rayleigh index is positive over the whole combustion area in all the unstable cases. A significant supersonic flame front structure of the first-order spinning mode was found in a cylindrical chamber, which means that a detonation wave could exist in the cylindrical chamber without a center body. The pressure and heat release rate oscillations at the pressure node are nonnegligible although their amplitudes are lower than those at the pressure antinode in the first-order standing mode with an annular chamber. Besides, the dominant frequency of pressure and heat release rate oscillations at the pressure node is twice as high as that at the pressure antinode.

High-order accurate schemes are employed to numerically simulate the interaction of a supersonic jet and a co-directional supersonic inflow. A double backward-facing step model is proposed to investigate the interaction between the jet shear layer and the supersonic inflow shear layer. It is found that due to the interaction of the shear layer, a secondary jet is injected into the recirculation zone at the intersection of the two shear layers. The secondary jet produced by the interaction of the two shear layers has a periodicity because of shear layers interaction. The distinction in the shape of double backward-facing steps will induce changes in the period of the secondary jet. The analysis and discussion of the periodicity of the secondary jet are mainly focused in this letter.

The distribution of film cooling effectiveness in supersonic mainstream of circle-rectangular Convergent-Divergent transition has been numerically investigated under different pressure ratios. The shock wave exerted superior influence on film cooling. In supersonic main flow, extra compression waves formed in upstream of the film holes, resulted by the obstruction of the multiple cooling jets. This exerted extra pressure to the boundary-layer, induced adverse pressure gradient, and led to weakening of the film flow attachment ability and decreasing of local cooling effectiveness. Bow oblique shock wave occurred in front of holes, the two oblique bow shaped low pressure zones formed on both sides of the hole, and low cooling effectiveness zones appeared accordingly. The inefficient region at the leading edge of the hole destroyed the film developing between holes, decreased the cooling effectiveness accumulation in the rear part. The decrease of hole incline angle caused an increase of cooling effectiveness, which reduced reverse velocity gradient caused by shock wave in the boundary-layer and improved film attachment. The influence of main flow pressure ratio to film cooling was also investigated, and found with increasing of the ratio, the influence will became even significant.

Nowadays, coherent and conventional supersonic jets are widely used in electric arc furnace (EAF) steelmaking processes. Generally, these jets are installed in the EAF oven wall with a tilt angle of 35?45?. However, limited studies have been conducted on the impact characteristics of these inclined supersonic jets. This study developed an optimized theoretical model to calculate the penetration depth of inclined coherent and conventional supersonic jets by combining theoretical modeling and numerical simulations. The computational fluid dynamics results are validated against water model experiments. A variable k is newly defined to reflect the velocity variation, which is related to the jet exit at the jet free distance. The results of the optimized theoretical model show that the lance height and lance angle influence the penetration depth of the inclined supersonic jet. At the same lance angle, the penetration depth decreases with the increase in the lance height. Similarly, it decreases with the decrease in lance angle at the same lance height. In addition, the penetration depth of an inclined coherent supersonic jet is larger than that of an inclined conventional supersonic jet under the same conditions. An optimized theoretical model can accurately predict the penetration depths of the inclined coherent and conventional supersonic jets.

This study experimentally investigated the effects of repetitive laser energy deposition using a supersonic intake model with a central conical compression surface and a 19×19 mm^2 cross-sectional duct in a Mach 1.92 indraft wind tunnel: especially in the subcritical and buzz modes. A single-pulse energy deposition was observed to suppress the flow separation at the compression surface by “sweeping” a shock wave system with the thermal bubble generated by the energy deposition. The duration of the sweeping effect was approximately 160 μs in the subcritical mode and 100–420 μs in the buzz mode. Furthermore, repetitive deposition of laser pulse energies was observed to moderate instabilities in both modes, and it increased the pressure recovery by as much as 8%; also, the occurrence of buzz was delayed, thereby widening the stable, subcritical regime. In both modes, there was a threshold value for the laser pulse repetition frequency fL, thd, which corresponded to the duration of the sweeping effect (e.g., fthd=6 kHz in subcritical mode). Below this frequency, the increase in the pressure was proportional to the repetition frequency; whereas above fthd, the effect per single pulse was reduced., Published Online:26 Oct 2019

The turbulence–chemistry interaction model with finite-rate chemistry is a common model for solving supersonic turbulent combustion and takes into consideration the interaction between turbulent mixing and chemical reactions as well as finite chemistry reaction instead of the fast chemistry assumption. However, not only the density and viscosity but also the chemical reactions are affected by the compressibility of the flow field with an increasing Mach number. Considering engineering applications, the compressibility correction was introduced to two recent turbulence–chemistry interaction models with finite-rate chemistry, the Partially Stirred Reactor (PaSR) model and Unsteady PaSR (UPaSR) model, in a Reynolds-averaged Navier–Stokes framework. Numerical simulations of two typical supersonic combustors showed that the interaction between the turbulence and combustion was intensive within complex supersonic chemical reaction flow and could be described by the fine-scale structure volume fraction. The distributions of temperature, pressure, velocity and components somewhat downstream of fuel injection areas were most obviously improved by the presented models. Moreover, the increase in computational time consumption by the compressibility correction was less than 2%. It was found that the Compressible PaSR (C-PaSR) model and the UPaSR model show better consistency with experimental results than the traditional PaSR model.

The sudden expansion of flow in a supersonic flow regime has gained relevance in the recent pasts for a wide run of applications. A number of kinematic as well as geometric parameters have been significantly found to impact the base pressure created within the suddenly expanded stream. The current research intends to create a predictive model for base pressure that is established in the abruptly extended stream. The artificial neural network (ANN) approach is being utilized for this purpose. The database utilized for training the network was assembled utilizing computational fluid dynamics (CFD). This was done by the design of experiments based L27 Orthogonal array. The three input parameters were Mach number (M), nozzle pressure ratio (NPR) and area ratio (AR) and base pressure was the output parameter. The CFD numerical demonstrate was approved by an experimental test rig that developed results for base pressure and used a nozzle and sudden extended axisymmetric duct to do so. The ANN architecture comprised of three layers with eight neurons in the hidden layer. The algorithm for optimization was Levenberg-Marquardt. The ANN was able to successfully predict the base pressure with a regression coefficient R2 of less than 0.99 and RMSE=0.0032. The importance of input parameters influencing base pressure was estimated by using the ANN weight coefficients. Mach number obtained relative importance of 47.16% claiming to be the most dominating factor.