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

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 thermal environment of four-engine liquid rocket exhaust plume impinging on the flame deflector with different impingement and uplift angles is analysed. The supersonic exhaust gas impinging model was established by using the compressible, multi-component, Reynolds-Averaged Navier-Stokes (RANS) equations with the finite volume method. A comparison between the numerical results and experimental data in the literature is made, which verified the validity and accuracy of the numerical model. Additionally, the flow fields of the four-engine rocket impinging on the flame deflector under different impingement and uplift angles are simulated. The results show that high temperatures on the deflector surface mainly occur on the impingement point or the cambered surface. A large impingement angle causes the reverse flow intensity to increase whilst a small angle causes the exhaust gas to deflect a little, a suitable uplift angle can smoothly guide the exhaust gas away from the deflector that the best thermal environment of the deflector channel appears at an impingement angle of 25°and an uplift angle of 5°. This study demonstrates that our model can effectively simulate the impinging flow field, and can be of use for the design of the flame deflector under the multi-engine rocket exhaust gas impact.

In the duplex steelmaking process, the oxygen flow rate is suppressed to reduce the increasing rate of the temperature in the molten bath, resulting in severe dynamic conditions. To improve the mixing effect of the molten bath, a Laval nozzle structure designed for combination gas has been proposed. In this research, five types of Laval nozzle structure have been built based on the combination gas content, and both numerical simulations and experiments are performed to analyze the flow field of the supersonic jet. The axial velocity and oxygen concentration were measured in the experiment, which agreed well with the numerically simulated data. The results show that both initial axial velocity and potential core length increase with the flow rate of combination gas. Further, applying a higher N2 flow rate could improve the oxygen utilization rate at different ambient temperatures, but this issue increases the oxygen utilization rate, however, the latter can be reduced at higher ambient temperatures.

The numerical methods for computing the stability derivatives of the aircraft by solving unsteady sensitivity equations which was proposed in our previous papers was extended to solve three-dimensional problems in this paper. Both the static and dynamic derivatives of the hypersonic blunt cone undergoing pitching oscillation around a fixed point were computed using the new methods. The predicted static derivative and dynamic derivative were found to be in reasonable agreement with the experimental data. For the present method, it is possible to distinguish the components of dynamic derivatives caused by different state parameters. It is found that $$ {C}_{m_{\dot{\alpha}}} $$ and $$ {C}_{m_q} $$ are usually of opposite signs and tend to eliminate each other, which makes $$ {C}_{m_{\dot{\alpha}}}+{C}_{m_q} $$ being much smaller than its individual components. Another feature of this method is that the moment of pressure derivatives proposed in the present paper can be used to predict the contribution of each part of the blunt cone to the overall stability quantitatively. It is found that the head region is crucial for the static stability and the body region contributes most to the dynamic stability.

Tandem blade technology has become an effective method to break the load limit of conventional aerodynamic configurations. To expand the application range of tandem blades, the supersonic tandem blade flow characteristic was studied and an optimal design was conducted by using a computational fluid dynamics (CFD) solver, with an inflow Mach number of 1.2. The main conclusions follow: (1) the tandem blade loss is difficult to control because of the complicated flow structures with supersonic inflow, (2) the forward blade loss dominates the tandem blade overall loss in the whole operating conditions, and (3) the tandem blade profile was optimized by considering the aerodynamic interaction between forward and aft blade. The numerical simulation results show that the total pressure loss declines by 20% at the design point, and the incidence range increases by about 0.5&deg

During the converter steelmaking process, the presence of supersonic oxygen jets can provide oxygen to high-temperature metal baths that promotes chemical reactions in the bath, accelerates the smelting rhythm, and facilitates a uniform distribution of the ingredients in the bath. In this paper, a computational fluid dynamics (CFD) model with combustion reactions is established and compared to the results of combustion experiment. This paper studies the behavior and fluid flow characteristics of supersonic oxygen jets under different environmental compositions under a steelmaking temperature of 1873 K. This validated CFD model can be used to investigate the effect of furnace gas on supersonic oxygen jet characteristics during the converter steelmaking process. The results indicate that the composition of furnace gas has an impact on the characteristics of the oxygen jet. Specifically, as the carbon monoxide (CO) volume fraction increases, the high velocity region of supersonic oxygen jet increases, and the high temperature and the high turbulent kinetic energy regions expand.

During the electric arc furnace steelmaking process, the coherent jet technology was widely used to protect the kinetic energy of the supersonic oxygen jet and achieve better mixing effects. Comparing with the conventional oxygen lance, the coherent lance could increase the surface area of impaction cavity, resulting in a better stirring effect and higher reaction rate. However, there was limited research about the effect of restriction structure for the coherent lance tip on the flow field characteristic of the main oxygen jet. In this research, three kinds of restriction structures have been investigated by numerical simulation and combustion experiment at room and high ambient temperature conditions. Then an optimum restriction structure would be tested in a 75 t electrical arc furnace steelmaking process to verify its metallurgical property.

An optimization method is introduced to design the aerodynamic parameters of a dual-spin two-dimensional guided projectile with the canards for trajectory correction. The nose guidance component contains two pairs of canards which can provide lift and despin with the projectile for stability. The optimal design algorithm is developed to decide the profiles both of the steering and spinning canards, and their deflection angles are also simulated to meet the needs of trajectory correction capabilities. Finally, the aerodynamic efficiency of the specific canards is discussed according to the CFD simulations. Results that obtained here can be further applied to the exterior ballistics design.

The supersonic mixing field induced by a novel wall-mounted cavity having a three-dimensional shape is investigated computationally. In the computation, the Reynolds-averaged Navier-Stokes (RANS) equations are solved to obtain the steady state solution. The resulting pattern of limiting streamlines is compared with the previous result of oil-flow surface visualization. The comparison shows that the pattern of limiting streamlines agrees well with the oil flow pattern not only inside the cavity but also around the injector. The computational jet-penetration heights are also compared with the experimental heights measured previously. The comparison shows that both heights agree well near the injector. Such agreements imply that the flows in the cavity and around the injector can be reproduced well by the present numerical simulation. The detailed flow structure is investigated using the computational results. It is found from the results that a vortex having a three-dimensional shape is produced in the cavity and that the shear-layer spanning the cavity deflects upward near the central plane of the duct owing to the upward flows induced by the vortex. It is also found that owing to the upward shear-layer deflection the jet discharged from the injector is protected from the primary flow having large momentum. As a result, the jet penetrates highly into the primary flow.

In the present paper, we discuss the possibility of a simplified three-dimensional unsteady simulation of plasma-assisted combustion of gaseous fuel in a supersonic airflow. Simulation was performed by using FlowVision CFD software. Analysis of experimental geometry show that it has essentially 3D nature that conditioned by the discrete fuel injection into the flow as well as by the presence of the localized plasma filaments. Study proposes a variant of modeling geometry simplification based on symmetry of the aerodynamic duct and periodicity of the spatial inhomogeneities. Testing of modified FlowVision $k-\varepsilon$ turbulence model named «KEFV» was performed for supersonic flow conditions. Based on that detailed grid without wall functions was used the field of heat and near fuel injection area and surfaces remote from the key area was modeled with using of wall functions, that allowed us to significantly reduce the number of cells of the computational grid. Two steps significantly simplified a complex problem of the hydrocarbon fuel ignition by means of plasma generation. First, plasma formations were simulated by volumetric heat sources and secondly, fuel combustion is reduced to one brutto reaction. Calibration and parametric optimization of the fuel injection into the supersonic flow for IADT-50 JIHT RAS wind tunnel is made by means of simulation using FlowVision CFD software. Study demonstrates a rather good agreement between the experimental schlieren photo of the flow with fuel injection and synthetical one. Modeling of the flow with fuel injection and plasma generation for the facility T131 TSAGI combustion chamber geometry demonstrates a combustion mode for the set of experimental parameters. Study emphasizes the importance of the computational mesh adaptation and spatial resolution increasing for the volumetric heat sources that model electric discharge area. A reasonable qualitative agreement between experimental pressure distribution and modeling one confirms the possibility of limited application of such simplified modeling for the combustion in high-speed flow.

The numerical simulations of flow over a spinning finned projectile at angles of attack ranging from $ 4^\circ $ to $ 30.3^\circ $ in supersonic conditions were carried out to investigate the flow mechanism of the Magnus effect. The finite volume method, a dual-time stepping method, and a $ \gamma - {\rm R}{{\rm e}_\theta} $ transition model were combined to solve the Reynolds-averaged Navier-Stokes (RANS) equations. The validation of temporal resolution, grid independence, and turbulence models were conducted for the accuracy of the numerical method. The numerical results were in certain agreement with archival experimental data. A comparison of the transient lateral force and time-averaged Magnus force between the body of finned projectile and the nonfinned body, the projectile fin and single fin was given. The key lies in the analysis of the reasons for the production of the Magnus force. The simulation provided a profound insight into the flow structure and revealed the following. The fin leading edge shock contributes to the unsteady interference on body lateral force, while the time-averaged body Magnus force is similar to that of the nonfinned body. At $ \alpha = 30.3^\circ $ , the shielding effect of body on crossflow weakens the time-averaged body Magnus force induced by asymmetrical flow separation, the magnitude of which is reduced to the value at $ \alpha = 8^\circ $ . The leeward separation vortices and the resistance on wingroot flow are responsible for the nonlinear interference of the projectile body on fin Magnus force at different angles of attack. When the low pressure region of the vortex core is equivalent to the size and position of fin, leeward separation vortices contribute more the time-averaged Magnus force and induce high frequency variation to the transient fin lateral force.

International audience; Fast technologies for numerical simulation of high-speed flows in ducts, developed in TsAGI, are described. The examples are presented of the application of experimental data, obtained at T-131 wind tunnel, for validation of the developed numerical technologies: 1) validation of 2.5D and 3D calculations of flow in the elliptic combustor with hydrogen supersonic combustion that was studied within HEXAFLY-INT international project; 2) validation of 2D and 2.5D calculations of flow in high-speed model combustor duct with step-like expansion. Preparation of new series of experiments, oriented on validation of turbulent combustion models, is described.

Рассматриваются процессы, сопровождающие вдув сверхзвуковой струи газа в расширяющуюся часть сопла, применительно к созданию управляющих усилий в ракетных двигателях. Разрабатывается подход к многопараметрической оптимизации геометрической формы сопла и параметров вдува струи в сверхзвуковой поток, основанный на применении численной модели турбулентного течения вязкого сжимаемого газа. В качестве параметров оптимизации используются степень нерасчетности вдуваемой струи, угол наклона сопла вдува к оси основного сопла, удаление сопла вдува от критического сечения основного сопла и форма выходного сечения сопла вдува. Проводится сравнение результатов расчетов, полученных для различных конфигураций системы подвода вдуваемого газа. Делаются выводы о влиянии входных параметров задачи на коэффициент изменения тяги сопла., The processes accompanying the supersonic jet injection into the divergent part of a nozzle are considered for the creation of control forces in rocket engines. An approach to the multiparameter optimization of the nozzle's geometric shape and the parameters of jet injection into the supersonic flow based on the application of a numerical model of the turbulent flow of a viscous compressible gas is proposed. The optimization parameters include the pressure ratio in the injection flow and nozzle flow, the angle of inclination of the nozzle to the axis of the main nozzle, the distance from the injection nozzle to the critical section of the main nozzle, and the shape of the injection nozzle outlet section. The numerical results obtained for various configurations of the trust control system are compared. A number of conclusions concerning the effect of input parameters on the trust coefficient are made., ВЫЧИСЛИТЕЛЬНЫЕ МЕТОДЫ И ПРОГРАММИРОВАНИЕ: НОВЫЕ ВЫЧИСЛИТЕЛЬНЫЕ ТЕХНОЛОГИИ, Выпуск 2 2018

International audience; A three-dimensional supersonic turbulent flow with symmetric normal injection of circularjets from the channel walls is numerically simulated. The initial Favre-averaged Navier–Stokesequations closed by the k–ω turbulence model are solved by an algorithm based on an ENO scheme.The mechanism of the formation of vortical structures due to the interaction of the jet with the freestream is studied for jet to crossflow total pressure ratios ranging from 3 to 50. It is known fromexperiments reported in the literature that, for n 10, mixing of the jet with the high-velocity flowleads to the formation of a pair of vortices and of an additional separation zone near the wall behindthe jet. It is demonstrated that the present numerical results are consistent with such findingsand that the pressure distribution on the wall ahead of the jet in the plane of symmetry is also inreasonable agreement with available experimental data.

The supersonic isentropic inlet is a type of inlet that is easy to be adopted to hypersonic and supersonic flying vehicle configuration and suitable to work over a large range of flying states. The inlet has a sound performance in terms of total pressure recovery and flow coefficient owing to its compact design to decrease the negative influences of shock waves. However, it still remains a tough task to design the inlet for the time being. It is here introduced in the paper the concept of expansion waves to simplify the design process of isentropic inlet. It is known to all that the air captured by the inlet will gradually decrease its speed due to influence of the countless numbers of compression waves. However, as we observe in the opposite direction, the air will accelerate its speed gradually in expansion waves flow field. One thing will remain the same in the preceding two flow situations, namely, streamline. Based on the idea, the profile traced in expansion waves flow field can yield the compression surface of isentropic inlet. Expansion waves are phenomena that are very common in supersonic flows. Certain types of expansion waves in a steady uniform parallel two-dimensional flow, which are called Prandtl-Meyer waves, can be analyzed by simple equations. With the joint solution of the MOC equations and Prandtl-Meyer equations, the supersonic expansion waves flow field, serving as the base flow field of designed inlet, can be obtained precisely. Next, the introduction of bilinear interpolation method allows the calculation of each point on a streamline starting from a given initial point. Finally, a streamline-traced sketch, derived from the above base flow field, will be the desired inlet forebody. The following two-dimensional numerical simulation conducted accordingly indicates the actual performance well satisfy the designed state. The fact that the designed inlet has a good performance in terms of total pressure recovery means that the design method is practical and feasible, and broadens applications of design method of such inlets.

International audience; Plasma spray physical vapor deposition aims to substantially evaporate powders in order to produce coatings with various microstructures. This is achieved by powder vapor condensation onto the substrate and/or by deposition of fine melted powder particles and nanoclusters. The deposition process typically operates at pressures ranging between 10 and 200 Pa. In addition to the experimental works, numerical simulations are performed to better understand the process and optimize the experimental conditions. However, the combination of high temperatures and low pressure with shock waves initiated by supersonic expansion of the hot gas in the low-pressure medium makes doubtful the applicability of the continuum approach for the simulation of such a process. This work investigates (1) effects of the pressure dependence of thermodynamic and transport properties on computational fluid dynamics (CFD) predictions and (2) the validity of the continuum approach for thermal plasma flow simulation under very low-pressure conditions. The study compares the flow fields predicted with a continuum approach using CFD software with those obtained by a kinetic-based approach using a direct simulation Monte Carlo method (DSMC). It also shows how the presence of high gradients can contribute to prediction errors for typical PS-PVD conditions.

To achieve a detailed understanding of underexpanded supersonic jet structures influenced by afterburning and other flow conditions, the underexpanded turbulent supersonic jet with and without combustions are investigated by computational fluid dynamics (CFD) method. A program based on a total variation diminishing (TVD) methodology capable of predicting complex shocks is created to solve the axisymmetric expanded Navier–Stokes equations containing transport equations of species. The finite-rate ratio model is employed to handle species sources in chemical reactions. CFD solutions indicate that the structure of underexpanded jet is typically influenced by the pressure ratio and afterburning. The shock reflection distance and maximum value of Mach number in the first shock cell increase with pressure ratio. Chemical reactions for the rocket exhaust mostly exist in the mixing layer of supersonic jet flows. This tends to reduce the intensity of shocks existing in the jet, responding to the variation of thermal parameters.

Steam&ndash, water injectors have been widely applied in various industrial fields because of their compact and passive features. Despite its straightforward mechanical design, the internal two-phase condensing flow phenomena are extremely complicated. In present study, a numerical model has been developed to simulate steam&ndash, water interfacial characteristics in the injectors based on Eulerian&ndash, Eulerian multiphase model in ANSYS CFX software. A particle model is available for the interphase transfer between steam and water, in which a thermal phase change model was inserted into the model as a CFX Expression Language (CEL) to calculate interphase heat and mass transfer. The developed model is validated against a test case under a typical operating condition. The numerical results are consistent with experimental data both in terms of axial pressure and temperature profiles, which preliminarily demonstrates the feasibility and accuracy of particle model on simulation of gas&ndash, liquid interfacial characteristics in the mixing chamber of injector. Based on the dynamic equilibrium of steam supply and its condensation, interfacial characteristics including the variation of steam plume penetration length and steam&ndash, water interface have been discussed under different operating conditions. The numerical results show that steam plume expands with steam inlet mass flow rate and water inlet temperature increasing, while it contracts with the increase of water inlet mass flow rate and backpressure. Besides this, the condensation shock position moves upstream with the backpressure increasing.

This paper deals with the vitiation effects of test air on the scramjet performance in the ground combustion heated facilities. The primary goal is to evaluate the effects of H 2 O and CO 2 , the two major vitiated species generated by combustion heater, on hydrogen-fueled supersonic combustor performance with experimental and numerical approaches. The comparative experiments in the clean air and vitiated air are conducted by using the resistance heated direct-connected facility, with the typical Mach 4 flight conditions simulated. The H 2 O and CO 2 species with accurately controlled contents are added to the high enthalpy clean air from resistance heater, to synthesize the vitiated air of a combustion-type heater. Typically, the contents of H 2 O species can be varied within the range of 3.5%–30% by mole, and 3.0%–10% for CO 2 species. The total temperature, total pressure, Mach number and O 2 mole fraction at the combustor entrance are well-matched between the clean air and vitiated air. The combustion experiments are completed at the fuel equivalence ratios of 0.53 and 0.42 respectively. Furthermore, three-dimensional (3D) reacting flow simulations of combustor flowpath are performed to provide insight into flow field structures and combustion chemistry details that cannot resolved by experimental instruments available. Finally, the experimental data, combined with computational results, are employed to analyze the effects of H 2 O and CO 2 vitiated air on supersonic combustion characteristics and performance. It is concluded that H 2 O and CO 2 contaminants can significantly inhibit the combustion induced pressure rise measured from combustor wall, and the pressure profile decreases with the increasing H 2 O and CO 2 contents in nonlinear trend; simulation results agree well with experimental data and the overall vitiation effects are captured; direct extrapolation of the results from vitiated air to predict the performance of actual flight conditions could result in over-fueling the combustor, possible inlet un-start and inappropriate combustion mode transition. The detailed analysis and discussion are presented and the research conclusions are summarized.

資料番号: PA1110036000

The impingements of supersonic oxygen jets on molten steel and slag in an 85 t top-blown convertor bath were studied and the flow field of molten bath was analyzed. The results show that with the lance height rising from 1.2 m to 1.8 m, the maximum velocity of liquid steel that 0.2 m under the molten steel surface decreases from 0.48 m/s to 0.27 m/s. However, the maximum velocity of liquid steel that 0.8 m under the molten steel surface increases from 0.06 m/s to 0.09 m/s. Therefore, High lance position has a positive effect on the homogeneous distribution of liquid steel velocity. The influence of oxygen flow rate on molten steel's flow field was analyzed in this paper. As shown in the result, when the intensity of oxygen supplied increasing from 235 m3·(h·t)-1 to 333 m3·(h·t)-1, the circulating direction of molten steel changes inversely.

The paper presents results of numerical simulation of laminar and turbulent weakly ionized plasma flows around model bodies under MHD interaction. Numerical investigations have been carried out of supersonic MHD flow of weakly ionized nitrogen plasma around model bodies (spherically blunted cylinder, truncated cylinder, and cone-cylinder lay-outs) at conditions of experiments on the Big Shock Tube (BST) of the Ioffe Institute of Russian Academy of Sciences. The purpose of the investigations was verification of a mathematical model and algorithm as well as analysis of predominant factors determining the MHD impact on the flow structure and thermal load on the model. The effect of the magnetic field induced by the coil installed in the model on the plasma flow was investigated and the efficiency of the MHD interaction was estimated. To enhance the efficiency of the MHD interaction a surface electric discharge was arranged between an electrode installed on the model nose and a ring coaxial electrode installed in the vicinity of the cone-cylinder conjugation. In the magnetic field induced by the coil the discharge rotates producing a domain of high plasma electric conductivity near the cone surface. Comparison of the predictions with experimental data obtained on the BST was presented. Thus the considered system is typical muliphysics and multiscale one involving interaction of supersonic flow with the electromagnetic field, description of large eddies evolution, turbulence model for subgrid scale, and effects of localized electric discharge. Investigations showed that at conditions under study the predominant factors of MHD impact on weakly ionized plasma flow around the body determining the flow structure are both the ponderomotive force and the Joule heating.

The paper conducts a numerical study of the air current in the supersonic multi-shock air intake (MAI) of the direct-flow propulsion jet engine (DFPJE) with various Mach numbers of the flight by methods of computing hydromechanics to define the throttle and high-speed characteristics. The one-parametrical model of the Spallarta-Almares turbulence, and also polynomial dependence of specific heat on temperature is used. The main attention is paid to creation of a computational model in a package of applied software and to specifying the boundary conditions, to define a factor which influences the final result. The paper offers a way to calculate a flow in the diffuser passage allowing to model stationary real processes by changing a type of boundary condition at the MAI output and values of the static pressure. The numerical calculation results, presented as distribution fields of the Mach number of airflow, clearly demonstrate a changing position of the oblique shock system and direct closing shock versus the flight speed. For the considered options of the flight Mach numbers (2; 2.5; 3) the throttle and high-speed characteristics are drawn up using, as a result of numerical modelling, the restoration coefficient values of a total pressure and a mass-flow coefficient depending on the value of a total pressure at the diffuser output.A framework of the optimum-operating mode for the considered MAI is visually shown. The presented model has been proved to be adequate for the specified boundary conditions and technique to perform calculation. The resultant throttle and high-speed characteristics can be further used for the external ballistic design and also as basic data for determining the working parameters of the DFPJE flow passage, in general.

Transverse jet flow field under supersonic inflow is simulated numerically for studying the characteristic of fuel transverse jet and fuel mixing in scramjet combustion chamber. Comparison is performed between simulated results and the results of references and experiments. Results indicate that the CFD code in this paper is applicable for simulation of transverse jut flow field under supersonic inflow, but in order to providing more effective numerical predictive method, CFD code should be modified through increasing mesh density and adding LES module.

The objectives of the present Action Group AD/AG-48 of the 'Group for Aeronautical Research and Technology in EURope' (GARTEUR) on 'Lateral jet interactions at supersonic speeds' is to deeply analyse the effects of the hot-gas jet by the use of CFD simulations. At the end, the main goal is to define the most appropriate similarity parameters for wind-tunnel tests using cold gases. Experiments from DLR Cologne and ONERA/MBDA-France on generic missile models are used in order to achieve numerical simulations of the interaction between a lateral jet and a supersonic cross flow of a missile. These experiments using either a cold-air jet or a hot-gas jet, the latter produced by the combustion of a solid propellant, are well documented and the data are used for the validation of the numerical simulation using either an ideal gas or a multi-species fluid, respectively. The objective of the present work is only the validation of the numerical simulations by comparison with the DLR and ONERA/MBDA-France experiments. The action group members use the same computational meshes for the comparison of the numerical results got from different CFD codes solving 3D, turbulent, RANS equations.