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Pouring water into burning oil, which can trigger acute reaction, is an extremely complicated process. The details can not be captured by experiments. In this paper, the process was reproduced by numerical simulation. The extremely dangerous phenomenon was deeplyanalyzed and the physical chemistry process after pouring water into burning oil was concisely discussed. Simplified numerical models for the complicated process were established and further numerically tested using open-source FDS fire dynamic modelling software to examine its feasibility. The results indicate that oil gas produced by splashing oil droplets disperses quickly in the vast room along with the flow of hot air. The oil gas can be delivered to the position that is far away from where the splashing oil droplets can reach during an extremely short duration, which makes the dramatic expansion of the flame. However, unwanted damage can be caused by the violent increasing of radiation heat flux of irruptive flame. The results of the simulation can give us deep understanding about the process, so the process may be rationally restrained or utilized in the future. [ABSTRACT FROM AUTHOR]

For the situation that the pitching moment of a civil fuselage-mounted aircraft shows to pitch up, the paper studied a new type of improvement measures-to truncate the inner slat. The numerical simulation method was used to study the aerodynamic characteristics of the basic landing configuration and the improvement configuration. The flow characteristics of the two configurations under the high angle of attack are given, and the flow mechanism of the improvement measures is analyzed. The influence of the improvement measures on the intake of the engine is also evaluated. The results show that: the low-speed pitching moment shape depends on the separation position of the wing; separation from the inside wing makes the aircraft "rise"; separation from the outer wing makes the aircraft "head",and separation from the inside wing makes the aircraft "bow"; the pitching moment provided by the horizontal tail determines the total pitching moment; the main reason for the aircraft "bow" is that the separation from the inside wing reduces the downwash of the horizontal tail from the wing; the improvement measures do not affect the engine intake of the fuselage-mounted aircraft when the angle of attack is not greater than the stall angle of attack.the improvement measures do not affect the engine intake of the fuselage-mounted aircraft when the angle of attack is not greater than the stall angle of attack.the improvement measures do not affect the engine intake of the fuselage-mounted aircraft when the angle of attack is not greater than the stall angle of attack. [ABSTRACT FROM AUTHOR]

Sonic boom is a special aero-acoustic phenomenon for supersonic aircraft, which contains both the fields of aerodynamics and nonlinear acoustics, so it is of important academic significance and application prospect. The present paper introduces its basic conception and major characteristics, the research history of its generation, development and evolution, and its harmfulness; then the numerical simulation method, wind tunnel experiment method and flight test method of sonic boom researches are emphatically presented, and also including the recently proposed sonic boom inhibition and low boom optimization design methods ; Finally, the current technique difficulties and challenges of sonic boom prediction are summarized. [ABSTRACT FROM AUTHOR]

In order to get an accurate simulation of the complex three-dimensional unsteady flows in turbomachine, an in-housed unsteady turbomachinery simulation program is developed based on a petaflops high performance computing platform. The full annulus grid simulation method has been developed. The conservative variable interpolation method for sliding interfaces and the high efficient massive parallel computation scheme are discussed in detail. A single stage compressor, a 1.5 stage turbine, and a 3.5 stage high-loaded compressor were simulated to verify and validate the program. The parallel efficiency of the program was also tested. The maximum grid number of the testcases is approximately 200 million, and the number of CPU cores is up to 4096.The results show that aerodynamic performance computed by this program agrees well with experimental data. The full unsteady computation method adopted in this paper guarantees variables' continuity at the sliding interface, weakens nonphysical entropy generation, and yields a reasonable flowfield. The grid partition methods by METIS and the parallel program scheme of MPI contribute to the program with a good load balance and a high parallel efficiency. This program has good expansibility and adaptability to the complex problem of large scale and satisfies the need of complex application flow analysis and exquisite turbomachiney flow filed study.The grid partition methods by METIS and the parallel program scheme of MPI contribute to the program with a good load balance and a high parallel efficiency. This program has good expansibility and adaptability to the complex problem of large scale and satisfies the need of complex application flow analysis and exquisite turbomachiney flow filed study.The grid partition methods by METIS and the parallel program scheme of MPI contribute to the program with a good load balance and a high parallel efficiency. This program has good expansibility and adaptability to the complex problem of large scale and satisfies the need of complex application flow analysis and exquisite turbomachiney flow filed study. [ABSTRACT FROM AUTHOR]

In this paper, 2-Dimension freely falling plates with different mass distributions are simulated by numerically solving the equations of motion coupling with the Navier-Stokes equations. With moving mesh, the motions related to different parameters are studied. The trajectories are identified and the unsteady dynamics are given by considering different dimensionless moments of inertia, and the initial falling angle and centroid positions. The results show that, except the three known classical trajectories:fluttering, tumbling, steady descent, we also observe two new motions. These motions are asymmetrical-fluttering and flutter-tumbling trajectories. They are caused by the centroid position deviating from the center of the plate and the dimensionless moment of inertia in the range of 0~0.95. These motions turn to the tumbling when the centroid position is beyond 0.13L. It indicates that the dimensionless moment of inertia and centroid position are important parameters in freely falling plates, but the physical mechanisms are different in falling motions. The present results can provide academic reference in study of low Reynolds number problems. [ABSTRACT FROM AUTHOR]

Numerical results from the first aeronautical computational fluid dynamics credibility workshop (AeCW-1) are summarized. AeCW-1 is organized by the CFD professional committee of the Chinese society of aerodynamics, in conjunction with the 1st Chinese conference of aerodynamics. A total of 26 data sets were submitted from 14 different institutions. The premise of AeCW-1 is to assess the state-of-the-art of modern CFD technology in China, identify areas needing additional research and development, promote the progress of CFD verification and validation, and provide technical support to the development of large airliner. The CHina-transport 1 (CHN-T1) as the subject geometry in wing-body-tail form, which is a single-aisle airliner developed by CARDC and represents a modern transonic commercial transport airplane. The CHN-T1 has been tested in CARDC 2.4 m×2.4 m transonic wind tunnel (FL-26) and 2.4 m×2.0 m European transonic wind tunnel (ETW). The AeCW-1 organizing committee provides a series of structured and unstructured grids, and the official test cases include the grid convergence study, the buffet study and the study of Reynolds number effect. The influence of the wind tunnel model support system and static aeroelastic deformation on the numerical results are the main concerns. This paper presents statistics analyses of the submitted numerical results and their comparisons with the experimental data from FL-26 as well as some data from ETW. These analyse lead to enables us to give some thoughts and suggestions for the future research activity. [ABSTRACT FROM AUTHOR]

The jet tab thrust vector control technology discussed in this paper has outstanding advantages in less thrust loss, faster response, stronger control forces, compared with jet vane technology, which meet the needs of future air combat for advanced air-to-air missile. In order to meet the requirements of control rate modeling and optimization of jet tab thrust vector control system, conical nozzle with spoilers in three different shapes, which are round, rectangle and sector, are numerically simulated. The effects of spoiler geometry parameters on the thrust vector performance are analyzed. The design method of reducing thrust loss of spoiler is presented. The results indicate that thrust vector angle and thrust loss ratio increases with increasing insertion depth of the spoiler. For the rectangular spoiler, thrust loss ratio decreases with the increase of spoiler width; and for the sector spoiler, thrust loss ratio decreases with the increase of the design parameter l. When the spoiler inserted depth and inserted area is constant, rectangle spoiler has the maximum thrust vector angle and thrust loss ratio, while round spoiler has the minimum. [ABSTRACT FROM AUTHOR]

国家重点研发计划" 数值飞行器原型系统开发" （ 2016 YFB0200700) 的支持下, 超大规模气动弹性数值模 拟软件的开发工作已经正式进入实施阶段。本文简要介绍了软件的研制进展情况, 主要包括超大规模结构网格生 成技术、基于超大规模结构网格的动网格技术、超大规模气动弹性数值模拟软件的确认以及下一步研究工作计划 等方面。实现了百亿量级高质量静态结构网格生成及数值模拟, 十亿量级结构网格的高质量自动变形; 开展了超 大规模气动弹性数值模拟软件静态结构网格生成、结构网格肖动变形、气动/ 结构数据传递等主要功能模块的并行 效率与并行一致性测试工作; 结合CRM 机翼/ 机身/ 平尾构型和AGARD445.6 机翼两个典型三维静气动弹性问题 和颤振问题, 通过与相应试验结果的对比, 开展了千万量级网格规模气动弹性数值模拟软件的确认工作. The aero-elastic software for ultra-scale simulations has been carried out for one and half year. This software is supported by "development of prototype system of numerical vehicle" project (2016YFB0200700) of national key R&D program of China. The progress briefly summarized in the present paper mainly includes ultra-scale grid technology, dynamic grid technology based on ultra-scale structured grid, the validation of the ultra-scale aero-elastic software, and the next-step plan. A high quality static structured grid with ten trillions points is constructed, and the corresponding numerical simulation is accomplished. Moreover, a high quality automatic deformation of billions structured grid is fulfilled. The ultra-scale parallel efficiency and parallel consistency of the main functional modules involved in the aero-elastic software is tested. This software includes the static structured grid construction module, automatic structured grid deformation module, and aeronautic/structure data transformation module. By comparing with corresponding experimental data, the aero-elastic software is validated with a ten millions structured grid for typical three-dimensional static aero-elastic and flutter examples, including the CRM wing/body/horizontal tail configuration and the AGRAD 445.6. [ABSTRACT FROM AUTHOR]

In order to study high-temperature non-equilibrium effects for Mars entry and predict precisely aerothermal environments and aerodynamic characteristics of Mars entry capsules, the numerical methods are developed to simulate the thermo-chemical non-equilibrium flow by considering the chemical reaction, vibrational excitation and thermo-chemical non-equilibrium effects, and surface catalytic and different wall temperature conditions are also taken into account. The methods are validated by comparing the blunt cylinder and MESUR capsule computational results with the corresponding reference and experimental data, and the aerothermal environments and aerodynamic characteristics for typical Mars entry capsules are also studied in this paper by the numerical methods. The results show that:1) under the computational conditions, there are severe chemical non-equilibrium and some thermal non-equilibrium in the flow-fields of Mars entry capsules; 2) the thermodynamic temperature model, reacting model, and surface catalysis conditions have great influence on surface heating rates; 3) the influence of gas models on aerodynamic characteristics is significant, while the influences of wall temperature conditions and surface catalysis on aerodynamic coefficients are weak. [ABSTRACT FROM AUTHOR]

This paper investigated the influence of rudder gap size on rudder efficiency of different rudder on flying wing aircraft using numerical simulation method. Result indicates that the rudder efficiency of inboard and outboard elevon is reduced because of rudder gap, and the bigger the gap size the more rudder efficiency decreases. Compared with that in no gap condition, the efficiency of split rudder is higher when rudder gap exists. The inboard and outboard elevon efficiency decrease caused by rudder gap can be attributed to the fact that the airflow of lower surface go through the gap to the upper surface, leading to the reduction of the pressure differences between the upper and lower surface, so as to block the main flow. The mechanism of split rudder efficiency increase at high deflection angle can be explained by the fact that high pressure airflow at the lower surface go through the gap and inject into reflux zone at upper surface. This flow phenomenon results in the reduced reflux range. [ABSTRACT FROM AUTHOR]

Flight load calculation has great importance to flap design of transport aircraft. Having been used for decades, traditional method is not sufficient for further design and type certification. According to the results of flight test (FT) and wind tunnel test (WT), it has been discovered that the design load of flap is not conservative in cruise and spoiler deflection configuration. For both situations, flap load is distinctly larger than traditional result, which will cause redesign works and certification issues. By analyzing test data, and using CFD tools, key reasons to both problems have finally been found. For cruise configuration, pressure coefficient of flap cabin was often set to zero approximately in previous design, however, in real conditions, the airproof measure between flap and main wing trailing edge often fails in spite of elaborately design and manufacture;moreover, the spoilers installed before the flap may drift and deflect lead to a flight load increment. As to the other problem, the spoiler deflection in high-lift configuration can result in an upwash to the flap. As a result, the flap load, especially for small flap deflection configuration, increases significantly, which may be even greater than maximum flap load with non-deflection spoiler. Improved methods, which have been proved in the latest airliner design program, are given in this paper to correct the traditional flight load calculation. The key point of new methods is to simulate the pressure distribution in flap cabin more accurately for the first problem, and estimate flap load increment caused by spoiler deflection based on maximum normal force of different flap deflections for the other problem. The results, based on new methods, comply with physical rules and real flight conditions more relevantly. [ABSTRACT FROM AUTHOR]

Theory study of viscous interaction effect is reviewed, and aerodynamics data of a complex hypersonic wing-body configuration is obtained by numerical simulation, then the viscous interaction model of longitudinal aerodynamic force coefficient in the condition of perfect gas is established. The quantitative uncertainty of the prediction by viscous interaction model, which is in the form of relative normal distance, is also presented. The results demonstrate that the means combining the theory of viscous interaction parameter and numerical simulation is practicable to assess the viscous interaction effect for hypersonic complex configurations. The viscous interaction model which associated viscous interaction parameter and angle of attack is established, with the ability of correlating varied hypersonic aerodynamics data in different attitudes, Mach numbers and angles of attack; The method in this paper is highly efficient and economic to obtain aerodynamics data during engineering design process. The results using Euler solver efficiently and corrected by presented viscous interaction model can be used to establish massive aerodynamics database for aircraft design in the future. [ABSTRACT FROM AUTHOR]

An in-house three-dimensional unstructured-grid flow solver based on a mixing plane method was used to simulate the flow field of the compressors. Pitchwise averaging was used at the rotor/stator interface for information transport, making the unsteady problem to become a "qusi-steady" one. Thereafter the computational gird needs only one flow passage, which greatly reduced the amount of the calculation. Two different transonic compressors, NASA rotor 67 and stage 35, were tested in this paper. The overall performance maps and spanwise profiles of several aerodyamic parameters were obtained from the calculation, and the results were compared to the experiment data. The results presented here including pressure ratio, temperature ratio and adiabatic efficiency are in good agreement with the experiment data, indicating the applicability and reliability of the method. [ABSTRACT FROM AUTHOR]

Flow separation occurs over the compression corners generated by deflected control surfaces on hypersonic vehicles. This phenomenon is coupled with shock/shock interactions, shock/boundary layer interactions and real gas effects. In order to design control surfaces of hypersonic vehicles and insure flight stability, it is important to predict this complex flowfield features on hypersonic vehicles. In this paper, a numerical study was conducted to assess the effects of real gas on the local flow separation of lifting body rudder. The effects of Mach number, wall temperature and flight altitude on separated flowfield features were highlighted. To this end, several numerical results derived for perfect gas, equilibrium gas and chemical nonequilibrium reacting gas were provided and compared. The calculation results show that dissociation reaction occurs within boundary layer and reduces the temperature when real gas effects is considered. Viscosity of gas within boundary layer becomes smaller, the loss of kinetic energy decreases and this strengthens the ability to overcome the adverse pressure gradient, so the flow becomes harder to separate. Real gas effects significantly decreases separation zone, and change the separation/attach shock wave position and strength. As a result of shock/shock and shock/boundary layer interactions, pressure and heat flux distribution near the separation zone are changed by real gas effects at the same time. From the comparison of equilibrium gas and perfect gas result at different altitude, a conclusion can be drawn that the flow separation becomes more sensitive to real gas effects at higher altitude. Wall temperature is an important factor to determine separated flowfield features. As wall temperature increases, separation zone size becomes larger as well as heat flux on the rudder decreases. As much number increases, the separation zone size decreases and the differences between real gas and perfect gas become larger, real gas effects is more significant. [ABSTRACT FROM AUTHOR]