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Aircraft icing is formed dynamically and appears with porous structure in microcosmic view. This structure directly influences the physical properties of ice, and varies with different icing conditions. In order to know this influence law better, the paper focuses on the difficult about question how micro features changed quantitatively at different temperatures. At the beginning, an information extraction approach is presented based on the image segmentation method. Secondly, some assumptions arc given following the quantitative results of ice in microcosmic view, and then the 3D mathematical model is built based on them. In order to measure distribution characteristics of air pores in the model quantitatively, a computational density is proposed. Finally, take micro-images of ice formed in icing wind tunnel as objects, the feasibility and rationality of the proposed information extraction method and mathematical model are demonstrated. Pores distribution and computational density of ice at different temperatures are also discussed, and the results show highly consistent with the macroscopic viewpoints. Through this paper, it provides quantitative basis for deep understanding of physical phenomenon and properties, and a new way to ice study in microscopic view can be given. [ABSTRACT FROM AUTHOR]

Based on the existing sequential function method for one-dimensional and two-dimensional surface heat flux identification, this paper presents a combination of neural network and sequential function method considering the real-time difficulty of three-dimensional identification. In this paper, a neural network is designed to correct the three-dimensional effects based on the results of one-dimensional identification, which can get accurate real-time results of peak heat flux. Furthermore, PSO algorithm is introduced to optimize the initial weights and thresholds of the neural network in order to obtain a better model. It can be seen from the test results that the method presented for the peak heat flux is accurate with errors less than 4%.The method avoids time complexity of three-dimensional identification and has good noise immunity and stability at the same time. [ABSTRACT FROM AUTHOR]

Smart morphing aircraft has already been a hot point of the aeronautics and aerospace research, which is one of the most possible domains that should bring the revolutionary influences to the development of the aeronautics and aerospace technologies. At first the necessaries about the smart morphing aircraft are pointed in this paper:1) As the scopes of the altitude and the velocity of future aircrafts increase widely, the conventional rigid configuration may not satisfy the requirements of the aerodynamic and flight characteristics under different flight conditions; 2) For the multiple flight missions and tasks achieved by a single plane, it may be necessary for aircraft to alter the aerodynamic configuration in performing different missions; 3) In order to improve the overall aerodynamic performance of the existing aircraft, it is required to maintain excellent aerodynamic and flight performance through adjusting aerodynamic shape in all flight stages. This paper introduces the meanings of "morphing" and "smart" in the modern sense of smart morphing aircraft, in which "morphing" refers to the continuous deformation of different spatial scales (local, distributed, global) and time scales, covering a wide range. According to the deformation scale and function, it can be divided into three categories:local deformation (small deformation), distributed deformation (medium-scale deformation), and integral deformation (large-scale deformation). It can be divided into two types according to the way of realization:mechanical deformation and deformation based on smart material structure. It is also pointed out that the so-called "smart" in this field is basically confined to a single field such as smart materials or structures and intelligent control, which is far from ideal smart morphing. So this paper focuses on the basic scientific issues and key technologies involved in morphing technology. Secondly, from the first Wright Brothers powered aircraft controlled by flexible deformable wings in 1903 to the variable swept wing technology represented by F14 in the 1960s and 1970s, to the successful demonstration of continuous variable rear-edge curvature flight on Gulf Stream Ⅲ aircraft in recent years, the development process of the morphing aircraft is systematically introduced in this paper. Thirdly, the key problems and research progress of morphing aircraft are systematically introduced from the point of view of the key technology and the bottleneck of deformable aircraft design. From the design point of view, the main problems include requirement analysis and conceptual research of the smart morphing aircraft, design technology of overall and subsystem of smart morphing aircraft. From the perspective of basic science and bottleneck technology, the main problems lie in following aspects:aerodynamics, flight dynamics and flight control of morphing aircraft, morphing structure, drive and morphing control. Fourthly, the connotation of smart morphing aircraft, morphing technical indicators, morphing materials and structures, and cost-effectiveness analysis is discussed. Finally, the future development of smart morphing vehicle technologies is prospected, and it is pointed out that smart morphing aircraft technology is spiral development. On the one hand, we need to carry out extensive and systematic research on basic theory and key technology, starting from the foundation; on the other hand, we need to sort out the background requirements of one or several types of morphing aircraft from the engineering point of view, in order to tract the orderly and rapid development of this field. [ABSTRACT FROM AUTHOR]

As an experimental technique for fluid mechanics, particle image velocimetry (PIV) can extract global and quantitative velocity field from images. With the development of artificial intelligence, designing PIV method based on deep learning is quite promising and worth studying. First, the authors in this paper introduce the optical flow neural network which is proposed in the computer vision community. Second, a dataset including particle images and the ground-truth fluid motions is generated to train the parameters of the networks. This leads to a deep neural network for particle image velocimetry which can provide dense motion estimation (one vector for one pixel) efficiently. The featuring of particle image extracted by the neural network is also investigated in this paper. It is found that feature matching improves the accuracy of estimation. The proposed network model is firstly evaluated by a synthetic image sequence of turbulent flow. A jet flow experiment is also carried out in this paper to validate the practicability. The experimental results indicate that compared with the traditional cross-correlation and optical flow methods, the proposed deep neural network has advantages in accuracy, spatial resolution as well as efficiency. [ABSTRACT FROM AUTHOR]

Amphibious aircraft studies are inadequate in domestic. Systematic technical guidance is insufficient regarding the high-lift devices of amphibious airplane. In order to grasp the basic principle and design technology for the high-lift devices of amphibious aircraft, which satisfy the aerodynamic and hydrodynamic design requirement, special design methods have been systematically investigated in this paper. Technical characteristics and design features are compared between amphibious aircraft and land-based aircraft based on domestic and international high-lift device types and design features. On the basis of the special operating environment, operating mode, and hydrodynamic characteristics of amphibious aircraft, various factors affecting the designing of high-lift devices of amphibious aircraft are discussed for the first time. These factors include splash, wave-resistance, and so forth. The high-lift device designs for amphibious aircraft should meet the general design principles of land-based aircraft, and satisfy the special technical requirement of aquatic environment and operating mode. The aerodynamic configuration of amphibious aircraft is closely related to hydrodynamic configuration. The design of high-lift devices for amphibious aircraft is significant different from that for land-based aircrafts. The main sensitive parameters in the design progress of high lift system are pointed out, and the correlation is discussed between trailing-edge high-lift devices design and anti-waves capability. Finally, some suggestions are given for the design principles of the high-lift devices of amphibious aircraft. Based on a lot of water/wind tunnel tests and CFD results of hydrodynamic/aerodynamic characteristics one high performance flap scheme of a civil large amphibian aircraft has been developed according to the present design principle. A water tunnel test shows that, there is no splash on the high-lift devices, and the airplane has the ability in resisting given height wave. This design is completed by solving the contradiction between aerodynamic configuration and hydrodynamic configuration. The design principle of high-lift devices for amphibious aircraft summarized in this paper is feasible, and has a high application worth. Moreover, it can be used as a reference material of the amphibious aircraft high-lift devices design. [ABSTRACT FROM AUTHOR]

The objective of this paper is to develop a boundary layer transition detect technique based on IR technique for laminar wing glove flight test. The main factors of flight environment affecting the technique were analyzed, including optical path, wing skin surface processing method, and heating etc., and then the technique solutions for these influence factors were proposed. A flight test was designed to validate the IR transition detection method, and a classic civil aircraft was used as the test platform. The flight altitude was from 5 km to 7 km, and the flight Mach number was from 0.5 to 0.65. To analyze the effects of wing heating system and solar radiation on IR map, IR images acquisitions were carried out under conditions that the wing heating system was pulse-on and pulse-off. Furthermore, a shadow on the measurement region was built to utilize the azimuth between the airplane and sun. Several flights were carried out to show the effects of the leading edge pollutants on laminar boundary layer. A transition position determination method based on statistical analysis was proposed to avoid misjudgment of the transition position from turbulent wedges induced by leading edge pollutants. Standard circle roughness elements were used to validate IR measurement results. The test results show that the SNR can be increased significantly by wing skin surface heating. Three sets of wing boundary layer transition position under flight conditions are provided, and the Reynolds number is larger than 13 million. Compared with fixed transition, the error of transition position measured from IR image is less than 0.5% chord length. The transition detection method proposed in this paper is appropriate for the laminar wing glove flight test. [ABSTRACT FROM AUTHOR]

In recent twenty years, with the development of very flexible aircraft such as High-Altitude-Long-Endurance aircraft. geometric nonlinear aeroelasticity has become a research hotspot. The geometric nonlinear aeroelastic problems of very flexible aircraft are gradually highlighted, resulting in new challenges for aeroelasticity. This paper presents the aerodynamic modelling methods mainly used in geometric nonlinear aeroelastic analysis and engineering application, targeting flexible aircraft with high aspect ratio wings. The characteristics, research status and the application of strip theory, panel methods and Computational Fluid Dynamics (CFD) methods applied in static and dynamic aeroelastic analysis are introduced. By analyzing and summarizing the research development and status at home and abroad for geometric nonlinear aeroelasticity, this paper also proposes some conclusions and suggestions on aerodynamic modelling of high aspect ratio wing under large deformation as a reference to researchers in both aeroelastic theoretical and engineering fields. [ABSTRACT FROM AUTHOR]

Flight test is an important way to obtain the aerodynamic parameters of reentry capsule under real flight conditions. The results of aerodynamic parameters identification play an important role in the evaluation of CFD and wind tunnel experimental data, and in the improvement of aerodynamic design. For the lunar-earth high-speed reentry, the aerodynamic environments are complex, the aerodynamic predictions are difficult, and the results are very uncertain. Therefore, it is very necessary to give the deviations of aerodynamic identification results. Otherwise the aerodynamic data are useless for aircraft design, so the deviations analysis is very important step. In this paper, the aerodynamic parameters estimation methods for lunar-earth high-speed reentry capsule are studied firstly, and the main factors on influencing estimation deviation are analyzed. On this basis, the deviation ranges of main aerodynamic parameters are computed by Monte Carlo method. The results show that the measurement errors of atmospheric parameter, acceleration, attitude angle, altitude, velocity are main factors on influencing estimation deviatons. The deviation ranges of trim angle of attack are very small, and the deviation ranges of lift-drag ratio and aerodynamic coefficients are larger at very high altitude because of the measurement errors of acceleration and atmospheric density. The methods of uncertainty analysis using modified Cramer-Rao to calculate the deviation of pitch moment derivatives and control force are feasible. The results of deviation in this paper can provide the basis for the design and analysis of reentry capsule. [ABSTRACT FROM AUTHOR]

In this review paper, we give a brief introduction of the past and present of the research in aeroacoustics and flow-induced noise. The main attention is focused on acoustic analogy, from whch the main mechanssms of nojse generated aerodynamrcally for representative set-ups can be explained. We also bHefly jntroduce the related computatoonal schemes and testing methods. Overall, we hope ths paper will help to claiy some misconceptions in aeroacoustjcs, especially those from beginners. Finally, we propose some possible important research directions for the future. [ABSTRACT FROM AUTHOR]

Ludwiig tubi tunnii is 已1 1 1 0 乂已0 1 friquintly in hypirsonic studiis dui to its。03--已€16已似 property and also high flow quality. Howivir, supirsonic wind tunnils basid on Ludwiig tube principle arc scarci tli hypcrsonic Ludwiig tubes arc spriading widely all over the world these days, restricting the experimental studies of supersoncc flows in a cost-effjcient way. Ths paper reviewed the design of a vaiiant of conveniional hypersonic Ludwieg tube, i.e., the so-called tandem nozzle supersonjc wind tunnel, taking the reconstrucioon of Mach 6 hypersonic Ludwieg tube tunnel from Technical University of Braunschweig as example. By inserting an additional Laval nozzle and a settiing chamber, tandem nozzle supersonic wind tunnel js capable to produce supersonic How lower than that of oilginal hypersonic Ludwieg tube. In ths paper, we introduced the aerodynamic desggn piinciples of tandem nozzle supersonic wind tunnel first; thereafter, the desggn for each components of tandem nozzle supersonic wind tunnel, such as the iirst and second Laval nozzles, the settiing chamber, were reviewed and elaborated ； next, the development and possible appiicadons of tandem nozzle supersonic wind tunnel were gjven. Tandem nozzle supersonic wind tunnel can extend the operaHon capablty of oHginal hypersonic Ludwieg tube faclty to supersonic low region w k h a reladvely low piice, while tt inheis the advantages of the hypersonic Ludwieg tube. T H s behavjor makes it suitable for fundamental research of supersonic flow in academic institutions and universities. [ABSTRACT FROM AUTHOR]

Landing gear is one of the most important aeroacoustic noise sources during modern large aircrafts takeoff and landing. This paper some research achievements and latest developments made through wind tunnel experiments, flyover experiments and numerical simulations in the aeroacoustic investigation field of landing gear in the last decades. This paper a brief overview of four aspects including the noise generation mechanisms, the noise reduction techniques, the similarity rule of wind tunnel aeroacoustic experiments and the prediction methods for landing gears. Landing gear broadband noise can be generally categorized into two parts is the flow separation induced noise and the other is the interaction noise between upstream and downstream components. Moreover, the resonant phenomenon appeared in some cavity configurations can also generate tonal noise. Some active and passive noise reduction methods such as fairing and plasma actuators which can suppress the flow separation and interaction phenomena are used to reduce landing gear noise. Finally, prediction of further research on landing gear noise is presented. [ABSTRACT FROM AUTHOR]

Firstly, the flight tests and its application of scaled aircraft models at home and abroad are introduced in this paper. The study topics are summarized as follows： the demonstration and validation of aerodynamic configurations, the flight test of aerodynamic forces, the flight test of dangerous boundaries, as well as the demonstration and validation test with new concepts and technologies. Secondly, this paper analyzes the key technologies of model flight test such as powered flight with autonomous control, quick model structure design and manufacture, model power system design and test, flight controller design and test, high precision measurement, and aerodynamic coefficients identification. Some results of CARDC (China Aerodynamics Research and Development Center) research on these key technologies are shown. Finally, the potential development directions of model flight test are prospected. [ABSTRACT FROM AUTHOR]

This paper briefly analyzes the advantage of Modem Design of Experiments (MDOE) contrast to One Factor At a Time (OFAT) test method, and tries to find the key influence factors of MDOE. A new design method of experiments is presented based on aerodynamic models. The research shows that aerodynamic model is one of the key elements of MDOE. In order to prove that, this paper first gives a common trigonometric series model for a kind of axial symmetry vehicle with electrical cable cover. Then, taking a certain flight vehicle as an example, this paper researches the effects of trigonometric series and response surface aerodynamic models on MDOE, according to saturation D-optimal rule. This research shows that aerodynamic model play an important role in MDOE and the new design method of experiments is effective. Some valuable conclusions have been reached. [ABSTRACT FROM AUTHOR]

During reentry process, aerodynamic shape of a capsule changes as thermal protection material ablates and recesses. It may have negative reaction on reentry performance. Studies on recession influence of Soyuz-shaped return capsule at single ballistic phase have been already carried out in a previous paper. In order to further investigate the influence of the recession, ablation models and heat mechanisms are used to calculate the material ablating of the Soyuz-shaped return capsule. Four recessed configurations were obtained at different reentry times, and the maximal recessed values of the ablation are 2.3 mm, 7.8 mm, 9.8 mm and 12.5 mm, respectively. Based on these data, numerical simulation was employed to study the effect of the recession on the Soyuz-shaped return capsule. The aerodynamic characteristic was calculated and analyzed for the initial configuration and the four different recessed configurations, and then compared with that in the previous study. The results indicate that the recessed value of the ablation increases along the trajectory. The ablation and recession reduce trim angle of attack, while the absolute value increases. They also lead to the increment of the trim lift-drag ratio and the forward movement of the longitudinal pressure center. The effect of the ablation becomes more intensive as the recession value increases. The variation of the trim angle of attack presents linear relation to the recessed value of the ablation. The present results have a qualitative agreement with those of the CEV in references, and also validate the conclusions in the previous paper. [ABSTRACT FROM AUTHOR]

The accuracy of flow field has a significant impact on the details of vortex turbu-lence and aeroacoustic calculation results. LES method is used for high lift device flow field simu-lation, and Ffcows Williams-Hawkings integral surface is used for far-field noise calculation. In this paper, numerical results are validated comparing with test results of flow around cylinder, comparison shows that LES method used in this paper can capture unsteady flow like vortex shedding and detachment, and provide accurate near field flow solution for the following far field noise calculation, acoustic analogy based on FW-H equation can solve far-field noise accurately and efficiently. Based on this, the flow characteristics, far-field characteristics and flow speed in-fluence of high lift device noise generation are numerically studied in this paper. Results show that the slat noise is mainly caused by the unstable turbulence between slat and main element of high lift device and slat trailing edge vortex shedding;while the flap noise is mainly caused by small scale high frequency and large scale low frequency vortices generated around the flap. [ABSTRACT FROM AUTHOR]

Lifting body configuration generally has weaker directional stability, hence an aerodynamic stability augmentation method is presented. In this method, the after-body?s side edges were cut to curved form, making the surface flow from expanding to compressing, the side aerodynamic force was drawn back, and the directional pressure center was drawn back simultaneously so the directional static stability was improved. In this paper, one single cone lifting body configuration was analyzed by this method，and the computational fluid dynamics method was applied in the analysis work. The analysis results showed that less cutting scope could improve the directional static stability remarkably under the moderate attack angle, but the longitude static stability, the lift to drag ratio and the lateral static stability would decrease for different levels in the same time. The main reason was that the compressing effect was decreased for the expanding effect on the side face, and the expanding effect was produced by the upwards flow coming from downside. In this paper, the lateral-directional stability influence by the side edge cutting was presented too. This method improves the directional static stability without any further directional stable rudder, it may become one of the choices for those correlative configuration design and optimization. [ABSTRACT FROM AUTHOR]

The existing MDOE methods based on the parametric model cannot meet the tunnel test requirements of advanced vehicle because of strong nonlinear aerodynamic behaviors. It is necessary to develop the MDOE method based on the non-parametric model instead. The formal design of experiments method of MDOE for wind tunnel tests is researched in this paper. The applicability of Latin Hypercube Sampling and Uniform Design in wind tunnel tests is compared through virtual experiments. And several optimizing methods based on Uniform Design are developed, which can improve the quality of design results to meet the requirement of wind tunnel tests. Studies show that, Uniform Design is more robust and uniform than Latin Hypercube Sampling for the MDOE method based on the non-parametric model, the optimization including boundary complement, density adjustment of sampling points and repetition points design can utilize the priori information to the design, the formal design of experiments method developed in the paper requires less sampling points than the OFAT, which are only 66.7% in the demonstration, and samples the intensive non-linear changing law adequately and exactly. The studies on the formal design of experiments method of MDOE in wind tunnel tests laid the foundation for further study of the MDOE method based on the non-parametric model. [ABSTRACT FROM AUTHOR]

The high-efficiency and high-accuracy of the information translation are the basis and guarantee of the unsteady CFD simulation of flowfield around the rotor. In this paper, a set of predetermined boundary embedded grid method has established by combing a new predetermined boundary embedded strategy with the proposed "Reverse Boundary，，(RB) and "Local Direct-Map" (LDP) technology. RB method which is dynamically adjusted to avoid nonlinear area through pre-estimation hole boundary overcomes the disadvantages of Top-Map method. The LDP method overcomes the contradictory problem of Inverse-map resolution and computation efficiency of the traditional Top-Map embedding method with Inverse-map. The distribution of boundary hole and embedded time under different handling characteristics are numerically tested, it is demonstrated that the predetermined boundary embedded method proposed in the paper is more robust and can be 16. 7 times as efficient as the traditional Top-Map embedding method. In order to further verify the validity of the simulation method for predicting the nonlinear flowfield of rotor with complex motions, a method suitable for the calculation of helicopter rotor flowfield based on compressible Reynolds averaged Navier-Stokes (RANS) is established. The Caradonna- Tung and UH-60A rotor in the hover and forward flight are numerically simulated, calculated results agree well with experimental data，it is concluded that the predetermined boundary embedded grid method can be used for the numerical simulation of high-performance rotor efficiently. [ABSTRACT FROM AUTHOR]

When a high-speed train passes through a tunnel, the pressure waves are produced, and lead to some problems, such as the passenger comfortableness and fatigue strength of train. At home and abroad, a lot of numerical and experimental studies have been conducted on tunnel pressure waves. Based on the understanding of the formation mechanism of the tunnel pressure waves, and the foreign scholars' method for the pressure waves in the plain tunnel based on characteristic wave in recent years, the paper develops the calculation method of wave superposition for tunnel pressure waves, where the tunnel pressure wave is supposed to be superposed by some characteristic waves. The result of this method is in good agreement with foreign field test result. Subsequently, using the method of wave superposition, the paper simulates the cases that have different blockage ratio, train speed and length of tunnel. By the comparison between these results and the calculation ones by the method based on one-dimensional unsteady compressible non-homentropic flow model and the method of characteristics of generalized Riemann Variables, the proposed method can be available to the prediction for pressure waves with the train speed of 150~400 km/h, the blockage ratio of 0.09~0.18 and the tunnel length of 0~20 km. This calculation method is more conducive to the understanding of the mechanism of pressure waves, avoids the complex process of numerical solution, and requires lower computational efforts. However, it cannot be used to calculate the gradient of pressure because the effect of gas-flow is not take into account. [ABSTRACT FROM AUTHOR]

In this paper, a method of aerodynamic heating prediction for hypersonic multi-body separation process is presented. The inviscid outflow solution during multi-body separation and the engineering method for aerodynamic heating are combined, and then the structure heating coupling effect is considered to calculate the temperature distribution of thermal protection systems. For assumed layout of multi-body separation aircraft, this method accomplishes the calculation of temperature distribution characteristics and heat flux distribution characteristics during the separation, and analyses the influence of multi-body separation on aerodynamic heating. The research results show that the computing technology developed in this paper can provide technical support for hypersonic aircraft design both on the thermal protection systems and aerodynamic heating characteristics analysis. [ABSTRACT FROM AUTHOR]

Aerodynamic data usually can be obtained through three ways; wind tunnel test, numerical calculation and flight test, each method have both advantages and disadvantages. To achieve more complete and more accurate aerodynamic data, deep comparison and acquisition of data from different ways, together with data fusion technique could be used. In the first part of this paper, the background, development, present situation and basic idea of data fusion in aerodynamics research field was introduced. Secondly, the method to establish data fusion rules was given, and two kinds of practical data fusion algorithms were presented, i. e., the fusion algorithm based on uncertainty bounds analysis and the algorithm based on aerodynamic mathematical modeling. And some engineering applications of these two algorithms were also introduced. At the last part of this paper, the application prospect of aerodynamic data fusion was introduced. This investigation indicates that the integrity and the accuracy of the aerodynamic data could be improved by using data fusion methods. However the utilization of aerodynamic data fusion needs not only the mathematics of the fusion algorithm but also the researcher's experience, there may exist some subjective factors in the fusion results, choice of fusion method depends on the practical problems. [ABSTRACT FROM AUTHOR]

In order to conduct the drag reduction optimization design of install parameters of nacelle, multi-zone FFD technique is established based on NURBS (Non-Uniform Rational B-Spline) basis function in this paper. The composite frames are founded through choosing the boundary condition of FFD shape and lattice reasonably. Then, it can realize the FFD parameterized design of the multi-control frames for complex shape in different zones. This multi-control frame FFD technique is used to conduct drag reduction design of the install position of certain airliner nacelle. The aerodynamic optimization design system is established by combining FFD technique, random inertia weight PSO arithmetic and Kriging surrogate model. It process the aerodynamic optimization design by taking horizontal install position and horizontal torsion angle of certain airliner nacelle as design variables, and the results show that the optimized nacelle has lower drag and better aerodynamic characteristic in certain attack angle. Therefore, the optimization design system based on multi-zone FFD technique established in this paper is reasonable and practical. [ABSTRACT FROM AUTHOR]

The hypersonic inward turning inlet has characteristics of high mass capture capability and compression capability, meanwhile it has direct influences on the design of the flight vehicle control ability and stability, therefore the design of forebody/inlet integration is of critical importance. In the paper, axi-symmetrical basic flow fields with controlled pressure distribution used for the designing of hypersonic inward turning inlet was established using the rotational method of characteristics. In combination of the streamline tracing technique, five forebodies/inward turning inlets were designed. The influences of different radial and circumferential positions on the inlet and forebody performances in longitudinal and course were studied numerically. The results show that the forebodies/inlets generate large aerodynamic forces and moments which influence both the flight vehicle control ability and stability. With the capture line being off from the center body, the performances of the flight vehicle static instability both longitudinal and lateral are improved; when the capture line changes in the circumferential direction, the degree of longitudinal static stability remains nearly unchanged; when the capture line is symmetrical about the longitudinal plane, mass flow ratio increases significantly compared to those with asymmetrical capture lines, but the penalty is the deterioration of the performance of course static stability. These conclusions presented in this paper have valuable contributions to the design of flight vehicles with high performance of handling and stability. [ABSTRACT FROM AUTHOR]

Air-breathing hypersonic flight vehicles are highly airframe/propulsion integrated, especially the outer and inner flows of the flight vehicles are very complicated and interact strongly with each other, considerable difficulties should be solved rationally in ground testing. Therefore, it is of great necessity to conduct researches on the related experimental methodology. This paper presents a brief analysis of main aerodynamic issues concerned with air-breathing hypersonic flight vehicles as well as associated experimental testing requirements, in which three typical types of experimental testing are focused on, they are performance testing of airframe/propulsion integrated vehicles, experiment of boundary layer transition with trips, and arc-heated facility testing of sharp leading edges. Besides, an overview of wind tunnel research methodologies both in China and abroad is included. This paper also provides critical similarity parameters that need to be ensured for the three types of experiments mentioned above, followed by an influence analysis of those key parameters that are difficult to achieve in ground testing. Based on the testing capability of the current test facilities, three typical wind tunnel experimental methods have been put forward, along with an effective method of acquiring accurate aerodynamic performance of the airframe/propulsion integrated vehicles in particular. The research materials presented should provide a useful guideline on the aerothermal and aerodynamic experimental research on hypersonic flight vehicles. [ABSTRACT FROM AUTHOR]

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]

In this paper, an entropy analysis and design optimization methodology is combined with multi-objective airfoil shape optimization to demonstrate the impact of entropy generation on aerodynamic designs. Model equations for the calculation of the local entropy generation rate in turbulent flows are presented by extending the Reynolds-averaging procedure to the entropy balance equation. We use the CST parametric method to model the airfoil configuration and combine the CFD solver with multi-objective optimization algorithm respectively, to compute aerodynamics quantities such as lift-drag ratio and entropy generation rate. We obtain the optimized airfoils that maximize lift-drag ratios and minimize entropy generation rates as optimization objectives. Moreover, the present method is compared with the traditional method. From the multi-objective solutions, a designer can select a set of airfoil shapes with low relative entropy generation, small actuation cost, and improved aerodynamic performance at the design conditions. The optimized results unfold that, the proposed method is capable of realizing better performance in practical aerodynamic optimization, with the features of strong global convergence and improved aerodynamic performance in the increase of lift-drag ratio and the decrease of entropy generation rate, compared with the initial population. [ABSTRACT FROM AUTHOR]

According to the design requirements related to high endurance factor, short takeoff and landing, deforming resistance of MALE UAV, the investigation of the design methodology on multi-constraints two-element airfoil is conducted based upon the earlier studies on the one-element airfoil. In this paper, the geometry of the two-element airfoil is generated with the method of controllable conic curves accompanied with the control points, and the parameters of control curves, slot and spin axis position are optimized using Genetic Algorithms. The results indicate that* compared with the ellipse cutting method, the present method using more control parameters is more flexible to the earlier airfoil with more reasonable pressure distributions. Moreover, the endurance factor, take-off lift and lift-to-drag ratio of the designed slot airfoil are improved substantially. In addition, flaps with wide angle deflection play the role of deceleration during landing. The designed slot airfoil satisfies multiple demands and the present method is suitable for the MALE UAV design. [ABSTRACT FROM AUTHOR]

Mars probe usually is an obtuse vehicle with a spherical-crown/inverted-cone configuration. For this kind of vehicles, the aft-body flow is complex, such as large scale separation and reattachment, shock/vortex interaction. Consequently, the dynamic characteristics are also extremely complex and become one of the key problems during the design process. In this paper, the approximate specific heat ratio method is used to simulate the atmospheric environment of Mars for ballistic range Mars probe test model ( MSL model). Based on the verification of the numerical method with in-house software FLY3D-Mars, the variation of pitching dynamic characteristics with specific heat ratio are numerically studied. The results show that the experimental configuration is dynamically stable in all the investigated ranges. As the specific heat ratio decreases, the compressibility becomes stronger, the dynamic derivative reduces and the dynamic stability can be improved accordingly. The reason for the change of dynamic stability with specific heat ratio is that the compressibility of the flow is different under the conditions with different specific heat ratios, which is consistent with that of ballistic range test. [ABSTRACT FROM AUTHOR]

As ihc firsi modeling flight test in China aiming at fundamental hypersonic aerodynamics problems, the real flight of MF 1 is absolutely successful, but there are also some differences between designed trajectory and real uncontrolled flight trajectory. In this paper, the cause of trajectory bias between flight test and prcflight design is studied by theoretical analysis and flight simulations firstly. Results show that, wind distribution measurement errors and predicted aerodynamic errors should he the main reasons for that trajectory bias. Then, the flight trajectory of the MF-1 is reconstructed using flight overload and angular rate data, and the whole flight trajectory data including angle of attack and angle of slide are obtained. It is also shown that, the wind measurement has significant effects to the accurate estimation of angle of attack and angle of slide if there is no direct measure of flow angle. [ABSTRACT FROM AUTHOR]

The problem of acoustic radiation restricts the performance of aero-engine. To further reveal the acoustic radiation mechanism of mixing layer under supersonic flow, and enhance the understanding, this paper studies the acoustic characteristics of supersonic mixing layer. The near field calculation is investigated by numerical simulation of parabolized stability equations (PSE), and an integral theory based on stability theory is applied to predict the far field acoustic pressure. The intensity and direction of Mach wave are calculated and in good agreement with previous results. The influence of the flow temperature at the upper and the lower sides is discussed in detail. The results show that, the slow mode dominates when the temperature ratio is less than 1, while the fast mode dominates when the temperature ratio is greater than 1. [ABSTRACT FROM AUTHOR]

Shock wave boundary layer interaction prefers to induce boundary layer separation, and it's necessary to analyze the flow structure, local high pressures and heat flux, and the separation bubble size. In this paper, we propose an analytical model for this typical flow structure, and the polar theory is used for auxiliary analysis. Based on the FIT. classical shock relationship, and fining formula of separation length, the analytical model can be solved. With the analytical model, we can quickly gain the flow structure and parameters which compares well with numerical simulation. Also, we use this model to analyze the height of separation bubble. The height increases with increasing flight altitude and deflection angle for incident shock wave, and it first decreases slightly then increases as incoming Mach number increases. [ABSTRACT FROM AUTHOR]

With the great improvement in computer technology and numerical methods, computational fluid dynamics has progressed significantly. Even though it is easy and fast lo obtain discrctized results via numerical simulations, the validity and accuracy of the results need to be carefully verified and validated. Grid convergence study is one of the important approaches in verification process; however, traditional methods cannot realize consistent refinement on general unstructured grids, causing difficulties in grid convergence study on general unstructured grids. In order to solve this problem, the downscaling accuracy testing method (DS method) was adopted in this paper. The DS method was successfully verified by comparing with traditional mesh refinement method and analytical stencil analysis for three cases: a) gradient reconstruction accuracy on isotropic grids; b) gradient reconstructive accuracy on anisotropic grids; c) flow simulation accuracy of a manufactured solution on isotropic grids. The DS method was implemented for accuracy tests on anisotropic stretched grids commonly used in the simulation of viscous flows, and the results were consistent with predictions' which demonstrate the advantage of the DS method on accuracy verification with general unstructured grids. [ABSTRACT FROM AUTHOR]

Research and development of the technology for near space hypersonic flying vehicles have drawn more and more attention. In the article "New Problems of Aerodynamics". Zhou and Zhang showed that, for near space flying vehicles, the existing aerodynamics is inadequate and one of the new problems is how to take the local rarefied gas effect into consideration. They also analysed hypersonic boundary layer of near space flying vehicles. Based on the research. Chen and Zhao studied the local rarefied gas effect for strong shear flow, and proposed a non-dimensional parameter Zh to characterize the rarefied gas effect. They indicate that the viscosity coefficient in conventional continuum model should be modified according to the value of Zh. In this paper, in a similar way, the rarefied gas effect in heat conduction problem is investigated by Direct Simulation Monte Carlo (DSMC). A characteristic parameter ZhT is proposed based on the analysis of the velocity distribution functions of molecules, and the correction on thermal conductivity depending on the parameter Zh r is obtained. A hypersonic flow of argon over a two dimensional cylinder is simulated by the proposed model, and the heat flux on the surface is examined. [ABSTRACT FROM AUTHOR]

Supercooled large droplet (SLD) is beyond the envelope described in the Appendix C of FAR 25. When the droplets impact on the surface of airfoil, they will run downstream and freeze to shape of ridge. SLD runback ice may endanger the safety of the aircraft significantly, but the aerodynamic effect of icing on vehicle performance is not completely understand. The numerical simulation of SLD icing is carried out using Reynolds Stress Model (RSM) in this paper. The computational maximum lift coefficients and stalling angles are agree with Lee's experimental data on NACA 23012m, the applicability of RSM in analysis of flow separation resulting from runback ice is valialated. The impact degree of runback ice is quite different according to airfoil geometry, and the supercritical airfoils and tail airfoils are specially analyzed in large civil aircraft. The results demonstrate that runback ice influences the flow separation and significant deteriorates aerodynamic performance of supercritical airfoils. For the tail airfoils, the aerodynamic impact of SLD icing is not so severe, and the size of separation region is diminished at higher angles of attack. The contents of this paper have reference values in aircraft design and airworthiness certification for large civil aircraft. [ABSTRACT FROM AUTHOR]

Checked maneuver is one of the most important loads checking cases stated in CCAR-25 for civil aircraft design. Different understanding of corresponding regulations leads to remarkable deviations in load calculations, which greatly affect accurate aircraft loads prediction. From inherent connections between GJB and CCAR-25, an novel operation law for checked maneuver is proposed in this paper. B operation during rapid pitching from GJB is first used, angular accelerations under loads factor of 1. 0 and 2. 5 are then checked and adjusted correspondingly until both get satisfied. Current method in this paper can guarantee B operation during rapid pitching for military aircraft, and in the meantime meet requirements for angular accelerations in CCAR-25, which eliminate a long existed disagreement in checked maneuver engineering calculation and will benefit future commonality design between civil and military aircrafts. Homemade computer programs implementing this method have been developed and applied successfully in trunk airliner project. [ABSTRACT FROM AUTHOR]

This paper applies the methods of CFD simulation of fluid flow, which are validated by wind tunnel results, to analyze the influences of installation location of nacelle chine on the lift performance of landing configuration from the change of spatial vortices. The numerical simulation results show that because of the influences of nacelle and pylon, the end of slats and cuts of main wing left by the deflection of slats, will induce a series of vortices in high angle of attack (α), which could bring separation in the upper surface of main wing easily. The vortex induced by nacelle chine can depress these vortices and low speed region of flow field above the main wing behind nacelle so that the stalling α and maximum lift coefficient will be improved. Regarding the landing configuration in this paper, its stalling α and maximum lift coefficient are increased by 2° and 0.15 respectively by the installation of nacelle chine. The vortex induced by nacelle chine will move downwards once the chine is put back or down, which leads to stronger restraint effects to other vortices and lower speed region. In order to increase the stalling a and maximum lift coefficient of landing configuration, both of the location and strength of vortex induced by nacelle chine must be taken into consideration in the design process. Because of some geometry restraints and adopting above investigation results, the location of nacelle chine is redesigned through the tradeoff study for the location and strength of space vortex. The CFD result shows that for the redesigned location of chine, geometry restraint is satisfied and the loss of maximum lift coefficient of landing configuration is less than 0. 015, which still meets the design requests. [ABSTRACT FROM AUTHOR]

Waverider is considered to be an ideal configuration for hypersonic vehicles in the near-space, but due to the serious aero-heating and facture limitations, its infinite sharp leading edge should be blunted which would result in a remarkable loss of aerodynamic performance and have a great effect on the practical value of waverider configuration. In this paper, one blunt method is proposed based on the typical aero-heating characteristics of waverider. Furthermore, numerical technology is used to study the flow characteristics of blunt waverider, including the mechanism of bluntness impact and the feasibility of this new blunt method. An experiment has been done in FD-14A hypersonic shock wind tunnel to obtain the flow structure and heat flux distribution in the nose region of waverider model. The results show that bluntness induces a detached bow shock around the nose region and the phenomenon of spillage along the leading edge, which therefore results in an increment of drag and a reduction of lift and lift-to-drag ratio. However, the serious aero-heating is limited in the nose region for blunt waverider, smaller blunt radius can be used for most of the leading edge. Owing to the improvement of aerodynamic performance and negligible influence on thermal protection system, the blunt method proposed by the present paper is proved to be feasible for hypersonic waverider design. Under the studied conditions, the lift-to-drag ratio increases about 3.0%. If the blunt method is optimized, waverider configuration is still one of the most important candidates for hypersonic flight in the near-space. [ABSTRACT FROM AUTHOR]

Aiming at aircraft aerodynamic optimization design problems with multi-requirements including multi-designing points, multi-objects and multi-constraints, a new method for optimization design is proposed in this paper; fast optimization design with multi-requirements. The method is established on the partial correlation and linear regression theories which originated from statistics. According to the analysis of partial correlation between design variables and design requirements, some of the requirements are transformed into the constraints of variables through the transformation model established by the linear regression theory, and the optimization result shows that the simpler the optimization model is, the shorter time will be costed in the optimization design. Two multi-requirements optimization design examples have been investigated in this paper, in which the RAE2822 and HSNLF(1)-0213 are as the initial foils respectively, and the optimization result by the method in this paper are compared with that of the Pareto multi-object method, which verifies the validity and reliability of the method in this paper. [ABSTRACT FROM AUTHOR]

Based on the present semi-empirical model of numerically predicting aerodynamic noise from wind turbine, effects of .several important parameters on aerodynamic noise from wind turbine are studied in the paper. To better simulate vortex-shedding noise near the tip of blade, tip loss correction is added in the present semi-model at the domain of blade tip. Using the noise predicting model developed in the paper, four different important parameters including blade airfoil, tip pitch angle, rotor angular speed and tail edge bluntness are mainly discussed in the paper, where the pattern of noise level with the frequency and A-weighted total sound power level and sound pressure level are also analyzed in detail. These influential patterns of the different aerodynamic parameters can provide technical references for the design and manufacture of low noise wind turbine, which helps to the study on the reduction of aerodynamic noise from wind turbine. [ABSTRACT FROM AUTHOR]

The hybrid Cartesian grid/gridless method is developed to simulate viscous flows. The flow field is basically discretized by Cartesian grids, and local gridless treatment is embedded in the region adjacent to the body. In order to simulate the thin boundary layer of the viscous flows interested in this paper, the cloud of anisotropic points is implemented in the local gridless region. A cloud overlap-free procedure is proposed which enables the hybrid method to treat complex flows involving close-coupled bodies. The unstructured Cartesian grid needed for the hybrid method is generated based on the quadtree data structure, which can help to control the quality of the clouds at the interface between the grid region and the gridless region. The resulting hybrid method is applied to solve two-dimensional compressible NS equations. The viscous flows past a NACA0012 airfoil and a RAE2822 airfoil are first simulated, and the numerical results obtained present a good agreement with available experimental data. The paper ends with a numerical example of the viscous flow past a two-element airfoil, which demonstrates the ability of the present method for treating more complex flows over multi-bodies. [ABSTRACT FROM AUTHOR]

Cycloidal propeller is a new device composed of two or more blades, which can provide thrust in all directions. It has become a research hotspot in the field of aeronautics research in foreign countries recent years. In this paper, domestic and foreign progress of research on the aerodynamics performance of cycloidal propeller is summarized, which include theoretical analysis, numerical simulation and experiments. It is found that the precision of the current aerodynamic force computation models for cyclorotor are still not enough and cannot be used widely, thus there is need to develop a theoretical model which can consider the dynamic stall, induced drag and downwash effect; and there is little discussion about the mechanism of unsteady aerodynamics of cyclorotor. The numerical analysis method of cyclorotor is still developing and need to be combined with PIV experiments in future to explain the flow mechanism of the cyclorotor. The aerodynamics efficiency of cycloidal propeller is also been discussed, it is found that the conclusions of cycloidal propeller's aerodynamics efficiency in papers are not the same, but the recent experimental data has shown it can be higher than conventional lift device, such as rotor. Finally, some further research topics on the aerodynamics performance are proposed. [ABSTRACT FROM AUTHOR]

An immersed boundary-lattice Boltzmann method based on non-uniform quadtree grids is proposed in present paper. In order to retain continuous physical quantity, equal interval interpolation in temporal and spatial space is used on the different grid interface. Because quadtree grids have identical spatial size in the same grid level and identical temporal size in the neighbor temporal level, equal interval interpolation retains second-order accuracy in temporal and spatial space, has simple computational process, need less resource, and achieves utilization of arbitrary boundary refined grids in the lattice Boltzmann method. To take advantage of the merit of the Cartesian grid, the velocity correction method is adopted in immersed boundary, which ensures the non-slip boundary compared to the direct forcing method or the momentum exchange method in the conventional immersed boundary-lattice Boltzmann method. At the same time, the large-eddy simulation model is applied to immersed boundary-lattice Boltzmann method. In this paper, the proposed model combines with quardtree grids data structure, boundary treatment with large eddy simulation to simulate the flows with high Reynolds number around the obstacle. To validate these models, the incompressible flows around a circular cylinder with Re = 200, 1000, 3900 are simulated, and the simulation results agree well with those of other methods. [ABSTRACT FROM AUTHOR]

The CLEAN arithmetic based on Spatial Source Coherence is developed and then applied to the wind tunnel test in the paper. The mechanism for this method is introducing a clean beam to instead the main lobe of beam picture iteratively, and getting rid of side lobe for background noise of wind tunnel or microphones distribution in phase array. The data from test in wind tunnel is optimized with this method. The results show that it improves the resolving power of low frequency noise and increases the signal-to-noise of the aero-acoustics test greatly. In the paper, coefficient K for judging the performance of phase array is also introduced, which including the effect to the resolving power both for frequency and distance from array to sound source plane. To some point, it is helpful to improve the efficiency for the coefficient K inducted. [ABSTRACT FROM AUTHOR]

The most traditional methods about computation of the noise are half-experience. This paper is used acoustic analogy which is in combine with computational fluid dynamics. The approach is achieved by two step procedure. In the first step, the unsteady flow field is computed in a near field domain comprising all acoustic sources via a URANS solver. In the second step, the FW-H integration approach with a permeable data surface is used to extend the dynamic near-field to the acoustic far-field. Compared with the traditional half-experience method, this approach only has small amount of calculation, and accurate result can be obtained. The most important, it can compute nonlinear noise and be used easily. This paper calculates the noise of the rod-airfoil by the use of aeroacoustics integral formulation, employing Gauss integral formula and constructing discrete scheme in time. In the time domain and frequency domain, this paper calculates respectively the acoustic pressure changes and sound pressure level changes over time steps in the observation. The result agrees well with the experiment data. Furthermore, the paper researches the effect of airfoil wake flow on precision. [ABSTRACT FROM AUTHOR]

The supercooled large droplet (SLD) has a diameter more than 50µm with a complicated icing mechanism, and it hasn't been prescribed in FAR25, which forms a significant flight safety risk. Since there are some defects on the dynamic models for icing simulation, in this paper a multifactor-effected breakup (MEB) model is introduced, and its effect on SLD icing simulation is validated. The paper indicates that MEB model can improve SLD icing simulation positively. Based on this model, the droplet dynamic effect are also analyzed on droplet-collecting-efficiency and SLD icing characteristic of changing diameter-distribution, impinging-characteristic. [ABSTRACT FROM AUTHOR]

Wind tunnel tests of wind loads acting on the roof of Shenzhen new railway station were conducted in a boundary layer wind tunnel at Hunan University, for the conditions without train and with trains inside the station. In addition, the effects of wind speed amplify in the station were also investigated. This paper presents the results of the mean and RMS wind pressure coefficients, and comparisons of the maximum mean negative wind pressure coefficients and fluctuating wind pressure coefficients for the case of without train and with trains were made. Furthermore, the characteristics of probability density functions of wind pressures measured from typical pressure taps were analyzed, and the peak factors of that were estimated under a probability significance level of 99.999%. On the other hand, the results obtained from numerical simulations were also presented in the paper, and the results were compared with the wind tunnel test data. The outcomes of the combined experimental and numerical study indicate that: (1) The maximum mean negative wind pressure coefficients on the roof occur at the windward leading edge region, where the maximum fluctuating wind pressure coefficients occur also in this region; (2) There are some differences of the maximum mean negative wind pressure coefficients and RMS wind pressure coefficients under conditions with different number of trains inside the station, but such effects on the overall pressure distributions on the whole roof are negligible; (3) There are clearly negative skewed distributions for some pressure taps at the windward leading roof edge and much longer negative tails are observed, which follow Non-Gaussian distributions; (4) Wind flow was found to be congregated inside the holes at the structure, and the maximum wind speed amplify coefficient is up to 1.34; And (5) the results obtained from the numerical simulations are in good agreement with these of the wind tunnel test, illustrating that the Large Eddy Simulation (LES) approach and Discretizing and Synthesizing Random Flow Generation (DSRFG) technique are effective tools for numerical simulations of wind effects on the complex long-span roof structures. [ABSTRACT FROM AUTHOR]

With the complex interaction between wind and structure, 1000kV steel tube frame belongs to wind-sensitive structures and the wind load becomes the main problem of the structure design. Based on the engineering background of a large steel tube frame for 1000kV UHV transmission lines with a 64m high, 54m span. This paper presents a deterministic-stochastic hybrid model to simulate a downburst. Wood's velocity profile for the vertical profile and Holmes' empirical model are utilized to generate the mean wind. The fluctuation is generated by stationary Gaussian stochastic process. A precise finite element model was established and the vibration characteristics of structure have been obtained, then the wind-induced vibration responses of structure were obtained in the time domain. The characteristics of mean displacement, RMS displacement and RMS acceleration at different nodes of structure have been analyzed in detaile. Meanwhile, the wind-induced vibration coefficient at important nodes of frame for different wind environment were also studied and compared, respectively. Through the research, some characteristics of wind-induced dynamic responses of 1000kV steel tube frame were revealed, the procedure and the results presented in this paper can be applied for the design of large steel tube frame structures. [ABSTRACT FROM AUTHOR]

The application of proper orthogonal decomposition (POD), which is an efficient method of data analysis for two dimensional airfoil transonic flow field reconstruction, has been proved successfully in this paper. The reduced order model (ROM) for flows has been constructed by using the snap shots from the knowledge of computational data or experimental data of flows. Based on this ROM, it is possible to reconstruct the entire flow field from the gappy data accurately in the range of snap shots. Using this approach, another application which is the two dimensional inverse airfoil design based on POD/ROM has been demonstrated successfully. This methodology demonstrates a great simplification for the inverse airfoil design problem. In this paper, the cases of transonic flow field reconstruction of RAE 2822 airfoil and the inverse design of NACA 63212 and Korn airfoil have been simulated. These cases prove that the method of flow field reconstruction and inverse airfoil design based on POD is efficient and utilizable in the range of transonic. [ABSTRACT FROM AUTHOR]

In order to design the variable pitch of large-scale wind turbine quickly and accurately, this paper applied the PSO algorithm to the aerodynamic pitch design and established the mathematical model of the design and calculation. With the 1.5MW large-scale wind turbine for pitch design calculation, the results show that the design calculation method proposed in this paper is feasible and practical. [ABSTRACT FROM AUTHOR]

The flow field and test condition in anechoic wind tunnel impacted acoustic propagation, and impacted noise sources identification. According to classical Beamforming algorithm for static source identification, making use of acoustic propagation law in flow field and shear layer of anechoic wind tunnel, this paper develops an aerodynamic noise sources identification method based on microphone arrays in anechoic wind tunnel, then numerical emulation and numerical proof are proceed. Numerical proof results indicate flow field and shear layer of anechoic wind tunnel impacted results of noise sources identification obviously, the influence can be corrected by making use of the computation method that introduced in this paper. [ABSTRACT FROM AUTHOR]