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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]

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]

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]

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]

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]

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]

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]

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]

Nowadays, developing aviation jet engines are the main driving force for transport vehicles. The study on electric propulsion concepts with distributed propellers and distributed duct fans for the transport vehicles has already begun, and the new propulsion is the core competence of future transport aviation aircraft. In this paper, the actuator disk method is used to verify individual propeller firstly, and the calculated thrust and torque agree well with the test results. Then, based on Reynolds averaged N-S equations and combined with the SA turbulence model, the wing slipstream effects on of distributed propellers with different rotating directions under the condition of low speed and high thrust at low altitude are analyzed by replacing the distributed propellers with actuator disk. Finally, the wing slipstream effects of single propeller with positive and negative rotating directions are studied, especially the effects on the lift and drag. The results show that the lift and drag of the aircraft with slipstream effects are both greater than those without slipstream effects for four types of distributed propeller rotating direction. The lift is closely related to the number of suction peaks on the upper and lower surfaces, and the types of distributed propeller rotating direction directly determine the number of wing suction peaks, especially the wing tip rotating direction. When adjacent blades rotate in the opposite directions, the blades in the middle position of the rotation shaft are both in the upward and downward states, leading to strengthened or weakened wing suction of the leading edge in the middle position compared with that in the same rotating direction. Moreover, as the wingtip propeller rotates in the direction opposite to the wingtip vortex rotating, the lift increases and the drag decreases. [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]

The paper presents a methodology for calculating the unsteady aerodynamic forces characteristics and gust response of the unmanned air vehicle (UAV) model using computational fluid dynamics (CFD) tools, and the efficiency of gust alleviation based on direct force control are also studied. Based on unsteady Euler equations, dynamic responses in vertical gust flow perturbation are investigated. The grid velocity method, which is used to model unsteady flow via grid movement, is introduced to simulate the gust influence. This approach avoids numerical instabilities and decouples the step change in the angle of attack from a pitch rate. And the dynamic chimera grid method is used to simulate the rudder movement of the gust alleviation effect. Numerical results for the NACA0006 airfoil are validated by comparison with analytical results for two dimensional indicial responses. The initial and final lift values of the indicial responses closely match the exact analytical values as given by piston theory and quasi steady thin airfoil theory respectively. The solutions at small times match quite closely the exact solution which is known only for small times given by linear theory. The application of the same method to the UAV model reveals important characteristics of the three-dimensional vertical discrete gust response for the One-minus-cosine gust profile and gusts cause step changes in the angle of attack, and important characteristics of the gust response and resulting flow features are explained. Furthermore, the flow perturbations under dynamic rudder motion are investigated using moving chimera grid technique. Our study shows that with certain designed motions, the flapping rudder can effectively alleviate the freestream gust fluctuation and ensure a stable flight. However, no single parameter can determine whether the flapping rudder can reduce the gust effect. Finally, unsteady aerodynamic loads are computed for the wing of model with the designed rudder motion, the significance of the gust propagation effects is shown. It is found that the direct calculation of the gust response using CFD gives quite accurate results and provides a rich database. In general, this paper shows that the well-designed rudder movement can effectively alleviate the freestream gust fluctuation and ensure high flight qualities. [ABSTRACT FROM AUTHOR]

Finite diameters of seven-hole probes will bring systematic errors in the measurement in shearing flow. In this paper, the calibration of the coefficient matrix of the three-dimensional seven-hole probe in steady flow field is carried out by numerical simulation. It is studied that several factors, such as velocity gradient, probe diameter and roll angle, cause systematic error in the measurement of parallel shearing flow field by numerical simulation. In the experiment of seven-hole probe measurement of the shearing flow field, it is determined the spatial resolution of the seven-hole probe in shearing flow fields. "Critical diameter" is defined and the error distribution of seven-hole probes is established in this paper in the measurement of shearing flow fields under the allowance of error. [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]

Hypersonic aerodynamic heating characteristics for an aerospace vehicle at Ma = 8.0 and Ma = 10. 2 was calculated by computational fluid dynamics (CFD) method and engineering rapid computational method based on in-viscid surface streamlines. The Roe flux difference scheme(FDS) was employed in the CFD method. For the engineering rapid computational method, the in-viscid surface streamlines were determined by the Euler equations on the Cartesian grid, and the surface heating distribution can be computed by the integration along the in-viscid surface streamlines according to the Eckert reference enthalpy theory. Both laminar and turbulent flows were simulated in the present study. Comparisons were carried out between the engineering rapid computational results, the CFD results, and the experimental results. The results show that the heating distribution can be well predicated for the airframe and the wing of the aerospace vehicle by the CFD method and the rapid engineering method developed in this paper. The error of the aerodynamic heating distribution at the stagnation point is less than 5%. [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]

High order accuracy numerical simulations for flows of high lift trapezoidal wing configuration were performed by solving Navier-Stokes equation together with SST two-equation turbulence model. In the study, WCNS-E-5 scheme was utilized in the discretization of convection terms, while 6th order central difference scheme for viscous term. Grid convergence result was obtained, which was in good agreement with test data, and WCNS scheme was less dependent on mesh density than 2nd order scheme. Compared with other computational results obtained by typical CFD solvers, both the force coefficients and the surface pressure distributions given in this paper shown better accordance with the test data, and more reasonable prediction of flow structure near stalling was obtained. As revealed from these results, WCNS scheme took higher accuracy order and was more suitable for simulation of flow around high lift configurations than 2nd order scheme. The capability and the robustness of the program to simulate complex high lift configurations were confirmed. [ABSTRACT FROM AUTHOR]

Grid fin, also known as lattice control, is a kind of unconventional aerodynamic lifting and control surfaces that consists of an outer frame supporting an inner grid of intersecting planar surfaces with small chord. In this paper an efficient meshing strategy is developed to generate high quality point-to-point multi-block structured grid for such a complex geometry. The N-S equations are solved numerically using the finite volume method on the herein generated multi-block structured grid, adopting multi-grid method to accelerate the convergence. With these strategies the flow around a grid-finned missile is simulated at subsonic, transonic and supersonic states, with free stream Mach number ranging from 0. 5 to 3. The simulation results show excellent agreement with the experiment data for the aerodynamic forces and moments. Detailed flow field is also obtained that exhibits the complex flow properties inside the grid frame. The results provide an effective means for accurate simulation of grid fin missiles flow fields. [ABSTRACT FROM AUTHOR]

The two-temperature or multi- temperature approach used in the study of nonequlibrium flow cannot describe the distribution at vibrational energy levels and assumes Boltzmann distribution under vibrational temperature instead. Detail study of the characteristics of nonequilibrium process with state-to-state model may inspire the improvement of the two- temperature or multi-temperature approach. In this paper, the state-to-state model is used to simulate the thermo-chemical nonequilibrium process of a stationary closed N2/N mixture system. The time evolution of the density of N and N2 at different levels and the relaxation of vibrational distribution under different controlling conditions (temperature, pressure or density) and different initial conditions are analyzed, the characteristics and contribution of each micro processes are investigated. The results show that ; (1) The vibrational-translational energy exchange processes are the important processes which drive the vibrational distribution to Boltzmann distribution, while the vibrational-vibrational energy exchange processes only affect the characteristic of the transition course. (2) The departure of population distribution from Boltzmann distribution prevails in thermo-chemical nonequilibrium flow, and the evolution of the vibrational distribution for dissociation regime and that for recombination regime are different. (3) The results obtained from the study of thermo-chemical process with state-to-state model would help to thorough comprehension of the nonequilibrium phenomenon, and may provide reference for the improvement of thermo-chemical model used in the simulation of hypersonic nonequilbrium flow. [ABSTRACT FROM AUTHOR]

Computational fluid dynamics (CFD) gradually plays an important role in the evaluation of the wind tunnel design project. This paper is about the evaluation results of the slotted test section of 0.6m x 0.6m continuous transonic wind tunnel including alternative design and flow field quality, especially comparing full slot with half slot. It presents the control equations and boundary conditions for numerical simulation firstly. Then the details of numerical results about the high subsonic and low supersonic flow characteristic are given. The length of velocity uniformity region and velocity deflection angle meet the design accuracy and that proved the rationality. The half slot project was better than the full one at high subsonic. Low supersonic field could be achieved when air was bleed from plenum chamber. However, more bleed air rate, shorter length of velocity uniformity region. Full slot project had less disorder on flow of the plenum chamber than the half. Further investigation on the flow mechanism of re-entry flaps indicated that the flow was unsteady one significantly. [ABSTRACT FROM AUTHOR]

In order to design thermal protection system, it is important to predict heat transfer rate precisely on the hypersonic vehicles. In present paper, considering chemical reaction, interal energy excitation and thermo-chemical non-equilibrium effects of high temperature air, the thermo-chemical non-equilibrium flow-fields are simulated using the numerical code AEROPH_Flow. An aero physical numerical code system for Hypersonic vehicles, the distributions of heat transfer rate is obtained, and the analysis is done for the influence of the chemical reaction model, thermodynamic non-equilibrium model, surface temperature, surface material catalytic action on the computational results. The results show that thermo-chemical non-equilibrium models have a more important influence on the computational results, it is important for the selection of the correct air reaction model and thermal non equilibrium model according to the flight environment and thermo-chemical non-equilibrium mechanism of the flow field around the hypersonic vehicles; the heat transfer rate isn't always the highest for lower surface temperature, and it become very complicated for the variation law of heat transfer rate along with the surface temperature under the condition of high Mach number or thermo-chemical non-equilibrium flow, so the surface temperature may as well be fixed choosed according to the distribution of the real flight condition or different constant for computing exactly heat transfer rate on the hypersonic vehicles. [ABSTRACT FROM AUTHOR]

A laminar separation bubble appears generally on the Eppler387 airfoil at low Reynolds number. In the present paper, control of the laminar separation bubble over the Eppler387 airfoil at low Reynolds number using suction or blowing is investigated computationally employing the Fluent solver. Numerical simulations are carried out with a range varieties of the key control parameters of jet locations, jet angles and jet velocity ratios. Numerical simulations showed that employing the gamma-theta transition model coupling the SSTk-w turbulence model can predict the location and extent of the laminar separation bubbles accurately; blowing/suction control is effective in suppressing the laminar separation bubble over the airfoil and increasing the lift-to-drag radio at low Reynolds number obviously. The results with larger suction velocity ratios are better at the fixed suction jet locations while the smaller blowing velocity ratios results are better at the fixed blowing jet locations. Furthermore, suction on the airfoil's upper surface is more effective to control laminar separation bubbles than blowing control. [ABSTRACT FROM AUTHOR]

The investigation on the aerodynamic characteristics of a high speed ramjet propelled projectile with the design Mach number of 4 was performed by the method of numerical simulations. The internal and external drag acting on the projectile in passive flight phase was obtained through calculations. Variations of the drags were discussed at the range of Mach 4 to Mach 3. The results show that the inlet studied in current paper can restart at the design Mach number; the maximum total pressure recovery of the inlet is about 0. 25 at the sub-critical operational condition. When the passive flight Mach number decreases from 4 to 3, the internal drag coefficient changes scarcely, it is nearly 0. 06. While the external drag coefficient increases from 0. 16 at the Mach number of 4 to 0. 21 at the Mach number of 3. The total drag coefficient ranges from 0. 23 to 0. 27 while the Mach number decreases from 4 to 3. The scale of friction drag is about 40% in internal drag and 20% in external drag. [ABSTRACT FROM AUTHOR]

A method of designing icing detection system for wind turbine is presented in this paper. The main idea of the approach is to integrate icing sensors on an extra prober and to predict ice on wind turbine according to the prober icing. Basic steps for design are proposed. Numerical arithmetic used for design configuration and shape of the icing prober is given. The arithmetic is composed of the Multiple Reference Frame (MRF) method to calculate flowfield of air, a Lagrangian method to compute droplet trajectories and a technique for fast computing ice accretion. Icing prober configuration for a 1.5MW horizontal axis wind turbine is then obtained with the approach. The state of wind turbine icing can be indicated by the prober in real time. All these achievements build a good base for future research. [ABSTRACT FROM AUTHOR]

In this paper, based on the fixed hypersonic inlet with rectangular section, a new concept of variable geometry hypersonic inlet that the tip can turn and move back off to increase the mass capture ratio is presented, and operating in a large Mach number range from 4 to 6.5. The flow in the hypersonic inlet was analyzed with numerical analyzer, and the inlet performance comparisons were made between the variable geometry hypersonic inlet and fixed inlet at different Mach number of inflow. The results indicated that the performance of the variable geometry inlet as much as fixed one under design Mach number, and the performance of the variable geometry inlet was better than fixed inlet under non design point. All of that indicate that the variable geometry hypersonic inlet was satisfying. [ABSTRACT FROM AUTHOR]

For controlling the flow separation on the wing at high angle of attack, an innovative technique for a finites-span wing stall control, flow deflector, is studied in the present paper. Wind tunnel experiments are conducted to validate the reliability of the numerical simulation. The comparison between them indicates that the flow deflector can control the stall of a wing and restrain the flow separation effectively at high angle of attack. Deep research of flow direction changing trend and boundary layer are carried out to investigate the mechanism of flow deflector. The results indicate that the flow deflector increase the angle of flow around the leading edge of the wing and decrease the angle of the coming flow to make it deflect to the upside of the wing. The flow deflector can suppress the three dimensional effects and maintain the two dimensional flow characteristics to some extent. Furthermore, the flow deflector can easily change the velocity profile in the boundary layer and make it more "fuller". It decreases the shape factor and increases the stability of the boundary layer to restrain the flow separation. [ABSTRACT FROM AUTHOR]

This paper is the numerical investigation reports about field characteristic of fan section. First, it presents the control equations and boundary conditions for numerical simulation. Then the details of numerical results are given about velocity characteristics of fan section in 0.55mX0.4m low turbulence aero- acoustics wind tunnel. Conclusion could be got that numerical methods used could give rational results comparing with experiment data. Then, basing on the flow around rotor and stator, the flow mechanism inside fan section was analyzed. At last different fan models were study based on the numerical results. These parameters imposed influence on performance and flow mechanism of fan including the distance between rotor and stator, stator swept and leaned. Stator leaned could significantly reduce or eliminate rotor wake and tip separation, and could produce a more uniform velocity profile at exit. Conclusions state the importance and feasibility of CFD in aerodynamical design of fan section in wind tunnel. [ABSTRACT FROM AUTHOR]

CFD method has the advantage of well simulating the flow over wind farm of complex terrain, compared with the method of using the software WAsP with linear model, thus becoming the development direction of assessment of wind resources in complex terrain. In order to use the CFD package FINE/TURBO and standard k-ε turbulence model with wall function in the wind resources assessment and micrositing of wind farm in the complex terrain, in this paper, Island Nanao had been simulated in wind directions at intervals of 30 degrees, under the assumption of neutral condition, to investigate the influence of inlet condition on the results. The simulation results show that the wind speed-up and intensity don't change essentially under different inlet conditions. Based on this law, with whole year wind data, an assessment method with post processing has been built, providing a strong reference for engineering practice. [ABSTRACT FROM AUTHOR]

Airfoils of larger thickness are needed to improve the rigidity of the inflatable wing, while their aerodynamic efficiency is generally not great. A new aerodynamic configuration of double row wings applied to low-speed aircraft is discussed in this paper, and based on this, a new configuration which adds proper installation angle to the posterior wing is discussed. Moreover, the effect of the relative location of the double wings as well as the airfoil thickness on the row double-wing configuration is also studied. Numerical simulation result indicates that high-pressure zone around stagnation point at the leading edge of the posterior wing increases the pressure under the anterior wing, which will improve the lift and lift-drag ratio of aircraft obviously at a range of small and middle angles of attack. At a 4° angle of attack, the lift-drag ratio is improved by an extent of 62.8%, while 2° installation angle to the posterior wing is able to further improves the lift-drag ratio of the wing-body by 70.5%. Meanwhile, aerodynamic performance of the double-wing configuration is sensitive to the relative location of the double wings. The thicker is the airfoil, the smaller is the camber, and the more efficient is the row wing configuration studied. The overall effect shows that the aerodynamic configuration researched provides a new approach to the design of inflatable aircrafts. [ABSTRACT FROM AUTHOR]