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Copyright of Journal of Ballistics / Dandao Xuebao is the property of Journal of Ballistics Editorial Department and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Rotating nozzle profile enables the supersonic variable Mach number wind tunnel to continuously adjust the Mach number of the experimental area during a single operation，which is important for studying the aerodynam ic problems in the maneuvering process of the aircraft and the starting process of the inlet. In the process of controlling the nozzle profile rotation，whether the flow field parameters can change linearly is an important criterion to measure the performance of supersonic variable Mach number wind tunnel. Therefore，the linear variation law of flow field parameters in the experimental area of variable Mach number wind tunnel is analyzed in this paper，and a numerical simulation model is established using the spring smoothing dynamic grid technique to verify whether the flow field parameters in the experimental area meet the linear variation law when the nozzle is located at the corresponding position of Mach number 3.041~3.215. The results show that the linear change of the flow field parameters in the wind tunnel experimental area is achieved by controlling the rotation of the nozzle profile，and the dynam ic calculation results are in good agreement with the expected linear change law of the flow field parameters in the experimental area. In the linear change process of flow field parameters with different accelerations，the deviation between the average parameters of the experimental area and the expected parameters at each time is less than 0.13%.

The process from test starting to supersonic flow state established in an outlet-opening wind tunnel，i. e. the starting problem of supersonic flow，has become a recognized challenge. In order to establish the feasible flow field starting method，and lay the foundation for the use of outlet-opening wind tunnel，three-dimensional numerical simulation research is carried out，which is based on a certain wind tunnel，and taken the panel blade of supersonic compressor as the research object. The failure reason of supersonic flow field starting on experimental condition is analyzed，and three flow field starting schemes are established. The results show that the starting failure reason is the formation of a strong normal shock wave at the leading edge of the cascade. Increasing the inlet total pressure alone cannot establish the supersonic flow. The supersonic cascade flow field can be started by increasing the overflow gap width on lower wall side，but the number of effective cascade passages is decreased and the boundary layer thickness is increased. Maintaining the width of overflow gap on the upper and lower walls at more than 1.0 pitch，the setting upper and lower supersonic walls in front of cascade and conducting suction can effectively start the supersonic flow field. The maximum fluctuation of outlet Mach number is 0.01，the maximum fluctuation of outlet flow angle is 0.09°. The periodicity can meet the experimental requirements.

The back pressure at the outlet of supersonic cyclone separator usually causes aerodynamic shock wave in Laval nozzle, and the interaction of supersonic flow, shock wave and phase transition is very complicated. In order to explore the supersonic flow properties of wet gas under the action of shock wave, study was performed on the methane-water vapor mixtures. Then, a mathematical model for the supersonic condensation and flow of wet natural gas was established. Meanwhile, the effect of the outlet back pressure and inlet pressure on the supersonic condensation property of wet natural gas was analyzed to evaluate the performance of the supersonic cyclone separator under the action of shock waves. The simulation results show that the shock waves can cause the rapid increase in pressure and temperature of water vapor in a short time, leading to the condensed droplets re-evaporate and the liquid mass fraction at outlet reduces sharply. Besides, as the inlet pressure increases, the location of nucleation approaches the throat of Laval nozzle gradually, and the limit nucleation rate and the droplet number decrease accordingly. However, the liquid mass fraction increases due to the increasing of growth rate and sizes of droplets. Moreover, the nozzle loses its function of liquefaction when the back pressure is higher than 25 kPa (i.e., the pressure ratio is higher than 0.58), and then the nozzle cannot operate normally. Therefore, it is necessary to control the reasonable pressure ratio to ensure the dehydration efficiency of the supersonic separator with back pressure.

Copyright of Lubrication Engineering (0254-0150) is the property of Editorial Office of LUBRICATION ENGINEERING and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Prediction and limitation of inlet pressure are important for supersonic vehicles. Numerical simulation on high speed inlet flow is conducted, and pressure characteristics with different Mach numbers, attack angles, yaw angles and excess air coefficients is analyzed. Based on dimensionless and decoupled method, empirical formula for high speed inlet pressure is fitted and verified by flight results. The results show that, with the Mach number increasing and excess air coefficient decreasing, the pressure in inlet increases. Attack angle and yaw angle have little influence on inlet pressure in normal work envelope. The inlet pressure estimated by empirical formula is in good agreement with measured data in flight tests, which means that the formula established in this paper has a high precision.

In order to obtain the flow characteristics of sub-super-sonic mixing layer including velocity distribution, pressure distribution and development of mixing layer, experimental and numerical investigations were conducted. PIV technique was employed to measure the two-dimensional velocity distribution in the experiment while the standard k-ω turbulent considering the effect of compressibility was adopted to simulate the flow characteristic of mixing layer. The Mach number of subsonic stream and supersonic one was 0.11 and 1.32, respectively. The results show the flow of mixing layer is temporally transient. The interface between two streams lies initially as an approximately line segment; afterward, it becomes wrinkled and distorted; finally, it breaks up. The mixing layer develops linearly along streamwise direction in the time averaged velocity field with a growth rate of 0.135. The velocity and total pressure distributions in the mixing layer are self-similar.

The aerodynamic heating effect of the aircraft during supersonic flight has a great impact on the structural strength and thermal protection design. The aerodynamic thermal environment in the real state needs to consider the coupling of the external flow field and the structure and the influence of the boundary conditions of the inner wall surface. The S-A turbulence model is used to solve the Navier-Stokes equation. The coupling of the flow field and the solid wall is used to realize the coupled numerical analysis of the external flow field and the structural field. The coupling numerical study of the supersonic flow aerodynamic heating of three different airfoils is carried out, and the aerodynamic thermal effects of the inner wall surface of the airfoil under different thermal boundary conditions are compared. The results show that different airfoils have aerodynamic thermal effects similar to aerodynamic forces. The boundary condition of the inner wall considering convective heat transfer is the closest to reality; considering the fuel tank of the wing is full, the heat flux at the leading edge of the 3D wing can be up to 4 200 W/m2.

Copyright of Chinese Journal of High Pressure Physics is the property of Chinese Journal of High Pressure Physics Editorial Office and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Journal of National University of Defense Technology / Guofang Keji Daxue Xuebao is the property of NUDT Press and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

A method for rapid computation of supersonic rotary derivatives is developed. It is based upon perturbation relation between rotary derivatives and normal velocity on surface due to angular velocity according to Riemann invariant and Euler equations solution using Cartesian meshes. The resulting formula of supersonic rotary derivatives depends only on steady Euler solution of the basic flow,therefore,all supersonic rotary derivatives of interest can be obtained at the same time after obtaining the CFD steady-state Euler solutions. The international dynamic derivatives standard model ″Basic Finner″ is studied as an example to validate the method. The calculating results are in agreement well with experimental data and reference data appeared in open literature. It is proved that the proposed method has less runtime,higher accuracy and better practicability in engineering,and can be used to rapidly calculate the supersonic rotary derivatives of flying vehicles with complicated configurations due to using Euler solutions of Cartesian meshes [ABSTRACT FROM AUTHOR]

Based on Euler-Euler two-fluid model, the gas and liquid flow control equations were established. Combined with the theory of condensation nucleation and droplet growth, the condensation properties of CO2 in a Laval nozzle were simulated. The results showed that both the mathematical models and numerical calculation methods used could reflect the process of gas condensation in the Laval nozzle accurately. The condensation wave of CO2 was not obvious due to its small latent heat of condensation. As the gas flow into the nozzle, the Mach number increased while pressure and temperature decreased, the undercooling increased up to 30 K, and then declined quickly to about 5 K after the occurrence of condensation. The process of nucleation showed a sharp both in time and space. Wilson point is about 2. 21 mm from the throat, the nucleation rate increased sharply from 0 to 2. 04 X 1021 m-3 • s-1 , while the number of droplets reached the order of 1015 instantly. After the formation of condensation core, gas molecules under certain degree of undercooling accumulated and liquefied on the droplets surface, causing droplet radius as well as humidity to increase rapidly. After the process of nucleation, the existing coagulation cores kept growing with the mathematical models and numerical calculation methods used could reflect the process of gas condensation in the Laval nozzle accurately. The condensation wave of CO2 was not obvious due to its small latent heat of condensation. As the gas flow into the nozzle* the Mach number increased while pressure and temperature decreased, the undercooling increased up to 30 K, and then declined quickly to about 5 K after the occurrence of condensation. The process of nucleation showed a sharp both in time and space. Wilson point is about 2. 21 mm from the throat, the nucleation rate increased sharply from 0 to 2. 04 X 1021 m-3 • s-1 , while the number of droplets reached the order of 1015 instantly. After the formation of condensation core, gas molecules under certain degree of undercooling accumulated and liquefied on the droplets surface, causing droplet radius as well as humidity to increase rapidly. After the process of nucleation, the existing coagulation cores kept growing with the droplet radius increasing to 1. 46 X 10-7 m and the humidity up to 0. 093 5 at the exit of the Laval nozzle. [ABSTRACT FROM AUTHOR]

Copyright of Electronic Components & Materials is the property of Electronic Components & Materials and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Copyright of Journal of Vibration Engineering is the property of Nanjing University of Aeronautics & Astronautics and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

The supersonic flutter of aeroelastic multibody systems was investigated. Based on the theory of multibody dynamics and the piston theory for unsteady aerodynamic forces,the governing differential algebraic equations (DAEs) for the aeroelastic multibody system were established. The supersonic flutter analysis was conducted through solution of the eigenvalue problem in the form of DAEs to discuss the stability of the perturbed motion and the equilibrium of the multibody system. The approach was used for the supersonic flutter problems of a plate wing and a wing with control surfaces. The flutter speed of the wing was obtained with the control surfaces in different positions. The agreement between the simulation results and the experimental ones shows that, the present method has high computational accuracy, and is suitable for the flutter analysis of engineering structures consisting of multiple components. [ABSTRACT FROM AUTHOR]

In order to solve the questions of degenerate parameter estimation and performance recovery control of the engine components in supersonic condition, based on the improved Kalman filter, an approach of three-dimensional difference is proposed to achieve the parameter estimation. Performance recovery control increases the outer control loop based on the conventional inner speed control, and the outer control loop consists of thrust estimation model and outer controller. A thrust estimation model is designed based on the least squares support vector regression. The optimal inputs of model are selected through the thrust feature selection algorithm. The accuracy is improved greatly compared with the conventional non-selected thrust estimation model. Fuzzy PI controller is designed to achieve the purpose of degenerating engine performance recovery by adaptively adjusting the speed of the inner loop control. Finally, the numerical simulation of supersonic state verifies the effectiveness of the above programs. [ABSTRACT FROM AUTHOR]

Water vapor in natural gas often leads to the decrease of gas calorific value per unit volume; it also reduces the flow area of pipeline. Moreover, the CO2 and H2S will corrode the pipeline when dissolved in liquid water. According to the containing water problems in high pressure natural gas during development, a roof-mounted type supersonic cyclone flow natural gas dehydration unit was designed in combination with the fluid mechanics and engineering thermodynamics principle. Based on the research status at home and abroad, the supersonic natural gas condensate flow numerical model was built, including multicomponent two-phase expansion flow model and water vapor condensing model. The flow characteristics of each work section of supersonic natural gas dehydration unit were numerical analyzed, and the device internal pressure, temperature, Mach number, vapor nucleation rate, moisture distribution were obtained. According to the results of numerical simulation, optimization design was carried out to improve the efficiency of separation. [ABSTRACT FROM AUTHOR]

Using the characteristic boundary conditions and high-order finite difference schemes, direct numerical simulation(DNS) was performed to investigate the effect of phase and frequency of the periodic blow and suction disturbance which was used to trigger the by-pass transition of the flat plate boundary layer with free stream Ma∞=2.25, Re∞=2.5×107/m (635 000/inch). The numerical results show that the randomness of the disturbance phase is not the necessary condition to trigger the transition, and with the deterministic disturbance phase, the average location of transition moves backward when the disturbance frequency increases. However, if the frequency is bigger than a critical value, the transition in the boundary layer will not happen. A DNS database was established under the condition of deterministic disturbance phase. The Van Direst velocity agrees well with the wall and logarithm law of incompressible flat plate boundary layer, and the turbulence intensities coincide with the incompressible flat plate experimental datum.The averaged thermodynamic variables are very close to that of reference results. These verify the present DNS results. Then, the statistics of thermodynamic variables fluctuation and Reynolds stress are extracted from the DNS database. Compared with the relative references' DNS results, the disturbance, boundary conditions and numerical methods have observably effect on the second order statistics and the averaged transition position. [ABSTRACT FROM AUTHOR]

This study evaluates the application of Menter's γReoθ transition model with a RANS solver for transition prediction on super-sonic and hypersonic flows. The transition onset position and transition length are approached by utilizing two transport equations for local Reynolds number and intermittency. The present work expands upon the transition prediction methodology by validating the model against wide range of Reynolds number, three-dimensional flows in the form of cones at angle of attack and bluntness effects. Test cases including Mach 3.5 cone with free stream unit Reynolds numbers from 2.8E7 to 7.8E7, Mach 5.91 flared cone and cone with 0 to 4deg angle of attack and Mach 7. 16 cone with 3 nose radius have been performed. Comparisons for skin friction coefficient and heat transfer with experimental data are presented. Results show that this model predicts the correct trends with respect to angle of attack and bluntness effects. It is found that the agreement between simulation and measurements varies better from small Reynolds number to large cases and However there still are few limits in this very model in predicting supersonic or hypersonic transition process. [ABSTRACT FROM AUTHOR]

With CFD-based local piston theory, expressions have been derived for the unsteady aerodynamic forces acting on the fuselage and wing. The flutter characteristics of a wing-fuselage complete vehicle have been studied. The flutter boundary results by CFD-based local piston theory agree well with those by the CFD/CSD direct coupling method, demonstrating that this approach is suitable and efficient for the flutter analysis of a wing-fuselage complete vehicle at high angles of attack. [ABSTRACT FROM AUTHOR]

A mathematical model of the adjustable ejector was established and the flow of the adjustable ejector was simulated by FLUENT on different spindle positions. The experimental results indicate that the axial pressure of an adjustable ejector increases linearly in the constant area mixing section, sharply in the beginning of diffuser section and then slowly in the other sections of diffuser. The velocity of flow in the mixing section differs significantly at the axis and boundary. The movement of spindle into the nozzle makes the position of shock wave upstream and motive flow rate decreases. [ABSTRACT FROM AUTHOR]

In order to carry out an overall rapid calculation and analysis on supersonic gas bearings flow field, the original simplified model was greatly improved. The flow passage of externally pressurized gas thrust bearing with central supply hole was built. Based on the fixed flow rate under the choke condition, calculation formulas were derived for flow velocity, Mach number, Reynolds number and press distribution in the flow passage to calculate the flow field. The result presents that, shock wave position, flow velocity, Mach number and Reynolds number are related to the supply press, the beating clearance and the diameter of supply hole. Compared with experimental results, it is found that, with the increase of supply press, the press force and the maximum raising press position obtained by the improved simplified model are basically consistent with experimental results, and the stress in the flow field can be reflected in essence. However, the shock wave at the entrance in the simplified model is assumed as the normal shock and gas viscousness is neglected, so the calculated minimum press deviates widely from the experimental result. [ABSTRACT FROM AUTHOR]