2014 Vol. 28, No. 2

Display Method:
Liao Daxiong, Huang Zhilong, Cheng Zhenhua, Tang Gengsheng
2014, (2): 1-6,20. doi: 10.11729/syltlx20130102
Abstract(181) PDF(44)
Abstract:
With the development of air transportations,detail-optimized designs of advanced aircrafts demand aerodynamic data under the flight Reynolds number conditions.Large-scale cry-ogenic wind tunnels,such as ETW and NTF,are the best ground testing facilities to obtain air-craft flow characteristics in the real flight conditions.To facilitate the development of large-scale high Reynolds number wind tunnels,the achieving means and types are summarized,the current developing status is discussed,the key technologies and solutions are analyzed in depth for design methodologies and construction concerns.Finally,the future designs and constructions of large-scale continuous cryogenic wind tunnel in China are prospected.
Luo Xisheng, Wang Xiansheng, Chen Mojun, Zhai Zhigang
2014, (2): 7-13,26. doi: 10.11729/syltlx20140015
Abstract(109) PDF(5)
Abstract:
A simple method of generating gas cylinder is proposed by using the soap film tech-nique.The formed interface is free of supporting mesh and the initial condition can be well con-trolled .Experiments are then carried out in a horizontal shock tube for a planar shock wave inter-acting with a two-dimensional or three-dimensional cylinder using high-speed schlieren technique. It is found that the schlieren images have fewer disturbing waves and the evolving interfaces are more symmetric comparing with the results in literature.Special attention is given to the three-dimensional effects caused by the minimum surface feature on the interface evolution and it is in-dicated that the interface evolving rate of three-dimensional cylinder is slower than that of two-di-mensional cylinder.This work will facilitate more understanding about the three-dimensional effect on the development of the Richtmyer-Meshkov instability.
Zhang Lei, Zhang Ruoling, Xiao Shide, Zhang Xiangwen, Le Jialing
2014, (2): 14-20. doi: 10.11729/syltlx20120131
Abstract(60) PDF(5)
Abstract:
The flow and heat transfer of n-decane under supercritical pressure were studied ex-perimentally using an electrically heated tube according to the conditions in a regenerative cooling scramjet.The fuel was heated in a 1.5 mm inner diameter tube of 1Cr18Ni9Ti,and the total length of the tube was 1300mm.The fuel mass fluxes were 0.93g/s,1.24g/s and 1.86g/s in the three runs respectively.Seventeen type K and 0.3mm outer diameter thermocouples were spot welded directly to the outer surface of the tube to measure outer wall temperatures.The pressure of n-decane was varied from 4.0 to 4.3MPa,and the fuel temperature was varied from 335 to 870K in the experiments.The Reynolds number ranged from about 800 to 70000 and Prandtl number ranged from about 0 to 1 5 .Blank tube heating experiments were used to calibrate the tube thermal loss rate at different temperatures.At the positions within 0~0.2m and 1.2~1.3m along tube,inner wall heat flux was smaller due to additional thermal loss by the joint metal blocks near the inlet and outlet.Especially within 1 .2~1 .3m which was near the outlet,the tube temperatures were much higher than those within 0~0.2m near the inlet,so the inner wall heat flux drops steeply down.In the tests the data within 0.2~1.2m along tube were chosen for stud-y.The heat transfer correlations of n-decane in laminar,transition and turbulent flow regions were determined using the method of least squares curve fitting.The calculated outer wall tem-peratures were compared with experimental values,which confirmed the validity of the presented correlations .
Li Xiaolu, Tang Kai, Lei Ming
2014, (2): 21-26. doi: 10.11729/syltlx20130058
Abstract(91) PDF(5)
Abstract:
This paper introduces the latest developments about flight tests of the vertical-fin buffeting from domestic and overseas.Deduces and analyses theoretically on the flight test meth-od for buffeting flight test are done.In domestic,the analysis of buffeting mostly focuses on the-oretical derivation and wind tunnel test up to now.But in actually flight,the result usually has some difference with theoretical derivation and wind tunnel test.In order to find the variation of buffeting in flight test,six acceleration-sensors on the vertical-fin are installed,including tip of leading edge,middle of leading edge and tip of trailing edge both in left and right vertical-fin.In flight test,wind-up-turn flight method is used at pressure altitude of 10000 meters,Mach from 0.65 to 0.95.From the test,the buffeting response of the vertical-fin in different Mach numbers is got.Then by the analytical method of root-mean-square (RMS),time-frequency analyses and power-spectrum-density (PSD),establishing the relation between buffeting response and either angle of attack (AOA)or frequency,given the initial angle of attack at different Mach numbers, we can find that the buffeting response is in accord with the variation in strouhal equation.The response mainly is in low-order model frequency of the vertical-fin,and this key frequency chan-ges with flight state and measuring position.The vertical-fin buffeting response increases with the increase of both the angle of attack and Mach number,and the angle of attack influences more.After plane exceeds initial buffeting angle of attack in flight,the buffeting response of the after edge in the tip of the vertical-fin is significantly greater than that of the leading edge with the increase of the angle of attack.
Li Feng, Gao Chao, Zheng Borui, Wang Yushuai
2014, (2): 27-31,44. doi: 10.11729/syltlx20130197
Abstract(66) PDF(5)
Abstract:
In view of the technical requirements for the research on high speed flow control by plasma actuation,a set of plasma flow control system in high speed wind tunnel experiments was established through the special model and experimental installation design,sealed insulation alignment and multi-layer electromagnetic shielding techniques,then the technical specification and operation strategy were proposed for high speed flow control test by plasma actuation,and the control law of plasma actuation on a two element airfoil flow was explored.After using these techniques,the insulation,sealing alignment problems of high voltage cable were solved and the induced voltage between model and experimental installation decreases by more than 90%.The wind tunnel test results show that this system operates stably and the experimental data is relia-ble.The flow of Ma= 0. 2 can be controlled effectively by plasma actuation.The flow separation of NACA0012 airfoil is weakened significantly,and its lift increases and drag decreases.In addi-tion ,the critical stall angle of attack increases by 2°and the maximum lift coefficient increases by 4% due to the actuation.As a result,the overall aerodynamic performance of NACA0012 airfoil is improved.The results of this study provide an important reference and technical support for the further research on high speed flow control by plasma actuation.
Liu Li, Wang Xiaoqing, Chen Pengzhen, Chong Jinsong
2014, (2): 32-38. doi: 10.11729/syltlx20130111
Abstract(904) PDF(5)
Abstract:
The measurement of weak current on water surface under wind wave environment is a significant procedure for studying the interaction between current and wind waves,which is very important for the microwave remote sensing of internal waves,submarine topography and eddy etc.Traditional measurement methods can not work under wind wave environment,and then make it hard to study the interaction effect between current and wind waves.A new method for measuring weak current under wind wave environment based on linear array CCD is proposed in this paper.The wave tank experiment with surface current caused by internal waves is carried out to prove the validity of this method.Then the experimental data show that the measuring precision of water surface weak current velocity is better than 0.3cm/s.The experiment indicates that this method can be used to measure the weak current under wind wave environment and research the interaction between current and wind waves.
Ma Binghe, Wang Yi, Jiang Chengyu, Li Yanbing, Wang Leitao
2014, (2): 39-44. doi: 10.11729/syltlx20140006
Abstract(101) PDF(8)
Abstract:
The fluid temperature difference between practical operation and its original calibra-tion procedure can cause an unacceptable error for micro thermal shear stress sensors.Thus,the temperature correction is necessary when the working water temperature is different from the original calibration temperature.By analyzing the temperature dependence of the overheat ratio and coefficients of the sensor's behavioral equation,a correction method is presented.Laminar flow experiments with changeable temperatures are carried out.The results show that the influ-ence of water temperature can be minimized,and the relative errors of sensor output at water temperature of 25℃ and 28℃ can be falling to 0.82% and 0.83% from 23.7% and 37.1%,re-spectively.
Ma Changyou, Hou Minjie, Ling Daijun, Xing Xiaolong
2014, (2): 45-50,58. doi: 10.11729/syltlx20130038
Abstract(59) PDF(7)
Abstract:
The tracer particles selecting and sowing was very important in PIV technology ap-plied to blow-down and high subsonic plane cascade flow field measurement.In this paper,tracer partides flow was effectively mixed with the main flow by using high pressure atomized particle generator and broadcaster installed in front of the steady pressure section,and we successfully obtained 2-D velocity fields of a compressor cascade channel and wake at zero incidence angle and inlet Mach number from 0 .2 to 0 .8 .In order to validate reliability of PIV results in the cascade flow field,the PIV results were compared with the results of the numerical simulation and three hole wake probe in the same aerodynamic conditions.The results indicated that 2-D velocity vec-tor field acquired by the PIV was closer to the numerical simulation results,and reasonably re-flected the flow structure of the blade passage;the PIV data got good agreement with the three hole wake probe data in the aspect of the velocity magnitude and flow angle at mainstream area of the cascade outlet.But the coincidence of flow angle at wake separation area was slightly bad. The main reason was that the flow angle in this area was beyond the scope of probe calibration. In addition,in order to obtain accurate PIV measuring results,it was very important to avoid contamination of the tracer particles to the window during the measurement.PIV measurement technique proposed in this paper could also be used for continuous cascade wind tunnel.
Li Shizhu, Cai Xiaoshu, Yu Jianfeng, Li Dianxi, Gao Yang, Li Junfeng, He Naibo
2014, (2): 51-58. doi: 10.11729/syltlx20130108
Abstract(66) PDF(5)
Abstract:
The probe used for transonic flow measurement should be calibrated in a wind tun-nel with Mach number from subsonic to transonic.The Mach number at the outlet of the slotted nozzle is changeable due to the self-adaptive effect of the nozzle in different backpressure.There-fore,the wind tunnel equipped with the slotted nozzle may be operated from subsonic to ultrason-ic for calibrating the transonic probe.For studying the performance of the slotted nozzle with wet steam as the working medium and optimizing its structure,detail numerical simulation is carried out by solving 3-D N-S equations and the realizablek-εturbulence model.The numerical results show that converging curve、divergent section length and slot size may affect flow field character-istic of the nozzle seriously.In a certain range of backpressure and inlet pressure ratio,there are optimal convergent curve,divergent length and slot size.According to the results of the numeri-cal simulation,a wind tunnel equipped with the slotted nozzle is developed.The Mach number of the tunnel with wet steam as the working medium may be continuously varied from zero to 1 .6 . The experimental results show that the flow at the outlet of the nozzle with such optimal struc-ture are uniform and stable in a wide range of Mach number from zero to supersonic.It is quali-fied to meet the requirements of transonic probe calibrating.
Yu Shice, Jiang Jianqun, Lou Wenjuan, Sun Bingnan
2014, (2): 59-64,68. doi: 10.11729/syltlx20120179
Abstract(90) PDF(5)
Abstract:
Large boundary-layer wind tunnel is necessary equipment for wind engineering re-search.In the background of the development of Zhejiang University ZD-1 boundary layer wind tunnel,the key technical problems in the design of the aerodynamic and vertical structure of the large recirculating boundary layer wind tunnel are introduced in detail.Test section of ZD-1 wind tunnel is designed as 4m width and 3m height,the maximum wind speed is designed as 55m/s, and the turbulence intensity is designed as not higher than 0 .5%.With the contraction ratio of 4∶1 ,single test section and single loop aerodynamic profile are used in aerodynamic design,e-quivalent diffusion angle of 0.22°is arranged in the test section,and the shape of corner vane is treated specially.The combination of steel structure and concrete structure is used in the vertical wind tunnel structure design.The results of flow field calibration show that the maximum wind speed of wind tunnel reaches 55m/s,the maximum energy ratio reaches 2.0,the hydrodynamic stability coefficient in master test area for wind speed of 40m/s is less than 0.5%,the maximum axial static pressure gradient is 0.0044,and the maximum turbulence intensity is 0.47%.It is showed that all the indexes meet the design requirements,and some indicators reach the standard of aviation wind tunnel.It is worth stressing that setting axial diffusion angle in the test section can reduce axial static pressure gradient,and the dynamic pressure nonuniformity coefficient con-tour shows the characteristics of horizontal distribution in the test section,which confirms that the uniformity of dynamic pressure for horizontal direction is better than vertical direction.Verti-cal structure design plays a certain role for improving the level of horizontal uniformity of the test section,which provides a reference for the design and construction of similar wind tunnels.
Wu Wentang, Hong Yanji, Ye Jifei, Jiang Guanlei
2014, (2): 65-68. doi: 10.11729/syltlx20130055
Abstract(85) PDF(9)
Abstract:
Impinging jets widely exist in many engineering fields,such as a short takeoff ver-tical landing (S/VTOL),as well as rocket lunching.It’s important to know the shock wave structure of impinging jets.In a simple rainbow schlieren system,the knife edge in the cut-off plane is replaced by a tricolor filter with a background color center and two colors separated by central band.The direction of the filter can usually be adjusted so that the density gradients in different direction can be examined.The rainbow schlieren apparatus requires no laser source, the optical alignment is simple,and it is tolerant to minor mechanical imperfections and/or vibra-tions.In this paper,the rainbow schlieren method is used in impinging jets experiments at differ-ent range and pressure ratio.Some perfect experimental results are gotten.Three kinds of impin-ging j ets are discussed and found that the size and shape of mach disk are depended on the pres-sure ratio.The result shows that when the distance between nozzle and ground plate is relatively large,the structure of impinging jet is similar to free jet,and the jet district near the ground plate is weakly.As the distance become closer,the jet region near the ground plate becomes stronger and a big stagnation bubble appears in the recirculation flow region at a particular pres-sure ratio.
Zhu Chao, Yao Feng, Chen Dejiang, Zhou Wei, Li Zeyu
2014, (2): 69-72. doi: 10.11729/syltlx20130043
Abstract(81) PDF(7)
Abstract:
The auto-ignition with low-pressure argon stream used to ignite the arc heated wind tunnel is investigated,in order to overcome the disadvantage of the conventional ignition method, which has a long preparing time,poor reliability and safety in operation.The effects of electrode gap,gas supply way and ignition sequence on ignition performance were examined experimental-ly.Arc heater with the length of several meters was ignited first by breakdown of argon between short electrode gaps at trigger voltage and then arc elongate to anode with the close of switches in sequence .It was indicted that arc heated wind tunnel could be easily ignited with low-pressure ar-gon stream by the way of step trigger.When the electrode gap is 25mm,the ignition performance is the best,whose start time is the shortest and start arc current is the lowest.Ignition gas argon should be injected in the rare of arc heater,which could reliably ignite arc heated wind tunnel and have several advantages in operation and maintenance.And a good ignition sequence used in arc heated wind tunnel was also suggested.
Lan Shengwei, Liu Sen, Li Yi, Huang Jie
2014, (2): 73-78,104. doi: 10.11729/syltlx2014pz34
Abstract(78) PDF(10)
Abstract:
Spacecraft breakup model is used to describe the distributions of number,size,are-a-to-mass ratio and velocity increment of debris produced in spacecraft fragmentation due to ex-plosion or impact.It is important for space debris environment modeling,impact risk assessment and emergency breakup events analysis.The basic forms and features of NASA standard breakup model which is widely used currently are introduced.The shortcomings of this model revealed in typical satellite impact events are analyzed.The main research advances on spacecraft breakup modeling aimed to improve and revise the current model in recent years are introduced.These re-search advances include the small satellite impact tests conducted by Kyushu University,the cube satellite impact tests and numerical simulations performed by Ernst-Mach Institute,and NASA’ s new satellite impact project started in 2011,in which a high fidelity satellite named DebrisSat will be manufactured and impacted.The spacecraft breakup model in China is focused on the studies performed by China Aerodynamics Research and Development Center (CARDC).A series impact tests of simulated satellites in different size and mass are conducted and numerical analysis for more than twenty fragmentation cases are finished.A new spacecraft impact breakup model named CARDC-SBM is constructed based on these test and numerical results.The differences be-tween CARDC-SBM model and NASA standard breakup model are presented in this paper.The problems and disadvantages in these current breakup model studies are analyzed,then the future research focused,such as the effect of impact conditions and the influence of different space mate-rials,is indicated.
Niu Wenxia, Huang Jie, Ke Fawei, Liang Shichang, Jian Hexiang, Liu Sen
2014, (2): 79-84. doi: 10.11729/syltlx2014pz38
Abstract(70) PDF(3)
Abstract:
In order to study the kinetic energy weapon damage effect to concrete targets, based on the experiment and numerical simulation method,the tungsten alloy projectile hyperve-locity impacting concrete building structure was studied and concrete building structure damage characteristics were discussed under the condition of hypervelocity impact.The weight of tung-sten alloy projectile used in the test is 35g,the ratio of which length to diameter is 5,and which impact velocity is 2.5km/s;the concrete compressive strength of building structures target is 34.4MPa.Studies had found that the top plate and bottom plate of the structure target were pen-etrated without the occurrence of disintegration under the experimental conditions.In addition, reinforcement in the target body had no significant effect on the test results.Calculated and ex-perimental results showed that speed of projectile fell from 2.5km/s to 2.0km/s~2.3km/s after impacting 40mm concrete slab.The debris cloud from concrete structure top plate had considera-ble speed.Some of them had strong killing effect on the staff with 23°expansion angle,while distribution range was greater than the perforation diameter on the bottom plate.
Chen Hong, Zhou Zhixuan, Huang Jie
2014, (2): 85-89. doi: 10.11729/syltlx2014pz29
Abstract(107) PDF(7)
Abstract:
To study the influence of attack angle on the process of tungsten rod oblique pene-trating aluminum plate,numerical simulation is carried out.In the simulation,SPH method is a-dopted,with Shock Equation of State and Steinberg strength model.The tungsten rod has twodifferent sizes,which areΦ1 0mm×2 3mm andΦ1 0mm×46mm.The thickness of aluminum plateis 1 0mm.The impact velocity is 2km/s,the impact angle is 60°.The attack angle consists ofpitch angle and yaw angle,the influences of each are studied separately.The pitch angle variesfrom-90°to 90°,and the yaw angle varies form 0°to 90°.The rotation and bending of the tung-sten rod are analyzed,the result shows that:the critical attack angle exists in the oblique pene-tration,within the scope of the critical attack angle,the moment of tungsten rod can be neglec-ted,and the tungsten rod doesn’t rotate or bend during the penetration process;outside thescope of the critical attack angle,the moment of the tungsten rod can’t be neglected,which canbend or even break the tungsten rod.The residual mass and residual velocity of the tungsten rodafter penetration are analyzed,the result shows that:when the yaw angle is 0°,the residual masskeeps at a small range unless the rod is broken,and the tungsten rod with L/Dratio of 4.3 iseasier to be broken than the tungsten rod with L/Dratio of 2 .3 ;when the pitch angle is 0°,theresidual mass decreases linearly as the yaw angle varies from 20°to 80°;for the tungsten rod withL/Dration of 2 .3 ,the fitted residual velocity-attack angle curve at 0° yaw angle overlaps thecurve at 0°pitch angle,which means the yaw angle and the pitch angle has the same influence on theresidual velocity of the tungsten rod;when the yaw angle is 0°,the residual velocity-attack angle curve ofdifferentL/Dratios shares the same shape,thus a uniform empirical formula of the residual velocity ofdifferentL/Dratios is constructed,the constant in the empirical formula is obtained by curve fitting.
Ma Zhaoxia, Huang Jie, Shi Anhua, Su Tie, Liu Sen
2014, (2): 90-94. doi: 10.11729/syltlx2014pz27
Abstract(61) PDF(4)
Abstract:
The atomic emission spectrum of ricochet debris cloud,which was from aluminum plate subjected to hypervelocity impact by aluminum projectile,was obtained by transient spec-troscopy during hypervelocity impact test.The waveband range of spectrum which achieved in the test was 250~340nm.By analyzing the spectrum,six aluminum spectrum peaks were distin-guished from the emission spectrum and two coupling spectrum peaks were decoupled.With the aluminum spectrum constants and intensity of the six peaks,the ricochet debris cloud tempera-tures were diagnosed separately in different impact condition by using the multiple spectrum peaks method.It was found that the temperature would increase when the projectile diameter in-creased and it would increase with the impact velocity increasing,too.By analyzing the test re-sults of different projectile diameter,it was found that the ricochet debris cloud temperature was acutely sensitive to impact velocity.Based on the data achieved in the test,experiential formula describing relationships among ricochet debris cloud temperature,projectile diameter and impact velocity was obtained.Then the radiation characteristics of ricochet debris cloud were studied, every spectrum peak was integrated and the spectrum in the range of 250~340nm was integrat-ed,too.It was found that the spectrum peaks integral,the waveband integral of 250~340nm and impact kinetic energy behave in a linear fashion.The slopes of the spectrum peaks integral to impact kinetic energy were obtained and it could be considered as the radiation efficiency of every spectrum peak in hypervelocity impact.Finally,taking the experiential formula of ricochet debris cloud temperature into consideration,the relation between atom number in the ground state and impact parameter was deduced.The relations among ionization rate,gasification rate and impact parameters were explored.The conclusions would provide technical support for the further studies of the particular cloud temperature produced by hypervelocity impact and the impact parameters deducibility from impact radiation.
Jiao Dezhi, Huang Jie, Ping Xinhong, Xie Aimin, Luo Jinyang, Liu Sen
2014, (2): 95-98. doi: 10.11729/syltlx2014pz01
Abstract(92) PDF(11)
Abstract:
Two-hundred meter free flight ballistic range is the only one comprehensive ballistic range equipment in China that can carry out aerodynamics test,aero physics test,erosion and hypervelocity impact.In order to meet the demand of the hypervelocity aero craft development, the equipment has been upgraded since 2009 .The upgraded equipment includes launch system, the chamber and vacuum system,velocity measurement and control system,photogram and ori-entation system.The launch system will be equipped with 203mm and 120mm bore two-stage light-gas guns.The weight of the model that the launchers can launch are from 0.5kg to 30kg, and the model velocity is from 0.3km/s to 5km/s.The diameter of the tank will be increased from 1.5m to 3m.At the same time,The range will equipped with new vacuum system,which can realize the simulation of altitude from 0km to 80km.Velocity measurement and control sys-tem will meet the big model measured requirement.Except shadow and schlieren equipment,the position system of double eyes front light imaging,X-ray imaging and measurement system will be constructed.After upgrading,from small to big bore two-stage light-gas guns will be e-quipped in the launcher system of 200m free-flight ballistic range,and the test abilities of the range will be improved in aero-dynamic test,erosion,high and hypervelocity impact test.
Wen Xuezhong, Ke Fawei, Chen Ping, Ma Zhaoxia, Huang Jie
2014, (2): 99-104. doi: 10.11729/syltlx2014pz16
Abstract(72) PDF(5)
Abstract:
In order to improve the performance of the shield without increasing its weight or size,an N-shape configuration with an oblique middle wall is proposed in the study based on the fact that the ballistic limit of oblique impact is higher than that of normal impact.The perform-ance of the N-shape configuration and a parallel triple-wall configuration with the same areal den-sity are compared and analyzed by 3D numerical simulation (SPH )and hypervelocity impact tests.The hypervelocity impact tests under different impact conditions have been carried on.The test results,which are achieved under the condition that the diameter of projectile in the normal impact is 4.0mm and the impact velocity is about 3.0km/s,indicat that there is little difference of damage on rear plates between the two configurations.In another normal impact test with a projectile diameter of 5.5mm and an impact velocity of 4.8km/s,the damage difference become considerable,and there is a perforation with a diameter of 2.0mm in the rear plate of the triple-wall configuration but no perforation happens to the rear plate of the N-shape configuration under the same condition.Simulation with the same impact parameters as the tests is carried out.The results show that the critical proj ectile diameter of the triple-wall configuration is about 4 .1 mm and that of N-shape configuration is about 4.3mm,when the velocity of the normal impact is 3.0 km/s.The critical projectile diameter of triple-wall configuration is about 5.5mm and that of N-shape configuration is about 5.9mm under the normal impact with a velocity of 4. 8km/s.The simulation results are in accordance with the test results.These results showed that the perform-ance of N-shape configuration is improved by the oblique middle wall under the condition of nor-mal impact espercially with a higher velocity.