实验流体力学

2014, (1)

Abstract(48) PDF(4)

Abstract:

Yi Shihe, Chen Zhi, He Lin, Wu Yu, Tian Lifeng

2014, (1): 1-11. doi: 10.11729/syltlx2014ty01

Abstract(84) PDF(16)

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Supersonic/hypersonic flows involved with the high speed flight have received vast interests.Features of such flows like unsteadiness,intensive gradient and compressibility chal-lenge the ability of experimental methods.Nano-tracer planar laser scattering (NPLS)is a non-intrusive optic technique developed by the authors’group in 2005 ,which can visualize the tempo-ral-resolved flow structures in a cross-section of instantaneous three-dimensional supersonic flow with high spatiotemporal resolution.Many studies have demonstrated that NPLS is an effective technique for studying supersonic turbulence.In recent years,the authors have made great pro-gresses in supersonic turbulence study using NPLS.And several techniques based on NPLS have been developed,such as NPLS-based density technique (NPLS-DT),which helps to understand the density field of supersonic flows and obtain Reynolds stress distribution.This paper intro-duced NPLS technique for visualizing the fine structures of supersonic flows,and reviewed its ap-plications in supersonic boundary layer,shock/boundary layer interaction and so on.With its a-vailable application on measuring the flow static parameters such as Reynolds stress and turbulent kinetic energy,NPLS would contribute to the development of turbulence models for compressible flows .

Cai Xiaoshu, Zhou Wu, Yang Huinan, Yang Bin, Chen Jun, Su Mingxu, Li Junfeng

2014, (1): 12-20. doi: 10.11729/syltlx20130069

Abstract(63) PDF(7)

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In this paper,the research advances in the in-line measurement techniques for com-bustion and flow field at Institute of Particle and Two-Phase Flow Measurement,University of Shanghai for Science and Technology have been presented.The flame radiation temperature measurement technique with fiber-based spectrometer has been improved,the flame temperature measurement accuracy and range were increased.Moreover,the flame emissivity and fuel compo-sition can be obtained.Temperature and water vapor concentration for detonation flame were de-termined by combining radiation and absorption spectroscopy methods.The spray droplet tem-perature was measured based on laser-induced fluorescence.Concentration of polluting gas emis-sions(SO2 and NOx)were obtained by differential optical absorption spectroscopy (DOAS)meth-od interfered by high density of dust and water droplets.Hg continuous monitoring system with spectroscopic method based on saturated vapor theory was developed.The temperature field of high-temperature object was measured with combination of spectroscopy and imaging methods. The velocity vector field and granularity distribution of multiphase flow were studied with single-frame and single-exposure imaging method.The liquid-phase parameter (film thickness and tem-perature)and vapor-phase parameter (water vapor temperature above the liquid film)during the liquid film evaporation process were simultaneously measured by tunable diode laser absorption spectroscopy (TDLAS).

Tao Yang, Zhao Zhongliang, Wang Hongbiao, Yang Haiyong, Guo Qiuting

2014, (1): 21-25. doi: 10.11729/syltlx20120203

Abstract(52) PDF(5)

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The wing rock phenomenon induced by forebody asymmetric vortices is likely to oc-cur on modern aircrafts with slender forebody flying at large angles of attack.Over a model of a pointed ogive-cylindrical body with 30°swept wing at the fixed angles of attack in the wind-tun-nel,the effect of artificial perturbation’s location around the nose tip is studied through the free to roll test in the critical Reynolds numbers range.The test apparatus and test methods as well as data acquisition were presented too.The experiments are conducted in the 0.6m×0.6m section of transonic wind tunnel of CARDC.The encoder with the resolution of 0.088°which is coupled with the rotor is employed to record the free to roll motion.High precision mechanical bearings are used to minimize the friction in free to roll experiments.A spherical particle with the diameter of 0.2mm which is attached onto the rotatable tip of the model is as the artificial perturbation.Based on various type of experiments of forebody asymmetric vortices that influence the wing rock motion over the slender body with 30°swept regular wing,the relationships between tip perturbation and roll-ing oscillation induced by different types of forebody asymmetric vortices are studied.As the test result indicated,the manual perturbation on the nose tip could prevent the wing rock remarkably when it was placed on side direction and the control effect work on a wide range of attack and Mach numbers.The wing rock flow control mechanics are briefly analyzed too.

Zhang Hui, Fan Baochun, Chen Zhihua, Li Hongzhi

2014, (1): 26-30. doi: 10.11729/syltlx20130011

Abstract(49) PDF(6)

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The Lorentz forces generated by electromagnetic fields on the surface of the cylin-der in the electrolyte solution may modify the structures of the flow boundary layer effectively.In this paper,control of vortex-induced vibration by Lorentz force is investigated experimentally which aims for vibration suppression.Experiments are conducted in a rotating annular tank filled with a low-conducting electrolyte.A long metal sheet was used as a flexible support to mount the cylinder with a 2cm diameter.The cylinder with an electro-magnetic actuator mounted on the surface is placed into the electrolyte.Force measurements have been carried out by strain gages attached to a fixed beam to which the cylinder is suspended and flow fields are visualized by dye markers.The results show that since the upper vortex and lower vortex are shed at upper and lower maximum displacement related to the cylinder,the vortex street formed is composed of two parallel rows with an opposite signs of the vortices.With the application of symmetrical Lorentz force the vortexes of cylinder are suppressed which leads to the suppression of lift oscilla-tory and then the vibration of cylinder is suppressed.The vortex street formed is composed of two parallel rows which turn to the composition of negative and positive staggered vortices. When the Lorentz force is large enough,the flow on the surface of the cylinder is accelerated by the Lorentz force which causes the flow separation point to move downstream and disappear grad-ually.Then,the vortexes move downstream away from the cylinder and the vibration amplitude of cylinder begins to decrease.The vortex shedding is suppressed fully and the vibration is decay-ed rapidly.Finally,the vibration is suppressed fully so that the flow field is steady.

Wang Hanfeng, Zhang Yunping, Zou Chao

2014, (1): 31-37. doi: 10.11729/syltlx20130006

Abstract(62) PDF(6)

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The near wake flow and pressure distribution on the rear end of a 25°slant angle Ahmed model were investigated at the Reynolds number of 8.7×105 ,using Cobra probe,pres-sure scanner and oil film flow visualization technique.Deflectors with different width mounted on the top edge and both side edges of the slant were tested to compare their effects on the near wake and aerodynamic drag.It is found that there is a pair of symmetrical tailing vortices in the near wake,which induces strong downwash flow near wake centerline.The strength of tailing vortices is inherently correlated with the pressure distribution on rear end and aerodynamic drag of the model.A stronger trailing vortex corresponds to larger negative pressure on the slant and also a larger aerodynamic drag.The deflectors at both sides of slant with the width of 1% model length result in about 3 .0% increase of the aerodynamic drag.With the width increases to 2% and 3%of model length,deflectors at both sides of slant can noticeably suppress the separation bubble, and result in a drag reduction of 3 .5%and 7 .2%,respectively.Deflector at the top edge prevents flow reattachment on the slant,thus eliminates the separation bubble.It suppresses the tail vor-tices and reduces the aerodynamic drag more effectively than those at the side edges.The drag re-duction rate reaches 9 .3%,1 0 .7% and 1 0 .9% for the top edge deflectors with the width of 1%, 2% and 3% of model length.

Qin Yongming, Tian Xiaohu, Dong Jingang, Zhang Jiang

2014, (1): 38-43. doi: 10.11729/syltlx20130016

Abstract(85) PDF(4)

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The stage separation of tandem layout vehicle is a complicated problem,including shock interaction,flow separation,vortices,and so on.The phenomenon has a great effect on the aerodynamic characteristics and trajectory of the first and the second stage.The study of the stage separation between stages of tandem layout vehicle is indispensable.This paper analyses the aero-dynamic characteristics and flow mechanism of stage separation without j et by the wind tunnel test.The tests are conducted in the FD-06 Sub-Tran-Supersonic wind tunnel in CAAA.The test model is a two stages model with tandem layout,connected to a stage separation support system, the distance and angle between stages are variable in this system.The aerodynamic characteristics of the first and the second stage are obtained by two six-component strain-gauged balances.The aerodynamic characteristics of the first and the second stage are analysed at typical mach numbers when the two stages are in-line.The test results of the second stage and the first stage at Mach number 0.75 and 1.79 are obtained.The results indicate that the two stages interfere each other in the stage distance range of the test condition (L/D≤2 ).As for the second stage,the axial force coefficient(CA)is related to the stage distance,the base drag coefficient(CAB)is negative as a result of “afterbody effect”which is a CFD result of a simple tandem layout,there is low speed vortex flow between two stages,and this phenomenon has a good effect on the separation of the two stages.The normal force coefficient(CN)and the pitch moment coefficient (MZ)of the sec-ond stage is not affected by the stage distance.As for the first stage at different mach number, the forebody axial force coefficient(CAF),CN and MZ are all affected by the stage distance,espe-cially at large angles of attack,where is shaking off the effect of afterbody flow of the second stage,the value of these aerodynamic coefficients increase with the increase of stage distance.

Jiang Tao, Weng Chunsheng

2014, (1): 44-48. doi: 10.11729/syltlx20130024

Abstract(86) PDF(3)

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Pulse detonation engines (PDE)have received the widespread attention in recent years .However,there are many unresolved issues related to PDE.In order to achieve higher lev-els of thrust,PDEs require relatively small times to complete the liquid fuel atomization and mix-ture with air for increasing the operating frequency.But the nozzle location can have influence on the performances of atomization and mixing in PDE.To study the nozzle location on the perform-ances of PDEs,a gasoline/air two phases pulse detonation engine with a diameter of 80mm and a length of 1 .2 m was designed.PDEs consist of a mixing chamber,ignition and detonation cham-ber.And the venturi is installed in the mixing chamber.The experiments of atomization of the liquid fuel and deflagration to detonation transition(DDT)were carried out by installing a nozzle in three different locations of the venturi.The droplet size distribution in front of the ignition chamber measured by Malven laser mastersizer and signals measured by dynamic pressure sensor were analyzed.The experimental results show that the nozzle location has a significant influence on performances of the PDE.When the nozzle is installed in the throat of venturi,PDE has the best gasoline-air distribution(the sauter mean radius is less than 76.9μm),the highest peak pres-sure of detonation wave(3.5MPa at 30Hz,5MPa at 10Hz)and the highest operating frequency (30Hz).When the nozzle is installed in the expanding of venturi,the sauter mean radius is 85. 4μm,the peak pressure of detonation wave is 4.0MPa and the operating frequency is 20Hz. When the nozzle is installed in the entrance of venturi,PDE has the worst gasoline-air distribu-tion(the sauter mean radius is more than 93.4μm),the lowest peak pressure of detonation wave (0.9MPa at 15Hz,1.1MPa at 10Hz)and the lowest operating frequency(15Hz).In the experimental range,the operating frequency can be increased by improving the effect of atomization and mix-ing .The results offer reference for further research of deflagration to detonation transition mech-anism and atomization of the liquid fuel,design and operation of the pulse detonation engine.

Li Binbin, Gu Yunsong, Cheng Keming

2014, (1): 49-54. doi: 10.11729/syltlx20130020

Abstract(47) PDF(4)

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As an important part of the formation of synthetic jet,the design and performance of synthetic jet actuator determine the field and effect of applications of the synthetic jet.How to design it efficiently with low energy consumption to obtain high control efficiency would be the key of synthetic jet applied in active flow control.Based on the study of single beveled orifice syn-thetic jet actuator,the concept of beveled orifices synthetic jet arrays is proposed.The unsteady flow structure of the typical single slit,double slit,the three slit beveled orifices synthetic jet ar-rays were tested with the application of PIV phase-locking technology.And the impact of orifices spacing ratio variation on momentum transport of beveled orifices synthetic j et arrays along the wall was studied.The results showed that,compared with the single slit and double slit,a strong vortex structure and wall-attached jet were formed around the three slit beveled orifices, and the synthetic jet had stronger tangetial velocity distribution along the wall.The flow along the wall was closer to the wall in the three slit beveled orifices,and prone to introduce the high momentum of the fluid in the main flow into the boundary layer at the bottom,thereby increas-ing the underlying energy distribution of the boundary layer.The momentum transport proper-ties of beveled orifices synthetic j et arrays are mainly from the induced effect of vortex and wall-attached jet effect along the wall,and also the flow velocity gradient after coupling of the two effects.Jet orifices spacing ratio is an important parameter affecting the momentum transport of synthetic jet arrays.When the jet spacing ratio is small,the opposite direction of flow velocity gradient difference between the vortex induction and wall-attached j et would be affected by the impact of jet from the first beveled orifice,which will weaken the vortex strength around the sec-ond beveled orifice.Thereby the velocity gradient difference of the flow will be reduced after cou-pling of the vortex and wall-attached jet,which will weaken the induction between the vortex and wall- attached j et .The research results indicated that the beveled orifices synthetic j et arrays will have stronger transporting momentum along the wall when the spacing ratio is 3 .0 .

Jiang Zenghui, Song Wei, Jia Quyao, Chen Nong

2014, (1): 55-59. doi: 10.11729/syltlx20130027

Abstract(76) PDF(3)

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Quantitative analysis for angular oscillation is conducted,from the observations of pitch angleθand yaw angleΨof an aircraft in two flight tests.Aerodynamic parameter identifica-tion for Van der Pol equation θ··= C00θ+Cm1 (1-C11θ2 )θ· and ·Ψ·= C00Ψ+Cn1 (1-C22Ψ2 )Ψ· de-scribing angular oscillations in pitch and yaw planes respectively,and parameter fitting for the outer envelope curve ofθ-t and Ψ-tobservations,are both carried out.The fitting and parameter identification curves of two flight tests both show good agreement with observations,which means the fitting and parameter identification results are both reliable.All the parameters,inclu-ding Cm1 ,C11 ,Cn1 and C22 ,obtained from parameter identification are similar to those from fit-ting of outside envelope curves in the two flight tests,except a bit difference for C11 by the two methods for the first flight test.The conclusion can be made that the angular oscillations in pitch and yaw planes are the typical non-linear negative damping limit cycle type oscillation,for Cm1 ,C11 ,Cn1 and C22 are all positive while C00 is negative.And that indicates the coning oscillation of the aircraft is stimulated by the aerodynamic damping moment of non-linear negative damping limit cycle type.The typical expression of the aerodynamic damping moment of non-linear nega-tive damping limit cycle type can be acquired according to parameters above.Finally,it is verified that dynamic instability of aircrafts can be brought by the aerodynamic damping moment of non-linear negative damping limit cycle type.

Wang Bolan, Zong Xin, Gu Yunsong

2014, (1): 60-64. doi: 10.11729/syltlx20130032

Abstract(89) PDF(9)

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The performance of optics instrument was more and more affected by aerosols while it was flying with the aircrafts in the atmosphere.It has a quite important significance on the evaluation of optics working efficiency on the aircrafts that the particulates moved and pooled in complex vortex flow field.While the estimate of particulates in the atmosphere coupled with complex vortex field is always one of the most important problems in the research of aircraft envi-ronment.The complexity of aerodynamics and the diversity of atmosphere are always the main factors to restrict both the development of experiments and the foundation of the numerical mod-el.The distributing and velocity characteristic of particulate in complex vortex flow field was re-searched for the first time by advanced PIV technology in the low speed wind tunnel.Through the measurement to the vortex after the rudder,the instantaneous picture of distributing particu-late in the flow field was captured by PIV technology.The laser light scaning the flow field and exposure synchronization were used in the experiment.The instantaneous regularities of distribu-tion of particulate in complex vortex flow field was finally acquired through the image post pro-cessing technology,cross-correlation processing and gray scale calculating on original particulate picture.The result shows that by the principle of Mie scattering through the particulates in the laser light,the picture about the instantaneous complex vortex and particulate distribution can be captured by CCD camera,and by this it solved the limitation that traditional measurement equip-ments can not get the information of distribution of particulate in complex vortex flow in real time. In the flow field dominated by vortex,the particulate in atmosphere was pulled by centrifugal force,moved to balance and finally pooled to a loop around the core of the vortex.In this area,the coeffi-cient of particulate density is much larger than it in the free flow,and in the center of the loop,there is few particulate in the core of the vortex.

Xiang Guangwei, Xie Bin, Zhao Zhongliang, Wang Chao, Wang Jie

2014, (1): 65-69. doi: 10.11729/syltlx20130015

Abstract(83) PDF(6)

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The virtual flight test(VFT)is an effective test mothed for getting characteristics of unsteady aerodynamics when aircrafts maneuver and understanding the aerodynamics-move-ment nonlinear coupling mechanism.The virtual flight test balance development is a key tech-nique in 2.4m×2.4m transonic wind tunnel virtual flight test.We developed a special dedicated VFT strain-gauged balance to measure aerodynamic parameters for closed-loop control.The test model is long and thin because of being devided into two parts so that the balance design space is limited,and the design loads do not match extremely.The aerodynamic forces on the model’s two parts get worse due to this division.Computational Fluid Dynamics (CFD)results show that the maximum normal force on the aft half of the model reaches 12000N,and the pitch moment reaches 6000N·m.One challenging thing is that the wind tunnel test requires both measurement of forces on two separated model parts and the movement of those parts rolling around the axial shaft with low-friction together,so the normal balances don′t meet the requirements.Movement and measurement must be fully considered together.Through theoretical analysis and numerical simulation we optimized the structure of the new balance.A new assembly configuration with bearings and mandril is used for the balance,which is inside the balance and could rotate with the model together with low friction.A kind of serial twin-balance with the same earth end in the middle of the balance was measure forces on the two parts of the model.Eight components are set on four groups of sixteen beams which are located in four different sections.Different dimen-sions of these beams are optimized to solve the loads match problem and conflict of measurement and movement of model parts.During the course of design,the balance stress and strain analysis optimization was generated by finite element simulation method using software ANSYS,and then combined bridges are schemed.Both calibration and wind tunnel test results show that the balance meets all requirements of VFT research.

Zhan Huahai, Zhang Xu, L Zhiguo, Yu Wei, Wang Shexi, Chu Weihua

2014, (1): 70-74. doi: 10.11729/syltlx20120093

Abstract(59) PDF(8)

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In this paper the design of a single vector six-component wind tunnel balance cali-bration system is introduced,which can calibrate piezoelectric balances by multi component cali-bration method adopting “unload”weight to accomplish instantaneous loading.The weight hits some rigidity object in movement by invariableness speed to accomplish unloading.And the weight can become a single vector load by adj usting the posture of the balance and hanging posi-tion in loading device.It can meet the need of super high-speed wind tunnel experiments.The structure of the system consists of balance posture adjustment equipment,weight loading equip-ment and unloading equipment by collision,and so on.The measurement system and control sys-tem are also expounded in the paper.Because of the simple structure and less error resources,the equipment has advantages of high efficiency and high calibration precision.

Zhang Yongsheng, Yin Shibo, Liu Dan, Jia Yi, Lang Weidong

2014, (1): 75-79. doi: 10.11729/syltlx20130031

Abstract(68) PDF(3)

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Because the flow around afterbody of transport airplane were badly disturbed by using general rear sting support,ventral support became important support form for transport airplane with upswept afterbody in wind tunnel.There are three kinds of ventral support form according to the support point amount,i.e.single point,double point,and three point.In order to reduce the interference of ventral support,a new-style single point ventral support system was designed and developed in FD-09 low speed wind tunnel based on existing underside high angle of attack experimental mechanism.The single point ventral support system has been completed and put in applications in 2011.It was used for the test of transport airplane and unmanned aerial vehicle.The single point ventral support system was consist of angle of attack mechanism, connecting joint,ventral support sting,and internal balance.The characteristics of single point ventral support system included:the system was simple and convenient for practical application, the model could be easily designed,support interference was relatively stable.The mainly per-formance index of single point ventral support system included:the range of angle of attack was-40°~40°,the range of angle of yaw was-30°~30°,the control precision of angle was 0.05°.The forced transition of support sting was obtained by covering nylon nets.In order to measure sup-port sting interference,the mirror support sting was also covered nylon nets for forced transi-tion.The traditional measurement method for the distances between model and floor was re-placed by a new method using grating ruler displacement sensors in ground effect test,and the control precision and test efficiency were greatly improved.The test results showed that the re-peated test precision of single point ventral support system was higher,and some precision index was close to advanced precision index of GJB.The support interference of single point ventral support system were relatively stable,and the mostly interference were small.The support inter-ference of single point ventral support system was less than the support interference of double point ventral support system.

Jin Xin, Yuan Bing, Zhang Lizhen, Cai Guangping

2014, (1): 80-84. doi: 10.11729/syltlx20120101

Abstract(61) PDF(9)

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The folding wing has a dynamic unfolding process during aircraft&store separa-tion,which will cause an evident change in aerodynamic characteristics of the store and further result in difficulty of acquiring the correct separation characteristics of the store by using the con-ventional CTS WTT (Captive Trajectory Simulation Wind Tunnel Test).Therefore,a new test technique is put forward for CTS WTT of store with folding wing,and emphasizes on obtaining aerodynamic data of stores at wings spreading during aircraft&store separation.The technical difficulty involved in this research is to establish the simulation method for the dynamic unfolding process of the folding wing,and its key point is the formation principle for the aerodynamic force of the store,That is both under the aircraft disturbance flow condition with the same location and same attitude angle.The store’s aerodynamic force when the wing is unfolded is the aerodynamic force when the wing is folded added with the aerodynamic difference of free flow between the un-folding and folding statuses of the wing.The whole testing method and procedure are introduced, and details in the procedures of the ground bench test,the establishment of parameter database for the unfolding status of the folding wing,the establishment of aerodynamic database for differ-ent statuses of the folding wing,and the aerodynamic coefficient corrective calculation of trajecto-ry calculation in relevant tests were presented.Then the technique was validated with the folding wing weapon on some light fighter.

Jia Qing, Yang Wei, Yang Zhigang

2014, (1): 85-88. doi: 10.11729/syltlx20130029

Abstract(100) PDF(5)

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The support interference in wing in ground effect wind tunnel test was simulated and analyzed by using computational fluid dynamics(CFD)method.The wing model NACA 0012 was applied.Because of the symmetry the half model was simulated.The cases of separate sup-port,the wing with and without support were simulated separately.The steady imcompressible Navier-Stokes equation and realizableκ-εmodel were used.The semi-infinite computational do-main was calculated.The floor was divided into two different parts,one part was fixed using wall boundary condition,and the other part was moving belt which was set with moving wall bounda-ry condition.Aerodynamic forces of frame and wing with interference were studied by analyzing the aerodynamic interference and flow field interference separately,and we got some conclusions. On one hand,the interference between frame and wing was increased with the increase of the wing’s attack angle.On the other hand,if the interference from flowing air was ignored,the er-ror for lift force of wing was little while that for drag force was large.Then flows with and with-out interference were compared with each other in order to study the effect of interference on de-velopment of wingtip vortex and flow around the wing.Within the area around 0.5c far from wingtip the wingtip vortex was almost not affected.The viscous flow near steady ground outside the moving belt had some effects on the flow around wing.The current study evaluates the relia-bility of test results in wind tunnel test and provides reference for optimization of wing in ground effect wind tunnel test and experimental investigation of ground effect.

Xie Yan, Li Ping, Jiang Hong, Wang Ruibo, Xue Jiangping

2014, (1): 89-93. doi: 10.11729/syltlx20120182

Abstract(62) PDF(6)

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The three key technical problems have be met in developing the continuous sweep-ing angle of attack test technique in 2 .4m transonic wind tunnel.First,it is difficult to compen-sate transonic flow field quickly and accurately because of persistent disturbances.Second,it is difficult to filter and denoise test data because of overlapped frequencies of useful and disturbing signals.Third,it is difficult to synchronize test data because of actual signals being not synchro-nized.In order to solve these problems,a series of research and development were carried out. Quick compensation of flow field is achieved through optimizing PID(Proportional Integral Differ-ential)parameters and optimizing filter of total and static pressure signals.Test signals are de-noised through a combined filtering approach of software,hardware and wavelet.Cross-correla-tion function is used in synchronizing wind tunnel test signals.Based on above research work,the technique of continuous sweeping angle of attack for force measurement test in 2 .4m transonic wind tunnel was achieved.This technique is verified through J7 calibration model,and test re-sults show that the technical problems mentioned above are solved effectively.Mach number con-trol precision has reached ±0.002 and the results have good consistency with those of conven-tional pitch-and-pause testing mode.

2014, (1): 94-94.

Abstract(48) PDF(4)

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2014, (1): 95-95. doi: 10.3969/j.issn.1672-9897.2014.01.019

Abstract(43) PDF(3)

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2014 Vol. 28, No. 1