Abstract:
The emergence of evacuated tube maglev transportation makes it possible for ground ultra-high-speed rail transit. However, limited by the demand for high-power propulsion motors and low vacuum environment, it is difficult to carry out experimental research. In this paper, the numerical research on the aerodynamic layout of the magnetic track and motor is carried out in the preliminary design of the Dynamic Model Test Platform for Multistate Coupled Rail Transit. Based on the geometric structure of the dynamic model test platform, considering the actual arrangement of the motor platform and the permanent magnet track in the tube, the three-dimensional, compressible RANS method and SST k–ω turbulence model are used to calculate the three-dimensional flow field structure and the shock wave reflection, propagation law of the superconducting maglev train in the low-pressure tube at ultra-high speed. The influence of the rectangular channel on the aerodynamic loads of the train and the flow field in the tube is compared and analyzed. The differences of the pressure and velocity change trend at the bottom of the train, and the shock wave strength at the tail and the wake structure are mainly explored. It is found that the step of the magnetic track and the motor can cause more flow separation and shock reflection in the wake region, resulting in tail pressure fluctuations. When the rectangular channel exists, the shock wave intensity at the tail of the train decreases, the shock wave phenomenon is more obvious, the aerodynamic drag coefficient decreases by 8.855%, and the aerodynamic lift coefficient increases by 14.312%. The research results can provide reference for the design of the magnetic track and motor platform of the multi-state coupling rail transit dynamic model test platform.
[Abstract](17) [FullText HTML](6) [PDF 6180KB](2)
Abstract:
Aiming at the demand of two-dimensional distribution high-resolution measurement of temperature and water vapor concentration in non-uniform scramjet combustion chamber expansion section, advanced tunable diode laser absorption spectroscopy (TDLAS) reconstruction method has been developed. By increasing the number of water vapor absorption lines obtained by scanning the laser wavelength, the number of equations for solving the reconstruction problem correspondingly increased, combining the absorbance equations of all absorption spectra under all laser paths, constructing the optimization objective function with temperature and concentration as unknowns, and using the global optimization simulated annealing algorithm to reconstruct the temperature and water vapor concentration distribution. In the direct-connect scramjet combustion test, the orthogonal optical path layout is adopted, and the square optical mechanical structure with 16 measuring optical paths of 5 horizontal and 11 vertical channels is designed. TDLAS measurement system is assembled, and the time division multiplexed direct absorption detection method is adopted for 5 DFB lasers, with the measurement frequency of 4 kHz. Five water vapor absorption spectral lines (7467.77、7444.36、7185.60、7179.75 and 6807.83 cm) can be obtained at each measured optical path, the system has carried out thermometric validation by using high-temperature furnace on the laboratory, and the temperature measurement deviation is within 2.7%. In the test, the absorption spectrum data synchronously collected under 16 optical paths are processed offline, and the distribution data of temperature field and water vapor partial pressure under various states of ignition, combustion and flameout are obtained. The test results show that TDLAS multi-absorption measurement technology can realize accurate and stable reconstruction, and meet the engineering application requirements of complex combustion flow field diagnosis and bad working conditions.
[Abstract](31) [FullText HTML](6) [PDF 0KB](8)
Abstract:
The cylinder with a pointed spike and the spiked cylinder with aerodome were investigated under the condition of Ma = 2.2 incoming flow using a direct-connected wind tunnel and a high-speed schlieren system. The experimental results were statistically analyzed to investigate the unsteady flow field pulsation of the spiked cylinder under supersonic incoming flow. Based on the transient data, the typical structure and evolution of the flow field were first interpreted. The convergence of the residuals was then used to assess the dependability of the statistical results. Finally, the pulsation characteristics of the flow field were further analyzed in terms of the time-averaged and pulsating flow fields. The results show that there is unsteady pulsation in the spiked cylinder flow field under the condition of the supersonic incoming flow, which is more intense in the case of the cylinder with a pointed spike and attenuated in the case of the spiked cylinder with aerodome, demonstrating the suppression of unsteady pulsation in the flow field by the aerodome.
[Abstract](12) [FullText HTML](4) [PDF 8306KB](2)
Abstract:
In order to study the jet flow characteristics caused by the pitching hydrofoil, a three-dimensional shadow imaging system is utilized to measure the turbulent flow field. By comparing the results of particle image velocimetry, two-dimensional particle tracking velocimetry and three-dimensional particle tracking velocimetry, it is found that the pure pitch motion of the rigid symmetric NACA0012 airfoil at a fixed position in the static fluid would produce weak jets in two directions, accompanied by the generation of small-scale vortices. The results of velocity statistics show that when the amplitude of the hydrofoil rational angle is large, more obvious vortex structure and velocity change are produced. The study obtained the three-dimensional wake structure generated by the pitching hydrofoil movement, and found that there is also a symmetric vortex structure in the depth direction. The results show that the velocity component in the depth direction generated by the pitching hydrofoil movement can not be ignored under the limited airfoil aspect ratio.
[Abstract](24) [FullText HTML](5) [PDF 6477KB](5)
Abstract:
The operation principle of the maglev flight wind tunnel is to drive the model to move at high speed in a closed straight pipe through magnetic suspension. The maglev system is particularly important for accurate control of the acceleration/uniform/deceleration process of the model. This study made a comprehensive analysis of the four maglev systems, including normal conductive electromagnetic suspension (EMS), permanent magnet electrodynamic suspension (PM–EDS), high/low temperature superconducting electrodynamics suspension (HTS/LTS–EDS), and high temperature superconducting pinning levitation (HTS–PL). These several maglev systems were comprehensively analyzed from five aspects. EMS system could not meet the specification requirement of the maximum operating speed Ma=1.0, which could not be used as an alternative maglev system for the maglev flight wind tunnel. Based on the analytic hierarchy process (AHP) and grey relational analysis (GRA), a comprehensive decision-making model of the maglev system was established for the application scenario of the maglev flight wind tunnel. Results show that the HTS–EDS and HTS–PL system have better application potential in the maglev flight wind tunnel.
[Abstract](16) [FullText HTML](7) [PDF 7131KB](4)
Abstract:
On the 300 kW DC axial tube electrode arc heater, the UI characteristics and thermal efficiency of CO2 and air are measured by experiment, and the regression analysis is carried out by using the similarity criterion number. The unified relationship of electric and thermal characteristics that can be applied to the two media is obtained, and compared with similar heaters abroad. The results show that CO2 and air arc heaters have similar electric and thermal characteristics, under the same input parameters (arc current and gas flow); the total pressure of CO2 is 18% lower than that of air, but the arc voltage, enthalpy and thermal efficiency are 5.9%, 6.7% and 10.9% higher respectively; the regression errors of UI characteristics and thermal efficiency are −13.0%～11.4% and −33.0%～34.7% respectively. This relationship plays an important guiding role in the operation and commissioning of the high-power arc heater.
[Abstract](22) [FullText HTML](7) [PDF 7437KB](4)
Abstract:
The controllable dynamic behavior of ferrofluid droplets under the magnetic field can be used to realize directional transport of small droplets or bubbles in microfluidic devices, anti-icing, droplet condensation, mineral flotation and other fields. At present, the mechanism, influencing factors and regulation methods of the field-assisted wetting behavior of magnetic fluid on the superhydrophobic surface are not clear. The wetting behavior and droplet shape evolutions of water-based ferrofluid on a hydrophobic surface under an external magnetic field are studied experimentally. Under the vertical magnetic field, the effects of the magnetic induction intensity and ferrofluid droplet size on the droplet wetting behaviors are investigated, and the contact line diameter and contact angle of the droplet are measured experimentally. The experimental results show that the apparent contact angle of the ferrofluid droplets decreases from above 90° to below 90° under the action of the weak magnetic field. Under the magnetic field, the nanomagnetic particles in the magnetic fluid form a chain structure along the direction of the magnetic field line and the droplet contact angle changes. Through a scaling analysis, the theoretical relationship of the magnetic field and the contact angle is established and it successfully predicts our experimental results. The work is valuable for controlling the wetting properties of the ferrofluid droplets on the solid surfaces under the magnetic field.
[Abstract](40) [FullText HTML](22) [PDF 7495KB](5)
Abstract:
Aviation kerosene is expected to act as the primary coolant of advanced gas turbine engines. In such situations, the aviation kerosene would exist at subcritical conditions near the critical point or even at supercritical conditions. Correspondingly, it is of vital importance to study the nozzle internal flow and jet for the design of engine combustors. This paper focuses on the internal flow characteristics and jet characteristics under trans/supercritical conditions. The review shows that the existing researches of the trans/supercritical internal flow are mainly limited to small-molecular or simple fluids, constant cross-section pipes, and narrow conditional parameters. The location of phase change depends on thermodynamic characteristics, geometric configurations, and injection parameters. The mixing efficiency of the trans/supercritical jet is largely affected by thermodynamic characteristics. However, the research on trans/supercritical internal flow characteristics of hydrocarbon fuel inside constriction nozzle channels and jet characteristics based on relatively complex nozzle configurations remains to be further developed. Accurate thermodynamic models of supercritical aviation kerosene remain to be established. The deformation and breaking mechanism of the jet fluid interface as well as the jet mixing behavior remains to be captured through advanced optical diagnostic methods. The mixing characteristic parameters and their change laws remain to be summarized and described.
[Abstract](52) [FullText HTML](24) [PDF 7524KB](9)
Abstract:
In the conventional “Z” structure schlieren technique, due to the limitation of large-size optical element materials and processing technology, the size of the test field is usually less than 1 meter. In order to show the flow field of a large-scale model in a wind tunnel, the focusing schlieren technique is proposed to show the flow field in the 1.2 m test area. According to the imaging principle, the large size Fresnel lens are replaced by a matrix light source. After solving the key technologies such as the engineering design of large-size light source splicing, the development of large-diameter focusing lens and the production of high-definition imaging screen, two sets of focusing schlieren systems with the test field of view of 1.2 m × 1.2 m were established, and the schlieren images of the hypervelocity flow field with high sensitivity were obtained in the wind tunnel. The flow visualization with larger field is expected to be realized through the splicing of larger size light sources.
[Abstract](31) [FullText HTML](16) [PDF 10196KB](2)
Abstract:
We designed an underwater passive fluidic thrust vectoring nozzle. It can easily generate pressure difference on both sides of the primary jet to deflect the jet only by controlling the valves of the secondary flow channel. However, the nonlinear features in the control law of the thrust vectoring angle such as “sudden jump” and “hysteresis” limit the further application of this technology. In this research, the flow characteristics of the primary jet in different transverse sections of the nozzle were studied by the dye flow visualization technology and particle image velocimetry technology. We discovered flow structures such as shear layer vortices, trailing edge backflow, and separation bubbles. Three-dimensional flow structures were also observed, including the transverse flow in the near-wall region and the corner flow at the joint of two walls. The study of the interaction law between flow structures provides a physical model basis for solving the nonlinear problems such as jump and hysteresis of the thrust vectoring control law.
[Abstract](39) [FullText HTML](20) [PDF 7105KB](5)
Abstract:
The complex flow field structure and the interaction between vortex structures make the flying wing configuration aircraft prone to transverse uncommanded motion at a high angle of attack. To suppress the uncommanded motion, two sets of jet actuators are arranged on the vehicle using two existing active jet control techniques, the control effect of the actuators is verified through wind tunnel force measurement experiments, and the mutual coupling relationship between the two sets of jet actuators is clarified. A virtual flight experiment is conducted in the wind tunnel to capture the uncommanded motion of the flying wing configuration aircraft in the transverse direction, and two methods, PID and deep reinforcement learning, are applied to suppress the uncommanded motion in this kind of highly coupled and nonlinear problem. The wind tunnel experiments show that the deep reinforcement learning method is more effective in controlling the highly coupled and nonlinear problem, and the trained intelligent model can effectively suppress the transverse uncommanded motion of the flying wing configuration aircraft model.
[Abstract](29) [FullText HTML](20) [PDF 8779KB](2)
Abstract:
In flight testing, the aeroengine flight thrust is indirectly obtained by the gas generator method. In order to improve the calculation accuracy of the flight thrust, it is necessary to accurately obtain the characteristics of the exhaust system. The laboratory calibration test and numerical simulation research were carried out by using the large bypass ratio separated exhaust system scale model. The results show that: the core nozzle characteristics obtained by the two methods are consistent, and the values are close. When the maximum core nozzle pressure ratio is 1.44, the deviations of the mass flow and the thrust are 0.73% and 0.18%, respectively; the characteristics of the separated exhaust system obtained by the two methods have the same trend and close values. When the max bypass nozzle pressure ratio equals 1.46, the deviations of the mass flow and the thrust are 0.64% and 0.18%, respectively; when the physical model and geometric model of the large bypass ratio separated exhaust system are reasonably simplified, the characteristic deviations of the separated exhaust system obtained by the two methods are in good agreement.
[Abstract](36) [FullText HTML](21) [PDF 13166KB](2)
Abstract:
In order to meet the requirement of the double-inlet test in the 4 m magnitude low speed wind tunnel, a test method of the double-inlet test in the 4 m × 3 m low speed wind tunnel of LSAI of CARDC was proposed. According to the method, a model is supported by one pole, and each inlet mass flow is simulated and controlled by an ejector with a digital pressure regulating valves system. In this method, the range of AOA is −10°～90°, the range of AOS is −45°～45°, the simulating maximum of the double-inlet mass flow is 2.9 kg/s and 1.4 kg/s. To validate the method, a double-inlet test was completed in the low speed wind tunnel. The test results show: the model is less affected by pipeline aerodynamics. The independent model and ejector support mechanism meets various model support requirements. Double inlets mass flow simulation and control are completely independent, which meets the requirement of studying interactions for double inlets.
[Abstract](79) [PDF 0KB](1)
Abstract:
Considering the fact that there are more error sources in the measured heat flux with thin-film gauges when the two-stage approach is applied to determine the thermal product and resistance-temperature factor, the transfer method is applied to directly calibrate thin-film gauges, in which the thermal product and resistance-temperature factor are treated as the sensitivity coefficients. To get the consistent calibration results of different thin-film gauges fabricated in a batch, the sensitivity coefficients are divided by the resistance-temperature factors of the thin-film gauges, and then the correction sensitivity coefficients are consistent. With the transfer calibration technique, the calibration results of thin-film gauges show a good linearity with a relative expanded uncertainty below 6.5%, which is lower than that reported in other researches, in which the two-stage approach is used to calibrate thin-film gauges.
[Abstract](57) [PDF 0KB](0)
Abstract:
To study the start/unstart phenomenon of the air-breathing vehicle inlet caused by acceleration or deceleration, which also bring the problem of aerodynamic mutation on vehicle, the test technique of continuous varying Mach number was performed in the 1.2 m supersonic wind tunnel based on the flow mechanism of the 2D wedge shock wave. By developing the shock wave generator system, continuous varying Mach number was realized successfully in one wind tunnel test process. And it has also been confirmed that the technical method can make Mach number vary simply and quickly with high quality and precision. Through the flow-field calibration, the quality of the instantaneous flow-field in variable Mach number region meets the standards of GJB eligibly, and thus the field can be used for force and pressure test compliantly. In addition, the test of inlet starting characteristics was carried out and the dynamic process and critical state from start to unstart were captured, which indicates that the test results agree with the numerical simulation accurately. The test technique could provide effective support for supersonic air-breathing vehicle in aerodynamic performance prediction and study.
[Abstract](47) [FullText HTML](14) [PDF 0KB](2)
Abstract:
In cavity structure, complex flows and high-intensity noises appear under the high-speed condition, seriously affecting the aerodynamic characteristics and structural safety of the aircraft. Through the methods of the PIV technology and dynamic pressure measurement, the cavity with a length-depth ratio of 3 to 10 is experimentally investigated in the range of Mach number 0.4 to 0.8. The influences of the length-depth ratio and Mach number on the flow field structure in the cavity are emphatically analyzed, and the correlations between the noise intensity and the flow velocity are revealed. The results show that: as the length-depth ratio increases, the thickness of the shear layer in the cavity increases rapidly and expands into the cavity, leading the impact position on the cavity to move down from the back wall to the bottom, and causing the flow type in the cavity to change from open to closed. The increase of the Mach number inhibits the shear layer from expanding into the cavity and induces the main recirculation vortex to move back and the flow type to be open. The amplitude of the overall sound pressure level is positively correlated with the flow velocity in the back of the cavity.
[Abstract](41) [FullText HTML](26) [PDF 0KB](0)
Abstract:
The PSP（Pressure Sensitive Paint） measurements correction after the wind tunnel test is often implemented through the least square method, and it tends to neglect the chordwise and spanwise flow characteristics on the wing. By employing the IPS （Infinite Plate Spline） technique, a hybrid correction method and process for the PSP test results was presented. A full-span airplane model, with PSI （Pressure Scanner Instrument） and PSP on the upper and lower surfaces of the wing, was utilized to carry out the PSP test and PSI test simultaneously in a transonic wind tunnel with the 2.4 m test section. The pressure measurements were conducted at Ma = 0.735 and in the angle of attack range from −6.38° to 10.59°. Test results indicate that the PSP data demonstrate a good agreement with PSI data. However, the PSP test system with only one camera shows a poor capacity to capture the aerodynamic field at the wing leading edge. The hybrid correction method for PSP results has proven to an effective approach, and the corrected data reveal that the hybrid correction can better deal with the pressure distribution characteristics of the full wing.
[Abstract](47) [FullText HTML](23) [PDF 2455KB](14)
Abstract:
In a conventional hypersonic wind-tunnel, pressure fluctuations of the boundary layer on a 7-degree half-angle sharp cone are measured by surface sensors and are analyzed by the linear stability theory. The influences of unit Reynolds numbers and Mach number on the stability and transition position of the boundary layer are studied. The length of the test model is 800 mm and the radius of the head is 0.05 mm. Test unit Reynolds numbers range from 0.49×10 7m–1 to 2.45×107 m–1. Test Mach numbers range from 5 to 8. The angle of attack is 0°. The transition position and the energy spectrum distribution of the disturbance wave in the boundary layer are obtained by the quantitative infrared thermography and high frequency surface pressure fluctuation measurement techniques. The frequency and growth rate of the most unstable wave are analyzed by using the linear stability theory. The experimental results show that the fluctuating pressure signal with obvious characteristics of the unstable wave spectrum can be measured in the transition region. The frequency of the pressure fluctuation is close to that of the second mode instability analyzed by the linear stability theory, and the amplitude variation trend is also similar to that of the theoretical analysis. With the increase of the unit Reynolds number, the instability appears earlier, the dominant frequency is increased, and the transition onset moves forward. The unstable wave in the boundary layer contains the first and second modes. When the free-stream Mach number is equal to 5, the transition is caused by the first mode, and when the Mach number is above 6, the transition is attributed to the second mode.
[Abstract](42) [FullText HTML](27) [PDF 0KB](13)
Abstract:
For a lifting body model, the boundary layer transition infrared thermogram measurement experiment was carried out in the conventional hypersonic wind tunnel, and the influence of different unit Reynolds number and Mach number on the lifting body boundary layer transition was studied, which was compared with the calculation results of the eN method. The length of the experimental model is 800 mm, the unit Reynolds number is 0.46×107～3.94×107 m–1, the Mach number is 5～8, and the angle of attack is 0°. The transition position and transition front of the boundary layer on the surface of the model are obtained by the large-area infrared thermogram technology. The analysis of the experimental results shows that there are crossflow instability and the second mode transition in the boundary layer of the lifting body. As the unit Reynolds number increases, the crossflow transition effect increases, the temperature rise on the lower and upper surfaces of the model increases, the transition front moves forward, and the transition area expands; as the Mach number increases, the crossflow transition effect gradually weakens and the transition position moves downstream, and the transition area significantly shrinks back. Moreover, the transition N factor at different Mach numbers and unit Reynolds numbers are relatively close, but the N factors of the upper and lower surfaces are different. The lower surface is about 6, and the upper surface is about 2.5. The high-frequency second mode transition occurs in the side edge at high unit Reynolds numbers.
[Abstract](45) [FullText HTML](20) [PDF 0KB](12)
Abstract:
At present, the Sivells method is widely used for the design of the inviscid hypersonic axisymmetric nozzle contour. And then, the nozzle contour viscous correction is performed by solving the axisymmetric momentum equation. This design procedure is validated by nozzles in conventional hypersonic wind tunnels and shock wind tunnels, which are operated under high Mach number and high total pressure conditions. Meanwhile, there are few validation studies of this procedure under high Mach number and low total pressure conditions. In this study, the nozzle design procedure based on the Sivells method is used for Mach 6, 8, 10, and 12 nozzle contour design under the low total pressure condition. Furthermore, in order to analyze nozzle flowfields, numerical simulation and wind tunnel experiment are carried out. It can be found that, the flowfields in Mach 6 nozzle and Mach 8 nozzle are consistent with expectation and the jet flowfields are so good that are suitable for test. In contrast, there are some over-expanded areas in the flowfields of Mach 10 nozzle and Mach 12 nozzle, which results in higher Mach number than expectation in those areas. The jet flowfield quality of Mach 10 nozzle is better than that of Mach 12 nozzle. It can be concluded that, under the condition of low total pressure, the Sivells method still works well for Mach 6 nozzle and Mach 8 nozzle design. Meanwhile, the method is less effective when applied to the Mach 10 nozzle and Mach 12 nozzle design.
[Abstract](74) [PDF 0KB](1)
Abstract:
Droplet spreading on a surface is ubiquitous in a variety of applications including aerospace, industry, and agriculture. Majority of these impacts are oblique, while previous studies focused on orthogonal impacts. Oblique impacts cannot be understood directly by previous theories and/or models. Evolution of film formation following a droplet impacting an oblique surface is investigated experimentally. Evolution of the film shape is obtained under various inclination angles and Weber numbers. Based on a new theory of droplet spreading on oblique surfaces, evolution of the film shape is analyzed. It is found that the film shape at small inclination angles can be predicted reasonably, but the error between the predicted maximum lamella width along the inclination direction and the experimental data is relatively big at large inclination angles since the length of the upstream lamella is assumed as a constant in the theory. Modifications of the theory including more detailed analysis of the length of the upstream lamella lead to an analytical model which permits the theoretical determination of the maximum lamella shape. It is shown that the error between the predicted results and the experimental results can be reduced from 61.8% by the previous theory to 3.2%. The model provides a better prediction on the lamella shape at large inclination angles, and a more concise and accurate theoretical tool for engineering applications.
[Abstract](51) [FullText HTML](20) [PDF 0KB](16)
Abstract:
There is still a shortage of the experimental research of boundary layer transition in compressible flows nowadays due to the difficulty in measuring the turbulence intensity. Aiming at studying the influence of the turbulence intensity on supersonic boundary layer transition, a plate model is tested in a blow-down facility （FL-24y of CARDC） at Mach 3. The turbulence intensity of the flow is changed by adjusting the arrangements in the stabilization section of the wind tunnel, which covers a range from 0.82% to 1.63%. The turbulence intensity is measured by interferometric Rayleigh scattering, while the boundary layer transition is derived by infrared thermography. The CFD simulation of the plate model transition is conducted based on the γ-Reθ transition model. The results show that the transition onset position （Fonset） and transition end position （Flength） obtained by the experiment and the simulation agree well, with the maximum relative error coefficient of 2% in Fonset and of 5% in Flength, which provides support to gain a deeper insight into the boundary layer transition mechanism in supersonic flows.
[Abstract](928) [PDF 0KB](28)
Abstract:
Oblique detonation engine has great potential application in high flight Mach number airbreathing vehicles because of its higher thermodynamic efficiency and smaller size. The research about oblique detonation engine is renewed all over the world in recent years. However, all of the oblique detonation experiments are conducted with hydrogen fuel or ethylene. There is no experimental result about kerosene oblique detonation. In order to examine the application feasibility of kerosene oblique detonation engine, the experimental study on liquid RP3 aviation kerosene oblique detonation engine is conducted in JF-12 shock tunnel and the test time is about 50ms. The difficult issue for the initiation of kerosene oblique detonation is that the ignition delay time of kerosene-air is too long and the autoignition cannot occur in the combustor. A new forced detonation initiation method is put forth to deal with this key issue. The total temperature of JF-12 shock tunnel is 3800 K and the global equivalence ratio is 0.9, which replicates Mach 9 flight-equivalent condition. The steady-state oblique detonation is obtained successfully during the experiments, which demonstrates the application feasibility of kerosene oblique detonation engine.
[Abstract](72) [FullText HTML](42) [PDF 9559KB](7)
Abstract:
As a novel anti-icing technology, superhydrophobic electrothermal coupled surface anti-icing possesses an excel-lent anti-icing efficiency with low energy consumption. Based on the water droplet impact behaviors and the wetting characteristics of the superhydrophobic surface, a prediction model of the heat flow density of superhydrophobic electrothermal coupled surface anti-icing is developed according to the thermal balance theory of the icing surface. The experimental analysis of the superhydrophobic electrothermal coupled surface anti-icing is carried out in a low-speed icing wind tunnel. The results show that the difference between the theoretical anti-icing heat flux and the experimental results is less than 6%, which verifies the prediction model. The analysis of the experimental results and energy consumption shows that the superhydrophobic electrothermal coupled surface anti-icing effectively reduces the energy consumption compared with the electrothermal method. With the freestream velocity of 10 m/s, liquid water content of 1 g/m3, mean volume diameter of 65 μm, and temperature of −15 ℃, the superhydrophobic coating can effectively prevent the formation of backwater due to its wetting property. For dry and wet surface anti-icing, the superhydrophobic electrothermal coupled surface anti-icing method reduces the energy consumption by about 43% and 33% respectively compared with the electrothermal method.
[Abstract](45) [FullText HTML](8) [PDF 7176KB](9)
Abstract:
Planar Laser-induced Fluorescence （PLIF） was used to study flow and mixing characteristics in cross-shaped mixers with four chamber aspect ratios rr=0.5, 1.0, 1.5 and 2.0） at 10<Re<500. Results show that, there are four flow regimes in the mixers with different depths, including the segregated flow, steady engulfment flow, pulsation flow and unsteady engulfment flow. For the steady engulfment flow, the flow field is dominated by three co-rotating vortices for r<1.0, but the center and satellite vortices rotate in opposite directions for r≥1.0. For the pulsation flow, the center vortex shrinks and expands periodically, and the fluid oscillates throughout the chamber for r>1.0. For r=1.0 and 0.5, the shedding of vortex rings emerges downstream. For the unsteady engulfment flow, periodical vortex merging and breakup is observed for r=1.0. For r=0.5, vortex breakup is invisible, and instead, the center vortex merges with a satellite vortex periodically. For r>1.0, the center vortex experiences growth, deformation, and breakup processes. Mixing in cross-shaped mixers was evaluated by the time-averaged intensity of segregation （IOS）, and the mixing mechanism is revealed. An increase in chamber aspect ratios decreases the critical Reynolds number for the engulfment flow and pulsation flow, which causes the mixing enhancement in the chamber at low Re.
[Abstract](57) [FullText HTML](43) [PDF 7957KB](11)
Abstract:
The cloud field calibration of icing wind tunnels usually has the disadvantage of high instrument dependence. To solve this problem, this paper proposes a method for identifying the parameters of cloud fields in icing wind tunnels based on multi-modal fusion. This method takes the icing image of the test model together with the parameters such as the angle of attack, air velocity, air temperature, and icing duration of the model as input, extracts and fuses the two characteristic parameters, and takes the liquid water content （LWC） and the average volume diameter of water droplets （MVD） as the output to train the neural network model. And then the rapid identification of icing cloud parameters is realized. In order to verify the effectiveness and feasibility of the proposed method, the paper takes NACA0012 airfoil icing as the research object, develops the cloud field identification program of the icing wind tunnel, analyzes the influence of the fusion proportion, and obtains the best network model suitable for ice parameter identification. On this basis, simulation and experimental evaluation are carried out. The full scale error of the proposed method for LWC and MVD is less than 12%, which has high identification accuracy and good generalization performance, and can provide an important supplement for the identification of cloud fields in the icing wind tunnel.
[Abstract](55) [FullText HTML](23) [PDF 6746KB](5)
Abstract:
Aerodynamic performances of the axial compressor of the 0.6 m continuous transonic wind tunnel are tested under various pressure conditions, and the Reynolds number effects are studied experimentally. The lowest total pressure of the compressor inlet is about 3 kPa, and the corresponding Reynolds number is approximately 5×104. Test results show that the Reynolds number effects are significant when as Reynolds number is below the critical value, which is 5×105 in the compressor design. Compared to the large Reynolds number condition, the pressure ratio under the low Reynolds number condition reduces rapidly, while the surge margin changes slightly. The mechanical loss of the shaft becomes the major loss of the compressor as the operation pressure drops, and has a significant influence on the compressor efficiency. Additionally, the correlations of the pressure ratio and efficiency with Reynolds number, obtained by data analysis, can offer a useful reference for design and numerical simulation of the axial compressor under the at low Reynolds number condition.
[Abstract](61) [FullText HTML](25) [PDF 7850KB](6)
Abstract:
Study on crossing shock waves/transitional boundary layer interaction in the double vertical wedges configuration was carried out using wind tunnel tests and numerical calculations. The wind tunnel tests were carried out at Φ 600 mm pulse combustion wind tunnel. The Mach number of the free stream condition is 3.0, and the unit Reynolds number is 2.1×106 m−1. The schlieren images, wall pressure and wall heat fluxes were obtained during the tests. The results show that because of the adverse pressure gradient caused by the crossing shock waves, the separation of the laminar boundary layer was captured near the shock waves intersection point. And the transition from laminar to turbulent occurred rapidly in the interaction region. After installation of vertex generator devices or roughness devices, the boundary layer transition position moved to the upstream of the interaction region, the separation was effectively inhibited. And the heat fluxes in the interaction region declined obviously. The peak value of heat fluxe was reduced by more than 25%. The shock wave structures and wall pressure distributions obtained from tests and simulations agreed well, while the prediction heat fluxes were much larger than the test results. The comparison between the calculated results of the transition model and the turbulence model shows that the obviously larger turbulence viscosity is the main reason why RANS methods over-predict the heat fluxes in the unseparated interaction region.
[Abstract](123) [FullText HTML](52) [PDF 6712KB](20)
Abstract:
A kind of the small size Schmidt–Boelter gage was improved for measuring dynamic heat flux in the continuous variable attack angle test in the conventional hypersonic wind tunnel. The Schmidt–Boelter gage improved was statically calibrated and dynamically tested by the heat flux calibration devices. The test results show that the sensitivity coefficient is 57.67 μV·kW−1·m2, the response time is 27 ms, the cut-off frequency is 26 Hz and the gage range coverage is 1–130 kW/m2. Then the quantitative relation between the continuous variable attack angle velocity and the maximum test error was established based on the feature response time constant. And referring to the heat flux measured in the step variable attack angle test, the maximum velocity of the continuous variable attack angle supported by the gage was evaluated within a certain margin of error.
[Abstract](48) [FullText HTML](21) [PDF 7871KB](7)
Abstract:
[Abstract](69) [FullText HTML](37) [PDF 7519KB](4)
Abstract:
The dynamic derivatives are the a necessary parameters in the process of analyzing the stability of the aircraft and designing the control law, in order to meet the demand for obtaining high-precision dynamic derivatives data for large-scale aircraft. Aerodynamics Research Institute of Aviation Industry Corporation of China （AVIC） developed a dynamic derivatives test system with five kinds of oscillations in the 4.5 m × 3.5 m low-speed wind tunnel. The test system uses servo hydraulic swing motor and servo hydraulic cylinder as the driving components of the motion, and directly generates arbitrary waveform motion with the control of the servo valve. The driving mode of the system has the characteristics of small movement transmission gap, high movement control precision, and high automation. The scale of the test model is up to 2.5 m, with the wind speed v =30～60 m/s, the angle of attack α= −36°～36°, and the sideslip angle β= −40°～40°. The verification tests of the dynamic standard model and a wing-body model were carried out, and the test results show that the dynamic derivatives data obtained by the test system is reasonable, the accuracy of the repeatability test data is within 3%, and the test system can provide high-quality dynamic derivatives data for large-scale aircraft.
[Abstract](42) [FullText HTML](27) [PDF 8409KB](5)
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The traditional mechanical method of debugging the double-pass schlieren system exhibits the problems that the fine positioning of the working position of the spherical mirror mechanism cannot be ensured, and the optical paths cannot be completely coincided after passing through the flow field twice in the experimental application in the hypersonic low density wind tunnel. Here, a novel double-pass schlieren system based on visual feedback was developed. The system via absolute encoder instruction control the AC servo motor to adjust the position of the spherical mirror mechanism. Moreover, the pitch and left-right deflection of the spherical mirror can be adjusted by the schlieren image quality evaluation results provided by the machine vision system（visual information feedback）. The position control system of double-pass schlieren parts based on visual feedback realizes the automatic positioning closed-loop control of the double-pass schlieren spherical mirror mechanism, and ensures that the light paths overlap as much as possible after passing through the flow field twice to eliminate ghosting during imaging of the model flow field（the definition of the flow field image is improved by 2.2 times compared with that obtained by the traditional method）.
[Abstract](58) [FullText HTML](42) [PDF 6901KB](10)
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The continuous scan test method for the inlet of airplane was studied in the FL–13 wind tunnel of CARDC. The test methods and procedures were proposed and the test data processing methods were also provided. Inlet tests were performed in the FL–13 wind tunnel to compare the conventional test method with the continuous scan test method. The test results with the continuous scan test method have a good consistency with the conventional test method, which verifies the availability and feasibility of the continuous scan test method for the inlet in the low speed wind tunnel. The research results show that the continuous scan test method can raise the tests efficiency and acquire more test data for the inlet test in the wind tunnel.
[Abstract](79) [FullText HTML](20) [PDF 6978KB](10)
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A kind of total temperature probe with Iridium Rhodium Iridium thermocouple is developed for improving the total enthalpy measurement accuracy. The size parameters of each component are optimized based on the fluid-thermal coupling model of the probe. The reheating rate of the probe is not less than 0.9 after optimization. The calculation and test results show that the temperature of the thermocouple node rises slowly as the temperature of the thermocouple tail and the shielding case rises. This fact results in the temperature of thermocouple node changing according to the measurement time period. So the measurement time period of the total temperature value should be regulated and the total temperature value must be calibrated. Therefore, a comparison calibration method is proposed, in which the total temperature probe used in the supersonic flow field can be traced to the standard calibration device in the subsonic flow field by an arc chamber total probe developed. Finally, the total enthalpy measurement test based on the total temperature probe is carried out in the arc heated wind tunnel. And the uncertainty of the total enthalpy measurement is calculated according to the uncertainty evaluation method based on the precision limit and deviation limit. The test results show that the total temperature probe has a high total enthalpy measurement accuracy. The repeatability precision is about 3% and the uncertainty is 6.4% in this test.
[Abstract](108) [FullText HTML](46) [PDF 10309KB](4)
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Large area surface temperature measurement technology is of great significance in the field of wind tunnel temperature measurement. In order to meet the needs of measurement for higher surface temperature, it is urgent to develop new temperature sensing materials and new temperature measurement technology. Temperature measurement based on the fluorescence intensity ratio of the thermal coupling energy levels of rare earth ions is a new temperature measurement technology. In this work, a temperature sensitive luminescent material （YAG: Dy3+） was synthesized. The corresponding relationship between the temperature and the ratio of emission intensity of the thermal coupling energy levels of rare earth Dy3+ ions （4F9/26H15/2, 4I15/26H15/2） was investigated in the temperature range from 50 to 1000 ℃. Based on this material, a comparative experiment of two temperature measurements that the fluorescence intensity ratio measurement and the infrared thermometer is carried out. It is shown that the measurement results of the two technologies have a high degree of agreement, which proves that the temperature sensitive luminescent material （YAG: Dy3+） can be used for temperature measurement in the range of 50–1000 ℃.
[Abstract](60) [FullText HTML](32) [PDF 9730KB](10)
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The development of the wingtip vortex is an important factor for the flight safety and airport efficiency of the aircraft landing on the runway. The near-field characteristics of the wingtip vortex mainly determine the vorticity of the vortex in the landing phase. In this paper, a simplified model of A320 is used as the object to observe the near-field configuration of the wingtip vortex in a low-speed tunnel of 1 m × 1 m. It is found that the horizontal tail vortex rotates around the wingtip vortex, and the rotational angular velocity in different flow stations is different. By comparing the simulation results, it is found that the rotational angular velocity of the horizontal tail vortex around the wingtip vortex is basically consistent with the experimental results, indicating that the development of the wake vortex under different Reynolds numbers has certain similarity in the characteristics of the rotational angular velocity between two vortices.
[Abstract](104) [FullText HTML](55) [PDF 9140KB](5)
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As one of the most commonly used ultrasonic guided waves, Lamb wave has the characteristics of concentrated energy, wide propagation range and small probe volume. Its application can be extended to the field of ice detection. In order to explore the propagation law of Lamb wave in ice, this paper constructs a physical model based on the Lamb wave propagation research platform of piezoelectric ceramics, and takes COMSOL Multiphysics software as the calculation tool to simulate the propagation of Lamb wave in ice with different thickness and length. On this basis, the Lamb wave ice detection platform was built, and the Lamb wave propagation experiment of the iced aluminum plate was carried out. Combined with the numerical simulation and experimental results, the effects of temperature, ice geometric characteristics and liquid water on the propagation characteristics of Lamb wave are clarified. The results show that the lower the temperature, the faster the group velocity of Lamb wave propagation; In a certain range of ice thickness, the attenuation of piezoelectric voltage amplitude at the receiver increases with ice thickness; The time delay of Lamb wave B1 mode wave at the receiver increases linearly with the increase of ice length; Liquid water only affects A0 mode of Lamb wave, but has little effect on S0 mode. The experimental and numerical simulation results are in good agreement, which provides a theoretical reference for Lamb wave ice detection technology.
[Abstract](108) [FullText HTML](56) [PDF 7607KB](7)
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Ice accretion detection is an important means to ensure flight safety and an important issue in the field of aircraft anti-icing. In this paper, the method of identifying the boundary between the ice surface and the interior is discussed by using the infrared thermal wave detection technology. With a flash infrared thermal wave detection system established, regular ice accretion samples and ice accretion samples with internal boundary were made, the ice accretion detection experiments were carried out, and the data of the infrared thermal wave sequence were collected. In addition, the traditional algorithm based on the first-order differential operator and the second-order differential operator was exploited for processing the ice edge. A new boundary recognition method was proposed as well, which combined the gauss-Pierre-Simon Laplace pyramid algorithm and the area filtering algorithm. Then, the feasibility of the proposed algorithm to identify the boundary of the ice accretion surface was discussed and compared. The experiments and the image data processing methods show that the traditional algorithm can successfully recognize the outer boundary of ice accretion, but can not accurately recognize the internal boundary of ice accretion. The new fusion algorithm can effectively recognize the ice edge and the internal boundary, but the image noise is higher than that of the traditional algorithm. It can be concluded that the new fusion algorithm has some advantages in the detection of the irregular icing surface, and it is expected to provide a new research idea for icing detection in the field of aircraft anti-icing.
[Abstract](129) [FullText HTML](59) [PDF 8342KB](8)
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To achieve wide temperature domain, high precision and ultra low dew point in-situ on-line measurement in the cryogenic wind tunnel, a technology based on the laser absorption spectrum is developed. In the method, the principles of laser absorption spectroscopic technology for dew point measurement are analyzed firstly. Then the absorption spectroscopic selection, spectral parameter calibration and spectral signal processing are provided. The experiments are carried out on the low temperature platform and in the 0.3 m cryogenic wind tunnel, which are compared to the chilled-mirror dew-point hygrometer measurement. The experimental results show that the developed technology can achieve wide temperature domain, high precision and in-situ on-line dew point measurement. The measurement range is from –100 ℃ to 30 ℃, the error is less than 1 ℃, and the time is less than 1 s. It can be used for ultra low dew point in-situ on-line measurement in the cryogenic wind tunnel.
[Abstract](200) [FullText HTML](98) [PDF 9052KB](15)
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For the problem of the monolithic fairing separating from a hypersonic test demonstrator in a high dynamic situation, the reverse-thrust jets simulation method and wind tunnel force test model design have been developed, to meet the requirements of simulating the jets interaction effect and separation distance influence in the hypersonic wind tunnel. The fairing’s aerodynamic characteristics, including the jets interaction effect and the separation distance influence, were obtained by the strain balance in circumstances where the Mach number of the free-stream was 5 and the dynamic pressure was 33 kPa. The study indicates that the jets interaction effect dominates fairing’s aerodynamic characteristics in the separation process. The maximum coefficients’ variation of the normal force, axial force and pitching moment are 44.5%, 32.4% and 198.6% respectively. The pressure center moves forward obviously, making the fairing with designed static stability presents un-stability features in the minus attack angles. The influence of the separation distance on fairing’s aerodynamic characteristics becomes weaker as the separation distance increases. Using a small positive angle as the initial separation attack angle is helpful for the fairing maintaining a stable attitude, benefitting separation security during the separation process.
[Abstract](424) [FullText HTML](146) [PDF 9375KB](30)
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The combination of bionics and drag reduction technology has opened up an important research direction in the field of drag reduction, and has made a significant breakthrough. For better implementation to reduce the wind resistance effect, this paper studies the composite micro-nano drag reduction structure, according to the principle of bionics, through CFD simulation combined with the laser micro-nano fabrication technology. A combined model of drag reduction structure wad established. The flight vehicle air sensor head surface with bionic sharkskin composite micro-nano structures was manufactured by laser interfernce scanning on the basis of the bionic sharkskin scale structures, to further improve the drag reduction performance. Through the parallel simulation and wind tunnel test, the drag reduction mechanism was theoretically analyzed, and the composite structures were manufactured with a drag reduction rate of up to 10.3%.
[Abstract](75) [FullText HTML](95) [PDF 7426KB](5)
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Proper Orthogonal Decomposition（POD） is a reduced order modeling（ROM） method based on 2nd-order statics, which simplifies the investigated wind-pressure field in a new coordinate system formed by a set of orthonormal basis. This paper suggests a method of bi-weighted POD（which weights POD by area and at the same time by root-mean-square）, and applies this method to the modal reduction of pressure field around buildings. Firstly, we introduce the POD expansion in a mean-square method, which demonstrates that POD is the optimal choice of ROM in the mean-square sense. Furthermore, we modify the original POD by the bi-weighting-method to improve its capacity of identifying coherent structures with lower energy in pressure field. For the last part, the validity of bi-weighted POD is roughly examined by a case study which applies the method to the pressure field of a 5∶1 rectangular cylinder. It turns out that the modified POD method improves the ROM accuracy at the area associated with lower energy in a significant way. In the meantime, a wind-pressure field ROM constructed by bi-weighted POD captures vital information provided by the original wind-pressure field and is spatially accuracy-consistent.
[Abstract](176) [FullText HTML](69) [PDF 15431KB](19)
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Despite the decisive influence of various vortex structures of a jet in crossflow on the jet trajectory and scalar mixing, there are few studies related to the high-frequency dynamic characteristics of shear-layer vortexes during transportation. This paper focuses on the high-frequency flow field characteristic, the scalar concentration distribution and the formation and collapse process of the turbulent microstructure of the jet in crossflow with different nozzle diameters and velocity ratios using 40 kHz particle image velocimetry（PIV） and 20 kHz acetone planar laser induced fluorescence（PLIF）. The experimental measurements of the velocity and scalar field show that: increasing the velocity ratio promotes the expansion of the circulation zone behind the jet; in the near field of the jet trajectory, power law fitted velocity distribution and shear-layer vortex trajectory shows an exponentially decrease of the jet velocity, the shear-layer vortex strength and vortex motion frequency also show a downward trend, with the frequency of the shear-layer vortex on the windward side slightly lower than that on the leeward side; as the jet velocity increases, the frequency of the shear-layer vortex increases gradually, but the Strouhal number decreases.
[Abstract](173) [FullText HTML](65) [PDF 6260KB](10)
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During the reentry process of the miniaturized reentry vehicle, small asymmetry of its shape can be produced due to surface ablation, resulting in a small rolling moment. In order to obtain the high-precision micro-rolling moment measurement data of the ablation model of the miniaturized reentry vehicle in the hypersonic wind tunnel, and obtain the other five component aerodynamic data, a six component micro-rolling moment gas bearing balance was developed. The rolling moment design load of the balance is 0.02 N·m, and the axial force design load is 200 N, which are orders different from each other. The overall force measurement scheme of “4+2” balance is proposed, where the four component main balance elements cooperate with the two-component Mx-X elements to complete the extremely mismatched six component aerodynamic measurement. The results of the static calibration and the wind tunnel test show that the balance has good resolution and strong anti-interference ability, and is little affected by temperature. The measurement results of the rolling moment coefficient reach the order of 10–7. The developed gas bearing balance is little affected by the temperature and can be reused. It can measure the six components of aerodynamic data including the micro-rolling moment at the same time, which greatly improves the test efficiency and reduces the error caused by model disassembly.
[Abstract](154) [FullText HTML](59) [PDF 6600KB](4)
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In low speed flutter tests, flutter models with small damping modes start continuous vibration usually at low speeds without obvious flutter divergence. Therefore, it’s of some uncertainty on determing the critical flutter wind speeds by visual inspection or by “damping method （DM）” of modal parameter identification. Considering the similarity between the vibration phenomenon of a small damping modal flutter test and that of a fighter buffet test, a technique named “amplitude turning point method （ATPM）” similarly to that used in identifying buffet boundaries is proposed to determine the critical flutter wind speeds. The method is based on RMS of vibration amplitudes, the curves of normalized vibration RMS changing with wind speeds are drawn, and critical flutter wind speeds are determined according to the first turning points of curves. In a small damping modal flutter test, the method was applied in the test data processing of the engine hangers with variable parameters. Comparing the ATPM results with the DM results and the numerical results, the following conclusions are made: the results of three methods are in agreement, the ATPM results are more similar to the numerical results than the DM results, and the ATPM is concise and reliable, with good stability and applicability.
[Abstract](1928) [FullText HTML](102) [PDF 6347KB](24)
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In order to obtain the underwater drag reduction performance of the biopolysac-charide solution, the drag reduction characteristics of four biopolysaccharide solutions of guar gum, xanthan gum, tragacanth gum and locust bean gum were tested in the gravity circulating water tank experimental system. The influence law of the injection rate, Reynolds number and injection mass fraction on the drag reduction is shown. The results show that the four biopoly-saccharide solutions have significant spray drag reduction effects, and the locust bean gum solution has the highest drag reduction rate（14.3%）. At a constant Reynolds number, with the increase of the injection rate, the drag reduction rate of each polysaccharide solution increases significantly, and shows different trends after reaching the peak value of drag reduction. The drag reduction effect of the polysaccharide solution is better when the Reynolds number is small （<2.0×104）. With the increase of the Reynolds number, the drag reduction law of the polysac-charide solution shows differentiation. Excessive injection mass fraction would reduce the drag reduction effect of the polysaccharide solution, and increasing Reynolds number would cause the phenomenon of “peak shift” with the increase of the mass fraction. By introducing relative injection mass fraction, the effects of the injection rate, Reynolds number and injection mass fraction on drag reduction are coupled with each other. With the increase of relative injection mass fraction, the drag reduction rate of each polysaccharide solution increases first and then decreases. Finally, based on the injection spray mass fraction, the drag reduction law of the polysaccharide solution was explained preliminarily.
2023, 37(1): 1-2.
Abstract(39) HTML(20) PDF(17)
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2023, 37(1): 1-2.
Abstract(54) HTML(13) PDF(18)
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2023, 37(1): 1-8. doi: 10.11729/syltlx20220131
Abstract(73) HTML(25) PDF(20)
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After several generations of technological innovation and development, China's rail transit has made remarkable achievements. The adhesion of the wheel rail system limits the further high-speed development of rail transit, and the application of the maglev technology in rail transit arises at the historic moment. During the 13th Five-Year Plan period, China began to develop the high-speed maglev transportation system with a speed of 600 km/h. The operational speed of the high-speed maglev train is 600 km/h and the Mach number reaches 0.49, and the aerodynamic performance of the train deteriorates sharply. Due to the change of the operational environment (train-rail gap, throttling on both sides) of the high-speed maglev train, the aerodynamic characteristics of the high-speed maglev train are different from that of the high-speed wheel-rail train. The aerodynamic problem has become one of the key issues in the design and development of the high-speed maglev train. In the present paper, the technical challenges faced by the aerodynamic design of the high-speed maglev train are discussed, and the solutions of the aerodynamic design of the high-speed maglev train are proposed. Then the aerodynamic design schemes of the China’s 600 km/h high-speed maglev train are introduced, and the future research fields of aerodynamics of the high-speed maglev train are prospected.
2023, 37(1): 9-28. doi: 10.11729/syltlx20220084
Abstract(90) HTML(45) PDF(17)
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Based on the SST $k-\omega$ turbulence model and the IDDES method, a three-dimensional numerical model was used to simulate the transient state of an evacuated tube maglev transport system at 800 km/h in the choked (blockage ratio of 0.3 and 0.2) and unchoked (blockage ratio of 0.1) states. The accuracy of the numerical method was verified using transonic wind tunnel bump test data. Additionally, the significant coherent structure of the flow field was extracted based on the proper orthogonal decomposition, the region with the strong unsteady load on the train surface was identified, and its space-time evolution law was revealed. The results show that the load distribution on the upper surface of the train is similar to that of the Laval nozzle, and the difference in load distribution between the chocked/unchoked conditions is mainly in the divergent section. The load distribution on the lower surface of the train becomes complex due to the abrupt change in the cross-section of the bogie cavity. The difference in the interaction between the upwash and downwash flow leads to different locations of the temperature peaks under the choked/unchoked conditions. The strong unsteady pressure region on the train surface is mainly located at the bottom bogie and has a characteristic frequency of 14 Hz. The tail car shock wave is also an unsteady source under choked conditions. The first-order modes of the middle and tail car temperature loads reflect the heat accumulation process.
2023, 37(1): 29-35. doi: 10.11729/syltlx20220053
Abstract(120) HTML(44) PDF(15)
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Adding the aeronautic wing to the high-speed train equivalently reduces its weight through the lift force provided by the wing. Hopefully, the energy consumption of the high-speed train can be reduced. This provides a new concept for the high speed train design. The aerodynamic characteristics of the wing directly affect the weight reduction effects. Therefore, it is important to analyze the aerodynamic characteristics of the wing under different conditions for the design of the train lift wing. The kε model was used in this study for numerical simulation. Firstly, the influence of the connection rod between the wing and the train roof on the aerodynamic characteristics of the lift wing was analyzed. On this basis, the effects of design parameters such as the wing-roof height, the incoming flow velocity and the angle of attack on the aerodynamic characteristics of the wing were studied. The results shows that: the influence of the connection rod on the lift and drag of the wing is less than 3.7%. Due to the high-speed airflow induced by the leading edge of the train roof model, the air velocity impacting on the lift wing decreases with the increase of the flying height of the lift wing, and the lift force tends to decrease. Within 3 times of the chord length height, the maximum lift difference of different lift wings does will not exceed 3%. When the velocity of the incoming flow is up to 90 m/s and larger, the lift coefficient and the drag coefficient of the lift wing were close to near 1.62 and 0.61, respectively. As the angle of attack varies within 0° to 22°, the lift coefficients of the wing increase continuously. However, the lift coefficients decrease when the attack angle is above 22°.
2023, 37(1): 36-43. doi: 10.11729/syltlx20220096
Abstract(47) HTML(11) PDF(14)
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When a high-speed train enters a tunnel, an initial compression wave occurs and radiates to the outside of the tunnel to form a micro pressure wave when it propagates longitudinally along the tunnel to the exit. An experimental device for generating the initial compression wave by the instantaneous release of high-pressure air was built, and the experimental research on the compression wave generated by it was carried out. Firstly, the composition of the experimental device was introduced, and the pressure time history curve and formation mechanism in the tunnel were analyzed. Secondly, the influence of the parameters of the experimental device on the initial compression wave was drawn out. The subsequent attenuation process of the compression wave was studied at last. The experimental results show that the pressure fluctuation in the tunnel is mainly affected by the reflected wave at the tunnel portal. The amplitude, gradient and positive peak value of the initial compression wave can be adjusted by changing the relevant parameters of the experimental device. The attenuation period of the compression wave is the same under different initial pressures of the high-pressure chamber, but the larger the initial amplitude is, the faster the pressure decays in the same time period.
2023, 37(1): 44-52. doi: 10.11729/syltlx20220120
Abstract(22) HTML(10) PDF(9)
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The different pressure fluctuation caused by a high-speed train passing through tunnels of various length can cause different degrees of pressure comfort problems for passengers. The one-dimensional compressible unsteady non-isentropic flow model characteristic line method and the time constant method pressure tightness index model were used to study the pressure wave outside the train and the pressure change characteristics inside the train under two pressure tightness indexes when a single high-speed maglev train passes through the tunnel. The concept of the critical tunnel length of the high-speed maglev single line based on the pressure comfort standard was improved, and the influence of the train speed and train dynamic pressure tightness index on the critical tunnel length was studied. It is found that: under the condition of critical tunnel length based on the maximum negative pressure value of the external pressure, the maximum negative pressure value of the internal pressure is smaller. The maximum value of the maximum pressure change in each 1, 3, 10 and 60 s in the train increases first and then decreases with the increase of the tunnel length, and there is the critical tunnel length under pressure comfort constraints. The critical tunnel length at different train speeds is different. Except for per 10 s limit conditions, the critical tunnel length under different train dynamic pressure tightness indexes is approximately the same. When a 600 km/h single-train maglev train with a dynamic pressure tightness index of 83 s passes through a 100 m2 tunnel, the critical tunnel length based on the UIC660 comfort standard is 10–12 km. The research results of this paper have good reference value for the study of tunnel clearance area and train pressure tightness based on comfort standard, and for further improvement of the theoretical system of the critical tunnel length of the rail transit based on the tunnel pressure wave effect.
2023, 37(1): 53-63. doi: 10.11729/syltlx20220110
Abstract(42) HTML(24) PDF(17)
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When the high-speed maglev train enters a tunnel, a compression wave generated by it will induce the air flow at the exit of the tunnel to form an adjoint velocity. The three-dimensional, compressible, unsteady calculation method is used to simulate the process of the high-speed maglev train passing through a tunnel with different blocking ratios under different speeds. The features of the slipstream around the tunnel exit induced by the compression wave are analyzed, and the influence of the train speed and blocking ratio on the slipstream is ascertained. The results show that at the tunnel exit, the trend and the peak speeds of the slipstream induced by the compression wave have no apparent change in the direction of the train’s movement; the peak wind speed of the measuring point outside the tunnel exit gradually decreases in the longitudinal range of 25 m, and basically remains unchanged in the transverse range of 5 m. With the vehicle speed and blocking ratio increasing, the peak wind speeds inside and outside the outlet are raised obviously. When the train speed is 600 km/h and the blocking ratio is 17.04%, the maximum wind speed at 5 m outside the tunnel is up to 56 m/s. This conclusion is helpful to strengthen people's understanding about the harm of the slipstream induced by a train passing through a tunnel, and to provide references for protection against the slipstream in the railway tunnel and the safe operation of maglev train in the future.
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2016, 30(1): 28-42.   doi: 10.11729/syltlx20150069
[Abstract](268) [PDF 6594KB](37)

2016, 30(4): 7-13.   doi: 10.11729/syltlx20150112
[Abstract](312) [FullText HTML](97) [PDF 10428KB](9)

2016, 30(1): 1-14,27.   doi: 10.11729/syltlx20150159
[Abstract](356) [PDF 6603KB](17)

2017, 31(2): 1-11.   doi: 10.11729/syltlx20160129
[Abstract](454) [FullText HTML](160) [PDF 7434KB](60)

2016, 30(1): 81-90.   doi: 10.11729/syltlx20150037
[Abstract](229) [PDF 4386KB](13)

2016, 30(2): 67-74.   doi: 10.11729/syltlx20150091
[Abstract](396) [PDF 5952KB](24)

2018, 32(1): 64-70.   doi: 10.11729/syltlx20170099
[Abstract](356) [FullText HTML](212) [PDF 9648KB](24)

2016, 30(5): 55-60.   doi: 10.11729/syltlx20160026
[Abstract](370) [FullText HTML](167) [PDF 7981KB](13)

OH和CH2O平面激光诱导荧光（PLIF)同时成像技术在研究火焰结构和燃烧反应中间产物二维分布等方面能够发挥重要作用。OH的分布被用来表征火焰反应区的结构，而CH2O的分布则被用来显示火焰预热区的分布。利用OH和CH2O PLIF同时成像技术研究了甲烷/空气部分预混火焰的结构。从实验系统、光路调节、时序同步、OH A-X（1，0)扫谱、数据采集和处理等方面讨论了PLIF同时成像技术的实验方法。实验结果表明，OH和CH2O PLIF同时成像能够分别呈现甲烷/空气部分预混火焰反应区和预热区不同形状的瞬时结构；由于反应区在相邻位置的结合，在火焰中能够局部生成新的分裂的预热区。
2017, 31(5): 39-45.   doi: 10.11729/syltlx20170002
[Abstract](145) [FullText HTML](88) [PDF 10066KB](3)

2017, 31(3): 38-45.   doi: 10.11729/syltlx20170015
[Abstract](330) [FullText HTML](164) [PDF 1175KB](11)

2022, 36(2): 49-73.   doi: 10.11729/syltlx20210110
[Abstract](2914) [FullText HTML](445) [PDF 8779KB](445)
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