[Abstract](0) [PDF 0KB](0)
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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](4) [PDF 0KB](0)
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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](2) [FullText HTML](2) [PDF 0KB](0)
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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](2) [FullText HTML](2) [PDF 0KB](0)
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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:
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](2) [FullText HTML](2) [PDF 0KB](0)
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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.
[Abstract](2) [FullText HTML](2) [PDF 0KB](0)
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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](2) [FullText HTML](2) [PDF 0KB](0)
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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](3) [PDF 0KB](0)
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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](1) [FullText HTML](2) [PDF 0KB](0)
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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](7) [PDF 0KB](0)
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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](18) [FullText HTML](7) [PDF 9559KB](2)
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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](15) [FullText HTML](5) [PDF 7176KB](3)
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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](16) [FullText HTML](6) [PDF 7957KB](1)
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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](16) [FullText HTML](4) [PDF 6746KB](3)
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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](22) [FullText HTML](6) [PDF 7850KB](3)
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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](45) [FullText HTML](10) [PDF 6712KB](10)
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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](24) [FullText HTML](18) [PDF 7871KB](3)
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[Abstract](25) [FullText HTML](16) [PDF 7519KB](3)
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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](40) [FullText HTML](14) [PDF 7410KB](2)
Abstract:
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°.
[Abstract](19) [FullText HTML](12) [PDF 8409KB](4)
Abstract:
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](37) [FullText HTML](16) [PDF 6901KB](7)
Abstract:
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](39) [FullText HTML](8) [PDF 6978KB](6)
Abstract:
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](36) [FullText HTML](15) [PDF 10309KB](2)
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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](38) [FullText HTML](19) [PDF 9730KB](8)
Abstract:
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](47) [FullText HTML](24) [PDF 9140KB](3)
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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](76) [FullText HTML](35) [PDF 7607KB](5)
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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](84) [FullText HTML](33) [PDF 8342KB](6)
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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](133) [FullText HTML](57) [PDF 9052KB](12)
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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](228) [FullText HTML](86) [PDF 9375KB](25)
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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](48) [FullText HTML](45) [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](107) [FullText HTML](35) [PDF 15431KB](16)
Abstract:
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](111) [FullText HTML](35) [PDF 6260KB](8)
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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](100) [FullText HTML](35) [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](185) [FullText HTML](73) [PDF 9212KB](15)
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Flow past an open cavity has been studied extensively, while less attention has been paid to the effects of confinement due to side walls, which produces rich flow dynamics and noise characteristics. In this study, the effects of confinement on flow structures and noise radiation in three-dimensional rectangular cavities are investigated experimentally. The length and depth are fixed, and five ratios of width/length (W/L=0.1–0.5) are considered. The measurements are performed in an acoustic wind tunnel. The pressure oscillations are onset after the wind speed is greater than Ma 0.03. Once the wind speed is greater than or equal to Ma 0.20, the flow and noise radiation are dominated by the self-sustained oscillations corresponding to the second Rossiter’s mode. Furthermore, the present experiments show that the local pressure oscillations and noise radiation of this frequency can be weakened or even eliminated when W/L is equal to or less than 0.3 for the wind speeds of Ma 0.20 and Ma 0.25. The upstream OASPLs in the far field can be reduced by more than 3 dB when W/L decreases from 0.4 to 0.3 at Ma 0.20. By analyzing the surface pressure and TR-PIV（Time-Resolved Particle Image Velocimetry） results, it is found that the suppression of the tonal noise is closely related with the changes of the primary recirculation and some secondary vortical structures by decreasing W/L. In particular, the intensity of the primary recirculation is greatly weakened with strong confinement effects, and the feedback process is not strong enough to produce self-sustained oscillations.
[Abstract](179) [FullText HTML](45) [PDF 6680KB](16)
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Compared with the conventional pressure atomization and pneumatic atomization, effervescent atomization has the advantages of high efficiency, economy and environmental protection, which has attracted the attention of various fields. In this paper, a series of experiments were carried out on a variable nozzle internal swirling effervescent atomizer, and the effects of the working parameters, hole structure and mesh number on the flow and spray characteristics were discussed. The results show that flow characteristics of the atomizer with different hole structures are basically the same. The change of the air-liquid ratio results in the change of the liquid mass flow rate under the same working pressure. Flow characteristics are not affected by the cutting screen, while the addition of the cutting screen induces 3%–7% attenuation of the spay mass flow rate of under the same working condition, and the smaller the cutting screen aperture is, the greater the reduction of the spray mass flow rate is. The distribution of spray particles presents a single peak structure, and the median mean diameter of the spray decreases with the increase of the working pressure or air-liquid ratio. Under the same spray energy consumption, the special-shaped hole structure is more helpful to improve the spray performance. The cutting screen is beneficial to the spray performance, but the choice of the mesh aperture should be made based on the atomizer structure, the working pressure and other comprehensive judgment; in addition, the addition of the cutting screen reduces the axial velocity of the spray mainstream to a great extent.
[Abstract](1873) [FullText HTML](73) [PDF 6347KB](17)
Abstract:
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.
[Abstract](328) [FullText HTML](51) [PDF 6671KB](6)
Abstract:
A kind of spherical water-cooled calorimeter including the ball crown calorific body and the heat shield is developed. The fluid-thermal coupling model of the calorific body and test water is established. The water temperature distribution characteristics in the waterway and the influence of water temperature on the heat flux measurement result are analyzed based on the model and heat flux calibration test. The results show that the closer the water in the waterway is to the heated surface, the higher the water temperature is and the greater the radial temperature gradient is. The smaller the water mass flow rate is, the greater the radial and axial temperature gradients are. So the thermocouple should be kept away from the heated surface and closer to the central axis of the waterway when the water-cooled calorimeter is designed. And the water-cooled calorimeter should be calibrated and the appropriate water mass flow rate range needs to be determined before use. Finally, the test results show that the spherical water-cooled calorimeter can be used to measure the stagnation point heat flux accurately in the long-term arc-heated wind tunnel test with multiple heat flux states.
[Abstract](289) [FullText HTML](114) [PDF 7140KB](21)
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This article reports our recent experimental study of airfoil flow separation control by flexible serrated trailing edge. The experiments were conducted in a straight-type wind tunnel and a hot-wire anemometer was used to measure the velocity profile downstream of the two-dimensional airfoil. Multi-scale coherent structures within the separated shear layers are analyzed both in the time and frequency domains. The results show that the separation bubble thickness decreases by almost 5% of the chord length, the flexible serrated trailing edge vibrates and deforms adaptively and absorbs nearly 20% of the trailing edge shear layer’s energy, perturbation transmits to the leading edge shear layer, and thus the power spectral density decreases significantly in the lower and larger bandwidth to reduce the noise. The coherent structures’ frequency and amplitude also decrease notably, breaking and inhibiting the large vortex package’s transmission obviously in the separation bubble.
2022, 36(5).
Abstract(14) PDF(7)
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2022, 36(5): 1-7. doi: 10.11729/syltlx20210139
Abstract(37) HTML(11) PDF(13)
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Aiming at the problems such as high pressurization ratio, wide suction range and multi-stage parameter matching in large hypersonic wind tunnel, this paper carried out the design calculation and suction test of the multi-stage ejector system in Φ1.2 m hypersonic wind tunnel. Through single stage performance debugging and multistage combination performance debugging of the first, second and third stage ejector without wind tunnel main flow, the suction performance of three stage multi nozzle central ejector with different working parameters is obtained. The minimum static pressure in the test section is 660 Pa without the main flow. And the parameter matching principle of efficient operation of the multi-stage ejector is analyzed and summarized. The test results of the ejector coefficient with the main airflow in the wind tunnel are in good agreement with the theoretical calculation values, which verifies the feasibility of the aerodynamic design method for the multistage multi-nozzle ejector. The design results are reliable, which can provide technical reference for the design of the multistage ejector system in the hypersonic wind tunnel or other ground aerodynamic test equipment.
2022, 36(5): 8-15. doi: 10.11729/syltlx20210196
Abstract(28) HTML(9) PDF(10)
Abstract:
Icing wind tunnel is an important infrastructure for research on aircraft icing and anti-deicing in which refrigeration system realizes precise control of the airflow temperature in the wind tunnel by adjusting the suction pressure of the compressor unit. The suction pressure control and cooling methods affect the wind tunnel test efficiency. In this paper, aiming at accurate prediction of compressor suction pressure, the support vector regression （APSO–SVR） optimized by adaptive particle swarm algorithm is used to establish pressure prediction model to conduct pressure prediction and experimental research. In order to further improve the efficiency of icing wind tunnel testing, multi-layer perceptron （MLP） neural network is used to establish an analysis model to analyze the influence of test parameters on the cooling rate of wind tunnel. The results show that the average absolute percentage error （EMAP） between the predicted and test value of the compressor suction pressure is less than 4%, and the mean square error （EMS） is less than 0.003; the parameters affecting the wind tunnel cooling rate are mainly airflow density, test wind speed, compressor suction pressure and the initial temperature of the heat exchanger outlet. Among them, the compressor suction pressure has the most significant effect on it.
2022, 36(5): 16-23. doi: 10.11729/syltlx20210080
Abstract(29) HTML(8) PDF(10)
Abstract:
With the increase of the operation power and higher requirement on the fine measurement of the large-scale continuous high speed wind tunnel, higher performance of the wind tunnel heat exchanger is required, which is mainly reflected in the heat transfer and pressure loss performance, temperature field uniformity and flow field perturbation characteris-tics. Based on recent years' research, the key technologies and research results in the design of the large-scale continuous high speed wind tunnel heat exchanger are comprehensively discussed. Firstly, the required characteristics of the wind tunnel heat exchanger are analyzed, and the main factors affecting various performances are summarized. Then the high efficiency low-pressure loss design technology is introduced along with the temperature uniformity control technology and the airflow disturbance control technology. The principles of comprehensive trade-off of the heat transfer efficiency and pressure loss of the heat exchanger are presented. And the influences of the heat transfer core arrangement, cooling water flow and its inlet direction, cross section shape of the heat exchanger segment on the temperature field uniformity are discussed. Finally the flow disturbance characteristics downstream of the heat exchanger under different flow conditions and heat exchanger structures are analyzed.
2022, 36(5): 24-33. doi: 10.11729/syltlx20210175
Abstract(26) HTML(7) PDF(8)
Abstract:
In order to meet the need of the variable Reynolds number plane cascade tests for the advanced turbofan engine, the variable density plane cascade wind tunnel was designed to change the subsonic, transonic and supersonic flow efficiently, adjust the Mach number （Ma） and Reynolds number （Re） independently, combine the compressor and turbine cascade tests and possess the ability of heat transfer or cooling experiments. The general design project of the wind tunnel was put forward, and the ejector, the flexible nozzle as well as the test chamber design problems were introduced in details. The results of the flow field debugging and the typical cascade tests were analysed. The research reveals that the design technologies of the components satisfy the main functions of the variable density plane cascade wind tunnel. The Reynolds number can be as low as 3.1×105 m–1 so that the low Reynolds number experiments can be made easyly with the facility. The flow field debugging results satisfy the National military standard GJB 1179A—2012 of the specification for flow quality of high and low speed wind tunnels. It provides a key test platform for the study of the transonic, supersonic Mach number and low Reynolds number flow problems of the compressor and turbine cascades.
2022, 36(5): 34-42. doi: 10.11729/syltlx20210177
Abstract(33) HTML(5) PDF(14)
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In recent years, the engines for large commercial airplane are being developed closely, and the tests for main components and the whole engine are being conducted massively in China. For the purpose of supporting the development of low-emission combustors, the paper introduces the typical R&D path of a civil engine combustor, and mainly focuses on the test technologies of single cup, multiple cups and full annular combustors, in which the main test subjects, test process, instrumentation, key points and common issues have been summarized. At the same time, by making a comparison of test pressure/temperature, instrumentation, test abundance and perspectiveness between China and western countries and by summarizing of Chinese civil combustor test technology developments, the paper gives reference to the domestic researchers in the fields of low-emission combustor research, test facility construction, measure-ment enhancement and gap filling.
2022, 36(5): 43-51. doi: 10.11729/syltlx20210101
Abstract(95) HTML(71) PDF(9)
Abstract:
When the aircraft flies at a high angle of attack, the flow field on the leeward surface of the slender body evolves in a complicated manner, and asymmetric vortices appear, generating random lateral forces, which greatly affect the maneuverability and agility of the aircraft. In order to solve this problem, a sliding discharge plasma actuator with an along-stream layout was used to conduct wind tunnel experiments on a slender body model, combined with pressure measurement and particle image velocimetry（PIV）. The experimental results show that the actuation voltage of 10 kV is the threshold voltage at which the flow control starts to take effect. When the velocity of the incoming flow is 10 m/s（Re=0.8×105）, the angle of attack is 45°, and the actuation voltage is 16 kV, the best flow control effect can be achieved at the normalized pulse frequency of 1.96, the lateral force coefficient can be reduced by 83.48%. However, as the flow velocity increases, the flow control effect becomes weaker gradually and is expected to disappear at 26 m/s.
2022, 36(5): 52-56. doi: 10.11729/syltlx20210035
Abstract(138) HTML(36) PDF(23)
Abstract:
The flight test needs fast response and small transducer for heat flux measurement. The dual-junction thermocouple, which is characterized by fast response, small size and abundant measured information, is one of the best solutions for the temperature and heat flux measurement in the flight test. The principle, structure and measurement method of the dual-junction thermocouple are studied, and dual-junction thermocouples are used in a flight test for the model surface temperature measurement. The model front-surface and back-surface temperatures were measured simultaneously by the dual-junction thermocouples, and it is found that the measured back-surface temperature has a greater error. The response time of the back-surface measurement point is much longer than that of the front-surface measurement point, which is affected by the junction size and the insulating coating between the junction point and the model surface. At present there is still a lack of the corresponding heat flux estimation method for the dual-junction thermocouple.
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2016, 30(1): 28-42.   doi: 10.11729/syltlx20150069
[Abstract](228) [PDF 6594KB](30)

2016, 30(4): 7-13.   doi: 10.11729/syltlx20150112
[Abstract](256) [FullText HTML](76) [PDF 10428KB](6)

2016, 30(2): 67-74.   doi: 10.11729/syltlx20150091
[Abstract](293) [PDF 5952KB](17)

2016, 30(1): 1-14,27.   doi: 10.11729/syltlx20150159
[Abstract](299) [PDF 6603KB](15)

2017, 31(2): 1-11.   doi: 10.11729/syltlx20160129
[Abstract](360) [FullText HTML](129) [PDF 7434KB](39)

2016, 30(1): 81-90.   doi: 10.11729/syltlx20150037
[Abstract](177) [PDF 4386KB](9)

2016, 30(3): 98-103.   doi: 10.11729/syltlx20150107
[Abstract](127) [PDF 4151KB](1)

2018, 32(1): 64-70.   doi: 10.11729/syltlx20170099
[Abstract](209) [FullText HTML](127) [PDF 9648KB](16)

2016, 30(5): 55-60.   doi: 10.11729/syltlx20160026
[Abstract](243) [FullText HTML](110) [PDF 7981KB](9)

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(4): 8-15.   doi: 10.11729/syltlx20160207
[Abstract](272) [FullText HTML](99) [PDF 5274KB](12)

2022, 36(2): 49-73.   doi: 10.11729/syltlx20210110
[Abstract](2452) [FullText HTML](230) [PDF 8779KB](230)
Abstract: