CN112945501A - Laminar flow wing transition position measurement test method - Google Patents
Laminar flow wing transition position measurement test method Download PDFInfo
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- CN112945501A CN112945501A CN202110148022.2A CN202110148022A CN112945501A CN 112945501 A CN112945501 A CN 112945501A CN 202110148022 A CN202110148022 A CN 202110148022A CN 112945501 A CN112945501 A CN 112945501A
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64F—GROUND OR AIRCRAFT-CARRIER-DECK INSTALLATIONS SPECIALLY ADAPTED FOR USE IN CONNECTION WITH AIRCRAFT; DESIGNING, MANUFACTURING, ASSEMBLING, CLEANING, MAINTAINING OR REPAIRING AIRCRAFT, NOT OTHERWISE PROVIDED FOR; HANDLING, TRANSPORTING, TESTING OR INSPECTING AIRCRAFT COMPONENTS, NOT OTHERWISE PROVIDED FOR
- B64F5/00—Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided for; Handling, transporting, testing or inspecting aircraft components, not otherwise provided for
- B64F5/60—Testing or inspecting aircraft components or systems
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/06—Measuring arrangements specially adapted for aerodynamic testing
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M9/00—Aerodynamic testing; Arrangements in or on wind tunnels
- G01M9/08—Aerodynamic models
Abstract
The invention discloses a measurement test method for a transition position of a laminar flow wing, which comprises the following steps: processing a laminar flow wing model and a sample wafer based on a temperature-sensitive paint technology; step two, acquiring a calibration relation between the temperature and the light intensity of the sample wafer in a calibration cabin; step three, establishing a test scene; step four, determining static debugging parameters in the temperature-sensitive paint measuring system; step five, joint adjustment is carried out on the temperature-sensitive paint measuring system and the wind tunnel measurement and control system to obtain a reference image, a background image and a sequence test image; and step six, respectively acquiring test images of different states of the laminar flow wing, and performing post-processing on the test acquired images to obtain an air flow transition image on the surface of the laminar flow wing model. The invention provides a laminar flow wing transition position measurement test method, which creatively establishes a laminar flow wing test mode based on a temperature-sensitive coating and ensures that the test effect has the remarkable advantages of high resolution and accurate transition position judgment compared with the prior art.
Description
Technical Field
The invention belongs to the technical field of wind tunnel tests, and particularly relates to a test method capable of accurately measuring the transition position of a laminar flow wing model surface without damaging the model surface in a wind tunnel test.
Background
In order to achieve the aims of energy conservation, emission reduction and increase of range, the aerodynamic drag reduction technology of the aircraft becomes the object of the key research of aerodynamic designers, and laminar flow design becomes possible gradually along with the rapid advance of the design technology and the manufacturing process of the aviation industry. For civil airliners, the laminar flow wing design technology can reduce the friction resistance by about 30%, further improve the cruising efficiency by about 15%, has obvious pneumatic benefits, improves the pneumatic performance, and effectively reduces the fuel consumption, pollution emission and flight noise. In order to examine the design method of the laminar flow wing and the aerodynamic characteristics of the laminar flow wing, a wing transition position measurement or prediction method is needed, and transition position determination is one of key technologies of the design of the laminar flow aircraft. The method for determining the transition position of the laminar flow wing generally adopts two means, namely numerical simulation and wind tunnel test. Because the numerical simulation calculation amount is large and the precision is difficult to guarantee, the transition position of the laminar flow wing is generally measured by adopting a wind tunnel test method, so that the design method of the laminar flow wing is verified, and the quality of the design is evaluated.
The transition position can be effectively judged by measuring the temperature distribution by utilizing the characteristic that the temperature of the laminar flow area and the turbulent flow area of the model is different due to the different heat convection intensity of the laminar flow and the turbulent flow. In the conventional transition measurement method, a temperature sensor is arranged on the surface of the wing model, the temperature of the surface of the model is obtained through measurement of the temperature sensor, and then the transition position of the surface of the model is judged according to the temperature distribution. The method has a plurality of defects: firstly, because the sensor protrudes or sinks from the surface of the wing model to interfere with the surface flow field, the air flow velocity or the flow field structure on the surface of the wing is different, and further the surface temperature measurement error is caused, the sensor and the surface of the wing model are required to be strictly flush, and the installation difficulty of the sensor is high. Secondly, the temperature sensor is installed, the wing model needs to be opened and grooved at the installation position and the wiring channel, the model design and processing difficulty is increased, and the model design and processing cost is increased. Thirdly, on thin parts such as wing tips or wing trailing edges, temperature sensors cannot be installed due to insufficient thickness space, and transition position measurement cannot be performed on the regions. And fourthly, the temperature sensor belongs to a discrete point measurement method, only a plurality of points on the surface of the wing can be measured, the spatial resolution is low, and the accurate transition position of the laminar flow wing cannot be obtained. The method has the advantages of surface measurement, common measurement methods comprise a phase change thermal map, a temperature-sensitive liquid crystal, an infrared thermal map, a phosphorescence thermal map and the like, but the application range is limited due to inherent defects of the method, for example, the phase change thermal map and the temperature-sensitive liquid crystal can only be used as a semi-quantitative measurement technology, the spatial resolution of an infrared camera is not high, the influence on the transmissivity of a transparent material in a measuring light path is large, the phosphorescence thermal map technology is obviously limited by the surface film forming mode and process of an inorganic phosphorescence substance, and the like.
The conventional temperature sensor transition measuring method is high in installation difficulty, large in flow interference, high in model design and processing cost, only capable of carrying out point measurement and low in spatial resolution, and common optical measuring methods such as a phase change thermal map, a temperature sensitive liquid crystal, an infrared thermal map and a phosphorescence thermal map restrict the application range due to inherent defects of the optical measuring methods, so that the transition position of the laminar flow wing model cannot be accurately measured. The temperature-sensitive coating method has the advantages of surface measurement, high spatial resolution, low model design and processing difficulty and cost, no disturbance of incoming flow, accurate transition position judgment and the like, but no related test method is applied to measurement of the transition position of the laminar flow wing at present and is urgently needed to be established.
Disclosure of Invention
An object of the present invention is to solve at least the above problems and/or disadvantages and to provide at least the advantages described hereinafter.
To achieve these objects and other advantages in accordance with the purpose of the invention, a layer flow wing transition position measurement test method is provided, including:
processing a laminar flow wing model and a sample wafer based on a temperature-sensitive paint technology;
step two, acquiring a calibration relation between the temperature and the light intensity of the sample wafer in a calibration cabin;
step three, installing a laminar flow wing model, a camera, an excitation light source, a power supply of the excitation light source, a synchronous trigger and a data processing industrial personal computer in the wind tunnel to establish a test scene;
step four, determining static debugging parameters in the temperature-sensitive paint measuring system by adjusting lens parameter setting, CCD exposure time setting and image acquisition time sequence;
step five, performing joint debugging on the temperature-sensitive paint measuring system and the wind tunnel measurement and control system to obtain a reference image and a background image in the state that a light source is turned on or off before the wind tunnel is started and a sequence test image after the wind tunnel is started, the light source is turned on and a flow field is stabilized;
and step six, respectively acquiring test images of different states of the laminar flow wing, and performing post-processing on the test acquired images to obtain an air flow transition image on the surface of the laminar flow wing model.
Preferably, in step one, the laminar flow airfoil coupon processing is configured to include:
processing a laminar flow wing test model and an aluminum sample wafer, cleaning the surfaces of the model and the sample wafer, spraying a temperature-sensitive paint primer on the surfaces of the model and the sample wafer, after the spraying is finished, placing the model and the sample wafer in an oven for baking for 6 hours for curing at the temperature of 90 ℃, and polishing the primer coating of the cured model and the sample wafer by 1500-mesh abrasive paper until the surface roughness is less than 0.8;
cleaning the surfaces of the model and the sample wafer, spraying temperature-sensitive paint finish on the primer coating, air-drying and curing for 12 hours at normal temperature, and polishing the cured surface paint coating of the model and the sample wafer by 1500-mesh abrasive paper until the surface roughness is less than 0.8.
Preferably, in the second step, the calibration relation between the temperature and the light intensity is obtained by placing the sample wafer in a calibration chamber, irradiating the sample wafer with an excitation light source, adjusting the temperature in the calibration chamber, collecting light intensity images of the sample wafer at different temperatures by a camera, and post-processing the images.
Preferably, in the third step, the laminar flow wing model is installed in a side wall supporting mode, the laminar flow wing model is connected to the half-mode supporting mechanisms on the left side wall and the right side wall of the wind tunnel test section through a left supporting plate and a right supporting plate, the position of the laminar flow wing model is in the range of the axis of the wind tunnel and the uniform area of the flow field, and a camera can shoot the area;
the camera and the excitation light source are arranged in the wind tunnel upper parking chamber, and the fixing device of the camera and the excitation light source can adjust the positions of the camera and the excitation light source along the axis of the wind tunnel, in the front-back direction, the up-down direction and the left-right direction, so that the camera and the light source can be positioned in the range of the axis of the wind tunnel and the observation window of the wind tunnel upper parking chamber, and the shooting direction of the camera and the irradiation;
the power supply of the excitation light source is installed in the wind tunnel upper parking chamber, the synchronous trigger and the data processing industrial personal computer are installed on a working platform outside the wind tunnel parking chamber, the camera and the light source control line are connected with the synchronous trigger, and the camera data line and the synchronous trigger control line are connected with the industrial personal computer.
Preferably, in the fifth step, when the temperature-sensitive paint measuring system is in joint regulation with the wind tunnel measuring and controlling system, normal blowing conditions are simulated in the wind tunnel, the parking room is closed, and meanwhile, the observation windows on the two sides of the test section are shaded;
under a blowing state test of the laminar flow wing, a sequence test image is obtained after a wind tunnel is started and a flow field is stable, a wind tunnel measurement and control system transmits a starting signal to a synchronous trigger, the synchronous trigger simultaneously transmits working signals to a camera and a light source after receiving the signals, the light source starts to irradiate, the camera starts to collect, and the light source is closed after image sequence collection.
Preferably, in the sixth step, after the blowing test of one state of the laminar flow wing is finished, the model is baked by using a far infrared baking lamp or a hot air blower, and the blowing test of the next state can not be carried out until the surface temperature of the model is recovered to be uniform until all the state tests are finished.
Preferably, the temperature-sensitive coating performance parameters are configured to:
the excitation peak spectral wavelength is 400nm, the emission peak spectral wavelength is 615nm, the temperature sensitivity is more than 1%/K within the temperature range of 273K-333K, the photodegradation rate of the coating is less than 1%/min, the storage life is more than 3 months, and the upper limit of the applicable temperature range is more than 60 ℃.
Preferably, the performance parameters of the camera are configured to:
the dynamic range of the gray scale is at least more than 8 bits, the spatial resolution is more than 800 multiplied by 600 pixels, the backboard is used for refrigeration, and a 650nm high-pass filter is used.
Preferably, the performance parameters of the excitation light source are configured to:
the transmittance of the optical filter is more than 90%, and the optical filter has two irradiation working modes of pulse and continuous, wherein the light source control model is TTL, and the output power is 8-12W;
the performance parameters of the synchronous trigger are required to have at least 2 paths of outputs, and the control precision is less than 20 nanoseconds.
Preferably, in step three, the static debugging parameters are set as: the aperture of a camera lens is set to be 12, the exposure time of the camera is set to be 400ms, each car has 100 periods, each acquisition period is 600ms, the light source delays for 5ms after the synchronous trigger receives the trigger signal, and the camera delays for 150 ms.
The invention at least comprises the following beneficial effects: firstly, the temperature-sensitive coating is adopted to replace a conventional temperature sensor on a laminar flow wing model, so that the interference of the sensor on the flow field, which is protruded or sunken on the surface of the wing model, is avoided.
Secondly, the invention does not need to open holes and open slots on the wing model, thereby reducing the design, processing difficulty and processing cost of the model.
Thirdly, the surface of the laminar flow wing model can be measured, the spatial resolution is high, and the accurate transition position of the surface of the laminar flow wing model can be obtained.
Fourthly, the laminar flow wing transition position measurement test method is suitable for accurate measurement of the laminar flow wing transition position and design method verification of the laminar flow wing, and has popularization and application values.
Additional advantages, objects, and features of the invention will be set forth in part in the description which follows and in part will become apparent to those having ordinary skill in the art upon examination of the following or may be learned from practice of the invention.
Drawings
FIG. 1 is a schematic view of the temperature sensitive measurement system and laminar flow wing model installation of the present invention;
FIG. 2 is a schematic structural diagram of the temperature-sensitive coating of the present invention;
FIG. 3 is a schematic flow chart of the test method of the present invention;
fig. 4 is a result graph of transition of laminar flow wings in embodiment 1 of the present invention.
Fig. 5 is a graph illustrating a transition result of a laminar wing section according to embodiment 1 of the present invention.
Detailed Description
The present invention is further described in detail below with reference to the attached drawings so that those skilled in the art can implement the invention by referring to the description text.
The invention discloses a measurement test method for transition position of laminar flow wing, which uses a measurement device shown in fig. 1, wherein the measurement device comprises: the device comprises a laminar flow wing model 2, a temperature-sensitive coating, a camera 3, an excitation light source 4, an excitation light source power supply 9, a synchronous trigger 5 and a data processing industrial personal computer 6, wherein the laminar flow wing model, the temperature-sensitive coating, the camera 3, the excitation light source power supply 4, the excitation light source power supply 9, the synchronous trigger 5 and the data; the laminar flow wing model is a metal model, a temperature-sensitive paint primer and a temperature-sensitive paint finish coat are sequentially covered on the surface of the model from bottom to top, and the schematic diagram of the coating layout is shown in FIG. 2; the temperature-sensitive paint consists of a temperature-sensitive paint primer and a temperature-sensitive paint finish, wherein the temperature-sensitive paint primer is called a substrate emitting layer and is white primer containing silicon dioxide, and the white primer is sprayed on the surface of a model and plays a role in improving the surface adhesion of the model, enhancing the luminous intensity of probe molecules and isolating heat. The temperature-sensitive paint finish paint is called a polymer functional layer and contains temperature-sensitive probe molecules, wherein the probe molecules are trivalent europium fluorescent complexes and are main temperature-sensitive luminescent materials of the temperature-sensitive paint. The temperature-sensitive coating has an excitation peak spectral wavelength of 400nm and an emission peak spectral wavelength of 615nm, the temperature sensitivity is greater than 1%/K and the pressure sensitivity is extremely low within a temperature range of 273K-333K, the photodegradation rate of the coating is less than 1%/min, the storage life is more than 3 months, and the upper limit of the applicable temperature range is greater than 60 ℃; the camera is a scientific grade CCD camera, and has high signal-to-noise ratio and gray dynamic range, wherein the gray dynamic range is at least more than 8 bits, the spatial resolution is more than 800 multiplied by 600 pixels, the camera is refrigerated with a back plate, a 650nm high-pass filter is used, and lenses with different focal lengths can be installed according to the shooting distance and the shooting area. The excitation light source is an LED ultraviolet light source, the transmittance of the optical filter is more than 90%, the LED ultraviolet light source has two irradiation working modes of pulse and continuous, the light source control model is TTL, and the output power is 8W-12W. The synchronous trigger can set the period, time delay, pulse width and pulse number of pulse signals, is used for realizing the time sequence control of camera exposure and excitation light sources, requires at least 2 paths of output, and has the control precision smaller than 20 nanoseconds. The data processing industrial personal computer is connected with the synchronous trigger and the high-speed camera and used for setting parameters of the synchronous trigger, further controlling the time sequence of irradiation of the excitation light source and exposure of the camera, receiving the light intensity image of the surface of the laminar flow wing model shot by the camera, performing image post-processing and obtaining the required air flow transition result image of the surface of the laminar flow wing model.
The test method comprises the following steps as shown in figure 3:
a. processing a laminar flow wing test model and an aluminum sample wafer, grinding screw holes and other depressions on the surface of the model by putty, solidifying and polishing, cleaning the surfaces of the model and the sample wafer by ethanol or acetone, stirring a temperature-sensitive paint primer and a solvent until the temperature-sensitive paint primer and the solvent are uniformly dispersed, spraying the temperature-sensitive paint primer on the surfaces of the model and the sample wafer by a spray gun, placing the model and the sample wafer in an oven for baking at 90 ℃ for 6 hours for solidification after the spraying is finished, and polishing the model and the primer coating of the sample wafer by 1500-mesh sand paper until the roughness is less than 0.8.
b. Cleaning the surfaces of the model and the sample wafer, stirring the temperature-sensitive paint finish and a solvent until the temperature-sensitive paint finish is uniformly dispersed, spraying the temperature-sensitive paint finish on the primer coating, air-drying and curing the finish for 12 hours at normal temperature, and polishing the surface paint coating of the model and the sample wafer by 1500-mesh abrasive paper after the curing is finished until the roughness is less than 0.8.
c. Placing a sample in a calibration cabin, irradiating the sample by an excitation light source, adjusting the temperature in the calibration cabin, acquiring light intensity images of the sample at different temperatures by a camera, and performing post-processing on the images to obtain a calibration relational expression of the temperature and the light intensity, wherein the temperature-sensitive paint calibration coefficient relational expression is as follows:
wherein P is pressure, alphajiTo calibrate the coefficients, IrThe ratio of the light intensity of the temperature-sensitive image to the light intensity of the reference image in the two states of blowing and not blowing is shown, and T is the temperature.
d. A mode of side wall support is adopted, a laminar flow wing model is connected to a half-mode supporting mechanism of the left side wall and the right side wall of a wind tunnel test section through a left supporting plate and a right supporting plate, and the left supporting plate and the right supporting plate are connected and fastened with a left rotating window and a right rotating window through screws and pins. The model position should be in the range of the wind tunnel axis and the flow field uniform area and in the area that can be shot by the camera.
e. The camera and the excitation light source are installed in the parking chamber on the wind tunnel, the camera and the excitation light source are sequentially connected onto the parking chamber slide rail through the fast-assembling plate, the cradle head and the slide block, the position of the light source and the position of the camera can be adjusted in the front-back direction, the upper-lower direction and the left-right direction of the axis of the wind tunnel by moving the slide block along the slide rail, the accurate installation and fixation of the measuring equipment are achieved, and the camera and the light source can be located in the 7 ranges of the axis of the. In order to improve the excitation light irradiation intensity and the image resolution and reduce the image distortion, the shooting direction of a camera and the light source irradiation direction are adjusted to be perpendicular to the surface of the wing model as much as possible. In order to shorten the light source irradiation and camera shooting distance and avoid the influence of refraction and scattering of optical glass, an optical observation window of an upper dwelling room of a test section is disassembled, a special adapter plate is installed, a mounting hole is formed in the adapter plate, a camera and an excitation light source are inserted into the mounting hole and fixedly installed, and a camera lens is flush with a light source head and an upper wall plate 8 of a wind tunnel.
f. The device comprises a power supply for installing an excitation light source, a synchronous trigger and a data processing industrial personal computer, wherein the power supply for the excitation light source is installed in a wind tunnel upper parking chamber, the synchronous trigger and the data processing industrial personal computer are arranged on a working platform outside the wind tunnel parking chamber, airflow flows in the wind tunnel parking chamber, and the power supply needs to be fastened. The camera and light source control line is connected with the synchronous trigger, and the camera data line and the synchronous trigger control line are connected with the industrial personal computer. The GigE kilomega network cable, the water cooling pipe and the TTL trigger signal line are led out through cable holes in the side wall of the standing chamber, and the cables are fixed by strapping tapes.
g. And performing static debugging on the temperature-sensitive paint measurement system, wherein the static debugging comprises the steps of lens parameter setting, CCD exposure time setting, image acquisition time sequence determination and the like. Lens parameters include focal length and aperture, focusing is intended to make the image as sharp as possible, and aperture and CCD exposure time determine the image gray level. Since the depth of field is larger as the aperture is smaller, the aperture needs to be narrowed as much as possible within a reasonable CCD exposure time in order to improve the quality of the image edge. The camera lens aperture is set to 12, the CCD exposure time is set to 400ms, and under this parameter, the reference image gray scale light intensity is 9000, reaching the full camera range 2/3. The acquisition time sequence of the synchronous trigger is set to be 100 periods of each vehicle, each acquisition period is 600ms, the light source delays for 5ms after the synchronous trigger receives the trigger signal, and the camera delays for 150 ms.
h. And (4) performing joint adjustment on the temperature-sensitive paint measuring system and the wind tunnel measurement and control system, simulating a normal blowing condition, closing a parking room, and performing shading treatment on observation windows on two sides of the test section.
i. Before the wind tunnel is started, the light source is turned on, the camera collects 20 reference light images, the light source is turned off after the collection is finished, and the camera collects 20 background images.
j. The wind tunnel is started, after the flow field is stable, the wind tunnel measurement and control system transmits a starting signal to the synchronous trigger, the synchronous trigger receives the signal and simultaneously transmits working signals to the camera and the light source, the light source starts to irradiate, the camera starts to collect, and after the collection of the image sequence is finished, the light source is closed.
k. After one state test is finished, the model is baked by using a far infrared baking lamp or an air heater, and the next state blowing test can be carried out until the surface temperature of the model is recovered to be uniform until all the state tests are finished.
And l, post-processing the acquired images to obtain an air flow transition image on the surface of the laminar flow wing model shown in the figures 4-5. The temperature-sensitive paint has the characteristics of photoluminescence and thermal quenching of exciting light, wherein the photoluminescence characteristic means that the paint can emit light of another wavelength under the irradiation of the exciting light with a certain wavelength, and the thermal quenching characteristic means that the intensity of the emitted light of the paint is reduced along with the increase of the temperature. Based on the two characteristics of the coating, in the scheme, the temperature-sensitive coating is sprayed on the surface of the model, the coating is irradiated by exciting light with a specific wavelength, the photoluminescence and thermal quenching characteristics of the coating on the surface of the model are utilized, the luminous intensity of the coating is converted into the surface temperature of the model, the transition position of the airflow on the surface of the model is judged according to the temperature gradient according to the characteristic that the temperatures of a laminar flow area and a turbulent flow area of the model are different due to the difference of the thermal convection intensities of laminar flow and turbulent flow, the test effect, the test precision and the operation simplicity are ensured, meanwhile, the large-area continuous measurement on the surface of the model can be realized by taking the temperature-sensitive coating as a sensing method, a sensor is not required to be arranged on the surface of the model, the interference of the sensor protruding or sinking on the surface of the airfoil to a convection field is avoided, no hole or slot is, the method has the advantages that the spatial resolution is high, the accurate transition position of the surface of the laminar flow wing model can be obtained, meanwhile, by combining other steps of the scheme, a laminar flow wing test mode based on the temperature-sensitive coating is creatively established, the method is suitable for accurate measurement of the transition position of the laminar flow wing and verification of a design method of the laminar flow wing, the test effect is guaranteed, and the method has the remarkable advantages of high resolution and accurate transition position judgment compared with the prior art.
Example (b):
the test model of the embodiment is a laminar flow wing model with a sweep angle of 20 degrees and a chord length of 200mm, the camera is a scientific-grade CCD camera, the gray dynamic range is 14 bits, the spatial resolution is 1600 multiplied by 1200 pixels, the refrigeration is carried out by a backboard, the adopted lens is an 8mm fixed-focus lens, and the adopted filter is a 650nm high-pass filter. The wavelength of a main luminous peak of an excitation light source is 400nm, the transmittance of the optical filter is more than 90%, the excitation light is irradiated by a pulse mode and a continuous mode, the light source control model is TTL, the filtering combination mode is low pass plus narrow wave, and the output power is 8W-12W. The synchronous trigger can set the period, time delay, pulse width and pulse number of pulse signals to realize the time sequence control of camera exposure and excitation light source, and the time sequence control is a single-path input 8-path output, and the control precision is less than 10 nanoseconds. The temperature-sensitive paint has an excitation peak spectrum wavelength of 400nm and an emission peak spectrum wavelength of 615nm, the temperature sensitivity is greater than 1%/K, the photodegradation rate of the paint is less than 1%/min, the storage life is more than 3 months, and the upper limit of the applicable temperature range is greater than 60 ℃ in the temperature range of 273-333K. In this embodiment, a map result of the transition position of the surface of the laminar wing model under the condition of mach number 0.73 attack angle 0 ° shown in fig. 4 and a comparison curve result of the transition position of the central section of the laminar wing model under different mach numbers shown in fig. 5 are obtained. As can be seen from fig. 4 and 5, the accurate position result of the laminar flow wing model surface transition can be obtained by using the laminar flow wing transition position measurement test method of the present invention.
In order to better explain the scheme, the post-processing of the test acquisition image comprises the following steps:
s1, loading a laminar flow wing background image, a reference image and a test sequence image which are collected by a camera, selecting marking characteristic points, identifying marking points and positioning the marking points on the reference image and the test sequence image, and storing a coordinate file of the positioned marking points;
s2, registering the test sequence image to the position of the reference image according to the coordinate relation of the mark points, checking the registration precision, if the precision reaches the standard, storing the registered test sequence image, entering the step S3, and if the precision does not reach the standard, returning to the step S1;
s3, subtracting the background image from the reference image, subtracting the background image from the test sequence image, and performing pre-filtering on the reference image and the test sequence image after the background image is subtracted;
s4, clipping the reference image and the test sequence image according to the area of the wing model in the image to obtain the clipped reference image and the test sequence image;
s5, image filling is carried out on temperature-sensitive paint-free areas such as screw holes in the wing model, the areas outside the wing model are set as background areas, and a reference image and a test sequence image after filling are obtained;
s6, carrying out ratio processing on the reference image and the test sequence image to obtain a light intensity ratio sequence image, and carrying out post-filtering on the ratio sequence image;
s7, obtaining a wing surface temperature data sequence image through conversion according to the relationship between the light intensity ratio sequence image and the temperature-sensitive paint calibration coefficient;
s8, calculating to obtain a laminar flow wing surface thermal flow data image according to the temperature data sequence image obtained in the step S7;
s9, judging whether the airflow direction is B according to the light intensity ratio image obtained in S7 or the heat flow data image obtained in S8, as shown in the wing surface transition region and the transition position a of fig. 4-5.
The above scheme is merely illustrative of a preferred example, and is not limiting. When the invention is implemented, appropriate replacement and/or modification can be carried out according to the requirements of users.
The number of apparatuses and the scale of the process described herein are intended to simplify the description of the present invention. Applications, modifications and variations of the present invention will be apparent to those skilled in the art.
While embodiments of the invention have been disclosed above, it is not intended to be limited to the uses set forth in the specification and examples. It can be applied to all kinds of fields suitable for the present invention. Additional modifications will readily occur to those skilled in the art. It is therefore intended that the invention not be limited to the exact details and illustrations described and illustrated herein, but fall within the scope of the appended claims and equivalents thereof.
Claims (10)
1. A measurement test method for a transition position of a laminar wing is characterized by comprising the following steps:
processing a laminar flow wing model and a sample wafer based on a temperature-sensitive paint technology;
step two, acquiring a calibration relation between the temperature and the light intensity of the sample wafer in a calibration cabin;
step three, installing a laminar flow wing model, a camera, an excitation light source, a power supply of the excitation light source, a synchronous trigger and a data processing industrial personal computer in the wind tunnel to establish a test scene;
step four, determining static debugging parameters in the temperature-sensitive paint measuring system by adjusting lens parameter setting, CCD exposure time setting and image acquisition time sequence;
step five, performing joint debugging on the temperature-sensitive paint measuring system and the wind tunnel measurement and control system to obtain a reference image and a background image in the state that a light source is turned on or off before the wind tunnel is started and a sequence test image after the wind tunnel is started, the light source is turned on and a flow field is stabilized;
and step six, respectively acquiring test images of different states of the laminar flow wing, and performing post-processing on the test acquired images to obtain an air flow transition image on the surface of the laminar flow wing model.
2. The laminar wing transition position measurement test method of claim 1, wherein in step one, laminar wing sample processing is configured to include:
processing a laminar flow wing test model and an aluminum sample wafer, cleaning the surfaces of the model and the sample wafer, spraying a temperature-sensitive paint primer on the surfaces of the model and the sample wafer, after the spraying is finished, placing the model and the sample wafer in an oven for baking for 6 hours for curing at the temperature of 90 ℃, and polishing the primer coating of the cured model and the sample wafer by 1500-mesh abrasive paper until the surface roughness is less than 0.8;
cleaning the surfaces of the model and the sample wafer, spraying temperature-sensitive paint finish on the primer coating, air-drying and curing for 12 hours at normal temperature, and polishing the cured surface paint coating of the model and the sample wafer by 1500-mesh abrasive paper until the surface roughness is less than 0.8.
3. The method for testing position transition of laminar flow wing according to claim 1, wherein in step two, the calibration relation between temperature and light intensity is obtained by placing the sample in a calibration cabin, irradiating the sample with an excitation light source, adjusting the temperature in the calibration cabin, and acquiring light intensity images of the sample at different temperatures by a camera and post-processing the images.
4. The laminar flow wing transition position measurement test method according to claim 1, characterized in that in step three, the installation of the laminar flow wing model is in a side wall support mode, the laminar flow wing model is connected to the half-mode support mechanisms on the left and right side walls of the wind tunnel test section through the left and right support plates, the position of the laminar flow wing model is in the range of the wind tunnel axis and the flow field uniform region, and the region can be shot by the camera;
the camera and the excitation light source are arranged in the wind tunnel upper parking chamber, and the fixing device of the camera and the excitation light source can adjust the positions of the camera and the excitation light source along the axis of the wind tunnel, in the front-back direction, the up-down direction and the left-right direction, so that the camera and the light source can be positioned in the range of the axis of the wind tunnel and the observation window of the wind tunnel upper parking chamber, and the shooting direction of the camera and the irradiation;
the power supply of the excitation light source is installed in the wind tunnel upper parking chamber, the synchronous trigger and the data processing industrial personal computer are installed on a working platform outside the wind tunnel parking chamber, the camera and the light source control line are connected with the synchronous trigger, and the camera data line and the synchronous trigger control line are connected with the industrial personal computer.
5. The laminar flow wing transition position measurement test method according to claim 1, characterized in that in step five, when the temperature-sensitive paint measurement system is adjusted in conjunction with the wind tunnel measurement and control system, normal blowing conditions are simulated in the wind tunnel, the parking room is closed, and meanwhile, the observation windows on two sides of the test section are shaded;
under a blowing state test of the laminar flow wing, a sequence test image is obtained after a wind tunnel is started and a flow field is stable, a wind tunnel measurement and control system transmits a starting signal to a synchronous trigger, the synchronous trigger simultaneously transmits working signals to a camera and a light source after receiving the signals, the light source starts to irradiate, the camera starts to collect, and the light source is closed after image sequence collection.
6. The laminar flow wing transition position measurement test method according to claim 1, wherein in step six, after one state blowing test of the laminar flow wing is finished, a far infrared baking lamp or a hot air blower is used for baking the model, and after the surface temperature of the model is recovered to be uniform, the next state blowing test can be performed until all the state tests are finished.
7. The laminar wing transition position measurement test method according to claim 2, wherein the temperature-sensitive coating performance parameter is configured to:
the excitation peak spectral wavelength is 400nm, the emission peak spectral wavelength is 615nm, the temperature sensitivity is more than 1%/K within the temperature range of 273K-333K, the photodegradation rate of the coating is less than 1%/min, the storage life is more than 3 months, and the upper limit of the applicable temperature range is more than 60 ℃.
8. The laminar wing transition position measurement test method of claim 1, wherein the performance parameters of the camera are configured to:
the dynamic range of the gray scale is at least more than 8 bits, the spatial resolution is more than 800 multiplied by 600 pixels, the backboard is used for refrigeration, and a 650nm high-pass filter is used.
9. The method of claim 2, wherein the performance parameters of the excitation light source are configured to:
the transmittance of the optical filter is more than 90%, and the optical filter has two irradiation working modes of pulse and continuous, wherein the light source control model is TTL, and the output power is 8-12W;
the performance parameters of the synchronous trigger are required to have at least 2 paths of outputs, and the control precision is less than 20 nanoseconds.
10. The laminar flow wing transition position measurement test method according to claim 2, wherein in step three, the static debugging parameters are set as: the aperture of a camera lens is set to be 12, the exposure time of the camera is set to be 400ms, each car has 100 periods, each acquisition period is 600ms, the light source delays for 5ms after the synchronous trigger receives the trigger signal, and the camera delays for 150 ms.
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111473948A (en) * | 2020-05-29 | 2020-07-31 | 中国空气动力研究与发展中心高速空气动力研究所 | Measurement test device is twisted to low temperature transonic speed equipment TSP |
| CN111551341A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Low-temperature transonic equipment TSP transition measurement test method |
| CN111551344A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Surface treatment method for low-temperature transonic equipment TSP transition measurement test model |
| CN211954611U (en) * | 2020-05-29 | 2020-11-17 | 中国空气动力研究与发展中心高速空气动力研究所 | Measurement test device is twisted to low temperature transonic speed equipment TSP |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
-
2021
- 2021-02-03 CN CN202110148022.2A patent/CN112945501A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111473948A (en) * | 2020-05-29 | 2020-07-31 | 中国空气动力研究与发展中心高速空气动力研究所 | Measurement test device is twisted to low temperature transonic speed equipment TSP |
| CN111551341A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Low-temperature transonic equipment TSP transition measurement test method |
| CN111551344A (en) * | 2020-05-29 | 2020-08-18 | 中国空气动力研究与发展中心高速空气动力研究所 | Surface treatment method for low-temperature transonic equipment TSP transition measurement test model |
| CN211954611U (en) * | 2020-05-29 | 2020-11-17 | 中国空气动力研究与发展中心高速空气动力研究所 | Measurement test device is twisted to low temperature transonic speed equipment TSP |
| CN112304563A (en) * | 2020-10-30 | 2021-02-02 | 中国空气动力研究与发展中心超高速空气动力研究所 | Wind tunnel test method for researching influence of transition on aerodynamic characteristics of hypersonic aircraft |
Non-Patent Citations (1)
| Title |
|---|
| 黄辉 等: "基于温敏漆的边界层转捩测量技术研究" * |
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