CN112945502B - Laminar flow wing transition position measurement test system - Google Patents
Laminar flow wing transition position measurement test system Download PDFInfo
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- CN112945502B CN112945502B CN202110149382.4A CN202110149382A CN112945502B CN 112945502 B CN112945502 B CN 112945502B CN 202110149382 A CN202110149382 A CN 202110149382A CN 112945502 B CN112945502 B CN 112945502B
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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
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Abstract
The invention discloses a laminar flow wing transition position measurement test system, which comprises: the metal model for laminar flow wing test is arranged in the wind tunnel test section, and is provided with primer and finishing paint prepared by temperature sensitive paint; a camera and an excitation light source which are arranged in the parking chamber on the wind tunnel and are matched with the installation position of the metal model; the power supply module is connected with the excitation light source; the synchronous controller and the industrial personal computer are arranged outside the wind tunnel residence chamber; the industrial personal computer is configured to be in communication connection with the synchronous controller and the camera, the primer is configured to be a white primer containing silicon dioxide, and the finishing paint is configured to contain temperature-sensitive probe molecules of trivalent europium fluorescent complex. The invention provides a laminar flow wing transition position measurement test system, which aims at a laminar flow wing model based on temperature sensitive paint, realizes surface area surface measurement of the laminar flow wing model, has high spatial resolution, and can acquire an accurate transition position of the surface of the laminar flow wing model.
Description
Technical Field
The invention belongs to the technical field of wind tunnel tests, and particularly relates to a test device capable of realizing accurate measurement of a laminar flow wing transition position under the condition of not damaging a model surface in a wind tunnel test.
Background
In order to achieve the goals of energy conservation, emission reduction and range increase, the aerodynamic drag reduction technology of the aircraft is the object of the key research of aerodynamic designers, and with the rapid progress of the design technology and manufacturing technology of the aviation industry, laminar flow design is gradually possible. For civil airliners, the laminar wing design technology can reduce friction resistance by about 30%, further can improve cruising efficiency by about 15%, has obvious pneumatic benefit, improves pneumatic performance, and simultaneously effectively reduces fuel consumption, pollution emission and flight noise. In order to test the aerodynamic characteristics of the laminar wing design method and the laminar wing, a wing transition position measurement or prediction method is needed, and the transition position determination is one of key technologies of laminar aircraft design. The method for determining the transition position of the laminar 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 ensure, a wind tunnel test method is generally adopted to measure the transition position of the laminar wing, so that the design method of the laminar wing is verified, and the design quality is evaluated.
By utilizing the characteristic that the temperature of a laminar flow area and the temperature of a turbulent flow area on the surface of the model are different due to the fact that the different heat convection intensities of the laminar flow and the turbulent flow, the transition position can be effectively judged by measuring the temperature distribution. The traditional transition measuring device adopts a temperature sensor, and the specific method is that the 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 temperature distribution. The device has a plurality of defects: 1. because the sensor protrudes or is sunken on the surface of the wing model to interfere the surface flow field, the airflow speed or the flow field structure on the surface of the wing is different, and further, the surface temperature measurement error is caused, so that the sensor is required to be strictly flush with the surface of the wing model, and the installation difficulty of the sensor is high. 2. The installation temperature sensor needs to open holes and slots on the wing model at the installation position and the wiring channel, so that the model design and processing difficulty is increased, and the model design and processing cost is increased. 3. On thin parts such as wing tips and wing trailing edges, temperature sensors cannot be installed due to insufficient thickness space, so that transition position measurement cannot be performed on the areas. 4. The temperature sensor belongs to a discrete point measuring method, can only measure a plurality of points on the surface of the wing, has very low spatial resolution, and cannot acquire the accurate transition position of the laminar wing. The device has the advantages of surface measurement, common measurement devices comprise a phase change thermal diagram, a temperature-sensitive liquid crystal, an infrared thermal diagram, a phosphorescence thermal diagram and the like, but the application range is limited due to inherent defects of the common measurement devices, such as the phase change thermal diagram and the temperature-sensitive liquid crystal can only be used as semi-quantitative measurement technology, the infrared camera has low spatial resolution and is greatly influenced by the transmissivity of transparent materials in a measurement light path, the phosphorescence thermal diagram technology is obviously limited by the surface film forming mode and the technology of inorganic phosphorescence substances, and the like.
In a word, the traditional temperature sensor transition measuring device has high installation difficulty, large flow interference and high model design and processing cost, can only perform point measurement, has low spatial resolution, and the commonly used optical measuring devices such as a phase change heat map, a temperature sensitive liquid crystal, an infrared heat map, a phosphorescence heat map and the like restrict the application range due to inherent defects of the optical measuring devices, and cannot accurately measure the transition position of the wing model. The temperature-sensitive coating has the advantages of surface measurement, high spatial resolution, low model design processing difficulty and cost, no disturbance of incoming flow, accurate transition position judgment and the like, but is not applied to the transition position measurement of the laminar wing at present, the test method is still immature, and related measurement test devices need to be established urgently.
Disclosure of Invention
It is an object of the present invention to address at least the above problems and/or disadvantages and to provide at least the advantages described below.
To achieve these objects and other advantages and in accordance with the purpose of the invention, there is provided a laminar flow wing transition position measurement test system, comprising:
the metal model for laminar flow wing test is arranged in the wind tunnel test section, and is provided with primer and finishing paint prepared by temperature sensitive paint;
a camera and an excitation light source which are arranged in the parking chamber on the wind tunnel and are matched with the installation position of the metal model;
the power supply module is connected with the excitation light source;
the synchronous controller and the industrial personal computer are arranged outside the wind tunnel residence chamber;
the industrial personal computer is configured to be in communication connection with the synchronous controller and the camera, and the synchronous controller is configured to be in communication connection with the power supply module and the camera;
the primer is configured to employ a white primer comprising silica, and the topcoat is configured to include temperature sensitive probe molecules of trivalent europium fluorescent complexes;
the shooting direction of the camera is configured to be perpendicular to the upper surface of the wing model, the excitation light source is configured to adopt an LED ultraviolet light source, and the optical head irradiates perpendicular to the surface of the wing.
Preferably, the temperature sensitive paint performance parameter is configured to:
the excitation peak spectrum wavelength is 400nm, the emission peak spectrum wavelength is 615nm, the temperature sensitivity is more than 1%/K in 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 gray scale is at least 8 bits, the spatial resolution is 800 multiplied by 600 pixels, the refrigerating device with backboard uses 650nm high-pass filter.
Preferably, the performance parameters of the excitation light source are configured to:
the light filter has a transmittance of more than 90%, and has two irradiation modes of pulse and continuous, the light source control model is TTL, and the output power is 8W-12W.
Preferably, the performance parameters of the synchronous trigger require at least 2 outputs with a control accuracy of less than 20 nanoseconds.
Preferably, after spraying temperature-sensitive paint primer on the surface of a metal model for laminar flow wing test, placing the model in an oven, baking at 90 ℃ for 6 hours for curing, and polishing the cured model primer coating by using 1500-mesh sand paper until the surface roughness of the model is less than 0.8;
cleaning the surface of the metal model, spraying temperature-sensitive paint finish on the primer coating, air-drying and curing for 12 hours at normal temperature, and polishing the cured finish coating of the model by using 1500-mesh sand paper until the surface roughness is less than 0.8.
The invention at least comprises the following beneficial effects: firstly, the test system is constructed aiming at the laminar flow wing model based on the temperature sensitive paint, so that the surface area surface measurement of the laminar flow wing model is realized, the spatial resolution is high, and the accurate transition position of the surface of the laminar flow wing model can be obtained.
Secondly, in the laminar flow wing model in the test system, the temperature sensitive coating is adopted to replace a conventional temperature sensor, so that the interference of the sensor protruding or sinking from the surface of the wing model to a flow field is avoided.
The test system is suitable for accurately measuring the transition position of the laminar flow wing and verifying the design method 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 installation connection diagram of a laminar flow wing transition position measurement test device;
FIG. 2 is a schematic diagram of a temperature sensitive coating structure of the present invention;
fig. 3 is a graph of a transition result of the laminar flow wing in embodiment 1 of the present invention.
Fig. 4 is a graph showing a transition result of a laminar flow airfoil section according to embodiment 1 of the present invention.
Detailed Description
The present invention is described in further detail below with reference to the drawings to enable those skilled in the art to practice the invention by referring to the description.
It will be understood that terms, such as "having," "including," and "comprising," as used herein, do not preclude the presence or addition of one or more other elements or groups thereof.
It should be noted that, in the description of the present invention, the orientation or positional relationship indicated by the term is based on the orientation or positional relationship shown in the drawings, which are merely for convenience of describing the present invention and simplifying the description, and are not indicative or implying that the apparatus or element to be referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as limiting the present invention.
In the description of the present invention, it should be noted that, unless explicitly specified and limited otherwise, the terms "mounted," "provided," "engaged/connected," "connected," and the like are to be construed broadly, and for example, "connected" may be a fixed connection, may be a detachable connection, or may be an integral connection, may be a mechanical connection, may be an electrical connection, may be a direct connection, may be an indirect connection via an intermediary, may be a communication between two elements, and for one of ordinary skill in the art, the specific meaning of the terms in this disclosure may be understood in a specific case.
Fig. 1 shows an implementation form of a laminar flow wing transition position measurement test device according to the present invention, including:
the metal model 2 for laminar flow wing test is arranged in a wind tunnel 1 test section, a primer and a finish paint prepared by temperature sensitive paint are arranged on the metal model, the laminar flow wing model in the structure is a metal model, the surface of the model is sequentially covered with the temperature sensitive paint primer and the temperature sensitive paint finish paint from bottom to top, the coating structure schematic diagram is shown in figure 2, the temperature sensitive paint consists of the temperature sensitive paint primer and the temperature sensitive paint finish paint, the temperature sensitive paint primer is called a substrate emission layer, is a white primer containing silicon dioxide, is sprayed on the surface of the model, plays the roles of improving the cohesiveness of the surface of the model, enhancing the luminous intensity and thermal isolation of probe molecules, the temperature sensitive paint finish paint is called a polymer functional layer, the temperature sensitive probe molecules are trivalent europium fluorescent complex, the main temperature sensitive luminescent material of the temperature sensitive paint is covered on the surface of the model in a spraying mode, the temperature sensitive paint is baked for 6 hours at 90 ℃ in an oven, and after the curing is finished, the primer coating is polished by 1500 meshes until the roughness is less than 0.8; the finish paint is covered on the primer coating in a spraying manner, air-dried and cured for 12 hours at normal temperature, and after curing, the finish paint is coated with 1500-mesh sand paper until the roughness is less than 0.8; the laminar wing model is connected to half-mould supporting mechanisms on the left and right side walls of the wind tunnel test section through left and right support plates, and the left and right support plates are connected and fastened with the left and right rotating windows through screws and pins; the model position is in the range of the uniform area of the wind tunnel axis and the flow field and in the area which can be shot by the camera;
the camera 3 and the excitation light source 4 are arranged in the wind tunnel upper standing room and matched with the installation position of the metal model, in the structure, the camera is a scientific CCD camera, is installed and fixed in the wind tunnel upper standing room and used for collecting light intensity images of the laminar wing model coating, can be provided with lenses with different focal lengths according to shooting distances and shooting areas, the shooting direction of the camera is vertical to the upper surface of the wing model, the excitation light source is an LED ultraviolet light source, is installed and fixed in the wind tunnel upper standing room and used for exciting the model surface coating, and the optical head is vertical to the surface of the wing as much as possible;
the power module 5 is connected with the excitation light source and is used for providing working voltage for the excitation light source under the action of the synchronous trigger and controlling the excitation light source to be in a working state or to be in a non-working state by power failure;
the synchronous controller (synchronous trigger) is arranged outside the wind tunnel residence chamber and is used for setting the period, time delay, pulse width and pulse number of pulse signals, the synchronous controller (synchronous trigger) is used for realizing time sequence control of camera exposure and excitation light sources, the industrial personal computer for data processing is connected with the synchronous trigger and the camera and is used for setting parameters of the synchronous trigger, further controlling the time sequence of excitation light source irradiation and camera exposure, receiving the light intensity image of the surface of the laminar flow wing model shot by the camera, carrying out image post-processing, and obtaining a required airflow transition result image of the surface of the laminar flow wing model. The temperature sensitive paint has photoluminescence and excitation light heat quenching property, the photoluminescence property means that the paint can emit light with another wavelength under the irradiation of excitation light with a certain wavelength, and the heat quenching property means that the emitted light intensity of the paint is reduced along with the temperature rise. Based on the two characteristics of the paint, the temperature-sensitive paint can be sprayed on the surface of the model in principle, the paint is irradiated by excitation light with specific wavelength, the luminous intensity of the paint is converted into the surface temperature of the model by utilizing the photoluminescence and heat quenching characteristics of the paint on the surface of the model, according to the characteristic that the temperature difference between a laminar flow area and a turbulent flow area of the surface of the model is caused by the difference between the thermal convection intensity of laminar flow and turbulent flow, the transition position of airflow on the surface of the model can be judged according to the temperature gradient, so that in the scheme, the temperature-sensitive paint can be used as a sensing method to realize large-area continuous measurement on the surface of the model, a sensor is not required to be installed on the surface of the model, the advantages of surface measurement, high spatial resolution, low processing difficulty and cost of the model design, no disturbance of incoming flow, accurate transition position judgment and the like of the temperature-sensitive paint are fully utilized, and holes and slots are not required, the design, the processing difficulty and cost of the model are reduced, the advantages of high resolution, accurate transition position judgment and the like are realized, and a laminar flow test system based on laminar flow paint technology can be built through the cooperation of other equipment, so that the temperature-sensitive test system based on the laminar flow on the technology can realize the measurement of the transition position of the wing.
The excitation peak spectrum wavelength of the temperature-sensitive coating is 400nm, the emission peak spectrum wavelength is 615nm, the temperature sensitivity is more than 1%/K in the temperature range of 273K-333K, the pressure sensitivity is extremely low, 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 ℃;
the camera is a scientific CCD camera, has high signal-to-noise ratio and gray dynamic range, 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 backboard, and a 650nm high-pass filter is used, so that lenses with different focal lengths can be installed according to shooting distance and shooting area. The camera is arranged in the residence chamber on the wind tunnel and used for collecting the light intensity image of the coating on the surface of the model, and the fixing device can be used for adjusting the position along the front and back, up and down and left and right of the axis of the wind tunnel, so that the camera can be positioned in the range of the axis of the wind tunnel and the observation window of the residence chamber on the wind tunnel. In order to improve the resolution of the image and reduce the distortion of the image, the shooting direction of the camera is adjusted to be vertical to the upper surface of the laminar flow wing model.
The excitation light source is an LED ultraviolet light source, the transmittance of the optical filter is more than 90%, the 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 excitation light source is arranged in the residence chamber on the wind tunnel and used for exciting the surface coating of the model, and the fixing device can be used for adjusting the position along the front and back, up and down and left and right of the axis of the wind tunnel, so that the light source can be positioned in the range of the axis of the wind tunnel and the observation window 8 of the residence chamber on the wind tunnel. In order to improve the intensity of excitation light irradiation, the irradiation direction of the light source is adjusted to be perpendicular to the upper surface of the laminar flow wing model as much as possible. In order to shorten the light source irradiation distance and avoid the influence of refraction and scattering of optical glass, the resident chamber optical observation window on the test section is dismantled, a special adapter plate is installed, a mounting hole is formed in the adapter plate, an excitation light source is inserted into the mounting hole for fixed mounting, and the light source head is flush with the wind tunnel upper wall plate 9.
The synchronous trigger can set the period, time delay, pulse width and pulse number of the pulse signal, is used for realizing the time sequence control of the camera exposure and the excitation light source, requires at least 2 paths of output, and has control precision smaller than 20 nanoseconds.
The data processing industrial personal computer is connected with the synchronous trigger and the camera and is 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, and carrying out image post-processing to obtain a required airflow transition result image of the surface of the laminar flow wing model.
Example 1
The test model of the embodiment is a laminar wing model with a sweepback angle of 20 degrees and a chord length of 200mm, the camera is a scientific CCD camera, the gray dynamic range is 14 bits, the spatial resolution is 1600 multiplied by 1200 pixels, the back plate is used for refrigerating, the adopted lens is an 8mm fixed focus lens, and the adopted filter is a 650nm high-pass filter. The main peak wavelength of the excitation light source is 400nm, the transmittance of the optical filter is more than 90%, the excitation light is irradiated by pulse and continuous two modes, the light source control model is TTL, the optical filter combination form is low-pass + narrow wave, and the output power is 8W-12W. The synchronous trigger can set the period, time delay, pulse width and pulse number of the pulse signals, realizes the time sequence control of the camera exposure and the excitation light source, is 8-way output with single-way input, and has control precision smaller than 10 nanoseconds. The adopted temperature-sensitive coating has an excitation peak spectrum wavelength of 400nm, an emission peak spectrum wavelength of 615nm, a temperature sensitivity of more than 1 percent/K in a temperature range of 273-333K, a coating photodegradation rate of less than 1 percent/min, a shelf life of more than 3 months and an upper limit of a suitable temperature range of more than 60 ℃.
The specific process of the laminar flow wing transition position measurement test in the embodiment is as follows:
a. and processing a laminar flow wing test model and an aluminum sample wafer, wherein the test model is a laminar flow wing model with a sweepback angle of 20 degrees and a chord length of 200mm, and the sample wafer is a round aluminum sample wafer with a diameter of 3cm and a thickness of 2 mm. And (3) grinding the pits such as screw holes on the surface of the model, solidifying and polishing, cleaning the surfaces of the model and the sample by using ethanol or acetone, stirring the temperature-sensitive paint primer and the 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 by using a spray gun, placing the model and the sample in an oven at 90 ℃ for 6 hours for solidifying after the spraying, and polishing the primer coating of the model and the sample by using 1500-mesh sand paper until the roughness is less than 0.8 after the solidifying.
b. And cleaning the surfaces of the model and the sample, stirring the temperature-sensitive paint finishing coat and the solvent until the temperature-sensitive paint finishing coat and the solvent are uniformly dispersed, spraying the temperature-sensitive paint finishing coat on the primer coating, air-drying and curing the finishing coat at normal temperature for 12 hours, and polishing the model and the sample finishing coat by using 1500-mesh sand paper after curing until the roughness is less than 0.8.
c. And placing the sample in a calibration cabin, irradiating the sample by an excitation light source, adjusting the temperature in the calibration cabin, collecting light intensity images of the sample at different temperatures by a camera, and performing post-processing on the images to obtain a calibration relation between the temperature and the light intensity.
d. And the laminar wing model is connected to half-module supporting mechanisms on the left and right side walls of the wind tunnel test section through left and right support plates in a side wall supporting mode, and the left and right support plates are connected and fastened with the left and right rotating windows through screws and pins. The model position should be in the area of the wind tunnel axis and the flow field uniformity region and in the area that can be photographed by the camera.
e. The camera and the excitation light source are installed in the residence chamber on the wind tunnel, the camera and the excitation light source are connected to the sliding rail of the residence chamber through the quick mounting plate, the cradle head and the sliding block in sequence, and the position of the light source and the camera in the front and back, up and down and left and right directions of the axis of the wind tunnel can be adjusted by moving the sliding block along the sliding rail, so that the accurate installation and fixation of the measuring equipment are realized, and 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 residence chamber on the wind tunnel. In order to improve the intensity of excitation light irradiation and the image resolution, reduce image distortion, and adjust the shooting direction of a camera and the irradiation direction of a light source 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, a resident chamber optical observation window on a test section is dismantled, a special adapter plate is installed, an installation hole is formed in the adapter plate, a camera and an excitation light source are fixedly installed in the installation hole in an inserted mode, and a camera lens, a light source and an upper wall plate of a wind tunnel are flush.
f. The method comprises the steps of installing an excitation light source power supply, a synchronous trigger and a data processing industrial personal computer, wherein the excitation light source power supply is installed in a wind tunnel residence chamber, the synchronous trigger and the data processing industrial personal computer are placed on a working platform outside the wind tunnel residence chamber, airflow flows in the wind tunnel residence chamber, and the power supply needs to be fastened. 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. The GigE gigabit network cable, the water cooling pipe and the TTL trigger signal cable are led out through cable holes on the side wall of the residence chamber, and the cables are fixed by strapping tapes.
g. And performing static debugging of 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 determining and the like. The lens parameters include focal length and aperture, the purpose of focusing being to make the image as sharp as possible, and aperture and CCD exposure time determining the image gray level. Since the smaller the aperture, the larger the depth of field, and in order to improve the quality of the image edge, it is necessary to narrow the aperture as much as possible within a reasonable CCD exposure time. The aperture of the camera lens is set to be 12, the exposure time of the CCD is set to be 400ms, and the gray scale light intensity of the reference image is 9000 under the parameter condition, so that the full range of the camera is 2/3. The acquisition time sequence of the synchronous trigger is set to be 100 cycles each time, each acquisition cycle is 600ms, the light source is delayed for 5ms after the synchronous trigger receives the trigger signal, and the camera is delayed for 150ms.
h. And performing joint debugging of the temperature-sensitive paint measuring system and the wind tunnel measurement and control system, simulating normal blowing conditions, closing the resident chamber, and performing shading treatment on observation windows on two sides of the test section.
i. Before the wind tunnel is started, a light source is turned on, a camera collects 20 reference light images, after the collection is completed, the light source is turned off, and the camera collects 20 background images.
j. After the wind tunnel is started and 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 the light source is turned off after the image sequence collection is finished.
k. After one state test is finished, the far infrared baking lamp or the hot air machine is utilized to bake the model, and the next state blowing test can be performed after the surface temperature of the model is recovered uniformly until all the state tests are finished.
And I, performing post-processing on the test acquisition image to obtain a laminar flow wing model surface airflow transition image as shown in fig. 3 and 4.
And the method of post-processing the trial acquisition image is configured to include:
s1, loading a laminar wing background image, a reference image and a test sequence image acquired by a camera, selecting a marking characteristic point, identifying a marking point and positioning the marking point on the reference image and the test sequence image, and storing a positioned marking point coordinate file;
s2, registering the test sequence image to a reference image position according to the coordinate relation of the mark points, checking registration accuracy, if the accuracy meets the standard, storing the registered test sequence image, entering a step S3, and if the accuracy does not meet 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 pre-filtering the reference image after subtracting the background image and the test sequence image;
s4, cutting the reference image and the test sequence image according to the region of the wing model in the image to obtain the cut reference image and the cut test sequence image;
s5, image filling is carried out on temperature-sensitive paint-free areas such as screw holes in the wing model, areas outside the wing model are set as background areas, and filled reference images and test sequence images 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 comparison value sequence image;
s7, converting according to the light intensity ratio sequence image and the temperature sensitive paint calibration coefficient relation to obtain a wing surface temperature data sequence image;
s8, calculating to obtain a laminar flow wing surface heat flow data image according to the temperature data sequence image obtained in the S7;
s9, according to the light intensity ratio image obtained in the S7 or the heat flow data image obtained in the S8, under the condition that the airflow direction is judged to be B according to the light intensity ratio or the gradient change of the heat flow along the chord direction of the laminar wing, as shown in fig. 3 and 4, an accurate wing surface transition area and a transition position A can be obtained.
The above is merely illustrative of a preferred embodiment, but is not limited thereto. In practicing the present invention, appropriate substitutions and/or modifications may be made according to the needs of the user.
The number of equipment and the scale of processing described herein are intended to simplify the description of the present invention. Applications, modifications and variations of the present invention will be readily apparent to those skilled in the art.
Although embodiments of the invention have been disclosed above, they are not limited to the use listed in the specification and embodiments. It can be applied to various fields suitable for the present invention. Additional modifications will readily occur to those skilled in the art. Therefore, the invention is not to be limited to the specific details and illustrations shown and described herein, without departing from the general concepts defined in the claims and their equivalents.
Claims (6)
1. The utility model provides a position measurement test system is twisted to laminar flow wing which characterized in that includes:
the metal model for laminar flow wing test is arranged in the wind tunnel test section, and is provided with primer and finishing paint prepared by temperature sensitive paint;
a camera and an excitation light source which are arranged in the parking chamber on the wind tunnel and are matched with the installation position of the metal model;
the power supply module is connected with the excitation light source;
the synchronous controller and the industrial personal computer are arranged outside the wind tunnel residence chamber;
the industrial personal computer is configured to be in communication connection with the synchronous controller and the camera, and the synchronous controller is configured to be in communication connection with the power supply module and the camera;
the primer is configured to employ a white primer comprising silica, and the topcoat is configured to include temperature sensitive probe molecules of trivalent europium fluorescent complexes;
the shooting direction of the camera is configured to be perpendicular to the upper surface of the wing model, the excitation light source is configured to adopt an LED ultraviolet light source, and the optical head irradiates perpendicular to the surface of the wing;
the specific process of the laminar flow wing transition position measurement test comprises the following steps:
a. processing a laminar flow wing test model and an aluminum sample wafer, wherein the test model is a laminar flow wing model with a sweepback angle of 20 degrees and a chord length of 200mm, and the sample wafer is a round aluminum sample wafer with a diameter of 3cm and a thickness of 2 mm;
b. cleaning the surfaces of the model and the sample, stirring the temperature-sensitive paint finishing coat and the solvent until the temperature-sensitive paint finishing coat is uniformly dispersed, spraying the temperature-sensitive paint finishing coat on the primer coating, air-drying and curing the finishing coat for 12 hours at normal temperature, and polishing the model and the sample finishing coat by using 1500-mesh sand paper after curing until the roughness is less than 0.8;
c. placing the sample in a calibration cabin, irradiating the sample by an excitation light source, adjusting the temperature in the calibration cabin, collecting light intensity images of the sample at different temperatures by a camera, and performing post-processing on the images to obtain a calibration relation between the temperature and the light intensity;
d. the method comprises the steps that a side wall supporting mode is adopted, a laminar wing model is connected to half-module supporting mechanisms of the left side wall and the right side wall of a wind tunnel test section through left and right support plates, the left and right support plates are connected and fastened with left and right rotating windows through screws and pins, and the model position is in the range of a uniform area of a wind tunnel axis and a flow field and in a region which can be shot by a camera;
the method comprises the steps of installing a camera and an excitation light source, wherein the camera and the excitation light source are installed in a resident chamber on a wind tunnel, the camera and the excitation light source are connected to a slide rail of the resident chamber through a quick mounting plate, a cradle head and a slide block in sequence, the positions of the light source and the camera in front and back, up and down and left and right directions of an axis of the wind tunnel can be adjusted by moving the slide block along the slide rail, accurate installation and fixation of measuring equipment are realized, the camera and the light source can be positioned in the range of the axis of the wind tunnel and an observation window of the resident chamber on the wind tunnel, the shooting direction of the camera and the irradiation direction of the light source are adjusted to be perpendicular to the surface of a wing model as much as possible, the resident chamber optical observation window on a test section is removed, a special adapter plate is installed, the camera and the excitation light source are inserted into the installation hole, and a camera lens is flush with an upper wall plate of the wind tunnel;
f. installing an excitation light source power supply, a synchronous trigger and a data processing industrial personal computer;
g. performing static debugging of the temperature-sensitive paint measurement system, wherein the static debugging comprises lens parameter setting, CCD exposure time setting and image acquisition time sequence determination; under the parameter condition, the gray level light intensity of the reference image is 9000, the full range of the camera is 2/3, the acquisition time sequence of the synchronous trigger is set to be 100 cycles each time, each acquisition cycle is 600ms, the light source delay is 5ms after the synchronous trigger receives the trigger signal, and the camera delay is 150ms;
h. performing joint debugging of a temperature-sensitive paint measuring system and a wind tunnel measurement and control system, simulating normal blowing conditions, closing a resident chamber, and performing shading treatment on observation windows on two sides of a test section;
i. before the wind tunnel is started, a light source is turned on, a camera collects 20 reference light images, after the collection is completed, the light source is turned off, and the camera collects 20 background images;
j. after the wind tunnel is started and 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 the light source is turned off after the image sequence collection is finished;
k. after one state test is finished, baking the model by using a far infrared baking lamp or an air heater, and carrying out the next state blowing test after the surface temperature of the model is recovered uniformly until all the state tests are finished;
carrying out post-treatment on the test acquisition image to obtain a laminar flow wing model surface airflow transition image;
the method for post-processing the test acquisition image is configured to include:
s1, loading a laminar wing background image, a reference image and a test sequence image acquired by a camera, selecting a marking characteristic point, identifying a marking point and positioning the marking point on the reference image and the test sequence image, and storing a positioned marking point coordinate file;
s2, registering the test sequence image to a reference image position according to the coordinate relation of the mark points, checking registration accuracy, if the accuracy meets the standard, storing the registered test sequence image, entering a step S3, and if the accuracy does not meet 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 pre-filtering the reference image after subtracting the background image and the test sequence image;
s4, cutting the reference image and the test sequence image according to the region of the wing model in the image to obtain the cut reference image and the cut test sequence image;
s5, image filling is carried out on temperature-sensitive paint-free areas such as screw holes in the wing model, areas outside the wing model are set as background areas, and filled reference images and test sequence images 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 comparison value sequence image;
s7, converting according to the light intensity ratio sequence image and the temperature sensitive paint calibration coefficient relation to obtain a wing surface temperature data sequence image;
s8, calculating to obtain a laminar flow wing surface heat flow data image according to the temperature data sequence image obtained in the S7;
and S9, obtaining an accurate transition region and a transition position A of the wing surface under the condition that the airflow direction is B according to the light intensity ratio image obtained in the S7 or the heat flow data image obtained in the S8 and according to the gradient change of the light intensity ratio or the heat flow along the chord direction of the laminar wing.
2. The laminar flow wing transition position measurement test system of claim 1, in which the temperature sensitive paint performance parameters are configured to:
the excitation peak spectrum wavelength is 400nm, the emission peak spectrum wavelength is 615nm, the temperature sensitivity is more than 1%/K in 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 ℃.
3. The laminar flow wing transition position measurement test system of claim 1, wherein the performance parameters of the camera are configured to:
the dynamic range of gray scale is at least 8 bits, the spatial resolution is 800 multiplied by 600 pixels, the refrigerating device with backboard uses 650nm high-pass filter.
4. The laminar flow wing transition position measurement test system of claim 1, wherein the performance parameters of the excitation light source are configured to:
the light filter has a transmittance of more than 90%, and has two irradiation modes of pulse and continuous, the light source control model is TTL, and the output power is 8W-12W.
5. The laminar flow wing transition position measurement test system according to claim 1, characterized in that the performance parameters of the synchronous trigger require at least 2 paths of output and the control precision is less than 20 nanoseconds.
6. The laminar flow wing transition position measurement test system according to claim 1, characterized in that after the surface of a metal model for laminar flow wing test is sprayed with temperature-sensitive paint primer, the model is placed in an oven to be baked for 6 hours at 90 ℃ for curing, and the cured model primer coating is polished by 1500-mesh sand paper until the surface roughness of the model primer coating is less than 0.8;
cleaning the surface of the metal model, spraying temperature-sensitive paint finish on the primer coating, air-drying and curing for 12 hours at normal temperature, and polishing the cured finish coating of the model by using 1500-mesh sand paper until the surface roughness is less than 0.8.
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| CN114061894B (en) * | 2021-11-12 | 2023-12-19 | 中国空气动力研究与发展中心高速空气动力研究所 | Cavity model pulsating pressure and noise measurement system |
| CN114061895B (en) * | 2021-11-12 | 2023-12-19 | 中国空气动力研究与发展中心高速空气动力研究所 | Image acquisition method for measuring pulsating pressure and noise of cavity model |
| CN113945354B (en) * | 2021-12-14 | 2023-04-07 | 中国空气动力研究与发展中心超高速空气动力研究所 | Test method for identifying flow partition characteristics of acceleration section of expansion wind tunnel |
| CN114459728B (en) * | 2022-04-13 | 2022-06-24 | 中国空气动力研究与发展中心高速空气动力研究所 | Low-temperature-sensitive paint transition measurement test method |
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