CN112945502A - Laminar flow wing transition position measurement test system - Google Patents
Laminar flow wing transition position measurement test system Download PDFInfo
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- CN112945502A CN112945502A CN202110149382.4A CN202110149382A CN112945502A CN 112945502 A CN112945502 A CN 112945502A CN 202110149382 A CN202110149382 A CN 202110149382A CN 112945502 A CN112945502 A CN 112945502A
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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 system for transition position of laminar flow wing, comprising: the metal model for the laminar flow wing test is arranged at the wind tunnel test section, and primer and finish prepared by temperature-sensitive paint are arranged on the metal model; the camera and the excitation light source are arranged in the wind tunnel upper parking chamber and 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 parking chamber; the industrial personal computer is in communication connection with the synchronous controller and the camera, the primer is configured to be a white primer containing silicon dioxide, and the finish paint is configured to be temperature-sensitive probe molecules containing a 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 measurement of a surface area of the laminar flow wing model, has high spatial resolution, and can obtain 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 accurately measuring the transition position of laminar flow wings without damaging the surface of a model 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. The conventional 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 by measuring the temperature of the temperature sensor, and then the transition position of the surface of the model is judged according to the temperature distribution. The device has a number of drawbacks: 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 wing transition position can be measured by an optical method, the device has the advantage of surface measurement, the commonly used measuring devices 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 device, for example, the phase change thermal map and the temperature sensitive liquid crystal can only be used as a semi-quantitative measuring technology, the infrared camera has low spatial resolution, is greatly influenced by the transmissivity of a transparent material in a measuring light path, the phosphorescence thermal map technology is obviously limited by the surface film forming mode and process of an inorganic phosphorescence substance, and the like.
In a word, the conventional temperature sensor transition measuring device 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 devices 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 devices, so that the transition position of the wing model cannot be accurately measured. The temperature-sensitive coating 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 at present, the temperature-sensitive coating is not applied to the measurement of the transition position of the laminar flow wing, a test method is not mature, and a related measurement test device is urgently required 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 testing system is provided, comprising:
the metal model for the laminar flow wing test is arranged at the wind tunnel test section, and primer and finish prepared by temperature-sensitive paint are arranged on the metal model;
the camera and the excitation light source are arranged in the wind tunnel upper parking chamber and 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 parking chamber;
the industrial personal computer is configured to be in communication connection with a synchronous controller and a camera, and the synchronous controller is configured to be in communication connection with a power supply module and the camera;
the primer is configured to adopt a white primer containing silicon dioxide, and the finishing coat is provided with temperature-sensitive probe molecules containing trivalent europium fluorescent complexes;
the shooting direction of the camera is arranged to be perpendicular to the upper surface of the wing model, the excitation light source is arranged to adopt an LED ultraviolet light source, and the optical head irradiates perpendicular to the surface of the wing.
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 8W-12W.
Preferably, the performance parameters of the synchronous trigger require at least 2 outputs and the control precision is less than 20 nanoseconds.
Preferably, after the temperature-sensitive paint primer is sprayed on the surface of the metal model for the laminar flow wing test, the model is placed in an oven and baked for 6 hours at the temperature of 90 ℃ for curing, and the cured model primer coating is polished by 1500-mesh sand paper until the surface roughness 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 finish coating of the cured model by 1500-mesh abrasive paper until the surface roughness is less than 0.8.
The invention at least comprises the following beneficial effects: the test system is built for the laminar flow wing model based on the temperature-sensitive paint, measurement of the surface area of the laminar flow wing model is achieved, the spatial resolution is high, and the accurate transition position of the surface of the laminar flow wing model can be obtained.
Secondly, the laminar flow wing model in the test system adopts the temperature-sensitive coating to replace a conventional temperature sensor, so that the interference of the sensor on the flow field protruding out of or sunken in the surface of the wing model is avoided.
Thirdly, the test system 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, 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 installation and connection of the laminar flow wing transition position measurement test device of the present invention;
FIG. 2 is a schematic structural diagram of the temperature-sensitive coating of the present invention;
fig. 3 is a result graph of transition of laminar flow wings in embodiment 1 of the present invention.
Fig. 4 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.
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 is to be understood that in the description of the present invention, the terms indicating orientation or positional relationship are based on the orientation or positional relationship shown in the drawings, and are used only for convenience in describing the present invention and for simplification of the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus, should not be construed as limiting the present invention.
In the description of the present invention, it should be noted that, unless otherwise specifically stated or limited, the terms "mounted," "disposed," "sleeved/connected," "connected," and the like are used in a broad sense, and for example, "connected" may be a fixed connection, a detachable connection, an integral connection, a mechanical connection, an electrical connection, a direct connection, an indirect connection through an intermediate medium, and a communication between two elements.
Fig. 1 shows an implementation form of a laminar wing transition position measurement test device according to the present invention, which includes:
the metal model 2 for laminar flow wing test arranged at the test section of the wind tunnel 1 is provided with a primer and a finish prepared by temperature-sensitive paint, the laminar flow wing model in the structure is the metal model, the surface of the model is sequentially covered with the temperature-sensitive paint primer and the temperature-sensitive paint finish from bottom to top, the structural schematic diagram of the coating is shown in figure 2, the temperature-sensitive paint consists of the temperature-sensitive paint primer and the temperature-sensitive paint finish, the temperature-sensitive paint primer is called a substrate emission layer and is white primer containing silicon dioxide, the white primer is sprayed on the surface of the model and plays the roles of improving the surface caking property of the model, enhancing the luminous intensity of probe molecules and thermal isolation, the temperature-sensitive paint finish is called a polymer functional layer and contains temperature-sensitive probe molecules, the probe molecules are trivalent fluorescent complexes and are main luminous materials of the temperature-sensitive paint, and the primer is covered, baking the mixture in an oven at the temperature of 90 ℃ for 6 hours for curing, and polishing the primer coating by 1500-mesh sand paper after curing until the roughness is less than 0.8; the finish paint is covered on the primer coating in a spraying mode, air-dried and cured for 12 hours at normal temperature, and the finish paint is coated with 1500-mesh sand paper after curing is finished until the roughness is less than 0.8; the laminar flow wing model is connected to a half-mode supporting mechanism 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, 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 is in the range of the wind tunnel axis and the flow field uniform area and in the area which can be shot by a camera;
the camera 3 and the excitation light source 4 are arranged in the wind tunnel upper dwelling chamber and matched with the installation position of the metal model, in the structure, the camera is a scientific-grade CCD camera, is installed and fixed in the wind tunnel upper dwelling chamber and is used for collecting light intensity images of a coating of a laminar flow wing model, different focal length lenses can be installed according to shooting distance and shooting area, 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 and is installed and fixed in the wind tunnel upper dwelling chamber and used for exciting the coating of the surface of the model, and the optical head is irradiated in a way vertical to the surface of;
the power supply module 5 is connected with the excitation light source and 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 through power failure;
the synchronous controller (synchronous trigger) can set the period, time delay, pulse width and pulse number of pulse signals and is used for realizing the time sequence control of camera exposure and excitation light source, the industrial personal computer for data processing is connected with the synchronous trigger and the camera and is used for setting the parameters of the synchronous trigger, further controlling the time sequence of excitation light source irradiation and camera exposure, receiving the laminar flow wing model surface light intensity image shot by the camera and performing image post-processing to obtain the required laminar flow wing model surface air flow transition result image. The temperature-sensitive coating has the characteristics of photoluminescence and thermal quenching of exciting light, wherein the photoluminescence characteristic means that the coating 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 coating is reduced along with the increase of the temperature. Based on the two characteristics of the coating, the temperature-sensitive coating can be sprayed on the surface of the model in principle, 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 can be judged according to the temperature gradient according to the characteristic that the temperature of a laminar flow region and the temperature of a turbulent flow region are different due to the difference of the thermal convection intensity of laminar flow and turbulent flow, so that in the scheme, the large-area continuous measurement of the surface of the model can be realized by using the temperature-sensitive coating as a sensing method, a sensor is not required to be arranged on the surface of the model, the temperature-sensitive coating has the advantages of surface measurement, high spatial resolution, low model design and processing difficulty and cost, no disturbance of incoming flow, accurate transition, The processing degree of difficulty and cost, it has the resolution ratio height, the outstanding advantage such as position judgement accuracy of twisted, simultaneously through the cooperation of other equipment, has set up a laminar flow wing test system based on temperature sensitive paint technique, can realize the measurement of position is twisted to laminar flow wing.
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 within a 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 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 camera is installed in a parking chamber on the wind tunnel and used for collecting light intensity images of a coating on the surface of the model, and the fixing device can be used for adjusting the position of the camera along the axis of the wind tunnel in the front-back direction, the up-down direction, the left-right direction and the left-right direction, so that the camera can be positioned in the range of the axis of the wind tunnel and the observation window of the parking chamber. In order to improve the image resolution and reduce the image distortion, 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 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 excitation light source is arranged in the wind tunnel upper parking chamber and used for exciting the surface coating of the model, and the fixing device can be used for carrying out position adjustment along the axis of the wind tunnel in the front-back direction, the up-down direction and the left-right direction, 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 wind tunnel upper parking chamber. In order to improve the irradiation intensity of the exciting light, the irradiation direction of the light source is adjusted to be as vertical as possible to the upper surface of the laminar flow wing model. In order to shorten the irradiation distance of the light source and avoid the influence of refraction and scattering of optical glass, the optical observation window of the upper dwelling room of the test section is dismantled, a special adapter plate is installed, the adapter plate is provided with an installation hole, an excitation light source is inserted into the installation hole and fixedly installed, and a light source head is flush with the upper wall plate 9 of the wind tunnel.
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 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, and performing image post-processing to obtain the required image of the laminar flow wing model surface air flow transition result.
Example 1
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.
The specific process of the laminar wing transition position measurement test of the embodiment is as follows:
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 sweep angle of 20 degrees and a chord length of 200mm, and the sample wafer is a round aluminum sample wafer with the diameter of 3cm and the thickness of 2 mm. Grinding the pits such as screw holes on the surface of the model by using putty, solidifying and polishing, cleaning the surfaces of the model and the sample wafer by using ethanol or acetone, stirring the 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 using a spray gun, after the spraying is finished, placing the model and the sample wafer in an oven for baking at 90 ℃ for 6 hours for solidification, and polishing the primer coating of the model and the sample wafer by using 1500-mesh abrasive paper after the solidification is finished 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. The method comprises the steps of placing a sample wafer in a calibration cabin, irradiating the sample wafer by an excitation light source, adjusting the temperature in the calibration cabin, collecting light intensity images of the sample wafer at different temperatures by a camera, and carrying out post-processing on the images to obtain a calibration relation formula of the temperature and the light intensity.
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 the fixation of the measuring equipment are achieved, and the camera and the light source can be located in the range of the axis of. 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, the optical observation window of the upper dwelling room of the test section is disassembled, a special adapter plate is installed, the adapter plate is provided with a mounting hole, a camera and an excitation light source are inserted into the mounting hole for fixed installation, and the camera lens is flush with the light source head and the wind tunnel upper wall plate.
f. The device comprises 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 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 circulation transition image on the surface of the laminar flow wing model shown in fig. 3 and 4.
And the method of post-processing the trial acquisition image is configured to include:
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, obtaining an accurate wing surface transition region and transition position a according to the light intensity ratio image obtained in S7 or the heat flow data image obtained in S8, and determining that the airflow direction is B according to the light intensity ratio or the gradient change of the heat flow along the chord direction of the laminar flow wing, as shown in fig. 3 and 4.
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 (6)
1. A laminar flow wing transition position measurement test system is characterized by comprising:
the metal model for the laminar flow wing test is arranged at the wind tunnel test section, and primer and finish prepared by temperature-sensitive paint are arranged on the metal model;
the camera and the excitation light source are arranged in the wind tunnel upper parking chamber and 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 parking chamber;
the industrial personal computer is configured to be in communication connection with a synchronous controller and a camera, and the synchronous controller is configured to be in communication connection with a power supply module and the camera;
the primer is configured to adopt a white primer containing silicon dioxide, and the finishing coat is provided with temperature-sensitive probe molecules containing trivalent europium fluorescent complexes;
the shooting direction of the camera is arranged to be perpendicular to the upper surface of the wing model, the excitation light source is arranged to adopt an LED ultraviolet light source, and the optical head irradiates perpendicular to the surface of the wing.
2. The laminar wing transition position measurement test system of claim 1, 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 ℃.
3. The laminar wing transition position measurement testing system 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.
4. The system of claim 1, 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 8W-12W.
5. The system of claim 1, wherein the synchronous trigger performance parameter requires at least 2 outputs and a control accuracy of less than 20 ns.
6. The laminar flow wing transition position measurement test system of claim 1, characterized in that, after spraying the temperature sensitive paint primer on the surface of the metal model for laminar flow wing test, the model is placed in an oven and 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 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 finish coating of the cured model by 1500-mesh abrasive paper until the surface roughness is less than 0.8.
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