CN116337401A - Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method - Google Patents
Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method Download PDFInfo
- Publication number
- CN116337401A CN116337401A CN202310036889.8A CN202310036889A CN116337401A CN 116337401 A CN116337401 A CN 116337401A CN 202310036889 A CN202310036889 A CN 202310036889A CN 116337401 A CN116337401 A CN 116337401A
- Authority
- CN
- China
- Prior art keywords
- model
- test model
- wake
- plasma sheath
- aircraft test
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 230000005855 radiation Effects 0.000 title claims abstract description 109
- 230000035939 shock Effects 0.000 title claims abstract description 91
- 238000000034 method Methods 0.000 title claims abstract description 39
- 238000005259 measurement Methods 0.000 title claims abstract description 25
- 238000012360 testing method Methods 0.000 claims abstract description 190
- 230000001360 synchronised effect Effects 0.000 claims abstract description 78
- 239000000725 suspension Substances 0.000 claims abstract description 37
- 230000003287 optical effect Effects 0.000 claims abstract description 27
- 238000010438 heat treatment Methods 0.000 claims abstract description 17
- 238000002679 ablation Methods 0.000 claims abstract description 13
- 238000012545 processing Methods 0.000 claims description 56
- 238000004364 calculation method Methods 0.000 claims description 18
- 238000007906 compression Methods 0.000 claims description 15
- 238000001228 spectrum Methods 0.000 claims description 15
- 230000006835 compression Effects 0.000 claims description 14
- 238000005516 engineering process Methods 0.000 claims description 10
- 239000000463 material Substances 0.000 claims description 10
- 239000000956 alloy Substances 0.000 claims description 8
- 238000000926 separation method Methods 0.000 claims description 8
- 238000003384 imaging method Methods 0.000 claims description 6
- 239000007769 metal material Substances 0.000 claims description 6
- 238000011084 recovery Methods 0.000 claims description 6
- 230000003595 spectral effect Effects 0.000 claims description 6
- 229910000831 Steel Inorganic materials 0.000 claims description 5
- 229910052755 nonmetal Inorganic materials 0.000 claims description 5
- 239000010959 steel Substances 0.000 claims description 5
- 238000000691 measurement method Methods 0.000 claims description 4
- 239000006185 dispersion Substances 0.000 claims description 3
- 230000001105 regulatory effect Effects 0.000 claims description 3
- 239000007921 spray Substances 0.000 claims description 3
- 238000011160 research Methods 0.000 abstract description 7
- 230000001052 transient effect Effects 0.000 abstract description 4
- 238000013461 design Methods 0.000 abstract description 2
- 238000002474 experimental method Methods 0.000 description 2
- 238000004088 simulation Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 238000006073 displacement reaction Methods 0.000 description 1
- 238000010494 dissociation reaction Methods 0.000 description 1
- 230000005593 dissociations Effects 0.000 description 1
- 230000005611 electricity Effects 0.000 description 1
- 230000005284 excitation Effects 0.000 description 1
- 239000005357 flat glass Substances 0.000 description 1
- 239000011521 glass Substances 0.000 description 1
- 239000013307 optical fiber Substances 0.000 description 1
- 238000002310 reflectometry Methods 0.000 description 1
Images
Classifications
-
- 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
-
- 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
- G01M9/02—Wind tunnels
- G01M9/04—Details
-
- G—PHYSICS
- G01—MEASURING; TESTING
- G01N—INVESTIGATING OR ANALYSING MATERIALS BY DETERMINING THEIR CHEMICAL OR PHYSICAL PROPERTIES
- G01N21/00—Investigating or analysing materials by the use of optical means, i.e. using sub-millimetre waves, infrared, visible or ultraviolet light
- G01N21/17—Systems in which incident light is modified in accordance with the properties of the material investigated
- G01N21/25—Colour; Spectral properties, i.e. comparison of effect of material on the light at two or more different wavelengths or wavelength bands
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T90/00—Enabling technologies or technologies with a potential or indirect contribution to GHG emissions mitigation
Abstract
The invention discloses a system and a method for measuring the radiation characteristics of a plasma sheath and wake light of a high enthalpy shock tunnel test model, and belongs to the field of hypersonic aerodynamics tests. The method is characterized in that a large-size high-enthalpy shock tunnel is utilized, a suspension type heavy-model electromagnetic synchronous release system is adopted to realize free flight of a heavy model without interference wake, a heating device and related design are adopted to realize heavy-model heating so as to simulate an ablation model, a spectrometer system and a radiometer system are adopted to realize transient optical radiation characteristic test, and further measurement of optical radiation characteristics of a plasma sheath and the wake thereof is realized. The invention not only can research the light radiation characteristics of the plasma sheath in the front edge area of the model, but also can research the light radiation characteristics of the model trail, and can accurately measure the light radiation characteristics of the plasma sheath and trail of the ablation model.
Description
Technical Field
The invention relates to a system and a method for measuring the radiation characteristics of a plasma sheath and wake light of a high enthalpy shock tunnel model, belonging to the field of hypersonic aerodynamics test.
Background
When the high-speed aircraft flies at an ultra-high speed in a near space, surrounding gas is heated by shock waves or is blocked and decelerated by viscosity, so that the random movement energy of molecules is increased, high temperature of thousands DEG C is generated to cause the vibration excitation dissociation and even ionization of gas molecules, the hypersonic gas is caused to present real gas properties, and the aircraft and the air interact to generate physical phenomena of mechanics, thermal, optics and electricity. The physical phenomena of light radiation generated by aircrafts with different characteristics and flow fields thereof are not identical. The limitation of hypersonic flow physical phenomenon in knowledge causes inappropriate physical modeling, and further causes certain uncertainty in numerical calculation, so that research on hypersonic flow rules through experiments is very important.
At present, when the high enthalpy shock tunnel is used for researching radiation characteristic rules, a fixed supporting mode is generally adopted, the model is interfered by a supporting mechanism, the difficulty of measuring flow field parameters is improved, and the measurement precision and accuracy of the photoelectric characteristic of the wake of the plasma sheath of the model are affected. In addition, current measurements are only carried out for conventional models, lack of measurement for ablation models, and cannot accurately simulate the high Wen Moxing optical radiation characteristics under flight conditions.
Disclosure of Invention
The technical solution of the invention is as follows: the system and the method for measuring the light radiation characteristics of the plasma sheath and the wake of the high-enthalpy shock tunnel model are provided, and the problem that the light radiation characteristics of the plasma sheath and the wake of the high-speed aircraft are difficult to accurately measure is solved.
The technical scheme of the invention is as follows:
in a first aspect, the invention provides a shock tunnel plasma sheath and wake light radiation characteristic measurement system, which comprises a high enthalpy shock tunnel, a spectrometer system, a radiometer system, a suspension type heavy model electromagnetic synchronous release system, a pressure sensor, a synchronous acquisition system and a calculation processing system;
high enthalpy shock tunnel: for providing a flow field for an aircraft test model;
suspension type heavy model electromagnetic synchronous release system: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received, and the aircraft test model is recovered;
a pressure sensor: the device is arranged on a compression pipe and a shock pipe of the high enthalpy shock tunnel and is used for collecting pressure signals of the compression pipe and the shock pipe after the operation of the tunnel and sending the pressure signals to a synchronous collecting system;
synchronous acquisition system: after receiving the pressure signal, simultaneously sending a release signal to the suspension type heavy-model electromagnetic synchronous release system and sending a starting signal to the spectrometer system and the radiometer system;
spectrometer system: after the model plasma sheath and the wake spectrum information are started, the model plasma sheath and the wake spectrum information are collected and sent to a computing processing system; radiometer system: collecting optical radiation information of the model plasma sheath and the wake after starting, and sending the information to a computing processing system;
the computing processing system comprises: and performing data processing on the optical information and the radiation information to obtain the plasma sheath and the trail optical radiation characteristics of the aircraft test model.
Preferably, the suspension type heavy-model electromagnetic synchronous release system is positioned in a Gao Han shock tunnel test section and comprises a release model device fixed above an attack angle mechanism and a buffer capture model device fixed below the attack angle mechanism;
release model device: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received;
buffer capture model device: and the nondestructive recovery of the aircraft test model is realized.
Preferably, the release model device comprises a signal processing device, an electromagnetic valve, a power supply and an electromagnet; the power supply is connected with the electromagnet through the electromagnetic valve, the electromagnet is used for adsorbing the aircraft test model, and the signal processing device is connected with the electromagnetic valve; when the signal processing device receives the release signal, the electromagnetic valve is controlled to be disconnected, and the electromagnetic valve loses the magnetic force to release the aircraft test model;
the buffer capture model device is a flexible steel wire mesh.
Preferably, the aircraft test model is made of a metal material or a nonmetal material; if the aircraft test model is made of a nonmetallic material, an alloy material element is placed in the cavity of the aircraft test model.
In a second aspect, the invention provides a method for measuring the radiation characteristics of a plasma sheath and a wake of a shock tunnel, which comprises the following steps:
calibrating a spectrometer system and a radiometer system;
determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, so as to ensure that the aircraft test model does not move simultaneously with the wind tunnel in the wind tunnel operation process;
installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
after receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
after receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the model plasma sheath and wake spectrum information and sends the information to the computing and processing system; the radiometer system collects the optical radiation information of the model plasma sheath and the trail and sends the information to the calculation processing system; recovering the aircraft test model by the suspension type heavy model electromagnetic synchronous release system;
and the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
In a third aspect, the invention provides a method for measuring the radiation characteristics of a plasma sheath and a wake of a shock tunnel, comprising the following steps:
calibrating a spectrometer system and a radiometer system;
determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, so as to ensure that the aircraft test model does not move simultaneously with the wind tunnel in the wind tunnel operation process;
installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
heating the aircraft test model to a predetermined temperature with a heating device to simulate an ablation model;
starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
after receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
after receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the model plasma sheath and wake spectrum information and sends the information to the computing and processing system; the radiometer system collects the optical radiation information of the model plasma sheath and the trail and sends the information to the calculation processing system; recovering the aircraft test model by the suspension type heavy model electromagnetic synchronous release system;
and the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
Preferably, the heating device comprises a direct current power supply arranged at the top of the wind tunnel test section, an electrode connected with the direct current power supply and a driving device; when the air vehicle is heated, the driving device drives the electrode to move downwards to be in contact with the air vehicle test model, the direct-current power supply is started, the air vehicle test model is heated through the electrode, and after the air vehicle test model reaches the preset temperature, the driving device drives the electrode to move upwards to be separated from the air vehicle test model.
Preferably, the surface temperature of the aircraft test model is measured by adopting a thermal environment infrared heat map technology, and the measurement method comprises the following steps:
the method comprises the steps of performing field calibration on the thermal infrared imager by using a surface source black body, wherein the surface source black body is arranged on the central line of a wind tunnel test section during calibration, and the radiation surface of the surface source black body is positioned at the center of a field of view of the thermal infrared imager;
the calibrated thermal infrared imager acquires surface level values of the aircraft test model at different moments and sends the surface level values to the data processor;
the data processor fits a temperature-level curve of the thermal infrared imager according to the calibration data, and obtains the surface temperatures of the aircraft test model at different moments according to the surface level values acquired by the thermal infrared imager.
Preferably, the aircraft test model is made of a metal material or a nonmetal material; if the aircraft test model is made of a nonmetallic material, an alloy material element is placed in the cavity of the aircraft test model.
Preferably, the relative radiation intensity calibration is carried out on each wavelength measurement wave band of the spectrometer system by using a high-temperature blackbody furnace, so as to obtain the relative response rate of each pixel of the CCD in different measurement wave bands.
Preferably, during spectrum measurement, according to the wavelength range required to be measured, a required grating incident angle is calculated by using a dispersion equation, and the grating angle of the spectrometer is adjusted to meet the grating incident angle requirement.
Preferably, the radiometer system comprises an infrared radiometer, an infrared radiation imaging system, an ultraviolet radiometer and a visible light radiometer;
the infrared radiometer adopts a medium-temperature cavity type blackbody furnace for calibration; the infrared radiation imaging system is calibrated by a large-area blackbody furnace; the ultraviolet radiation meter and the visible light radiometer are calibrated by adopting a high-brightness broadband white light source.
Preferably, the length L of the aircraft test model is determined by:
L≤(10～20)tV ∞
wherein t is the effective test time of the high enthalpy shock tunnel, V ∞ Is the free flow velocity at the outlet of the spray pipe.
Preferably, after the aircraft test model is freely put into the wind tunnel stable flow field, a schlieren photo of the aircraft test model is obtained by utilizing a laser schlieren technology, the separation shock wave distance and wake of the aircraft test model are measured, whether the flow field flows through the model for more than two wake periods is judged, if so, the test is continued, and if not, the wind tunnel running state is regulated until the flow field flows through the model for more than two wake periods.
Preferably, the ratio between the separation shock distance Δ, the aircraft test model diameter D and the schlieren photo pixels is as follows:
wherein Pe1 is standing point shock wave separation distance pixel along the horizontal center line of the aircraft test model, and Pe2 is model diameter pixel judged and read along the vertical center line of the aircraft test model.
Preferably, the response time t of the suspension type heavy-model electromagnetic synchronous release system 1 Wind tunnel run time t 2 Time t required by sending pressure signals acquired by pressure sensors to synchronous acquisition system 3 Satisfies the following relationship
t 1 ＜t 3 -t 2 。
Preferably, the suspension type heavy-model electromagnetic synchronous release system is positioned in a Gao Han shock tunnel test section and comprises a release model device fixed above an attack angle mechanism and a buffer capture model device fixed below the attack angle mechanism;
release model device: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received;
buffer capture model device: and the nondestructive recovery of the aircraft test model is realized.
Preferably, the release model device comprises a signal processing device, an electromagnetic valve, a power supply and an electromagnet; the power supply is connected with the electromagnet through the electromagnetic valve, the electromagnet is used for adsorbing the aircraft test model, and the signal processing device is connected with the electromagnetic valve; when the signal processing device receives the release signal, the electromagnetic valve is controlled to be disconnected, the electromagnet loses magnetic force, and the aircraft test model is released;
the buffer capture model device is a flexible steel wire mesh.
The invention provides a high enthalpy shock tunnel model plasma sheath and wake light radiation characteristic measurement system and method, which has the beneficial effects compared with the prior art that:
according to the invention, a large-size high-enthalpy shock tunnel is utilized, a suspension type heavy model electromagnetic synchronous release system is adopted to realize free flight of an aircraft test model without interference wake, the interference problem of a model supporting mechanism on the wake is solved, a spectrometer system and a radiometer system are adopted to realize transient light radiation characteristic test, and the measurement precision and accuracy of the photoelectric characteristic of the model plasma sheath wake are ensured.
The heating device is designed to heat the aircraft test model to simulate the ablation model, and the accurate simulation of the high Wen Moxing light radiation characteristic under the flight condition is realized.
The invention not only can research the light radiation characteristics of the plasma sheath in the front edge area of the model, but also can research the light radiation characteristics of the model trail, and can accurately measure the light radiation characteristics of the plasma sheath and trail of the ablation model.
Drawings
FIG. 1 is a schematic flow chart of a method for measuring the radiation characteristics of a plasma sheath and wake light of a high enthalpy shock tunnel model;
FIG. 2 is a schematic diagram of a system for measuring the radiation characteristics of a plasma sheath and wake light of a high enthalpy shock tunnel model according to the invention; wherein (a) is a horizontal section and (b) is a vertical section;
the system comprises a 1-compression tube, a 2-shock tube, a 3-spectrometer system, a radiometer system, a 4-grating, a 5-calculation processing system, a 6-synchronous acquisition system, a 7-release model device, an 8-aircraft test model and a 9-buffer capture model device.
Detailed Description
The features and advantages of the present invention will become more apparent and apparent from the following detailed description of the invention.
The invention provides a high enthalpy shock tunnel model plasma sheath and wake light radiation characteristic measurement system and method, and provides a feasible thought verified by experiments for accurate measurement of pure air non-interference model plasma sheath and wake light radiation characteristics. The aircraft test model adopts a heavy model to ensure that the displacement of the model in a flow field is as small as possible, and the air flow does not influence the attitude angle of the model.
And measuring the optical radiation characteristics of the plasma sheath and the wake thereof by utilizing a large-size high-enthalpy shock tunnel and combining a heavy model free flight technology without interference wake, a heavy model heating technology and a transient optical radiation characteristic test technology. The invention not only can research the light radiation characteristics of the plasma sheath in the front edge area of the model, but also can research the light radiation characteristics of the model trail, and can accurately measure the light radiation characteristics of the plasma sheath and trail of the ablation model.
As shown in fig. 2, the high enthalpy shock tunnel model plasma sheath and wake light radiation characteristic measurement system comprises a high enthalpy shock tunnel, a spectrometer system, a radiometer system 3, a suspension type heavy model electromagnetic synchronous release system, a pressure sensor, a synchronous acquisition system 6 and a calculation processing system 5.
High enthalpy shock tunnel: the nozzle outlet size is 1.2 m-2.5 m, which is used for providing a flow field for the aircraft test model 8.
Suspension type heavy model electromagnetic synchronous release system: before the wind tunnel is started, suspending the aircraft test model 8 in the wind tunnel; after the wind tunnel is started, the aircraft test model 8 is released after the release signal is received.
A pressure sensor: the synchronous acquisition system is arranged on a compression pipe 1 and a shock tube 2 of the high enthalpy shock tunnel and used for acquiring pressure signals of the compression pipe 1 and the shock tube 2 after the operation of the tunnel and transmitting the pressure signals to the synchronous acquisition system 6.
Synchronous acquisition system 6: after receiving the pressure signal, simultaneously sending a release signal to the suspended heavy-model electromagnetic synchronous release system and sending a start signal to the spectrometer system and the radiometer system 3.
Spectrometer system: and after the starting, acquiring model plasma sheath and wake spectrum information, and sending the information to the computing and processing system 5.
Radiometer system: and after the start, collecting optical radiation information of the model plasma sheath and the wake and sending the information to the computing and processing system 5.
The computing processing system 5: and performing data processing on the optical information and the radiation information to obtain the plasma sheath and the trail optical radiation characteristics of the aircraft test model.
The suspension type heavy-model electromagnetic synchronous release system is positioned in a Gao Han shock tunnel test section and comprises a release model device 7 fixed above an attack angle mechanism and a buffer capture model device 9 fixed below the attack angle mechanism. Release model device: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received; buffer capture model device: and the nondestructive recovery of the aircraft test model is realized.
The release model device 7 comprises a signal processing device, an electromagnetic valve, a power supply and an electromagnet; the power supply is connected with the electromagnet through the electromagnetic valve, the electromagnet is used for adsorbing the aircraft test model, and the signal processing device is connected with the electromagnetic valve; when the signal processing device receives the release signal, the electromagnetic valve is controlled to be disconnected, the electromagnet loses magnetic force, and the aircraft test model is released. The buffer capture model device 9 is a flexible steel wire mesh.
The aircraft test model is made of a metal material or a nonmetal material; if the aircraft test model is made of a nonmetallic material, an alloy material element is placed in the cavity of the aircraft test model.
The system is used for measuring the model plasma sheath and the trail light radiation characteristics. The model plasma sheath and the wake light radiation characteristic measuring method can be used for a common model and an ablation model. Fig. 1 is a flow chart of a method for measuring the radiation characteristics of a plasma sheath and wake light of a high enthalpy shock tunnel model.
Specifically, the method for measuring the radiation characteristics of the plasma sheath and the wake light of the common model comprises the following steps:
(1) Calibrating a spectrometer system and a radiometer system;
(2) Determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
(3) Suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, and fixing the aircraft test model on an attack angle mechanism to ensure that the aircraft test model does not move simultaneously with a wind tunnel in the wind tunnel operation process;
(4) Installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
(5) Starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
(6) After receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
(7) After receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the spectrum information of the flow field around the model and sends the spectrum information to the calculation processing system; the radiometer system collects the light radiation information of the flow field around the model and sends the light radiation information to the computing and processing system; after the aircraft test model is released, recovering the aircraft test model by using a suspension type heavy model electromagnetic synchronous release system;
(8) And the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
The method for measuring the radiation characteristics of the plasma sheath and the wake light aiming at the ablation model comprises the following steps:
(1) Calibrating a spectrometer system and a radiometer system;
(2) Determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
(3) Suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, and fixing the aircraft test model on an attack angle mechanism to ensure that the aircraft test model does not move simultaneously with a wind tunnel in the wind tunnel operation process;
(4) Installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
(5) Heating the aircraft test model to a predetermined temperature with a heating device to simulate an ablation model;
(6) Starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
(7) After receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
(8) After receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the spectrum information of the flow field around the model and sends the spectrum information to the calculation processing system; the radiometer system collects the light radiation information of the flow field around the model and sends the light radiation information to the computing and processing system; after the aircraft test model is released, recovering the aircraft test model by using a suspension type heavy model electromagnetic synchronous release system;
(9) And the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
When the aircraft test model is heated to a preset temperature by using the heating device, if the model is made of a metal material, the power supply voltage directly heats the model; for an object incapable of being directly heated by resistance, an indirect resistance heating mode can be adopted, a special alloy material is made into an alloy material element, and the alloy material element is placed in a mold cavity.
Specifically, the heating device comprises a direct current power supply arranged at the top of the wind tunnel test section, an electrode connected with the direct current power supply and a driving device; when the air vehicle is heated, the driving device drives the electrode to move downwards to be in contact with the air vehicle test model, the direct-current power supply is started, the air vehicle test model is heated through the electrode, and after the air vehicle test model reaches the preset temperature, the driving device drives the electrode to move upwards to be separated from the air vehicle test model. The model moves downwards in the test and enters the range of the flow field; the electrode remains above the test section and does not enter the flow field. The DC power supply device is placed on the top of the test section.
The thermal environment infrared thermal image technology is adopted to measure the surface temperature of the aircraft test model, the infrared thermal imager observes the radiation of the blackbody through the infrared glass, so that the transmissivity and the reflectivity loss of the infrared window glass do not need to be corrected, and finally, the data processor gives a fitting curve of the temperature-level value of the infrared thermal imager. The specific measurement method is as follows:
(1) And (3) calibrating the thermal infrared imager on site by using a surface source black body, wherein the surface source black body is arranged on the central line of the wind tunnel test section during calibration, and the radiation surface of the surface source black body is positioned at the center of the field of view of the thermal infrared imager.
(2) And the calibrated thermal infrared imager acquires surface level values of the aircraft test model at different moments and sends the surface level values to the data processor.
(3) The data processor fits a temperature-level curve of the thermal infrared imager according to the calibration data, and obtains the surface temperatures of the aircraft test model at different moments according to the surface level values acquired by the thermal infrared imager.
When the optical radiation measurement is carried out in the test section, the adopted spectrometer system and the radiometer system need to be calibrated. The spectrometer system has different measuring wave band ranges, and the relative radiation intensity calibration is carried out on each wavelength measuring range of the spectrometer by using a high-temperature blackbody furnace (the temperature is more than 500 degrees), so that the relative response rate of each pixel of the CCD in different measuring wave bands is obtained. During test measurement, according to the wavelength range required to be measured, a required grating incident angle is calculated by using a dispersion equation, and the angle of the grating 4 is adjusted to meet the requirement of calculating the incident angle.
The radiometer system includes an infrared radiometer, an infrared radiation imaging system, an ultraviolet radiometer, and a visible light radiometer. Radiometers of different wavebands are calibrated differently. The infrared radiometer adopts a medium-temperature (300-500 DEG) cavity type blackbody furnace for calibration, and the infrared radiation imaging system uses a large area (the area is larger than 1 m) 2 ) The blackbody furnace is calibrated, and the ultraviolet radiation meter and the visible light radiation meter are calibrated by adopting a high-brightness (the brightness is more than 100 ten thousand illuminance) broadband white light source.
When the model is initially designed, numerical simulation is required to be carried out by utilizing a Navier-Stokes equation with a plurality of components and a plurality of temperatures, the time required for forming the model trail is initially obtained, and then the size of the model is judged. And evaluating the size of the test model according to the effective test time of the high enthalpy shock tunnel flow field calibration. When measuring the wake, the wake time needs to be more than 2 cycles. The test model length is determined by the effective run time and the nozzle outlet free flow velocity as shown in the following formula:
L≤(10～20)tV ∞
wherein t is the effective test time of the high enthalpy shock tunnel, V ∞ Is the free flow velocity at the outlet of the spray pipe.
After the aircraft test model is freely put into the wind tunnel stable flow field, a schlieren photo of the aircraft test model is obtained by utilizing a laser schlieren technology, the separation shock wave distance and wake of the aircraft test model are measured, whether the flow field flowing through the model is longer than the time for forming two wake periods or not is judged, if so, the test is continued, and if not, the wind tunnel running state is regulated until the flow field flowing through the model is longer than the time for forming two wake periods or not.
The ratio between the disjunct shock distance delta and the model diameter D and the schlieren photo pixels is as follows:
wherein Pe1 is standing point shock wave separation distance pixel along the horizontal center line of the model, and Pe2 is model diameter pixel judged and read along the vertical center line of the model.
And (3) constructing a suspension type heavy model electromagnetic synchronous release system, and suspending the model in the wind tunnel. The heavy model electromagnetic synchronous release system needs to be matched with the wind tunnel operation time for constraint, and the response time t of the heavy model electromagnetic synchronous release system 1 Wind tunnel run time t 2 And the time t required by the pressure sensor to acquire the pressure signal and send the pressure signal to the synchronous acquisition system 3 Satisfies the following relationship
t 1 ＜t 3 -t 2 。
And a release model device and a buffer capture model device are established, so that instantaneous release and lossless recovery of the heavy model are realized.
And (3) a series of high-frequency high-temperature high-pressure sensors are arranged at the corresponding positions of the compression pipe and the shock tube according to the sensor mounting interfaces reserved in the compression pipe and the shock tube. The pressure sensor is connected with the synchronous acquisition system, and after signal triggering, the acquired data is connected into the calculation processing system through an optical fiber. The frequency of the pressure sensor is more than 1MHz, the temperature is more than 2000K, and the pressure is more than 10MPa.
Aiming at the difficult problem that the optical radiation characteristics of a plasma sheath and a wake of a high-speed aircraft are difficult to accurately measure, the invention utilizes a large-size high-enthalpy shock tunnel, adopts a suspension type heavy-model electromagnetic synchronous release system to realize free flight of a heavy model without interference wake, adopts a heating device and related design to realize heavy-model heating so as to simulate an ablation model, adopts a spectrometer system and a radiometer system to realize transient optical radiation characteristic test so as to develop a quantitative measurement method of the optical radiation characteristics of the model of the large-size aircraft, and solves the difficult problem of accurate measurement of the plasma sheath and the wake of the model.
The method can not only study the light radiation characteristics of the plasma sheath in the front edge area of the model, but also develop the study of the light radiation characteristics of the model trail, and can accurately measure the light radiation characteristics of the plasma sheath and the trail of the ablation model, thereby being a better method.
What is not described in detail in the present specification belongs to the known technology of those skilled in the art.
Claims (18)
1. The shock tunnel plasma sheath and wake light radiation characteristic measurement system is characterized in that: the system comprises a high enthalpy shock tunnel, a spectrometer system, a radiometer system, a suspension type heavy model electromagnetic synchronous release system, a pressure sensor, a synchronous acquisition system and a calculation processing system;
high enthalpy shock tunnel: for providing a flow field for an aircraft test model;
suspension type heavy model electromagnetic synchronous release system: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received, and the aircraft test model is recovered;
a pressure sensor: the device is arranged on a compression pipe and a shock pipe of the high enthalpy shock tunnel and is used for collecting pressure signals of the compression pipe and the shock pipe after the operation of the tunnel and sending the pressure signals to a synchronous collecting system;
synchronous acquisition system: after receiving the pressure signal, simultaneously sending a release signal to the suspension type heavy-model electromagnetic synchronous release system and sending a starting signal to the spectrometer system and the radiometer system;
spectrometer system: after the model plasma sheath and the wake spectrum information are started, the model plasma sheath and the wake spectrum information are collected and sent to a computing processing system; radiometer system: collecting optical radiation information of the model plasma sheath and the wake after starting, and sending the information to a computing processing system;
the computing processing system comprises: and performing data processing on the optical information and the radiation information to obtain the plasma sheath and the trail optical radiation characteristics of the aircraft test model.
2. The shock tunnel plasma sheath and wake optical radiation characteristic measurement system of claim 1, wherein: the suspension type heavy-model electromagnetic synchronous release system is positioned in a Gao Han shock tunnel test section and comprises a release model device fixed above an attack angle mechanism and a buffer capture model device fixed below the attack angle mechanism;
release model device: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received;
buffer capture model device: and the nondestructive recovery of the aircraft test model is realized.
3. The shock tunnel plasma sheath and wake optical radiation characteristic measurement system of claim 2, wherein: the release model device comprises a signal processing device, an electromagnetic valve, a power supply and an electromagnet; the power supply is connected with the electromagnet through the electromagnetic valve, the electromagnet is used for adsorbing the aircraft test model, and the signal processing device is connected with the electromagnetic valve; when the signal processing device receives the release signal, the electromagnetic valve is controlled to be disconnected, and the electromagnetic valve loses the magnetic force to release the aircraft test model;
the buffer capture model device is a flexible steel wire mesh.
4. The shock tunnel plasma sheath and wake optical radiation characteristic measurement system of claim 1, wherein: the aircraft test model is made of a metal material or a nonmetal material; if the aircraft test model is made of a nonmetallic material, an alloy material element is placed in the cavity of the aircraft test model.
5. The method for measuring the radiation characteristics of the plasma sheath and the wake light of the shock tunnel is characterized by comprising the following steps:
calibrating a spectrometer system and a radiometer system;
determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, so as to ensure that the aircraft test model does not move simultaneously with the wind tunnel in the wind tunnel operation process;
installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
after receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
after receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the model plasma sheath and wake spectrum information and sends the information to the computing and processing system; the radiometer system collects the optical radiation information of the model plasma sheath and the trail and sends the information to the calculation processing system; recovering the aircraft test model by the suspension type heavy model electromagnetic synchronous release system;
and the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
6. The method for measuring the radiation characteristics of the plasma sheath and the wake light of the shock tunnel is characterized by comprising the following steps:
calibrating a spectrometer system and a radiometer system;
determining the size of an aircraft test model according to the effective test time of the high enthalpy shock tunnel flow field calibration and the size of a uniform area of the flow field, and designing the aircraft test model;
suspending the aircraft test model in a high enthalpy shock wind tunnel test section by using a suspension type heavy model electromagnetic synchronous release system, so as to ensure that the aircraft test model does not move simultaneously with the wind tunnel in the wind tunnel operation process;
installing pressure sensors at the corresponding positions of the compression pipe and the shock wave pipe, wherein the pressure sensors are connected with a synchronous acquisition system;
heating the aircraft test model to a predetermined temperature with a heating device to simulate an ablation model;
starting the wind tunnel, and sending the pressure signal acquired by the pressure sensor to the synchronous acquisition system;
after receiving the pressure signal, the synchronous acquisition system simultaneously sends a release signal to the suspended heavy-model electromagnetic synchronous release system and sends a starting signal to the spectrometer system and the radiometer system;
after receiving the release signal, the suspension type heavy model electromagnetic synchronous release system freely puts the aircraft test model into the wind tunnel stable flow field; the spectrometer system collects the model plasma sheath and wake spectrum information and sends the information to the computing and processing system; the radiometer system collects the optical radiation information of the model plasma sheath and the trail and sends the information to the calculation processing system; recovering the aircraft test model by the suspension type heavy model electromagnetic synchronous release system;
and the calculation processing system performs data processing on the acquired spectral information and the acquired radiation information to obtain the plasma sheath and wake light radiation characteristics of the aircraft test model.
7. The method for measuring the radiation characteristics of the plasma sheath and the wake of the shock tunnel according to claim 6, wherein the heating device comprises a direct current power supply arranged at the top of the test section of the tunnel, an electrode connected with the direct current power supply and a driving device; when the air vehicle is heated, the driving device drives the electrode to move downwards to be in contact with the air vehicle test model, the direct-current power supply is started, the air vehicle test model is heated through the electrode, and after the air vehicle test model reaches the preset temperature, the driving device drives the electrode to move upwards to be separated from the air vehicle test model.
8. The method for measuring the characteristics of the plasma sheath and the wake light radiation of the shock tunnel according to claim 7, wherein the surface temperature of the aircraft test model is measured by adopting a thermal environment infrared heat map technology, and the measuring method comprises the following steps:
the method comprises the steps of performing field calibration on the thermal infrared imager by using a surface source black body, wherein the surface source black body is arranged on the central line of a wind tunnel test section during calibration, and the radiation surface of the surface source black body is positioned at the center of a field of view of the thermal infrared imager;
the calibrated thermal infrared imager acquires surface level values of the aircraft test model at different moments and sends the surface level values to the data processor;
the data processor fits a temperature-level curve of the thermal infrared imager according to the calibration data, and obtains the surface temperatures of the aircraft test model at different moments according to the surface level values acquired by the thermal infrared imager.
9. The method for measuring the radiation characteristics of the plasma sheath and the wake of the shock tunnel according to claim 5 or 6, wherein the aircraft test model is made of a metal material or a nonmetal material; if the aircraft test model is made of a nonmetallic material, an alloy material element is placed in the cavity of the aircraft test model.
10. The method for measuring the radiation characteristics of the shock tunnel plasma sheath and the wake light according to claim 5 or 6, wherein the relative radiation intensity calibration is performed on each wavelength measurement wave band of the spectrometer system by using a high-temperature blackbody furnace, so as to obtain the relative response rate of each pixel of the CCD in different measurement wave bands.
11. The method for measuring the characteristics of the plasma sheath and the wake light radiation of the shock tunnel according to claim 10, wherein during the spectrum measurement, a required grating incident angle is calculated by using a dispersion equation according to a wavelength range required to be measured, and the grating angle of the spectrometer is adjusted so as to meet the requirement of the grating incident angle.
12. The shock tunnel plasma sheath and wake radiation characteristic measurement method of claim 5 or 6, wherein the radiometer system comprises an infrared radiometer, an infrared radiation imaging system, an ultraviolet radiometer, and a visible radiometer;
the infrared radiometer adopts a medium-temperature cavity type blackbody furnace for calibration; the infrared radiation imaging system is calibrated by a large-area blackbody furnace; the ultraviolet radiation meter and the visible light radiometer are calibrated by adopting a high-brightness broadband white light source.
13. The method of measuring the characteristics of the plasma sheath and the wake radiation of a shock tunnel according to claim 5 or 6, wherein the length L of the aircraft test model is determined by the following formula:
L≤(10～20)tV ∞
wherein t is the effective test time of the high enthalpy shock tunnel, V ∞ Is the free flow velocity at the outlet of the spray pipe.
14. The method for measuring the plasma sheath and wake light radiation characteristics of a shock tunnel according to claim 5 or 6, wherein after the aircraft test model is freely put into the stable flow field of the tunnel, a schlieren photo of the aircraft test model is obtained by utilizing a laser schlieren technology, the separation shock distance and wake of the aircraft test model are measured, whether the flow field flowing through the model is longer than the time for forming two wake periods is judged, if so, the test is continued, and if not, the running state of the tunnel is regulated until the flow field flowing through the model is longer than the time for forming two wake periods.
15. The method for measuring the characteristics of the plasma sheath and the wake optical radiation of the shock tunnel according to claim 14, wherein the ratio relationship among the separation shock distance delta, the aircraft test model diameter D and the schlieren photo pixels is as follows:
wherein Pe1 is standing point shock wave separation distance pixel along the horizontal center line of the aircraft test model, and Pe2 is model diameter pixel judged and read along the vertical center line of the aircraft test model.
16. According toThe method for measuring the characteristics of the plasma sheath and the wake light radiation of the shock tunnel according to claim 5 or 6, wherein the response time t of the suspension type heavy-model electromagnetic synchronous release system 1 Wind tunnel run time t 2 Time t required by sending pressure signals acquired by pressure sensors to synchronous acquisition system 3 Satisfies the following relationship
t 1 ＜t 3 -t 2 。
17. The method for measuring the characteristics of the plasma sheath and the wake light radiation of the shock tunnel according to claim 5 or 6, wherein the suspension type heavy model electromagnetic synchronous release system is positioned in a Gao Han shock tunnel test section and comprises a release model device fixed above an attack angle mechanism and a buffer capture model device fixed below the attack angle mechanism;
release model device: before the wind tunnel is started, suspending the aircraft test model in the wind tunnel; after the wind tunnel is started, the aircraft test model is released after the release signal is received;
buffer capture model device: and the nondestructive recovery of the aircraft test model is realized.
18. The method for measuring the characteristics of the plasma sheath and the wake optical radiation of the shock tunnel according to claim 17, wherein the release model device comprises a signal processing device, an electromagnetic valve, a power supply and an electromagnet; the power supply is connected with the electromagnet through the electromagnetic valve, the electromagnet is used for adsorbing the aircraft test model, and the signal processing device is connected with the electromagnetic valve; when the signal processing device receives the release signal, the electromagnetic valve is controlled to be disconnected, the electromagnet loses magnetic force, and the aircraft test model is released;
the buffer capture model device is a flexible steel wire mesh.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310036889.8A CN116337401A (en) | 2023-01-10 | 2023-01-10 | Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310036889.8A CN116337401A (en) | 2023-01-10 | 2023-01-10 | Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116337401A true CN116337401A (en) | 2023-06-27 |
Family
ID=86891990
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202310036889.8A Pending CN116337401A (en) | 2023-01-10 | 2023-01-10 | Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116337401A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117782510A (en) * | 2024-02-28 | 2024-03-29 | 中国航空工业集团公司沈阳空气动力研究所 | Wind tunnel protection device and method for continuous wind tunnel model release test |
| CN117782510B (en) * | 2024-02-28 | 2024-05-14 | 中国航空工业集团公司沈阳空气动力研究所 | Wind tunnel protection device and method for continuous wind tunnel model release test |
-
2023
- 2023-01-10 CN CN202310036889.8A patent/CN116337401A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117782510A (en) * | 2024-02-28 | 2024-03-29 | 中国航空工业集团公司沈阳空气动力研究所 | Wind tunnel protection device and method for continuous wind tunnel model release test |
| CN117782510B (en) * | 2024-02-28 | 2024-05-14 | 中国航空工业集团公司沈阳空气动力研究所 | Wind tunnel protection device and method for continuous wind tunnel model release test |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN103884449B (en) | A kind of NOZZLE ARC temperature non-contact measurement system based on fiber-optic transfer | |
| CN105241576B (en) | A kind of blast-furnace hot-air inner lining of furnace based on distribution type fiber-optic corrodes modeling method | |
| CN103558243B (en) | A kind of high-speed aircraft hot surface full field deformation measure device based on optical means | |
| Ni et al. | Lagrangian acceleration measurements in convective thermal turbulence | |
| Hubner et al. | Heat-transfer measurements in hypersonic flow using luminescent coating techniques | |
| CN113588204B (en) | Method for measuring interference characteristics of shock wave boundary layer of air inlet channel | |
| CN107101994B (en) | A kind of spectral emissivity measuring device of opaque material | |
| Fang et al. | Comparison of unsteady pressure-sensitive paint measurement techniques | |
| CN111473735B (en) | Device and method for measuring diameter and bow curvature of optical fiber preform on line | |
| CN109632867A (en) | It is a kind of for examine the hypersonic Burning corrosion resistance of material can pilot system and method | |
| CN111103297A (en) | Non-contact detection method and system for quality of building outer wall surface | |
| CN110207830A (en) | A kind of the imaging sensor caliberating device and scaling method in nonblackbody radiation source | |
| Yali et al. | Ultraviolet 3D digital image correlation applied for deformation measurement in thermal testing with infrared quartz lamps | |
| CN108256166A (en) | A kind of data processing method for thermo-mapping technique | |
| CN104697665B (en) | A kind of blast funnace hot blast stove temperature monitoring method based on distribution type fiber-optic | |
| He et al. | Simultaneous density and velocity measurements in a supersonic turbulent boundary layer | |
| CN116296227A (en) | Test method for measuring radiant heat flow on surface of high-enthalpy expansion wind tunnel model | |
| CN107941667A (en) | Hot environment Dual-Phrase Distribution of Gas olid multiparameter measuring device and method | |
| CN102967577A (en) | Biological aerosol transmitance testing arrangement based on Fourier transform infrared spectrometer | |
| CN116337401A (en) | Shock tunnel plasma sheath and wake light radiation characteristic measurement system and method | |
| CN211553794U (en) | Data measurement and control processing device of non-contact detection system for quality of building outer wall surface | |
| CN107014865A (en) | Exterior wall finish of building set force checking device | |
| Winter et al. | Emission spectroscopic measurements with an optical probe in the NASA Ames IHF arc jet facility | |
| Zhao et al. | An experimental study of aero-optical aberration and dithering of supersonic mixing layer via BOS | |
| CN111103296A (en) | Data measurement control processing device of non-contact detection system for quality of building outer wall surface |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |