CN114193790B - Forming method of reinforced shell cabin section of composite material with different resin systems - Google Patents
Forming method of reinforced shell cabin section of composite material with different resin systems Download PDFInfo
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- CN114193790B CN114193790B CN202111458053.4A CN202111458053A CN114193790B CN 114193790 B CN114193790 B CN 114193790B CN 202111458053 A CN202111458053 A CN 202111458053A CN 114193790 B CN114193790 B CN 114193790B
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- rib
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- cabin section
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- 239000002131 composite material Substances 0.000 title claims abstract description 27
- 229920005989 resin Polymers 0.000 title claims abstract description 27
- 239000011347 resin Substances 0.000 title claims abstract description 27
- 238000000034 method Methods 0.000 title claims abstract description 26
- 238000004804 winding Methods 0.000 claims abstract description 31
- 229920000049 Carbon (fiber) Polymers 0.000 claims abstract description 19
- 239000004917 carbon fiber Substances 0.000 claims abstract description 19
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims abstract description 19
- 238000000465 moulding Methods 0.000 claims abstract description 17
- 239000004643 cyanate ester Substances 0.000 claims abstract description 14
- 239000000463 material Substances 0.000 claims abstract description 13
- 239000003822 epoxy resin Substances 0.000 claims abstract description 12
- 229920000647 polyepoxide Polymers 0.000 claims abstract description 12
- 239000003292 glue Substances 0.000 claims abstract description 8
- 239000011248 coating agent Substances 0.000 claims abstract description 6
- 238000000576 coating method Methods 0.000 claims abstract description 6
- 239000000835 fiber Substances 0.000 claims abstract description 6
- 238000004519 manufacturing process Methods 0.000 claims abstract description 6
- 238000009966 trimming Methods 0.000 claims abstract description 6
- 230000003014 reinforcing effect Effects 0.000 claims abstract description 4
- 238000007711 solidification Methods 0.000 claims abstract description 4
- 230000008023 solidification Effects 0.000 claims abstract description 4
- 238000005498 polishing Methods 0.000 claims description 9
- 238000004364 calculation method Methods 0.000 claims description 8
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 claims description 6
- 229910052782 aluminium Inorganic materials 0.000 claims description 6
- 230000002787 reinforcement Effects 0.000 claims description 4
- 239000011159 matrix material Substances 0.000 claims description 3
- 238000007598 dipping method Methods 0.000 claims description 2
- XLJMAIOERFSOGZ-UHFFFAOYSA-M cyanate Chemical compound [O-]C#N XLJMAIOERFSOGZ-UHFFFAOYSA-M 0.000 description 3
- 239000011324 bead Substances 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 230000007547 defect Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000004593 Epoxy Substances 0.000 description 1
- 229910000831 Steel Inorganic materials 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 238000000227 grinding Methods 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000012943 hotmelt Substances 0.000 description 1
- 230000008092 positive effect Effects 0.000 description 1
- 238000002360 preparation method Methods 0.000 description 1
- 239000010959 steel Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
- 239000013585 weight reducing agent Substances 0.000 description 1
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
- B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
- B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/54—Component parts, details or accessories; Auxiliary operations, e.g. feeding or storage of prepregs or SMC after impregnation or during ageing
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/54—Component parts, details or accessories; Auxiliary operations, e.g. feeding or storage of prepregs or SMC after impregnation or during ageing
- B29C70/545—Perforating, cutting or machining during or after moulding
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3097—Cosmonautical vehicles; Rockets
-
- 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
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Landscapes
- Chemical & Material Sciences (AREA)
- Engineering & Computer Science (AREA)
- Composite Materials (AREA)
- Mechanical Engineering (AREA)
- Moulding By Coating Moulds (AREA)
Abstract
The invention relates to a molding method of a reinforced shell cabin section of composite materials with different resin systems, which comprises the following steps: step S1: designing and manufacturing a cabin section forming die; step S2: preparing prepreg to prepare winding glue; step S3: paving reinforcing areas at the front end and the rear end of the cabin section by using cyanate ester prepreg; step S4: winding by using carbon fiber to impregnate epoxy resin glue solution; step S5: coating the forming die by using a vacuum auxiliary material; step S6: trimming the ring rib and the reinforced area after solidification; step S7: laying a skin by using cyanate ester prepreg; step S8: co-curing and molding the epoxy resin ring rib and the cyanate ester skin autoclave; step S9: removing the cabin section forming die to obtain composite material reinforced shell cabin sections with different resin systems; the invention has the advantages that: the internal quality and the appearance surface precision of the product can be improved, the condition that the strength requirement of the cabin section is influenced by the buckling of the fiber at the position of the annular rib is avoided, the annular rib and the skin are made of different resin system composite materials, the process difficulty is reduced, and the internal quality of the annular rib of the cabin section is improved.
Description
Technical Field
The invention relates to the technical field of composite structural member forming processes, in particular to a method for forming a reinforced shell cabin section of a composite material with different resin systems.
Background
In recent years, the development of composite material industry has been rapid, and the urgent need for structure weight reduction has been mainly derived from industrial fields such as aerospace industry and high-speed rail industry. The composite reinforced shell structure is widely applied to various cabin sections such as a fairing, an interstage section, a transition section and the like of a carrier rocket, and has the main characteristics of high strength, high specific modulus, low cost, short production period and higher stability. However, with the continuous development of the composite material industry, the size of the cabin section is continuously increased, the materials are continuously updated, the structure is continuously optimized, and the requirements on the molding and manufacturing process of the cabin section with the reinforced shell structure of the composite material are continuously improved.
As a reinforcement shell cabin section of a composite material with different resin systems, the process of resin system selection, rib winding and trimming, prepreg laying and autoclave secondary curing involved in the molding process is a relevant target in molding, and a molding method of the reinforcement shell cabin section of the composite material with different resin systems is needed.
Disclosure of Invention
In view of the above problems, the invention aims to provide a molding method of a composite material reinforced shell cabin section with different resin systems, which is used for solving the problems that the molding quality of cabin section ribs is poor, the matching property of the ribs and the skin is poor, and the appearance quality of a product cannot meet the requirements, so as to overcome the defects in the prior art.
The invention provides a molding method of a reinforced shell cabin section of a composite material with different resin systems, which specifically comprises the following steps:
step S1: designing and manufacturing a cabin section forming die, wherein the cabin section forming die mainly comprises a core die, an aluminum split male die and a winding shaft;
step S2: preparing prepreg to prepare winding glue;
step S3: the cyanate ester prepreg is used for paving the front and rear end reinforcing areas of the cabin section, wherein after the preset layers are paved, vacuumizing and prepressing are carried out, so that the exhaust gas prepreg paving space is more compact;
step S4: the carbon fiber is used for dipping epoxy resin glue solution for winding, and the annular rib groove of the die is filled to be flush with the outer surface of the split male die and then continuously wound higher than the surface of the split male die; winding six groups of carbon fiber prepreg filaments along the annular rib grooves, filling the annular rib grooves of the die to be level with the outer surface of the split male die, continuing winding and 2-3mm (compression amount) higher than the surface of the split male die, and sequentially carrying out the above operations on all annular ribs;
step S5: coating a forming die by using a vacuum auxiliary material, and performing primary pre-curing by using an autoclave; coating the winding prepreg filaments by using vacuum auxiliary materials in sequence, and pre-curing under the condition of pressurization (0.15-0.3 MPa) at a low temperature (90-110 ℃) by using an autoclave;
step S6: trimming the ring rib and the reinforced area after solidification; polishing after the ring rib is pre-cured, firstly polishing the rib to be flush with the outer surface of the split male die, then calculating the difference value of the expansion amount of each rib and the forming die in the secondary curing process through an expansion amount calculation formula, and setting the thickness of the laid prepreg at the rib according to the difference value;
step S7: laying a skin by using cyanate ester prepreg, wherein after each laying set layer number, vacuumizing and prepressing are carried out;
step S8: curing and molding the epoxy resin ring rib and the cyanate ester skin autoclave;
step S9: and removing the cabin section forming die to obtain the composite material reinforced shell cabin sections with different resin systems.
As a preferred aspect of the present invention, in step S2, a prepreg is prepared with a T700-grade carbon fiber as a reinforcement and a cyanate resin as a matrix, and an epoxy resin-impregnated carbon fiber yarn is prepared.
As a preferred aspect of the present invention, in step S4, the method for calculating the carbon fiber winding thickness d includes:
d＝(n×s)/(b×V f ) (1)
wherein n is the number of yarn groups; s is the sectional area of the rib; b is the advancing amount of the winding trolley; v (V) f Is the fiber volume content.
In step S6, the bead grinding height and the bead prepreg laying thickness are calculated according to the thermal expansion amount. The expansion amount calculation formula is as follows:
ΔL＝α×L×T (2)
wherein alpha is the expansion coefficient; l is the total length of the material; t is the maximum temperature difference.
As a preferred aspect of the present invention, in step S8, the vacuum curing process conditions are: the vacuum degree is not more than-0.095 MPa, the curing temperature is 120-190 ℃ and the pressure is 0.3-0.5 MPa.
The invention has the advantages and positive effects that:
1. according to the invention, the outside envelope size of the cabin section is 3200mm multiplied by 2000mm, the height of the ribs is 35mm, composite materials with different resin systems are selected to be respectively used for forming the ribs and the skin, and polishing and repairing are carried out after the ribs are wound, cured and formed, so that the matching performance between the ribs and the skin is ensured to be improved during secondary curing. And then paving and forming the skin prepreg, and finally heating, pressurizing and curing by using an autoclave. And the problems that the molding quality of the ribs of the cabin section is poor, the matching property of the ribs and the skin is poor, and the appearance quality of the product cannot meet the requirements are solved.
2. According to the invention, the T700-grade carbon fiber/cyanate resin prepreg and the epoxy resin for winding are firstly prepared, the carbon fiber yarn is soaked with the epoxy resin for winding the ribs, and the cyanate resin system prepreg is used for the skin. On the premise of ensuring the temperature resistance and structural stability of the outer skin of the cabin section, the resin matrix of the rib is replaced by epoxy resin, so that the manufacturability is improved, the molding difficulty is reduced, the internal quality of the rib is improved, and internal defects such as bubbles are avoided.
3. According to the invention, the matching property of the rib positions and the skin is improved by trimming after the rib is pre-cured, and the laying process parameters are adjusted according to the thermal expansion difference of the steel/aluminum die material and the carbon fiber material, so that the appearance quality of the cabin section is good, the rib positions of the outer skin are smooth, and the buckling of the fiber is avoided to influence the overall structural strength.
3. After the ring rib is formed, the ring rib is trimmed through checking the thermal expansion matching relation between the forming die and the carbon fiber composite material, the wound process layer is polished, and the prepreg is laid at the ring rib position, so that the ring rib is matched with the skin during secondary curing. The method can improve the internal quality of the product and the precision requirement of the appearance surface of the product, avoid the influence of fiber buckling at the position of the annular rib on the strength requirement of the cabin section, and greatly reduce the process difficulty by adopting different resin system composite materials for the annular rib and the skin, and improve the internal quality of the annular rib of the cabin section.
Drawings
Other objects and attainments together with a more complete understanding of the invention will become apparent and appreciated by referring to the following description taken in conjunction with the accompanying drawings. In the drawings:
FIG. 1 is a partial view of a molding die for a composite reinforced shell section of a different resin system in an embodiment of the invention.
Figure 2 is a schematic illustration of the appearance of a composite reinforced shell pod product of different resin systems in an embodiment of the present invention.
FIG. 3 is a flow chart in an embodiment of the present invention.
Description of the drawings: a core mould 1 and an aluminum split male mould 2.
Detailed Description
In the following description, for purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be evident, however, that such embodiment(s) may be practiced without these specific details. In other instances, well-known structures and devices are shown in block diagram form in order to facilitate describing one or more embodiments.
Fig. 1-3 show overall schematic diagrams according to embodiments of the present invention.
As shown in fig. 1-3, the method for forming the reinforced shell cabin section of the composite material with different resin systems provided by the embodiment of the invention comprises the following steps:
step S1: designing and manufacturing a cabin section forming die; the cabin section forming die mainly comprises a core die 1, an aluminum split male die 2 and a winding shaft 3. Wherein the core mould 1 is positioned at the center of the forming mould; the aluminum split male die 2 is attached to the outer surface of the core die 1; the winding shaft 3 is arranged at flanges at the front end and the rear end of the core mold 1 and can be detached at any time.
Step S2: preparing prepreg to prepare winding glue; wherein the high-performance carbon fiber/cyanate ester prepreg adopts a hot-melt prepreg preparation technology.
Step S3: using cyanate ester prepreg to lay the front and rear end reinforcing areas of the cabin section; after the fixed layers are laid, vacuumizing and prepressing are carried out, so that the exhaust gas prepreg is more compact between the layers.
Step S4: using carbon fiber to impregnate epoxy resin glue solution for winding, filling the annular rib groove of the die to be flush with the outer surface of the split male die, and continuing winding and being higher than the surface of the split male die; and winding six groups of carbon fiber prepreg wires along the circumferential rib grooves, filling the circumferential rib grooves of the die to be level with the outer surface of the split male die, continuously winding the circumferential rib grooves and 2-3mm (compression amount) higher than the surface of the split male die, and sequentially performing the above operation on all the circumferential ribs.
Step S41: the calculation method of the carbon fiber winding thickness d comprises the following steps:
d＝(n×s)/(b×V f ) (1)
wherein n is the number of yarn groups; s is the sectional area of the rib; b is the advancing amount of the winding trolley; v (V) f Is the fiber volume content.
Step S5: coating a forming die by using a vacuum auxiliary material, and performing primary pre-curing by using an autoclave; wrapping the winding prepreg filaments by using vacuum auxiliary materials in sequence, and pre-curing by using an autoclave under pressure (0.15-0.3 MPa) at a low temperature (90-110 ℃).
Step 6: trimming the ring rib and the reinforced area after solidification; polishing after the ring rib is pre-cured, polishing the rib to be flush with the outer surface of the split male die, calculating the difference value of the expansion amount of each rib and the forming die in the secondary curing process through an expansion amount calculation formula, and setting the thickness of the laid prepreg at the rib according to the difference value.
Step 61: and calculating the rib polishing height and the rib prepreg laying thickness according to the thermal expansion amount. The expansion amount calculation formula is as follows:
ΔL＝α×L×T (2)
wherein alpha is the expansion coefficient; l is the total length of the material; t is the maximum temperature difference.
Step 7: laying up a skin using a cyanate ester prepreg; after the fixed layers are laid, vacuumizing and prepressing are carried out, so that the exhaust gas prepreg is more compact between the layers.
Step 8: curing and forming the epoxy resin ring rib and the cyanate ester skin in an autoclave;
step 81: the vacuumizing and curing process conditions are as follows: the vacuum degree is not more than-0.095 MPa, the curing temperature is 120-190 ℃ and the pressure is 0.3-0.5 MPa.
Step 9: and removing the mould to obtain the composite material reinforced shell cabin sections with different resin systems.
The foregoing is merely illustrative embodiments of the present invention, but the scope of the present invention is not limited thereto, and any person skilled in the art can easily think about variations or substitutions within the technical scope of the present invention, and the invention should be covered. Therefore, the protection scope of the invention is subject to the protection scope of the claims.
Claims (3)
1. The molding method of the reinforced shell cabin section of the composite material with different resin systems is characterized by comprising the following steps:
step S1: designing and manufacturing a cabin section forming die, wherein the cabin section forming die mainly comprises a core die, an aluminum split male die and a winding shaft;
step S2: preparing prepreg to prepare winding glue;
step S3: the cyanate ester prepreg is used for paving the front and rear end reinforcing areas of the cabin section, wherein after the preset layers are paved, vacuumizing and prepressing are carried out, so that the exhaust gas prepreg paving space is more compact;
step S4: the carbon fiber is used for dipping epoxy resin glue solution for winding, and the annular rib groove of the die is filled to be flush with the outer surface of the split male die and then continuously wound higher than the surface of the split male die; winding six groups of carbon fiber prepreg wires along the annular rib grooves, filling the annular rib grooves of the die to be level with the outer surface of the split male die, continuing winding and 2-3mm higher than the surface of the split male die, and sequentially carrying out the above operations on all annular ribs; the calculation method of the carbon fiber winding thickness d comprises the following steps:
d＝(n×s)/(b×V f ) (1)
wherein n is the number of yarn groups; s is the sectional area of the rib; b is the advancing amount of the winding trolley; v (V) f Is the volume content of the fiber;
step S5: coating a forming die by using a vacuum auxiliary material, and performing primary pre-curing by using an autoclave; coating the winding prepreg wires sequentially by using vacuum auxiliary materials, and performing pressurization and pre-curing at a low temperature by using an autoclave;
step S6: trimming the ring rib and the reinforced area after solidification; polishing after the ring rib is pre-cured, firstly polishing the rib to be flush with the outer surface of the split male die, then calculating the difference value of the expansion amount of each rib and the forming die in the secondary curing process through an expansion amount calculation formula, and setting the thickness of the laid prepreg at the rib according to the difference value; calculating the rib polishing height and the rib prepreg laying thickness according to the thermal expansion amount, wherein the expansion amount calculation formula is as follows:
ΔL＝α×L×T (2)
wherein alpha is the expansion coefficient; l is the total length of the material; t is the maximum temperature difference;
step S7: laying a skin by using cyanate ester prepreg, wherein after each laying set layer number, vacuumizing and prepressing are carried out;
step S8: curing and molding the epoxy resin ring rib and the cyanate ester skin autoclave;
step S9: and removing the cabin section forming die to obtain the composite material reinforced shell cabin sections with different resin systems.
2. The method for molding a reinforced shell section of a composite material with different resin systems according to claim 1, wherein in step S2, a prepreg is prepared by using a T700-grade carbon fiber as a reinforcement and a cyanate ester resin as a matrix, and an epoxy resin impregnated carbon fiber yarn is prepared.
3. The method for molding a reinforced shell section of a composite material with different resin systems according to claim 1, wherein in step S8, the vacuum curing process conditions are as follows: the vacuum degree is not more than-0.095 MPa, the curing temperature is 120-190 ℃ and the pressure is 0.3-0.5 MPa.
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| CN202111458053.4A CN114193790B (en) | 2021-12-01 | 2021-12-01 | Forming method of reinforced shell cabin section of composite material with different resin systems |
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| CN114193790B true CN114193790B (en) | 2024-02-20 |
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| WO2020119652A1 (en) * | 2018-12-11 | 2020-06-18 | 中南大学 | Composite-material forming and manufacturing apparatus based on microwave chamber |
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2021
- 2021-12-01 CN CN202111458053.4A patent/CN114193790B/en active Active
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| KR20050108834A (en) * | 2004-05-13 | 2005-11-17 | 에스케이케미칼주식회사 | Method of molding fiber-resin composite using plane heater |
| CN103963315A (en) * | 2014-05-29 | 2014-08-06 | 上海飞机制造有限公司 | Prepreg/resin transfer molding co-curing process method for composite materials |
| CN104608915A (en) * | 2015-01-30 | 2015-05-13 | 中国人民解放军国防科学技术大学 | Multilayer grating bearing cylinder and preparation method thereof |
| CN108859172A (en) * | 2018-03-19 | 2018-11-23 | 连云港神鹰复合材料科技有限公司 | A kind of preparation method of the integrally formed composite material shelter of precast framework formula |
| WO2020119652A1 (en) * | 2018-12-11 | 2020-06-18 | 中南大学 | Composite-material forming and manufacturing apparatus based on microwave chamber |
| CN109895418A (en) * | 2019-03-27 | 2019-06-18 | 成都联科航空技术有限公司 | A kind of processing method of abnormity hollow structure composite material parts molding core model |
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| CN114193790A (en) | 2022-03-18 |
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