CN114320607A - Non-rotary double-layer structure anti-icing fairing cap of aircraft engine - Google Patents
Non-rotary double-layer structure anti-icing fairing cap of aircraft engine Download PDFInfo
- Publication number
- CN114320607A CN114320607A CN202210014022.8A CN202210014022A CN114320607A CN 114320607 A CN114320607 A CN 114320607A CN 202210014022 A CN202210014022 A CN 202210014022A CN 114320607 A CN114320607 A CN 114320607A
- Authority
- CN
- China
- Prior art keywords
- fairing
- shell
- screen
- support plate
- icing
- 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
- 230000008093 supporting effect Effects 0.000 claims description 7
- 238000007789 sealing Methods 0.000 claims description 4
- 238000003466 welding Methods 0.000 claims description 3
- 238000009434 installation Methods 0.000 claims description 2
- 230000002265 prevention Effects 0.000 description 3
- 238000004080 punching Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 2
- 238000009825 accumulation Methods 0.000 description 1
- 238000005452 bending Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000008014 freezing Effects 0.000 description 1
- 238000007710 freezing Methods 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000011148 porous material Substances 0.000 description 1
Images
Abstract
The invention discloses a non-rotary double-layer structure anti-icing fairing of an aircraft engine, which comprises: the fairing comprises a fairing shell and a fairing screen, wherein the fairing shell and the fairing screen are welded and fixed in the middle, and an inner cavity for anti-icing gas to flow is formed between the fairing shell and the fairing; the front end of the fairing shell is fixedly welded with the upper edge of the outer end of the fairing lining, and an anti-icing airflow discharge outlet is formed between the front end of the fairing shell and the fairing lining; the lower edge of the inner end of the fairing lining is fixedly welded with the front end of the fairing screen, so that an inner cavity formed by the fairing shell and the fairing screen is communicated with the anti-icing airflow exhaust outlet. The fairing is connected with the fan casing into a whole through 1 central bolt, the anti-icing airflow enters the front edge of the fairing along an inner cavity formed by the fairing shell and the fairing screen, the outer surface of the shell is heated, and finally the anti-icing airflow enters a main flow passage through an exhaust outlet formed by the fairing lining and the fairing shell at the front edge of the fairing. The technical problem of non-rotating fairing anti-icing is solved.
Description
Technical Field
The invention relates to a non-rotary double-layer structure anti-icing fairing of an aircraft engine, belonging to the technical field of engines.
Background
The inlet of the aircraft engine is the most common icing part in the engine, main icing parts comprise a fairing, an inlet support plate, a fan blade and the like, icing directly causes the inlet of the engine to be blocked, the inlet flow of the engine is reduced by a light person, and the surge of the engine is caused by a heavy person; the accumulated ice on the surface of the blade can damage the aerodynamic shape of the blade, so that the efficiency of the fan is reduced; the ice accumulation or shedding can cause uneven rotation, which causes the vibration of the engine to rise; the fallen ice blocks can easily damage engine parts and even damage blades, thereby causing flight accidents. It follows that the icing problem has a great influence on the performance and operating stability of the engine, and therefore aircraft engines must be designed to be ice protected. At present, the fairing of most civil engines is a rotary fairing which rotates along with a rotor, the principle of air film heating is mainly adopted, and an exhaust hole is formed in the fairing and used for exhausting hot air to heat the fairing and preventing the fairing from freezing.
The front pivot of the fan is arranged on a certain type of engine and used for supporting and fixing the front end of a fan rotor, so that a fairing at the front end of a bearing cannot be prevented from being frozen by adopting a rotating fairing, and the fairing needs to be fixed and reasonably designed in order to meet the ice prevention requirement of an engine inlet.
Disclosure of Invention
In order to solve the technical problem, the invention provides a non-rotary double-layer structure anti-icing fairing of an aircraft engine.
The invention is realized by the following technical scheme.
The invention provides a non-rotary double-layer structure anti-icing fairing of an aircraft engine, which comprises:
a fairing shell belonging to the outer shell;
the fairing screen belongs to an inner shell, the fairing screen and the middle part of the fairing shell are welded and fixed, and an inner cavity for gas to flow is formed between the fairing screen and the fairing shell;
the front end of the fairing shell is welded and fixed with the outer end of the fairing lining sleeve, and a space is arranged between the front end of the fairing shell and the fairing lining sleeve to form an anti-icing airflow exhaust outlet;
the inner end of the fairing lining is fixedly welded with the front end of the fairing screen, so that an inner cavity formed by the fairing shell and the fairing screen is communicated with the anti-icing airflow exhaust outlet.
The outer part of the fairing screen is provided with a long stamping boss and a short stamping boss, and the fairing screen is welded with the fairing shell into a whole through resistance welding at the positions of the long stamping boss and the short stamping boss; under the supporting action of the long stamping boss and the short stamping boss, the fairing shell and the fairing screen jointly form a double-layer structure inner cavity for preventing ice gas from flowing.
The long stamping bosses and the short stamping bosses are all arranged at four positions of the circular array at uniform intervals.
The front end of the fairing shell is provided with a stamping boss, the fairing shell is fixedly welded with the outer end of the fairing lining through the stamping boss, and under the supporting action of the stamping boss, an anti-icing airflow discharge outlet is formed at the interval between the front end of the fairing shell and the fairing lining.
The stamping bosses are distributed at four positions of the circular uniform array at the front end of the fairing shell.
The fairing comprises a fairing shell and a fairing screen, and is characterized by further comprising a ring body, wherein the front end of the ring body is welded and fixed with the rear end of the fairing shell, the rear end of the ring body is welded and fixed with the rear end of the fairing screen, nineteen small holes are formed in the ring body to form an airflow channel, and airflow flows into an inner cavity formed by the fairing shell and the fairing screen.
The most front end of the fairing lining is provided with twelve small grooves which can be used for locking a locking washer during assembly, and the fairing and the fan case are connected into a whole through the connection action of a bolt and the locking washer.
A C-shaped plate is fixed on the inner surface of the rear end of the fairing screen; the plate is clamped with an air inlet support plate component.
The air inlet support plate component comprises a support plate, a support plate fairing arranged at the front part of the support plate, a sealing strip arranged at the rear part of the support plate through a wave spring, and an inlet rectifier blade tail arranged at the rear part of the support plate, wherein the upper part of the support plate is provided with a through hole which is communicated with an air passage formed by the support plate and an inner cavity of the support plate fairing;
the anti-icing air flow from the core machine enters the front casing air collection cavity and flows into the small hole through the air guide mounting seat.
The invention has the beneficial effects that: the anti-icing airflow enters the leading edge of the fairing along an inner cavity formed by the fairing housing and the fairing screen, passes through and heats the exterior of the housing, and finally enters the main flow passage through an exhaust outlet formed by the fairing liner and the fairing housing at the leading edge of the fairing. The anti-icing air-entraining amount of 0.6% of the flow of the core machine is enough to maintain the anti-icing surface in a dry state of more than 2 in a low altitude state, and the anti-icing effect of the fairing is remarkable. The technical problem of ice prevention of the non-rotating fairing is solved.
Drawings
FIG. 1 is a schematic cross-sectional view of a fairing of the present invention;
FIG. 2 is a schematic structural view of a fairing panel with long stamped bosses and short stamped bosses distributed thereon according to the present invention;
FIG. 3 is a schematic structural view of a fairing body with stamped bosses thereon according to the present invention;
FIG. 4 is a schematic structural view of the ring body, the plate and the positioning groove of the present invention;
FIG. 5 is a schematic view of the fairing lining, lock washer, bolt, and grove arrangement of the present invention;
FIG. 6 is a schematic structural view of the inlet leg assembly and cowl attachment of the present invention;
FIG. 7 is a schematic structural view of the intake plate assembly of the present invention;
in the figure: 1-a fairing cap; 2-fairing shell; 3-a fairing shield; 4-ring body; 5-a plate sheet; 6-fairing lining; 8-positioning grooves; 9-an air intake plate assembly; 10-a strut fairing; 11-a support plate; 12-a sealing strip; 13-wave spring; 14-inlet rectifier blade tail; 15-front case gas-collecting cavity; 16-a lock washer; 17-a bolt; 18-pores; 19-a small groove; 20-stamping a boss; 21-long stamping boss; 22-short stamped boss.
Detailed Description
The technical solution of the present invention is further described below, but the scope of the claimed invention is not limited to the described.
See fig. 1-7.
The invention relates to an aircraft engine non-rotary double-layer structure anti-icing fairing 1, comprising:
a fairing shell 2 belonging to the outer shell, as shown in fig. 1;
the fairing screen 3 belongs to an inner shell, the fairing screen 3 and the fairing shell 2 are welded and fixed in the middle, a long stamping boss 21 and a short stamping boss 22 are arranged on the outer portion of the fairing screen 3, the long stamping boss 21 and the short stamping boss 22 are arranged at four positions which are evenly spaced in a circular array, and the fairing screen 3 and the fairing shell 2 are welded into a whole through resistance welding at the four positions of the long stamping boss 21 and the four positions of the short stamping boss 22; under the supporting action of the long stamping boss 21 and the short stamping boss 22, the fairing shell 2 and the fairing screen 3 jointly form a double-layer structure inner cavity for anti-icing gas to flow, as shown in fig. 1 and 2;
the front end of the fairing shell 2 and the outer end of the fairing lining 6 are welded and fixed into a whole, and a gap is arranged between the front end of the fairing shell 2 and the fairing lining 6 to form an anti-icing airflow discharge outlet; the front end of the fairing shell 2 is provided with four punching bosses 20 which are circularly and uniformly arrayed, the fairing shell 2 and the fairing lining 6 are welded and fixed into a whole through the punching bosses 20, and under the supporting action of the punching bosses 20, a space is formed between the front end of the fairing shell 2 and the fairing lining 6 to form an anti-icing airflow discharge outlet, as shown in fig. 1 and 3;
the inner end of the fairing lining 6 and the front end of the fairing screen 3 are welded and fixed into a whole, so that an inner cavity formed by the fairing shell 2 and the fairing screen 3 is communicated with an anti-icing airflow discharge outlet, as shown in figure 1;
the front end of the ring body 4 and the rear end of the fairing shell 2 are welded and fixed into a whole, the rear end of the ring body 4 and the rear end of the fairing screen 3 are welded and fixed into a whole, nineteen small holes 18 are arranged on the ring body 4 to form an airflow channel, and airflow flows into an inner cavity formed by the fairing shell 2 and the fairing screen 3, as shown in figure 1;
the most front end of the fairing lining 6 is provided with twelve small grooves 19 which can be used for locking a locking washer 16 during assembly, and the fairing 1 and the fan casing are connected into a whole through the connection action of a bolt 17 and the locking washer 16, as shown in figures 1 and 5.
Specifically, during assembly, the bolt 17 is screwed down according to 60-60 degrees 15', then unscrewed and screwed into contact, finally screwed to 180 degrees +/-15 degrees, two radial parts of the lock washer 16 are bent into the small groove 19 of the fairing lining 6, the other two parts are bent into the groove of the head of the bolt 17, the bolt 17 is locked, the lock washer 16 is not allowed to have cracks, and the schematic diagram after bending is shown in fig. 5-6.
A C-shaped plate 5 is welded and fixed on the inner surface of the rear end of the fairing screen 3, as shown in figures 1 and 4; the sheet 5 is clamped with an air intake plate assembly 9, as shown in fig. 1 and 6.
The air inlet support plate assembly 9 comprises a support plate 11, a support plate fairing 10 arranged at the front part of the support plate 11, a sealing strip 12 arranged at the rear part of the support plate 11 through a wave spring 13, and an inlet rectifier blade tail part 14 arranged at the rear part of the support plate 11, wherein the upper part of the support plate 11 is provided with a through hole which is communicated with an air passage formed by the support plate 11 and an inner cavity of the support plate fairing 10, as shown in fig. 6 and 7;
the ring body 4 is provided with a positioning groove 8, the support plate 11 and the positioning groove 8 are positioned and aligned for installation, the lower part of the support plate 11 is provided with a through hole communicated with the small hole 18, and anti-icing air from the core machine flows through the air-entraining mounting seat, enters the front casing air-collecting cavity 15 and flows into the small hole 18, as shown in fig. 6 and 7.
As shown in fig. 1, 6 and 7, in operation, the anti-icing air from the core engine flows through the bleed air mounting into the front casing air collection chamber 15, then through the air inlet strut assembly 9 and heats the strut fairing 10, then flows through the nineteen small holes 18 at the junction of the strut 11 and the ring 4 into the double-layer cavity formed by the fairing housing 2 and the fairing screen 3, and flows along the double-layer cavity to the front edge of the fairing 1, during which the anti-icing air flows through and heats the housing exterior and finally flows through the discharge outlet formed by the fairing lining 6 and the fairing housing 2 at the front edge of the fairing 1 into the main flow passage. The anti-icing air-entraining amount of 0.6 percent of the flow of the core machine is enough to maintain the anti-icing surface in a dry state of more than 2 ℃ in a low altitude state, and the anti-icing effect of the fairing is remarkable. The technical problem of ice prevention of the non-rotating fairing is solved.
Claims (10)
1. A non-rotating double-deck ice protection cowl for an aircraft engine, comprising:
a fairing shell (2) belonging to the outer shell;
the fairing screen (3) belongs to an inner shell, the fairing screen (3) and the middle part of the fairing shell (2) are welded and fixed, and an inner cavity for gas to flow is formed between the fairing screen (3) and the fairing shell (2);
the front end of the fairing shell (2) is welded and fixed with the outer end of the fairing lining (6), and a space is reserved between the front end of the fairing shell (2) and the fairing lining (6) to form an anti-icing airflow exhaust outlet;
the inner end of the fairing lining (6) is welded and fixed with the front end of the fairing screen (3), so that an inner cavity formed by the fairing shell (2) and the fairing screen (3) is communicated with an anti-icing airflow exhaust outlet.
2. The non-rotating double structure ice-shedding cowl for an aircraft engine of claim 1, wherein: the outer part of the fairing screen (3) is provided with a long stamping boss (21) and a short stamping boss (22), and the fairing screen (3) is welded with the fairing shell (2) into a whole through resistance welding at the positions of the long stamping boss (21) and the short stamping boss (22); under the supporting action of the long stamping boss (21) and the short stamping boss (22), the fairing shell (2) and the fairing screen (3) jointly form a double-layer structure inner cavity for preventing ice from flowing.
3. The non-rotating double structure ice-shedding cowl for an aircraft engine of claim 2, wherein: the long stamping bosses (21) and the short stamping bosses (22) are arranged at four positions of the circular array at uniform intervals.
4. The non-rotating double structure ice-shedding cowl for an aircraft engine of claim 2, wherein: the front end of the fairing shell (2) is provided with a stamping boss (20), the fairing shell (2) is welded with the outer end of the fairing lining (6) through the stamping boss (20), and under the supporting action of the stamping boss (20), an anti-icing airflow discharge outlet is formed at the interval between the front end of the fairing shell (2) and the fairing lining (6).
5. The non-rotating double layer structural ice-shedding fairing of claim 4, wherein: the stamping bosses (20) are distributed at four positions of the circular array at the front end of the fairing shell (2).
6. The non-rotating double layer structural ice-shedding fairing of claim 4, wherein: the fairing shell comprises a fairing shell body (2) and is characterized by further comprising a ring body (4), wherein the front end of the ring body (4) is welded and fixed with the rear end of the fairing shell body (2), the rear end of the ring body (4) is welded and fixed with the rear end of the fairing screen (3), nineteen small holes (18) are formed in the ring body (4), an airflow channel is formed, and airflow flows into an inner cavity formed by the fairing shell body (2) and the fairing screen (3).
7. The non-rotating double layer structural ice-shedding fairing of claim 6, wherein: the most front end of the fairing lining (6) is provided with twelve small grooves (19) which can be used for locking a locking washer (16) during assembly, and the fairing (1) and the fan casing are connected into a whole through the connection action of a bolt (17) and the locking washer (16).
8. The non-rotating double layer structural ice-shedding fairing of claim 7, wherein: a C-shaped plate (5) is fixed on the inner surface of the rear end of the fairing screen 3; and the plate (5) is clamped with an air inlet support plate component (9).
9. The non-rotating double layer structural ice-shedding fairing of claim 8, wherein: the air inlet support plate component (9) comprises a support plate (11), a support plate fairing (10) arranged at the front part of the support plate (11), a sealing strip (12) arranged at the rear part of the support plate (11) through a wave spring (13), and an inlet rectifier blade tail part (14) arranged at the rear part of the support plate (11), wherein the upper part of the support plate (11) is provided with a through hole which is communicated with an air passage formed by the support plate (11) and the support plate fairing (10) at intervals.
10. The non-rotating double layer structural ice protection cowl according to claim 9, wherein: the anti-icing air flow guide device is characterized in that the ring body (4) is provided with a positioning groove (8), the support plate (11) and the positioning groove (8) are positioned, aligned and installed, the lower part of the support plate (11) is provided with a through hole communicated with the small hole (18), and anti-icing air flow from the core machine enters the front casing air collection cavity (15) through the air guide installation seat and flows into the small hole (18).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210014022.8A CN114320607A (en) | 2022-01-06 | 2022-01-06 | Non-rotary double-layer structure anti-icing fairing cap of aircraft engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210014022.8A CN114320607A (en) | 2022-01-06 | 2022-01-06 | Non-rotary double-layer structure anti-icing fairing cap of aircraft engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114320607A true CN114320607A (en) | 2022-04-12 |
Family
ID=81024656
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202210014022.8A Pending CN114320607A (en) | 2022-01-06 | 2022-01-06 | Non-rotary double-layer structure anti-icing fairing cap of aircraft engine |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114320607A (en) |
Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR1015993A (en) * | 1949-04-07 | 1952-10-29 | Canadian Patents Dev | Apparatus for preventing icing on the components of the intake device for gas turbo-engines |
| US4240250A (en) * | 1977-12-27 | 1980-12-23 | The Boeing Company | Noise reducing air inlet for gas turbine engines |
| US4738416A (en) * | 1986-09-26 | 1988-04-19 | Quiet Nacelle Corporation | Nacelle anti-icing system |
| US20090260341A1 (en) * | 2008-04-16 | 2009-10-22 | United Technologies Corporation | Distributed zoning for engine inlet ice protection |
| JP2011021506A (en) * | 2009-07-14 | 2011-02-03 | Ihi Corp | Gas turbine engine |
| US20110179765A1 (en) * | 2010-01-22 | 2011-07-28 | Lalli Anthony S | Jet engine shield and deicer |
| CN203476508U (en) * | 2013-08-21 | 2014-03-12 | 中国航空工业集团公司沈阳发动机设计研究所 | Anti-icing screw-plug assembly structure |
| CN103850800A (en) * | 2012-11-29 | 2014-06-11 | 中航商用航空发动机有限责任公司 | Air inlet anticer and civil turbofan engine |
| US20150003968A1 (en) * | 2012-12-28 | 2015-01-01 | United Technologies Corporation | Gas turbine engine with attached nosecone |
| CN108869094A (en) * | 2018-07-27 | 2018-11-23 | 清华大学 | Rotate detonation engine |
| CN111577463A (en) * | 2020-05-25 | 2020-08-25 | 中国航发沈阳发动机研究所 | Engine air inlet casing structure |
| CN111577462A (en) * | 2020-05-25 | 2020-08-25 | 中国航发沈阳发动机研究所 | Engine air inlet frame |
| CN212202252U (en) * | 2020-04-28 | 2020-12-22 | 江苏火龙动力科技有限公司 | Engine hood with rectification boss |
-
2022
- 2022-01-06 CN CN202210014022.8A patent/CN114320607A/en active Pending
Patent Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR1015993A (en) * | 1949-04-07 | 1952-10-29 | Canadian Patents Dev | Apparatus for preventing icing on the components of the intake device for gas turbo-engines |
| US4240250A (en) * | 1977-12-27 | 1980-12-23 | The Boeing Company | Noise reducing air inlet for gas turbine engines |
| US4738416A (en) * | 1986-09-26 | 1988-04-19 | Quiet Nacelle Corporation | Nacelle anti-icing system |
| US20090260341A1 (en) * | 2008-04-16 | 2009-10-22 | United Technologies Corporation | Distributed zoning for engine inlet ice protection |
| JP2011021506A (en) * | 2009-07-14 | 2011-02-03 | Ihi Corp | Gas turbine engine |
| US20110179765A1 (en) * | 2010-01-22 | 2011-07-28 | Lalli Anthony S | Jet engine shield and deicer |
| CN103850800A (en) * | 2012-11-29 | 2014-06-11 | 中航商用航空发动机有限责任公司 | Air inlet anticer and civil turbofan engine |
| US20150003968A1 (en) * | 2012-12-28 | 2015-01-01 | United Technologies Corporation | Gas turbine engine with attached nosecone |
| CN203476508U (en) * | 2013-08-21 | 2014-03-12 | 中国航空工业集团公司沈阳发动机设计研究所 | Anti-icing screw-plug assembly structure |
| CN108869094A (en) * | 2018-07-27 | 2018-11-23 | 清华大学 | Rotate detonation engine |
| CN212202252U (en) * | 2020-04-28 | 2020-12-22 | 江苏火龙动力科技有限公司 | Engine hood with rectification boss |
| CN111577463A (en) * | 2020-05-25 | 2020-08-25 | 中国航发沈阳发动机研究所 | Engine air inlet casing structure |
| CN111577462A (en) * | 2020-05-25 | 2020-08-25 | 中国航发沈阳发动机研究所 | Engine air inlet frame |
Non-Patent Citations (1)
| Title |
|---|
| 宣益民;连文磊;: "航空发动机整流罩新型防冰方法", 科学通报, vol. 61, no. 25, pages 2843 - 2850 * |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| JP4393797B2 (en) | Compressor bleed case | |
| US6017189A (en) | Cooling system for turbine blade platforms | |
| US4608819A (en) | Gas turbine engine component cooling system | |
| US8439639B2 (en) | Filter system for blade outer air seal | |
| KR101076834B1 (en) | Improved variable geometry assembly for turbochargers | |
| US3391904A (en) | Optimum response tip seal | |
| CN100419235C (en) | Internal anti-icing device of fanjet | |
| RU2569408C2 (en) | Face part of splitter of axial turbine machine with anti-ice device | |
| US4512712A (en) | Turbine stator assembly | |
| EP3382199B1 (en) | Nacelle for a wind turbine including a cooling circuit | |
| EP3543471B1 (en) | System for thermally shielding a portion of a gas turbine shroud assembly | |
| EP2375005B1 (en) | Method for controlling turbine blade tip seal clearance | |
| JP4005905B2 (en) | Nozzle blade bands for gas turbine engines | |
| US20090317244A1 (en) | Gas turbine engine with improved thermal isolation | |
| EP3063379A1 (en) | Radial seal with offset relief cut | |
| CN111911237A (en) | Turbine rotor with cooling air-entraining supercharging impeller | |
| US20180230841A1 (en) | Gas turbine seals | |
| US6920748B2 (en) | Methods and apparatus for operating gas turbine engines | |
| US20140064937A1 (en) | Fan blade brush tip | |
| CN114320607A (en) | Non-rotary double-layer structure anti-icing fairing cap of aircraft engine | |
| GB2144492A (en) | Stator assembly for bounding the flow path of a gas turbine engine | |
| US20180106161A1 (en) | Turbine shroud segment | |
| US10167740B2 (en) | Turbine engine having an element for deflecting foreign objects | |
| EP3170990B1 (en) | Outer airseal for gas turbine engine | |
| CN103620228A (en) | Easily adaptable compressor bleed system downstream of a vane platform |
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 |