CN114810217A - Turbine rotor blade - Google Patents
Turbine rotor blade Download PDFInfo
- Publication number
- CN114810217A CN114810217A CN202110112445.9A CN202110112445A CN114810217A CN 114810217 A CN114810217 A CN 114810217A CN 202110112445 A CN202110112445 A CN 202110112445A CN 114810217 A CN114810217 A CN 114810217A
- Authority
- CN
- China
- Prior art keywords
- blade
- tail
- film slit
- turbine
- cover plate
- 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
- 230000000694 effects Effects 0.000 claims abstract description 10
- 230000009471 action Effects 0.000 claims abstract description 5
- 238000005086 pumping Methods 0.000 claims abstract description 5
- 230000001154 acute effect Effects 0.000 claims abstract description 4
- 238000001816 cooling Methods 0.000 abstract description 24
- 239000007789 gas Substances 0.000 description 14
- 230000004048 modification Effects 0.000 description 3
- 238000012986 modification Methods 0.000 description 3
- 238000005192 partition Methods 0.000 description 3
- 238000013461 design Methods 0.000 description 2
- 230000002035 prolonged effect Effects 0.000 description 2
- 238000004088 simulation Methods 0.000 description 2
- 241001125840 Coryphaenidae Species 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 238000006243 chemical reaction Methods 0.000 description 1
- 239000000112 cooling gas Substances 0.000 description 1
- 230000007423 decrease Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000009826 distribution Methods 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- 230000002265 prevention Effects 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention aims to provide a turbine rotor blade which is provided with a cooling structure and can improve the blade tip cooling effect. The turbine movable blade comprises a blade tip and a blade tail edge cold air cavity positioned inside a blade body, wherein the blade tip comprises a bottom cover plate, a pressure side rib and a suction side rib which are enclosed to form a groove, an air film slit is arranged at the tail part of the groove bottom cover plate, the tail area of the pressure side rib is provided with a notch, the notch extends to the tail edge of the blade tip, the included angle between the center line of the air film slit and the surface of the groove bottom cover plate is an acute angle, and the air film slit communicates the blade tail edge cold air cavity with the tail area of the blade tip, so that under the action of the pressure of the blade tail edge cold air cavity and the pumping effect, cold air flows out of the blade tail edge cold air cavity through the air film slit and covers the tail area of the suction side rib.
Description
Technical Field
The present invention relates to a turbine blade, and particularly to a cooling structure for a turbine blade.
Background
Economy, reliability, safety and serviceability are important criteria for evaluating commercial aircraft engines. The gas turbine is a rotating machine which converts internal energy in high-temperature and high-pressure gas into mechanical energy, is one of the most critical components in the aircraft gas turbine, and the performance of the gas turbine has a direct influence on the performance of the aircraft gas turbine. The high-pressure turbine blade of the commercial aircraft engine works in a high-temperature, high-pressure and high-speed environment, bears very large thermal load and mechanical load, has large variation range of temperature and rotating speed, has certain elongation of the blade, and has certain variation range of blade tip clearance. Tip leakage loss accounts for about 30% of the total loss of the turbine stage, and the tip is high in temperature, difficult to cool and easy to ablate. Therefore, the high pressure turbine blade tips greatly affect the economics, reliability, safety, and serviceability of commercial aircraft engines.
The turbine movable blade tip region has circumferential and axial dominant pressure difference, the circumferential pressure difference causes high-pressure gas of a pressure surface to flow to a suction surface through a blade tip gap to form blade tip gap leakage flow, the leakage flow is seriously deviated from the size and the direction of the main flow speed of a flow channel, the thermal performance of the turbine gas is greatly influenced, the leakage flow and the leakage vortex also increase the difficulty of heat transfer and cooling near the blade tip of the turbine, and the unsteady performance of a downstream flow field is greatly influenced.
In order to meet the demand for increasing the temperature of the inlet gas of the turbine continuously, to avoid exposure of hot-end components to the high-temperature gas environment, and to prevent thermal fatigue damage failure due to severe thermal stress caused by insufficient or uneven cooling, advanced cooling techniques (film cooling and impingement cooling) have been developed rapidly under the condition of slow development of advanced materials.
Chinese patent specification No. CN207093147U describes a tip cooling structure of a turbine rotor blade, in which a pressure surface platform is formed at a position close to a tip of the blade, and is inwardly concave from the pressure surface side to the suction surface side, a film hole is arranged on the pressure surface platform, a partition plate is arranged on the pressure surface along the main flow direction of blade flow guiding, and a film hole is also arranged on the partition plate. A plurality of partitions divides the pressure surface platform 1 into a plurality of sections. The top of the diaphragm in the height direction is flush with the top surface of the blade tip, and the surface of the diaphragm on the pressure surface side is a continuation of the pressure surface of the blade. The outlet of the air film hole is positioned on the pressure surface of the clapboard.
Chinese patent specification No. CN207554113U describes another tip cooling structure for turbine moving blades, in which a tip boss may be provided with a cold air hole, cooling air is ejected from the cold air hole provided on the tip boss, an air film formed by impacting a casing covers the inner surface of the casing, and forms impact cooling and air film cooling for the casing and its area, and the cooling air film flowing to a pressure side blocks and exchanges heat with leakage flow, so as to reduce leakage flow to a certain extent, and reduce leakage flow to enable more gas to perform useful work, and effectively reduce the strength of leakage vortex system, thereby reducing leakage loss, and improving the aerodynamic performance and heat exchange performance of the turbine.
The blade tips of the turbine movable blades are exposed in high-temperature gas, and due to the difficulty of blade tip cooling design, the cooling effect of cold air flowing in from the blade roots to the blade tips after heat exchange with the blade bodies is obviously weakened, so that the blade tip structure is very easy to ablate. Therefore, how to further optimize the structure and cooling manner of such blade tips to obtain better leakage prevention and cooling effects becomes a matter of great concern.
Disclosure of Invention
The invention aims to provide a turbine rotor blade which is provided with a cooling structure and can improve the blade tip cooling effect.
The turbine movable blade comprises a blade tip and a blade tail edge cold air cavity positioned inside a blade body, wherein the blade tip comprises a bottom cover plate, a pressure side rib and a suction side rib which are enclosed to form a groove, an air film slit is arranged at the tail part of the groove bottom cover plate, the tail area of the pressure side rib is provided with a notch, the notch extends to the tail edge of the blade tip, the included angle between the center line of the air film slit and the surface of the groove bottom cover plate is an acute angle, and the air film slit communicates the blade tail edge cold air cavity with the tail area of the blade tip, so that under the action of the pressure of the blade tail edge cold air cavity and the pumping effect, cold air flows out of the blade tail edge cold air cavity through the air film slit and covers the tail area of the suction side rib.
In one embodiment, the gas film slit has a length to width ratio of between 1.5 and 8.
In one embodiment, the included angle between the center line of the air film slit and the upper surface of the blade tip is 10-50 degrees.
In one embodiment, the air film slit is an expanding slit in a flow direction of the cold air.
In one embodiment, the film slit divides the groove bottom cover plate into a cover plate body portion and a cover plate trailing edge portion, the cover plate trailing edge portion being configured as a ramp.
In one embodiment, the included angle between the slope and the center line is 0-40 °.
In one embodiment, the film slit has a front edge and a rear edge opposite the front edge, and the front edge or the rear edge forms an angle of 0 ° to 20 ° with the centerline.
In one embodiment, the film slit has a front edge and a rear edge opposite the front edge, the rear edge smoothly transitioning with the ramp.
The cooling of the tail edge of the blade tip is a difficult point of cooling the turbine movable blade, the air film hole is arranged in the limited space of the tail part of the movable blade tip, the air film hole is specifically arranged at the tail part of the blade top cover plate, the included angle between the center line of the air film hole and the surface of the cover plate at the bottom of the groove is an acute angle, the air film hole communicates the cold air cavity at the tail edge of the blade with the tail part area of the blade tip, and under the action of the pressure of the cold air cavity at the tail edge of the blade and the pumping effect, cold air flows out from the cold air cavity at the tail edge of the blade through the cold air hole and covers the tail part area of the rib tip at the suction side, so that the tail part of the blade tip of the turbine movable blade is cooled, the tail temperature of the tail part of the turbine movable blade tip is reduced, the heat exchange characteristic of the blade tip is improved, and the service life of the blade tip of the turbine movable blade is prolonged.
Drawings
The above and other features, properties and advantages of the present invention will become more apparent from the following description of the embodiments with reference to the accompanying drawings, in which:
FIG. 1 is a schematic illustration of a turbine bucket assembly;
FIG. 2 is a schematic view of a turbine bucket;
FIG. 3 is a schematic illustration of a tip of a turbine bucket partially in section;
FIG. 4 is a schematic illustration of another perspective of a tip of a turbine bucket partially in section;
FIG. 5 is a flow chart of an air film slit simulation;
fig. 6 is a partially enlarged schematic view of a portion corresponding to a circle 54 in fig. 3.
FIG. 7A is a graph of an auxiliary plane perpendicular to the centerline of the film slit intersecting the surface of the film slit in one embodiment.
FIG. 7B is a graph of the intersection of an auxiliary plane perpendicular to the centerline of the film slit with the surface of the film slit in another embodiment.
FIG. 7C is a graph of the intersection of an auxiliary plane perpendicular to the centerline of the film slit with the surface of the film slit in yet another embodiment.
FIG. 8 is a turbine blade tip trailing edge schematic view.
Detailed Description
The present invention is further described in the following description with reference to specific embodiments and the accompanying drawings, wherein the details are set forth in order to provide a thorough understanding of the present invention, but it is apparent that the present invention can be embodied in many other forms different from those described herein, and it will be readily appreciated by those skilled in the art that the present invention can be implemented in many different forms without departing from the spirit and scope of the invention.
It is noted that these and other figures which follow are merely exemplary and not drawn to scale and should not be considered as limiting the scope of the invention as it is actually claimed.
As shown in fig. 1, a plurality of turbine blades are mounted on the turbine blade assembly 100 around the engine axial direction. As shown in fig. 2, the turbine bucket includes a dovetail 1, a platform 2, and a blade body 3. The blade body 3 has a concave pressure surface 31 and a convex suction surface 32, the pressure surface 31 and the suction surface 32 extending between a leading edge 35 and a trailing edge 36 of the blade and between a blade root 4 and a blade tip 5, the pressure surface also being called a basin, the suction surface also being called a back, and the blade tip also being called a tip. The lobed profile, similar to dolphins, crescents, etc., gradually increases in thickness from a leading edge 35 to a maximum thickness and then gradually decreases to a trailing edge 36. The blade profiles with different design sections are stacked according to a certain linear rule to form a blade body 3, the surface of the blade body 3 is a space curved surface with three-dimensional characteristics, fluid flowing through the surface of the blade body generates different speed and pressure distributions due to the pressure surface 31 and the suction surface 32, the pressure difference on the surface of the blade drives the turbine movable blade, the turbine movable blade drives the turbine movable blade assembly 100, and therefore conversion from internal energy to kinetic energy is achieved.
FIG. 3 is an enlarged schematic view of the pressure side tip trailing edge region cut away according to one embodiment of the present invention, showing the tip 5, the blade trailing edge air cavity 60 inside the blade body. The blade tip 5 comprises a bottom cover plate 55, a pressure side rib 51 and a suction side rib 52 enclosing a groove. The blade tip tail splitting slit 56 is arranged at the tail edge, and from the practical point of view, the blade tip tail splitting slit 56 can not carry out effective air film covering on the tail part of the blade tip, so that the cooling effect is not ideal enough.
The cover plate 55 at the bottom of the groove is provided with an air film slit 57, the air film slit 57 is used for communicating the cold air cavity 60 at the tail edge of the blade with the tail area 54 of the blade tip, and under the action of the pressure of the cold air cavity 60 at the tail edge of the blade and the pumping effect, cold air in the cold air cavity 60 at the tail edge of the blade flows out along the air film slit 57 and covers the tail part 53 of the suction side rib of the blade tip. As shown in fig. 5 and 8, the trailing region of the pressure side rib 51 has a notch 510 extending over the trailing edge of the blade tip.
With continued reference to FIG. 5, FIG. 5 primarily shows the air film slit numerical simulation streamlines. Referring to fig. 8, the notch 510 exposes the tip suction side rib tail 53, the cool air 58 flowing out from the air film slit 57 can cover the tip suction side rib tail 53, and the cool air 58 flows into the main flow along the notch 510, thereby cooling the tip tail of the turbine blade, reducing the tip tail temperature of the turbine blade, improving the tip heat exchange characteristic, and prolonging the tip life of the turbine blade. The gap separates the tip pressure side rib 51 and the suction side rib 52 on the pressure side of the tip tail region 54 and does not affect the thickness of the pressure side rib 51 and the suction side rib 52, such that the strength of the tip tail edge region 54 is substantially unaffected by the gap.
As can be seen from FIG. 5, the air film slit only penetrates through the cover plate at the bottom of the groove, the strength of the two side ribs is not affected by the air film slit, and the coverage area of the cold air forming air film is larger and more uniform.
Tip clearance leakage flow exists in a blade tip area of a turbine rotor blade, the leakage flow is seriously deviated from the size and the direction of the main flow speed of the flow channel, the main flow rate is reduced by the leakage flow, the leakage flow does not do useful work basically, and the work of gas on the blade is reduced. Leakage flow blends with the main flow to form leakage vortices, which dissipate and affect the cascade exit flow angle, while both the leakage flow and the leakage vortices block the main flow path, both of which increase the aerodynamic losses of the turbine. For a modern high pressure turbine, this results in turbine stage losses of up to 30% of the total aerodynamic losses. The cooling gas is uniformly covered on the tip tail edge part by the scheme, so that the tip tail area of the turbine movable blade is cooled, the tip tail temperature of the turbine movable blade is reduced, the tip heat exchange characteristic is improved, and the tip life of the turbine movable blade is prolonged.
Fig. 6 shows an enlarged air film slit. The film slit is different from a hole, which can be understood as making an auxiliary plane perpendicular to the slit center line 575, and the pattern formed by the intersection of the auxiliary plane and the peripheral surface of the film slit is a rectangle, a rectangle with rounded edges, a racetrack shape, or the like, as shown in fig. 7A to 7C, wherein the ratio of the length a (maximum length) to the width b (maximum width) is generally between 1.5 and 8.
Preferably, as shown in the film slit direction of fig. 6, the angle θ between the center line 575 of the film slit 57 and the upper surface of the blade tip is in the range of 10 ° to 50 °.
Preferably, the gas film slit 57 is an expanding slit.
Preferably, the angle α between the seam edge 571 and the centerline 575 of the film slit 57 is in the range of 0 ° to 20 °.
Preferably, the included angle β between the hem 572 and the centerline 575 of the film slit 57 is in the range of 0 ° to 20 °.
As shown in fig. 4, the air film slit 57 divides the groove bottom cover plate 55 into a cover plate main portion 550 and a cover plate trailing edge portion 573, the cover plate trailing edge portion 573 being provided as a slope.
Preferably, the angle γ between the cover plate trailing edge portion 573 and the centerline 575 of the film slit 57 is in the range of 0 ° to 40 °.
Preferably, the juncture of the hem 572 and the cover plate trailing edge portion 573 is a smooth transition.
Although the present invention has been disclosed in terms of the preferred embodiment, it is not intended to limit the invention, and variations and modifications may be made by one skilled in the art without departing from the spirit and scope of the invention. Therefore, any modification, equivalent change and modification of the above embodiments according to the technical essence of the present invention are within the protection scope defined by the claims of the present invention, unless the technical essence of the present invention departs from the content of the present invention.
Claims (8)
1. The turbine movable vane is characterized in that an air film slit is arranged at the tail part of the groove bottom cover plate, a tail area of the pressure side rib is provided with a notch, the notch extends to the tail edge of the vane tip, an included angle between the center line of the air film slit and the surface of the groove bottom cover plate is an acute angle, and the air film slit is used for communicating the blade tail edge cold air chamber with the tail area of the vane tip, so that under the action of the pressure of the blade tail edge cold air chamber and the pumping effect, cold air flows out of the blade tail edge cold air chamber through the air film slit and covers the tail area of the suction side rib.
2. The turbine blade as claimed in claim 1, characterized in that the ratio of the length to the width of the film slit is between 1.5 and 8.
3. The turbine bucket of claim 1 wherein the centerline of the film slit makes an angle of 10 ° to 50 ° with the upper surface of the blade tip.
4. The turbine blade of claim 1 wherein the film slit is an expanding slit in the direction of cold gas flow.
5. The turbine bucket of claim 1 wherein the film slit separates the groove bottom cover plate into a cover plate body portion and a cover plate trailing edge portion, the cover plate trailing edge portion being configured as a ramp.
6. The turbine bucket of claim 5 wherein the angle between the ramp and the centerline is between 0 ° and 40 °.
7. The turbine bucket of claim 1 wherein the film slit has a leading edge and a trailing edge opposite the leading edge, the leading edge or the trailing edge having an angle of 0 ° to 20 ° with the centerline.
8. The turbine bucket of claim 5 wherein the film slit has a leading edge and a trailing edge opposite the leading edge, the trailing edge smoothly transitioning with the ramp junction.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110112445.9A CN114810217A (en) | 2021-01-27 | 2021-01-27 | Turbine rotor blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110112445.9A CN114810217A (en) | 2021-01-27 | 2021-01-27 | Turbine rotor blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114810217A true CN114810217A (en) | 2022-07-29 |
Family
ID=82524446
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110112445.9A Pending CN114810217A (en) | 2021-01-27 | 2021-01-27 | Turbine rotor blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114810217A (en) |
Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5720431A (en) * | 1988-08-24 | 1998-02-24 | United Technologies Corporation | Cooled blades for a gas turbine engine |
| CN1519458A (en) * | 2003-01-31 | 2004-08-11 | ���չ�˾ | Turbine blade |
| CN101131096A (en) * | 2006-08-21 | 2008-02-27 | 通用电气公司 | Flared tip turbine blade |
| US8801377B1 (en) * | 2011-08-25 | 2014-08-12 | Florida Turbine Technologies, Inc. | Turbine blade with tip cooling and sealing |
| CN207093147U (en) * | 2017-06-15 | 2018-03-13 | 中国航发商用航空发动机有限责任公司 | The blade tip cooling structure of aero engine turbine blades |
| CN207554113U (en) * | 2017-03-31 | 2018-06-29 | 中国航发商用航空发动机有限责任公司 | Aero-turbine rotor assembly and its blade |
| CN207829957U (en) * | 2017-12-14 | 2018-09-07 | 中国航发商用航空发动机有限责任公司 | Blade tip groove air film hole cooling structure |
| CN110030036A (en) * | 2019-05-10 | 2019-07-19 | 沈阳航空航天大学 | Seam gaseous film control structure is split in a kind of impact of turbine blade tail |
| CN110566283A (en) * | 2019-10-09 | 2019-12-13 | 西北工业大学 | Air film cooling structure for top of high-pressure turbine power blade |
-
2021
- 2021-01-27 CN CN202110112445.9A patent/CN114810217A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5720431A (en) * | 1988-08-24 | 1998-02-24 | United Technologies Corporation | Cooled blades for a gas turbine engine |
| CN1519458A (en) * | 2003-01-31 | 2004-08-11 | ���չ�˾ | Turbine blade |
| CN101131096A (en) * | 2006-08-21 | 2008-02-27 | 通用电气公司 | Flared tip turbine blade |
| US8801377B1 (en) * | 2011-08-25 | 2014-08-12 | Florida Turbine Technologies, Inc. | Turbine blade with tip cooling and sealing |
| CN207554113U (en) * | 2017-03-31 | 2018-06-29 | 中国航发商用航空发动机有限责任公司 | Aero-turbine rotor assembly and its blade |
| CN207093147U (en) * | 2017-06-15 | 2018-03-13 | 中国航发商用航空发动机有限责任公司 | The blade tip cooling structure of aero engine turbine blades |
| CN207829957U (en) * | 2017-12-14 | 2018-09-07 | 中国航发商用航空发动机有限责任公司 | Blade tip groove air film hole cooling structure |
| CN110030036A (en) * | 2019-05-10 | 2019-07-19 | 沈阳航空航天大学 | Seam gaseous film control structure is split in a kind of impact of turbine blade tail |
| CN110566283A (en) * | 2019-10-09 | 2019-12-13 | 西北工业大学 | Air film cooling structure for top of high-pressure turbine power blade |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| JP4876043B2 (en) | Flared tip turbine blade | |
| JP5711741B2 (en) | Two-dimensional platform turbine blade | |
| US10641107B2 (en) | Turbine blade with tip overhang along suction side | |
| CA2596777C (en) | Conformal tip baffle airfoil | |
| EP2148042B1 (en) | A blade for a rotor having a squealer tip with a partly inclined surface | |
| US7607893B2 (en) | Counter tip baffle airfoil | |
| CA2610090C (en) | Triforial tip cavity airfoil | |
| US8061989B1 (en) | Turbine blade with near wall cooling | |
| JPS58126402A (en) | Aerofoil which can be cooled | |
| CN110410158B (en) | Turbine rotor blade of gas turbine | |
| US7850428B2 (en) | Aerofoils | |
| US9353632B2 (en) | Aerofoil structure | |
| US10082031B2 (en) | Rotor of a turbine of a gas turbine with improved cooling air routing | |
| JPH07253003A (en) | Gas-turbine engine | |
| JP6159151B2 (en) | Turbine blade | |
| GB2438861A (en) | Film-cooled component, eg gas turbine engine blade or vane | |
| EP2597264A2 (en) | Aerofoil cooling arrangement | |
| EP2844840A2 (en) | Turbine airfoil trailing edge cooling slots | |
| CN112922674B (en) | Turbine blade with air film cooling groove | |
| CN114810217A (en) | Turbine rotor blade | |
| CN115263436A (en) | Transonic turbine rotor blade, turbine rotor and turbine | |
| JP5980137B2 (en) | Turbine blade | |
| US20130224019A1 (en) | Turbine cooling system and method | |
| CN116537886A (en) | Blade top structure and blade with improve cooling efficiency | |
| CN116085063A (en) | Active wind shadow air film cooling structure for improving end wall cooling performance |
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 |