CN115324657A - Turbine working blade shroud cooling structure - Google Patents
Turbine working blade shroud cooling structure Download PDFInfo
- Publication number
- CN115324657A CN115324657A CN202211244045.4A CN202211244045A CN115324657A CN 115324657 A CN115324657 A CN 115324657A CN 202211244045 A CN202211244045 A CN 202211244045A CN 115324657 A CN115324657 A CN 115324657A
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- blade
- shroud
- cooling
- outlet hole
- air outlet
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- 238000001816 cooling Methods 0.000 title claims abstract description 59
- 238000007789 sealing Methods 0.000 claims description 70
- 230000009545 invasion Effects 0.000 abstract description 3
- 239000007789 gas Substances 0.000 description 11
- 238000013016 damping Methods 0.000 description 8
- 239000002737 fuel gas Substances 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000000903 blocking effect Effects 0.000 description 1
- 239000000567 combustion gas Substances 0.000 description 1
- 239000000112 cooling gas Substances 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 210000000332 tooth crown Anatomy 0.000 description 1
- 238000011144 upstream manufacturing Methods 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
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- 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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
- F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a turbine working blade shroud cooling structure, comprising: the blade comprises a blade body and a blade cover, wherein a front cavity top air outlet hole and a rear cavity top air outlet hole are arranged at the blade cover at intervals, and the direction of a blade basin of the blade body pointing to a blade back is a first direction; and the flow guide boss is arranged on the blade shroud and positioned between the air outlet hole at the top of the front cavity and the air outlet hole at the top of the rear cavity, and the extending direction of the flow guide boss is in the same direction as the first direction. According to the embodiment of the invention, the guide boss is arranged, so that the backward flow of cold air along the axial direction can be inhibited, the invasion amount of gas to the surface of the blade shroud is controlled, the problem of weaker cooling capability of the blade shroud is solved, and the purposes of improving the cooling capability of the blade shroud of the turbine working blade of the aircraft engine, reducing the temperature of the blade shroud and improving the reliability of the turbine working blade and the engine are achieved.
Description
Technical Field
The specification relates to the technical field of aircraft engines, in particular to a turbine working blade shroud cooling structure.
Background
The turbine working blade is an important working part in an aircraft engine, and the shrouded structure is one of main measures for controlling the blade tip clearance of the turbine working blade and improving the performance of the turbine, and simultaneously limits the strength vibration of the turbine working blade. The shroud working blade belongs to a high-temperature and high-rotation component in an engine, bears the environment temperature of far high material allowable temperature and the action of high centrifugal force, becomes one of key points of the service life of a turbine blade and even the service life of the engine, and realizes the full cooling of the shroud is a main measure for improving the reliability and the service life of the turbine working blade and even the engine.
The blade crown is positioned at the tip of the turbine working blade and surrounds into an annular ring along the circumferential direction. The blade body of the working blade of the turbine gradually shrinks along the radial direction, the blade tip part is small, the blade bears the action of high temperature and high centrifugal force, the blade crown structure is simpler and easier, and the reliability of the blade is higher. Advanced aircraft engine turbine blades are typically of a transonic design, with high tip shroud pressure gradients, which present challenges to tip shroud cooling. A tooth crown structure with a sealing device is commonly adopted in modern aircraft engines, cold air in the blade body of a turbine working blade is guided to be used for cooling the blade crown, the blade crown is used as a cantilever structure under high centrifugal force, a complex cooling form is limited, the cold air outflow mainly flows backwards along the axial direction of the engine under high axial pressure gradient, and the sufficient cooling of the upstream and circumferential cantilever sections of the blade crown is difficult to realize.
Disclosure of Invention
In view of this, the embodiments of the present disclosure provide a turbine working blade shroud cooling structure, which achieves the purpose of efficient cooling of the shroud.
The technical scheme of the invention is as follows: a turbine rotor blade shroud cooling arrangement comprising: the blade comprises a blade body and a blade cover, wherein a front cavity top air outlet hole and a rear cavity top air outlet hole are arranged at the blade cover at intervals, and the direction of a blade basin of the blade body pointing to a blade back is a first direction; and the flow guide boss is arranged on the blade shroud and positioned between the air outlet hole at the top of the front cavity and the air outlet hole at the top of the rear cavity, and the extending direction of the flow guide boss is in the same direction as the first direction.
Furthermore, the flow guide boss comprises a connecting section and two vertical sections, the two vertical sections are arranged in a parallel interval and in a staggered mode, two ends of the connecting section are connected with one ends, close to the blade body, of the two vertical sections respectively, and the connecting section is located between the air outlet hole in the top of the front cavity and the air outlet hole in the top of the rear cavity.
Further, the blade shroud is provided with a blade shroud platform, cooling holes are formed in the blade shroud platform, one ends of the cooling holes are communicated with the upper portion of the blade shroud platform, and the other ends of the cooling holes are communicated with the inner cooling cavity of the blade body.
Furthermore, a front sealing tooth and a rear sealing tooth are arranged on the blade crown platform, the front sealing tooth and the rear sealing tooth are arranged at intervals, and a front cavity top air outlet hole and a rear cavity top air outlet hole are formed between the front sealing tooth and the rear sealing tooth.
Further, cooling holes are provided on the front side of the front obturating teeth.
Furthermore, preceding tooth of obturating and back tooth of obturating all set up towards the front side slope, and preceding tooth of obturating has the contained angle with the radial direction of blade body, and back tooth of obturating has the contained angle with the radial direction of blade body.
Furthermore, the included angle between the front sealing tooth and the radial direction of the blade body is 0-20 degrees, and the included angle between the rear sealing tooth and the radial direction of the blade body is 0-20 degrees.
Furthermore, a front sealing structure and a rear sealing structure are arranged on the blade shroud platform, the distance between the front sealing structure and the casing is a first gap, the distance between the rear sealing structure and the casing is a second gap, and the first gap is smaller than the second gap.
Further, the ratio of the second gap to the first gap is 1.2 to 3.
Compared with the prior art, the embodiment of the specification adopts at least one technical scheme which can achieve the beneficial effects that at least: according to the embodiment of the invention, the guide boss is arranged to inhibit the backward flow of the cold air along the axial direction, control the invasion amount of the gas to the surface of the blade shroud, improve the problem of weaker cooling capability of the blade shroud, and achieve the purposes of improving the cooling of the blade shroud of the turbine working blade of the aircraft engine, reducing the temperature of the blade shroud and improving the reliability of the turbine working blade and the engine.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required to be used in the embodiments will be briefly described below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and it is obvious for those skilled in the art to obtain other drawings without creative efforts.
FIG. 1 is a side view of a turbine rotor blade according to the present invention;
FIG. 2 is a top view of a turbine rotor blade according to the present invention;
FIG. 3 is a side perspective view of a turbine rotor blade according to the present invention.
Reference numbers in the figures: 1. a tenon; 2. a leaf body; 3. a leaf shroud; 10. a front cavity; 11. a rear cavity; 12. a case; 13. a front sealing tooth; 14. a rear sealing tooth; 16. an annular middle chamber; 18. the front end surface of the front sealing tooth; 19. the front end surface of the rear sealing tooth; 20. a front seal tooth front platform; 21. a cooling hole; 22. an air outlet hole is formed in the top of the front cavity; 23. an air outlet hole is formed in the top of the rear cavity; 24. a flow guide boss; 25. a leaf basin edge damping boss; 26. a leaf back side damping boss; 27. and (4) connecting the sections.
Detailed Description
The embodiments of the present application will be described in detail below with reference to the accompanying drawings.
The following description of the embodiments of the present application is provided by way of specific examples, and other advantages and effects of the present application will be readily apparent to those skilled in the art from the disclosure herein. It is to be understood that the embodiments described are only a few embodiments of the present application and not all embodiments. The present application is capable of other and different embodiments and its several details are capable of modifications and/or changes in various respects, all without departing from the spirit of the present application. It is to be noted that the features in the following embodiments and examples may be combined with each other without conflict. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present application.
As shown in fig. 1 to 3, an embodiment of the present invention provides a turbine working blade shroud cooling structure including a blade body and a guide boss 24. The blade body is provided with a blade shroud 3, a front cavity top air outlet hole 22 and a rear cavity top air outlet hole 23 are arranged at the position of the blade shroud 3 at intervals, and the direction of a blade basin of the blade body pointing to a blade back is a first direction; the flow guide boss 24 is arranged on the blade shroud 3 and is positioned between the front cavity top air outlet hole 22 and the rear cavity top air outlet hole 23, and the extending direction of the flow guide boss 24 is the same as the first direction.
According to the embodiment of the invention, the guide boss 24 is arranged to inhibit the backward flow of the cold air along the axial direction, control the invasion amount of the combustion gas to the surface of the blade shroud, improve the problem of weak cooling capability of the blade shroud, and achieve the purposes of improving the cooling capability of the blade shroud of the turbine working blade of the aircraft engine, reducing the temperature of the blade shroud and improving the reliability of the turbine working blade and the engine.
The flow guide boss 24 in the embodiment of the invention comprises a connecting section 27 and two vertical sections, the two vertical sections are arranged in parallel at intervals and in a staggered manner, two ends of the connecting section 27 are respectively connected with one ends of the two vertical sections close to the blade body, and the connecting section 27 is positioned between the front cavity top air outlet hole 22 and the rear cavity top air outlet hole 23.
The blade body consists of a tenon 1, a blade body 2 and a blade shroud 3. Blade body 2 is divided into leading edge and trailing edge along the axial, and blade body 2 is divided into blade basin limit and blade back limit along circumference, and blade body 2 is inside to have a cooling chamber, and the cooling chamber is divided into antechamber 10 and 11 two parts in back chamber along the axial.
Wherein, the blade shroud 3 is provided with the blade shroud platform, has seted up cooling hole 21 on the blade shroud platform, and the one end and the blade shroud platform top of cooling hole 21 communicate, and the other end and the inside cooling chamber intercommunication of blade 2 of cooling hole 21.
Preferably, the blade crown platform is provided with a front sealing tooth 13 and a rear sealing tooth 14, the front sealing tooth 13 and the rear sealing tooth 14 are arranged at intervals, and a front cavity top air outlet 22 and a rear cavity top air outlet 23 are positioned between the front sealing tooth 13 and the rear sealing tooth 14.
The front sealing teeth 13 and the rear sealing teeth 14 divide the blade shroud 3 into a front part, a middle part and a rear part along the axial direction, the front end faces 18 and 19 of the front sealing teeth and the rear sealing teeth are inclined forwards along the axial direction, the front sealing teeth 13 form an included angle with the radial direction of the blade body, and the rear sealing teeth 14 form an included angle with the radial direction of the blade body. The front platform 20 of the front obturating tooth is provided with the cooling hole 21, and the cooling hole 21 has circumferential inclination from the blade back to the blade basin direction. The cooling holes 21 are arranged in the circumferential inclination of the blade back to the blade basin, and the radial speed is reduced by utilizing the circumferential component of the outlet speed, so that cold air can flow along the circumferential direction.
Further, the included angle between the front sealing tooth 13 and the radial direction of the blade body is 0-20 degrees, and the included angle between the rear sealing tooth 14 and the radial direction of the blade body is 0-20 degrees.
The front sealing teeth 13 and the rear sealing teeth 14 are beneficial to locking cold air by adopting the structure, and simultaneously form vortex to play a role in enhancing cooling.
It should be noted that, in this embodiment, an included angle between the front sealing tooth 13 and the radial direction of the blade body is equal to an included angle between the rear sealing tooth 14 and the radial direction of the blade body, and the specific included angle may be selected according to different requirements.
An annular middle cavity 16 is formed between the front sealing tooth 13 and the rear sealing tooth 14 and the casing 12, a flow guide boss 24 is arranged between a front cavity top air outlet hole 22 and a rear cavity top air outlet hole 23, two sides of the flow guide boss 24 extend to the edge of a blade shroud, and two sides of the flow guide boss 24 are respectively connected with a blade basin side damping boss 25 and a blade back side damping boss 26.
The blade crown platform is provided with a front sealing structure and a rear sealing structure, the distance between the front sealing structure and the casing 12 is a first gap, the distance between the rear sealing structure and the casing 12 is a second gap, and the first gap is smaller than the second gap. In the present embodiment, the ratio of the second gap to the first gap is 1.2 to 3.
The cool air of the turbine blade of the aircraft engine enters an inner cavity of the blade body 2 from the bottom of the tenon 1 (the structure of the inner cavity is shown in figure 3), after the blade body 2 is cooled, a part of the cool air flows out of the blade body 2 (or does not flow out of the blade body), and a part (or all) of the cool air flows out of the top of the blade; the gas flows in from the front edge of the blade and flows out from the tail edge.
The lower surface of the blade shroud 3 and the edge plate at the root part of the blade body 2 form a main gas flow passage, and an annular sealing passage is formed by the upper surface of the blade shroud 3 and the casing 12. Gas enters the main flow channel and the sealing channel from the front part of the blade body 2, and airflow forms a pressure difference from a blade basin to a blade back in the main flow channel to push the working blades of the turbine to rotate to do work; the sealing channel has the function of limiting a large amount of gas to directly flow to the rear through the blade tip gap, and the work doing efficiency of the working blade of the turbine is improved.
According to the embodiment of the invention, under the condition that the sealed channel on the upper surface of the blade shroud 3 is limited to allow fuel gas to pass through, a small amount of cooling gas is utilized to reduce the temperature of the ambient airflow on the upper surface, improve the mixing uniformity and improve the cooling of the blade shroud.
According to the invention, residual gas discharged by cooling limited blade bodies is fully utilized, and the rotating action of the sealing channel structure of the blade shroud 3 and the rotating action of the working blades are finely analyzed, so that the cold air rotates in the circumferential direction and flows in the reverse direction, more vortexes are generated, the mixed cold air can stay in the sealing channel of the blade shroud for a longer time, and the blade shroud 3 is cooled.
The air flow which flows out of the cooling hole 21 is at a certain speed due to the fact that the pressure flow of the cold air is larger than the pressure of the gas, if the air flow is directly upward in the radial direction, the air flow is directly blown to the stator component casing 12, the pressure gradient at the gap formed by the stator component casing 12 and the sealing teeth is high, the gas speed is high, the cold air is directly carried to the downstream by the gas, and cooling cannot be achieved. The radial outflow speed of the cold air can be reduced by adopting the inclination angle from the blade back to the blade basin, and the backward outflow quantity of the cold air along the front sealing teeth 13 is correspondingly reduced; the circumferential component of the speed of the cold air outlet is utilized to make the cold air flow along the circumference; meanwhile, the working blades of the aero-engine rotate from the blade basin to the blade back, so that cold air flowing out of each blade flows in the opposite direction along easier rotation. Under the condition of the front sealing tooth 13, the front end face 18 of the front sealing tooth is inclined forwards, and under the blocking action of the front end face 18 of the front sealing tooth, the cold air can easily realize vortex at the front end face, and the cold air is fully in surface contact with the front platform 20 of the front sealing tooth, so that better cooling is realized.
For the annular middle cavity 16, the two sides of the blade basin and the blade back are the blade basin edge damping boss 25 and the blade back edge damping boss 26, which are weak links of the whole blade shroud 3, and the key for improving the reliability and the service life of the turbine working blade is to fully cool the blade and reduce the temperature.
According to the scheme of the invention, the gap between the front sealing teeth 13 is controlled to be used for sealing and amplifying the gap between the rear sealing teeth 14, the inflow of fuel gas is slightly increased under the condition of ensuring the gap between the front sealing teeth 13, the amount of cold gas at the top of the blade body 2 can be rapidly increased, so that the mixing temperature is obviously reduced compared with the temperature of the fuel gas, and the front sealing teeth 13, the rear sealing teeth 14 and the guide boss 24 are utilized to generate vortex, so that the residence time of the mixed low-temperature gas is increased, and the cooling of the part is realized.
The guide boss 24 is arranged, pressure distribution in the front and the back of the middle cavity can be adjusted, meanwhile, airflow flows to the downstream in the opposite direction under the rotation effect, when the airflow flows to the lower blade, due to the fact that the blade back edge is slightly behind the blade basin, air cooling flows to the next blade along the circumferential direction more easily, the air cooling flows farther, circumferential comprehensive cooling of the blade shroud 3 is achieved, the front sealing teeth 13, the rear sealing teeth 14 and the relative low-pressure area of the back of the guide boss 24 form more vortexes, and the blade shroud 3 is fully cooled. Particularly, the cooling of the blade basin side damping boss 25 and the blade back side damping boss 26 far away from the blade body 2 is effectively improved.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (9)
1. A turbine rotor blade shroud cooling arrangement comprising:
the blade comprises a blade body and a blade cover, wherein the blade body is provided with a blade cover (3), a front cavity top air outlet hole (22) and a rear cavity top air outlet hole (23) are arranged at the position of the blade cover (3) at intervals, and the direction of a blade basin of the blade body pointing to a blade back is a first direction;
and the flow guide boss (24) is arranged on the blade shroud (3) and is positioned between the front cavity top air outlet hole (22) and the rear cavity top air outlet hole (23), and the extending direction of the flow guide boss (24) is in the same direction as the first direction.
2. The turbine rotor blade shroud cooling structure as claimed in claim 1, wherein the guide boss (24) includes a connecting section and two vertical sections, the two vertical sections are spaced in parallel and are disposed in a staggered manner, two ends of the connecting section are respectively connected to one ends of the two vertical sections close to the blade body, and the connecting section is located between the front cavity top outlet hole (22) and the rear cavity top outlet hole (23).
3. The turbine working blade shroud cooling structure according to claim 1, characterized in that the shroud (3) is provided with a shroud platform, the shroud platform is provided with cooling holes (21), one end of each cooling hole (21) is communicated with the upper side of the shroud platform, and the other end of each cooling hole (21) is communicated with the internal cooling cavity of the blade body.
4. The turbine working blade shroud cooling structure according to claim 3, wherein the shroud platform is provided with front sealing teeth (13) and rear sealing teeth (14), the front sealing teeth (13) and the rear sealing teeth (14) are arranged at intervals, and the front cavity top air outlet hole (22) and the rear cavity top air outlet hole (23) are located between the front sealing teeth (13) and the rear sealing teeth (14).
5. Turbine rotor blade shroud cooling according to claim 4, characterized in that the cooling holes (21) are arranged on the front side of the front obturating tooth (13).
6. The turbine working blade shroud cooling structure according to claim 4, characterized in that the front sealing teeth (13) and the rear sealing teeth (14) are both arranged obliquely toward the front side, and the front sealing teeth (13) have an angle with the radial direction of the blade body, and the rear sealing teeth (14) have an angle with the radial direction of the blade body.
7. The turbine working blade shroud cooling arrangement as claimed in claim 6, characterized in that the leading obturating tooth (13) has an angle of 0 ° to 20 ° with the radial direction of the blade body, and the trailing obturating tooth (14) has an angle of 0 ° to 20 ° with the radial direction of the blade body.
8. The turbine rotor blade shroud cooling arrangement of claim 4, wherein the shroud platform is provided with a front seal structure and a rear seal structure, the front seal structure being spaced from the casing (12) by a first gap, the rear seal structure being spaced from the casing (12) by a second gap, the first gap being less than the second gap.
9. The turbine rotor blade shroud cooling structure of claim 8 wherein the ratio of said second clearance to said first clearance is 1.2 to 3.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211244045.4A CN115324657A (en) | 2022-10-12 | 2022-10-12 | Turbine working blade shroud cooling structure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211244045.4A CN115324657A (en) | 2022-10-12 | 2022-10-12 | Turbine working blade shroud cooling structure |
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| CN115324657A true CN115324657A (en) | 2022-11-11 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202211244045.4A Pending CN115324657A (en) | 2022-10-12 | 2022-10-12 | Turbine working blade shroud cooling structure |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117869014A (en) * | 2024-01-03 | 2024-04-12 | 中国航发湖南动力机械研究所 | Turbine casing with cooling structure and aircraft engine |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117869014A (en) * | 2024-01-03 | 2024-04-12 | 中国航发湖南动力机械研究所 | Turbine casing with cooling structure and aircraft engine |
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