CN112031878A - Turbine rotor blade apex double-wall structure - Google Patents
Turbine rotor blade apex double-wall structure Download PDFInfo
- Publication number
- CN112031878A CN112031878A CN202011219499.7A CN202011219499A CN112031878A CN 112031878 A CN112031878 A CN 112031878A CN 202011219499 A CN202011219499 A CN 202011219499A CN 112031878 A CN112031878 A CN 112031878A
- Authority
- CN
- China
- Prior art keywords
- wall
- turbine rotor
- rotor blade
- double
- turbine
- 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
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
- F01D5/20—Specially-shaped blade tips to seal space between tips and stator
Abstract
The invention belongs to the field of aero-engines and gas turbines, and particularly relates to a double-wall structure of a blade tip of a turbine rotor blade. The technical scheme of the invention is as follows: the utility model provides a turbine rotor blade apex double-layer wall structure, includes outer wall and inner wall, outer wall and inner wall setting are on the turbine rotor blade apex, the whole edge setting of turbine rotor blade apex is followed to the outer wall, the inner wall sets up in the outer wall, the outer wall with the inner wall is parallel track form and distributes, the outer wall with form recess one between the inner wall, the inner wall encloses into recess two. The double-wall structure of the blade tip of the turbine rotor blade provided by the invention effectively reduces the gas leakage amount of the blade tip clearance, and improves the turbine efficiency, thereby reducing the fuel oil loss of an engine and improving the thrust of an aircraft engine and a gas turbine.
Description
Technical Field
The invention belongs to the field of aero-engines and gas turbines, and particularly relates to a double-wall structure of a blade tip of a turbine rotor blade.
Background
When turbine rotor blades of aero-engines and gas turbines work, under the action of thermal load, centrifugal load, pneumatic load and the like, the blades are subjected to certain circumferential extension deformation, so that the diameters of the rotor blades and the outer ring of a stator cannot achieve zero clearance control, and a certain clearance is required to exist between the blade tips of the rotor blades and a casing. The blade tip clearance of the turbine rotor blade is one of important indexes influencing the performance of an aeroengine and a gas turbine, the too small blade tip clearance can cause the collision and abrasion of the turbine rotor blade and the outer ring of a casing, and the too large clearance can cause the large gas leakage amount and reduce the turbine efficiency. When the gas passes through the blade tip clearance, the gas flows from the pressure surface of the turbine rotor blade to the suction surface of the turbine rotor blade to form leakage flow, the leakage flow can block the main flow of the channel, the downstream instability is caused, the complexity of the heat exchange of the blade is increased, the pneumatic loss gas flows in the blade tip clearance, and the efficiency of the engine is reduced. The blade tip groove is a structure commonly applied to the blade tip of the existing turbine rotor blade, and along with the continuous rise of the front temperature of the turbine of the engine, the function of the blade tip groove in reducing the blade tip clearance leakage amount is smaller and smaller, so that the further improvement of the performance of the engine is limited.
Disclosure of Invention
The invention provides a double-wall structure of a blade tip of a turbine rotor blade, which effectively reduces the leakage rate of gas in a gap of the blade tip and improves the turbine efficiency, thereby reducing the fuel loss of an engine and improving the thrust of an aircraft engine and a gas turbine.
The technical scheme of the invention is as follows:
the utility model provides a turbine rotor blade apex double-layer wall structure, includes outer wall and inner wall, outer wall and inner wall setting are on the turbine rotor blade apex, the whole edge setting of turbine rotor blade apex is followed to the outer wall, the inner wall sets up in the outer wall, the outer wall with the inner wall is parallel track form and distributes, the outer wall with form recess one between the inner wall, the inner wall encloses into recess two.
Further, the blade tip double-wall structure of the turbine rotor blade is characterized in that the thicknesses of the outer wall and the inner wall are 4-6% of the thickness of the turbine rotor blade; the width of the first groove is 1-2 times of the thickness of the inner wall.
Further, the turbine rotor blade tip double-wall structure is characterized in that the height of the outer wall is 1-2 times of the width of the groove; the height of the inner wall is 90-110% of that of the outer wall.
When gas flows into the tip clearance through the pressure surface of the turbine rotor blade, low-speed vortex is formed in the first groove between the outer wall and the inner wall by airflow, the vortex is filled in the first groove between the outer wall and the inner wall, and the gas is prevented from flowing into one side of the suction surface of the turbine rotor blade from one side of the pressure surface of the turbine rotor blade through the tip clearance; the same low-speed vortex can be formed in the groove I on one side of the suction surface of the turbine rotor blade, so that gas leakage is prevented, the sealing effect is better, the fuel oil loss of an engine is reduced, and the thrust of an aircraft engine and a gas turbine is improved.
The invention has the beneficial effects that: the invention can reduce the gas leakage of the blade tip clearance of the turbine rotor blade, and the total pressure loss coefficient is reduced by 16 percent, thereby improving the turbine efficiency, reducing the fuel oil loss of the engine and improving the thrust of the aircraft engine and the gas turbine.
Drawings
FIG. 1 is a top view of a double-walled structure of a turbine rotor blade tip;
FIG. 2 is a perspective view of a double-walled construction of a turbine rotor blade tip;
FIG. 3 is a schematic view of a gas flow in a double-walled structure at the tip of a turbine rotor blade.
In the figure: 1. outer wall, 2 inner wall, 3 pressure surface, 4 suction surface, 5 airflow, 6 casing.
Detailed Description
As shown in fig. 1 and 2, the double-wall structure of the blade tip of the turbine rotor blade comprises an outer wall 1 and an inner wall 2, wherein the outer wall 1 and the inner wall 2 are arranged on the blade tip of the turbine rotor blade, the outer wall 1 is arranged along the whole edge of the blade tip of the turbine rotor blade, the inner wall 2 is arranged in the outer wall 1, the outer wall 1 and the inner wall 2 are distributed in a parallel track mode, a first groove is formed between the outer wall 1 and the inner wall 2, and a second groove is formed by the inner wall 2. The chord length of the blade is 100mm, the thickness of the blade is 50mm, the thickness of the outer wall 1 and the thickness of the inner wall 2 are 2mm, the width of the first groove is 3mm, and the height of the outer wall 1 and the height of the inner wall 2 are 5 mm.
As shown in fig. 3, when the gas flows into the tip clearance between the casing 6 and the blade through the turbine rotor blade pressure surface 3, the gas flow 5 forms a low-speed vortex in the first groove between the outer wall 1 and the inner wall 2, and the vortex fills the first groove between the outer wall 1 and the inner wall 2 to prevent the gas from flowing into the turbine rotor blade suction surface 4 side through the tip clearance from the turbine rotor blade pressure surface 3 side; the same low-speed vortex can be formed in the groove I on one side of the suction surface 4 of the turbine rotor blade, so that gas leakage is prevented, the sealing effect is better, the fuel oil loss of an engine is reduced, and the thrust of an aircraft engine and a gas turbine is improved.
Claims (3)
1. The utility model provides a turbine rotor blade apex double-layer wall structure, its characterized in that includes outer wall and inner wall, outer wall and inner wall setting are on turbine rotor blade apex, the outer wall sets up along the whole edge of turbine rotor blade apex, the inner wall sets up in the outer wall, the outer wall with the inner wall is parallel track form and distributes, the outer wall with form recess one between the inner wall, the inner wall encloses into recess two.
2. The turbine rotor blade tip double-wall structure of claim 1, wherein the outer wall and inner wall thickness is 4-6% of the turbine rotor blade thickness; the width of the first groove is 1-2 times of the thickness of the inner wall.
3. The turbine rotor blade tip double-walled structure of claim 2, wherein the outer wall height is 1-2 times a width of the groove; the height of the inner wall is 90-110% of that of the outer wall.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011219499.7A CN112031878A (en) | 2020-11-05 | 2020-11-05 | Turbine rotor blade apex double-wall structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011219499.7A CN112031878A (en) | 2020-11-05 | 2020-11-05 | Turbine rotor blade apex double-wall structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112031878A true CN112031878A (en) | 2020-12-04 |
Family
ID=73573167
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011219499.7A Pending CN112031878A (en) | 2020-11-05 | 2020-11-05 | Turbine rotor blade apex double-wall structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112031878A (en) |
Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5660523A (en) * | 1992-02-03 | 1997-08-26 | General Electric Company | Turbine blade squealer tip peripheral end wall with cooling passage arrangement |
| EP0916811A2 (en) * | 1997-11-17 | 1999-05-19 | General Electric Company | Ribbed turbine blade tip |
| CN1931475A (en) * | 2005-05-09 | 2007-03-21 | 斯奈克玛服务公司 | Method of manufacturing a hollow blade comprising a squealer tip and method of repairing such a blade |
| CN101191424A (en) * | 2006-11-30 | 2008-06-04 | 通用电气公司 | Turbine blade and turbine blade cooling system and methods |
| CN101960097A (en) * | 2008-03-05 | 2011-01-26 | 斯奈克玛 | Turbine blade and corresponding turbine and turbo machine with most advanced and sophisticated cooling |
| CN102434220A (en) * | 2010-09-15 | 2012-05-02 | 通用电气公司 | Abradable bucket shroud |
| CN202417609U (en) * | 2011-12-13 | 2012-09-05 | 河南科技大学 | Turbine cooling blade with blade tip leakage prevention structure |
| CN104033186A (en) * | 2013-03-05 | 2014-09-10 | 株式会社日立制作所 | Gas Turbine Blade |
| CN104838092A (en) * | 2012-12-17 | 2015-08-12 | 通用电气公司 | Robust turbine blades |
| CN106481364A (en) * | 2015-09-02 | 2017-03-08 | 通用电气公司 | Construction for turbine rotor blade end |
| WO2017119898A1 (en) * | 2016-01-08 | 2017-07-13 | Siemens Aktiengesellschaft | Turbine blade with multi-layer multi-height blade squealer |
| CN107075956A (en) * | 2014-11-04 | 2017-08-18 | 赛峰飞机发动机公司 | Turbo blade with end cap |
| CN108506049A (en) * | 2018-03-15 | 2018-09-07 | 哈尔滨工业大学 | Inhibit the ball basal edge column cavity leaf top of turbine tip clearance flow |
-
2020
- 2020-11-05 CN CN202011219499.7A patent/CN112031878A/en active Pending
Patent Citations (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5660523A (en) * | 1992-02-03 | 1997-08-26 | General Electric Company | Turbine blade squealer tip peripheral end wall with cooling passage arrangement |
| EP0916811A2 (en) * | 1997-11-17 | 1999-05-19 | General Electric Company | Ribbed turbine blade tip |
| CN1931475A (en) * | 2005-05-09 | 2007-03-21 | 斯奈克玛服务公司 | Method of manufacturing a hollow blade comprising a squealer tip and method of repairing such a blade |
| CN101191424A (en) * | 2006-11-30 | 2008-06-04 | 通用电气公司 | Turbine blade and turbine blade cooling system and methods |
| CN101960097A (en) * | 2008-03-05 | 2011-01-26 | 斯奈克玛 | Turbine blade and corresponding turbine and turbo machine with most advanced and sophisticated cooling |
| CN102434220A (en) * | 2010-09-15 | 2012-05-02 | 通用电气公司 | Abradable bucket shroud |
| CN202417609U (en) * | 2011-12-13 | 2012-09-05 | 河南科技大学 | Turbine cooling blade with blade tip leakage prevention structure |
| CN104838092A (en) * | 2012-12-17 | 2015-08-12 | 通用电气公司 | Robust turbine blades |
| CN104033186A (en) * | 2013-03-05 | 2014-09-10 | 株式会社日立制作所 | Gas Turbine Blade |
| CN107075956A (en) * | 2014-11-04 | 2017-08-18 | 赛峰飞机发动机公司 | Turbo blade with end cap |
| CN106481364A (en) * | 2015-09-02 | 2017-03-08 | 通用电气公司 | Construction for turbine rotor blade end |
| WO2017119898A1 (en) * | 2016-01-08 | 2017-07-13 | Siemens Aktiengesellschaft | Turbine blade with multi-layer multi-height blade squealer |
| CN108506049A (en) * | 2018-03-15 | 2018-09-07 | 哈尔滨工业大学 | Inhibit the ball basal edge column cavity leaf top of turbine tip clearance flow |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CA2548168C (en) | Turbine airfoil with variable and compound fillet | |
| US8133032B2 (en) | Rotor blades | |
| US9074483B2 (en) | High camber stator vane | |
| US9284845B2 (en) | Turbine airfoil tip shelf and squealer pocket cooling | |
| US10253637B2 (en) | Method and system for improving turbine blade performance | |
| US20100158696A1 (en) | Curved platform turbine blade | |
| US9175567B2 (en) | Low loss airfoil platform trailing edge | |
| US20130108433A1 (en) | Non axis-symmetric stator vane endwall contour | |
| US10677066B2 (en) | Turbine blade with airfoil tip vortex control | |
| CN110030036B (en) | Impact split-joint air film cooling structure of turbine blade tail edge | |
| CN105179022A (en) | Turbine blade of blade top rib wing structure | |
| JP2012052526A (en) | Shrouded turbine blade with contoured platform and axial dovetail | |
| EP2775097A2 (en) | Stator vane row | |
| EP3392459A1 (en) | Compressor blades | |
| CN102434223B (en) | Low-pressure stage final blade of large-flow air-cooled steam turbine | |
| CN104005796A (en) | Groove impairment structure and method of novel turbine blade grid end wall | |
| CN104196573A (en) | Low-pressure last stage blade of rotating speed-variable air cooling industrial steam turbine with steam exhaust area of 3.6 m2 | |
| CN112012801A (en) | Tooth point hook type labyrinth seal structure | |
| CN109707461B (en) | High-rotation-speed high-load final-stage movable blade for industrial steam turbine | |
| CN112031878A (en) | Turbine rotor blade apex double-wall structure | |
| CN203948136U (en) | The low pressure stage group exhaust stage blade of leaving area 3.6m2 variable speed air cooling industrial steam turbine | |
| CN203584475U (en) | Turbine runner sealing structure and aeroengine turbine structure | |
| CN103806946B (en) | Leaving area 2.1m2variable rotating speed industrial steam turbine low-pressure stage group exhaust stage blade | |
| CN115263436A (en) | Transonic turbine rotor blade, turbine rotor and turbine | |
| CN115289499A (en) | Hollow support plate of gas inlet of combustion chamber of gas turbine |
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 | ||
| RJ01 | Rejection of invention patent application after publication | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20201204 |