CN106761951A - The leading edge cooling structure and the engine with it of a kind of turbine rotor blade - Google Patents
The leading edge cooling structure and the engine with it of a kind of turbine rotor blade Download PDFInfo
- Publication number
- CN106761951A CN106761951A CN201710048924.2A CN201710048924A CN106761951A CN 106761951 A CN106761951 A CN 106761951A CN 201710048924 A CN201710048924 A CN 201710048924A CN 106761951 A CN106761951 A CN 106761951A
- Authority
- CN
- China
- Prior art keywords
- leading edge
- blade
- rotor blade
- inner chamber
- turbine rotor
- 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
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
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
- F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a kind of leading edge cooling structure of turbine rotor blade and the engine with it, it is related to technical field of engines.Spirality flow-disturbing rib (11) is provided with the leading edge inner chamber (1) of the leading edge cooling structure rotor blade shown of the turbine rotor blade, the spirality flow-disturbing rib (11) is set in leading edge inner chamber (1) along the short transverse of blade.Leading edge cooling structure of the rotor blade comprising turbine rotor blade as described above on the engine.The advantage of the invention is that：Spirality flow-disturbing rib is provided with leading edge inner chamber, heat convection area is increased relative to traditional horizontal flow-disturbing rib, the boundary layer separation and turbulent closure scheme near internal face are enhanced, promotes cold air to produce eddy flow in leading edge inner chamber, further the heat convection of reinforcing internal face and cold air.Enhanced heat exchange is carried out to blade inlet edge inner chamber, the cooling effect of front edge area has been improve, has played a part of to reduce blade inlet edge temperature while leading edge cold air consumption is not increased.
Description
Technical field
The present invention relates to technical field of engines, and in particular to a kind of leading edge cooling structure of turbine rotor blade and have
Its engine.
Background technology
Generally, gas-turbine unit is made up of compressor, combustion chamber and the big part of turbine three, and turbine part fires high temperature
The interior energy of gas is converted into mechanical energy.At present, the turbo blade of gas-turbine unit needs to be worked under the gas condition of high temperature,
Bear to exceed well over the temperature of matrix material fusing point, the temperature capability of first order turbo blade determines the whirlpool of gas-turbine unit
Temperature before wheel.Need to draw the cavity that the relatively low pressure-air of temperature enters blade from compressor under normal circumstances, as cooling
Gas is cooled down to turbo blade.
Multiple cooling ducts can be typically designed in blade cavity, the difference of the thermic load according to blade wall different zones
It is different, to design different cooling structures and blade is effectively cooled down, control blade wall surface temperature is in the range of use and tries one's best
Uniformly.At present, turbine rotor blade has applied the internal cooling stream of complexity, using cold air as few as possible, reaches to the greatest extent
Possible cooling effect high, so as to improve the overall efficiency of engine.
Turbine rotor blade blade can be divided into leading edge, exhaust side, leaf basin and blade back by region.For blade, leading edge
Region is directly facing high-temperature fuel gas direction of flow, and in the stationary point position of leading edge, vanes receive highest thermic load, it is therefore desirable to right
Front edge area carries out reinforcing cooling, reduces blade wall surface temperature.Horizontal flow-disturbing rib is generally set using gaseous film control or inner chamber
Mode to leading edge strengthen cool down.
In the prior art, generally there are two methods for reducing leading edge temperature：A kind of method is to set multiple rows of edge in leading edge
The radial skew angle of the air film hole of leaf high radial distribution, leading edge air film hole and blade is between 20 degree to 30 degree.Cold air enters preceding
After edge cavity, blade is flowed out by air film hole, edge surface forms film overcast in front of the blade, reduce the wall surface temperature of leading edge；In addition
A kind of method is that the horizontal flow-disturbing rib being distributed along leaf high radial is set in leading edge cavity.Cold air enters in leading edge cavity, from leaf
Root radially, finally by blade tip position discharge blade, carries out convection current cooling, horizontal flow-disturbing rib to blade tip to inwall during flowing
So that there is boundary layer separation during cold air flow, convection transfer rate, the heat exchange of enhancing cold air and internal face are improved.
Although above two method is all cooled down to front edge area to a certain extent, with turbine before turbine
Temperature it is constantly improve, under limited cold air consumption, easily there is the not enough problem of cooling in leading edge, it is therefore desirable to cold to leading edge
But structure is improved, and further improves cooling effect.
The content of the invention
It is an object of the invention to provide a kind of leading edge cooling structure of turbine rotor blade and the engine with it, to solve
The problem at least one certainly or at least in the presence of mitigation background technology.
The technical solution adopted by the present invention is：A kind of leading edge cooling structure of turbine rotor blade is provided, rotor blade
Spirality flow-disturbing rib is provided with leading edge inner chamber, the spirality flow-disturbing rib sets in leading edge inner chamber along the short transverse of blade
Put.
Preferably, the spirality flow-disturbing rib is set to multiple spiral circle segmental arcs.
Preferably, the side wall of the leading edge inner chamber includes the first side wall near trailing edge direction and away from trailing edge direction the
Two side walls, are provided with impact opening on the first side wall, the impact opening is used for transporting cold-air in the leading edge inner chamber and rushes
Hit in the second sidewall；Air film hole is provided with the second sidewall, the air film hole is used to make in the leading edge inner chamber
Cold air flow to blade outer surface.
Preferably, the air film hole comprises at least 3, and one of them is arranged at leading edge stationary point, and 2 are arranged on institute in addition
The both sides in leading edge stationary point are stated, the both sides refer respectively to be close to the side and the side near blade back of leaf basin.
Preferably, multiple spiral circle segmental arcs are separately positioned on the both sides in leading edge stationary point, and the both sides refer respectively to lean on
The side and the side near blade back of nearly leaf basin.
Preferably, the spiral circle segmental arc is integrally formed with the side wall of the leading edge inner chamber.
Preferably, multiple impact openings, short transverse of multiple impact openings in blade are provided with the first side
Set gradually；Multiple air film holes are provided with the second sidewall, multiple air film holes set successively in the short transverse of blade
Put.
Preferably, the second sidewall is provided with multiple air film holes in the same section in blade height direction, described many
The axis shape of any 2 air film holes has angle in individual air film hole, and any two air film hole is in the one end away from leading edge inner chamber
Distance of the distance more than the one end near leading edge inner chamber.
Preferably, the impact opening is set to bellmouth, and the small end opening of the bellmouth is towards the leading edge inner chamber of blade.
Present invention also offers a kind of engine, the rotor blade on the engine includes turbine rotor as described above
The leading edge cooling structure of blade.
The beneficial effects of the present invention are：The leading edge cooling structure of turbine rotor blade of the invention sets in leading edge inner chamber
Spirality flow-disturbing rib is put, heat convection area has been increased relative to traditional horizontal flow-disturbing rib, it is often more important that in enhancing
The boundary layer separation and turbulent closure scheme of near wall, promote cold air leading edge inner chamber produce eddy flow, further reinforcing internal face and
The heat convection of cold air.Enhanced heat exchange is carried out to blade inlet edge inner chamber, the cooling effect of front edge area has been improve, not increased
Play a part of to reduce blade inlet edge temperature while leading edge cold air consumption.And because the type of cooling improves the cold of leading edge
But effect, can reduce cold air consumption, so as to improve overall efficiency in the case where blade inlet edge wall surface temperature level is maintained.
Brief description of the drawings
Fig. 1 is the schematic cross-section of the leading edge cooling structure of the turbine rotor blade of one embodiment of the invention.
Fig. 2 is the schematic cross-section of the leading edge cooling structure in blade height direction of the turbine rotor blade shown in Fig. 1.
Fig. 3 is the enlarged drawing at A shown in Fig. 2.
Wherein, 1- leading edges inner chamber, 11- spirality flow-disturbing ribs, 12- the first side walls, 13- second sidewalls, 2- impact openings, 3- gas
Fenestra, 4- leaf basins, 5- blade backs.
Specific embodiment
To make the purpose, technical scheme and advantage of present invention implementation clearer, below in conjunction with the embodiment of the present invention
Accompanying drawing, the technical scheme in the embodiment of the present invention is further described in more detail.In the accompanying drawings, identical from start to finish or class
As label represent same or similar element or the element with same or like function.Described embodiment is the present invention
A part of embodiment, rather than whole embodiments.Embodiment below with reference to Description of Drawings is exemplary, it is intended to used
It is of the invention in explaining, and be not considered as limiting the invention.Based on the embodiment in the present invention, ordinary skill people
The every other embodiment that member is obtained under the premise of creative work is not made, belongs to the scope of protection of the invention.Under
Face is described in detail with reference to accompanying drawing to embodiments of the invention.
In the description of the invention, it is to be understood that term " " center ", " longitudinal direction ", " transverse direction ", "front", "rear",
The orientation or position relationship of the instruction such as "left", "right", " vertical ", " level ", " top ", " bottom " " interior ", " outward " are based on accompanying drawing institute
The orientation or position relationship for showing, are for only for ease of the description present invention and simplify description, rather than the dress for indicating or implying meaning
Put or element with specific orientation, with specific azimuth configuration and operation, therefore it is not intended that must be protected to the present invention
The limitation of scope.
As shown in Figure 1 to Figure 3, the leading edge cooling structure of a kind of turbine rotor blade, sets in the leading edge inner chamber 1 of rotor blade
Spirality flow-disturbing rib 11 is equipped with, spirality flow-disturbing rib 11 is set in leading edge inner chamber 1 along the short transverse of blade.The blade
Short transverse refers to the line direction from blade root to blade tip.
The leading edge cooling structure of turbine rotor blade of the invention is provided with spirality flow-disturbing rib in leading edge inner chamber, relatively
Increase heat convection area in traditional horizontal flow-disturbing rib, it is often more important that enhance boundary layer separation near internal face and
Turbulent closure scheme, promotes cold air to produce eddy flow in leading edge inner chamber, further the heat convection of reinforcing internal face and cold air.Before blade
Edge inner chamber has carried out enhanced heat exchange, improves the cooling effect of front edge area, is played while leading edge cold air consumption is not increased
Reduce the effect of blade inlet edge temperature.
In the present embodiment, spirality flow-disturbing rib 11 is set to multiple spiral circle segmental arcs.The rotation direction of spiral circle segmental arc can be with
It is clockwise, or counterclockwise, the part wall of whole leading edge inner chamber or inner chamber can be covered.In the present embodiment,
The spacing of two spiral circle segmental arcs of arbitrary neighborhood is equal in blade height direction, multiple spiral circle segmental arcs；Can manage
Solution, the spacing between two neighboring spiral circle segmental arc can also set according to actual conditions, for example, in an alternative implementation
In example, the leading edge according to blade can set the spiral circle segmental arc of different spacing, in leaf in the different zones in blade height direction
The leading edge temperature of piece region higher, can reduce the distance between two neighboring spiral circle segmental arc, increase heat convection area,
In the relatively low region of temperature, the distance between two neighboring spiral circle segmental arc can be increased.
It is understood that spirality flow-disturbing rib 11 may be arranged as a continuous spirality flow-disturbing rib, for example,
In another alternative, spirality flow-disturbing rib 11 extends upwardly to blade tip position from the blade root position of blade.Equally, according to
The leading edge of blade can set different pitch in the different zones in blade height direction, higher in the leading edge temperature of blade
Region, can reduce the pitch of spirality flow-disturbing rib 11, increase heat convection area, in the relatively low region of temperature, can increase
The pitch of spirality flow-disturbing rib 11.
In the present embodiment, the side wall of leading edge inner chamber 1 is included near the first side wall 12 in trailing edge direction and away from trailing edge side
To second sidewall 13, be provided with impact opening 2 on the first side wall 12, impact opening 2 is used in forward edge inner chamber 1 transporting cold-air and rushes
Hit in second sidewall 13；Air film hole 3 is provided with second sidewall 13, air film hole 3 is used to flow to the cold air in leading edge inner chamber 1
Blade outer surface.
In the present embodiment, air film hole 3 comprises at least 3, and one of them is arranged at leading edge stationary point, in addition 2 settings
In the both sides in the leading edge stationary point, the both sides refer respectively to be close to the side and the side near blade back 5 of leaf basin 4.
In the present embodiment, multiple spiral circle segmental arcs are separately positioned on the both sides in leading edge stationary point, the both sides difference
Refer near the side of leaf basin 4 and the side near blade back 5.The side wall one of the spiral circle segmental arc and leading edge inner chamber 1 into
Type.
In the present embodiment, multiple impact openings 2 are provided with first side 12, multiple impact openings 2 are in the height side of blade
To setting gradually；Multiple air film holes 3 are provided with second sidewall 13, multiple air film holes 3 set gradually in the short transverse of blade.
Cooling air enters leading edge inner chamber 1 by radially-arranged multiple impact openings 2, and impact leading edge stationary point position is corresponding interior
Wall, the heat exchange of cold air and inwall is strengthened using the mode of shock peening, and the aperture of impact opening 2 is particular value in design, is played
Throttling action, the cold air flow of leading edge inner chamber 1 can be adjusted into by adjusting aperture.A cold air part in leading edge inner chamber 1
Flowed out by the air film hole 3 of the radially-arranged air film hole 3 of the row for being arranged in leading edge stationary point and its both sides, one is covered in blade back overlying
Layer cooling air film, high-temperature fuel gas and blade isolation are come, and reduce outside wall surface temperature；Effect of another part cold air in centrifugal force
Under, continue to be flowed to blade tip direction in leading edge inner chamber 1.
In the present embodiment, second sidewall 13 is provided with multiple air film holes 3 in the same section in blade height direction, many
The axis shape of any 2 air film holes has angle in individual air film hole 3, and any two air film hole is in the one end away from leading edge inner chamber
Distance is conducive to cold air uniformly to form cold air film in leading edge outer surface more than the distance of the one end near leading edge inner chamber.
In the present embodiment, impact opening is set to cylindrical hole.It is understood that the impact opening can also be set
It is bellmouth, the small end opening of the bellmouth the advantage is that towards the leading edge inner chamber of blade, is conducive to improving cold air to the
The impact velocity of two side faces.
Present invention also offers a kind of engine, the rotor blade on the engine includes turbine rotor as described above
The leading edge cooling structure of blade.The cooling effect of leading edge is improve due to the type of cooling, blade inlet edge wall can maintained
In the case of temperature levels, cold air consumption is reduced, so as to improve overall efficiency.
It is last it is to be noted that：The above embodiments are merely illustrative of the technical solutions of the present invention, rather than its limitations.To the greatest extent
Pipe has been described in detail to the present invention with reference to the foregoing embodiments, it will be understood by those within the art that：It is still
Technical scheme described in foregoing embodiments can be modified, or which part technical characteristic is equally replaced
Change；And these modifications or replacement, do not make the essence of the essence disengaging various embodiments of the present invention technical scheme of appropriate technical solution
God and scope.
Claims (10)
1. the leading edge cooling structure of a kind of turbine rotor blade, it is characterised in that：
Spirality flow-disturbing rib (11) is provided with the leading edge inner chamber (1) of rotor blade, the spirality flow-disturbing rib (11) is in leading edge
Short transverse in inner chamber (1) along blade is set.
2. the leading edge cooling structure of turbine rotor blade as claimed in claim 1, it is characterised in that：The spirality flow-disturbing rib
(11) multiple spiral circle segmental arcs are set to.
3. the leading edge cooling structure of turbine rotor blade as claimed in claim 2, it is characterised in that：The leading edge inner chamber (1)
Side wall include near trailing edge direction the first side wall (12) and the second sidewall (13) away from trailing edge direction, the first side wall
(12) it is provided with impact opening (2) on, the impact opening (2) is for the interior transporting cold-air of the leading edge inner chamber (1) and impacting in institute
State in second sidewall (13)；Be provided with air film hole (3) on the second sidewall (13), the air film hole (3) for make it is described before
Cold air in edge inner chamber (1) flows to blade outer surface.
4. the leading edge cooling structure of turbine rotor blade as claimed in claim 3, it is characterised in that：The air film hole (3) is extremely
3 are included less, one of them is arranged at leading edge stationary point, in addition 2 both sides for being arranged on the leading edge stationary point, the both sides point
Do not refer near the side of leaf basin (4) and the side near blade back (5).
5. the leading edge cooling structure of turbine rotor blade as claimed in claim 4, it is characterised in that：Multiple spiral arcs
Section is separately positioned on the both sides in leading edge stationary point, and the both sides refer respectively to be close to the side of leaf basin (4) and near blade back (5)
Side.
6. the leading edge cooling structure of turbine rotor blade as claimed in claim 5, it is characterised in that：The spiral circle segmental arc with
The side wall of the leading edge inner chamber (1) is integrally formed.
7. the leading edge cooling structure of turbine rotor blade as claimed in claim 6, it is characterised in that：The first side (12)
On be provided with multiple impact openings, multiple impact openings set gradually in the short transverse of blade；On the second sidewall (13)
Multiple air film holes are provided with, multiple air film holes set gradually in the short transverse of blade.
8. the leading edge cooling structure of turbine rotor blade as claimed in claim 7, it is characterised in that：The second sidewall is in leaf
Multiple air film holes are provided with the same section of piece short transverse, the axis shape of any 2 air film holes in the multiple air film hole
Have angle, any two air film hole the one end away from leading edge inner chamber distance be more than near leading edge inner chamber one end away from
From.
9. the leading edge cooling structure of turbine rotor blade as claimed in claim 8, it is characterised in that：The impact opening is set to
Bellmouth, the small end opening of the bellmouth is towards the leading edge inner chamber of blade.
10. a kind of engine, it is characterised in that：Rotor blade on the engine is included such as any one of claim 1 to 9 institute
The leading edge cooling structure of the turbine rotor blade stated.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710048924.2A CN106761951A (en) | 2017-01-23 | 2017-01-23 | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710048924.2A CN106761951A (en) | 2017-01-23 | 2017-01-23 | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN106761951A true CN106761951A (en) | 2017-05-31 |
Family
ID=58941479
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710048924.2A Pending CN106761951A (en) | 2017-01-23 | 2017-01-23 | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106761951A (en) |
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108150224A (en) * | 2017-12-21 | 2018-06-12 | 西安交通大学 | A kind of eddy flow is the same as impacting cooling structure inside the turbine blade being combined |
| CN110773501A (en) * | 2019-12-02 | 2020-02-11 | 南京航空航天大学 | Device and method for removing blocking holes of thermal barrier coating of turbine blade by pulsating porous water jet |
| CN111485956A (en) * | 2020-04-24 | 2020-08-04 | 中国航发湖南动力机械研究所 | Turbine guide cooling blade |
| CN112178692A (en) * | 2020-10-27 | 2021-01-05 | 西北工业大学 | Longitudinal corrugated cooling structure with L-shaped impact orifice plate |
| CN112302727A (en) * | 2020-11-23 | 2021-02-02 | 华能国际电力股份有限公司 | Turbine blade leading edge cooling structure |
| CN112610284A (en) * | 2020-12-17 | 2021-04-06 | 东北电力大学 | Gas turbine blade with spiral band |
| CN112867844A (en) * | 2019-01-17 | 2021-05-28 | 三菱动力株式会社 | Turbine rotor blade and gas turbine |
| WO2021104002A1 (en) * | 2019-11-29 | 2021-06-03 | 大连理工大学 | Curvilinear exhaust slit structure for trailing edge of turbine blade |
| CN114046180A (en) * | 2021-11-02 | 2022-02-15 | 西北工业大学 | Combined hole air film cooling structure utilizing rotational flow |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
| CN115247575A (en) * | 2022-05-12 | 2022-10-28 | 中国航发四川燃气涡轮研究院 | Spiral turbine blade cooling unit and cooling structure |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6139258A (en) * | 1987-03-30 | 2000-10-31 | United Technologies Corporation | Airfoils with leading edge pockets for reduced heat transfer |
| US20050025623A1 (en) * | 2003-08-01 | 2005-02-03 | Snecma Moteurs | Cooling circuits for a gas turbine blade |
| CN104088673A (en) * | 2008-11-07 | 2014-10-08 | 三菱重工业株式会社 | Vane for turbine |
| CN205743994U (en) * | 2016-05-18 | 2016-11-30 | 中航商用航空发动机有限责任公司 | Flow-disturbing rib structure and turbo blade for turbo blade |
-
2017
- 2017-01-23 CN CN201710048924.2A patent/CN106761951A/en active Pending
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6139258A (en) * | 1987-03-30 | 2000-10-31 | United Technologies Corporation | Airfoils with leading edge pockets for reduced heat transfer |
| US20050025623A1 (en) * | 2003-08-01 | 2005-02-03 | Snecma Moteurs | Cooling circuits for a gas turbine blade |
| CN104088673A (en) * | 2008-11-07 | 2014-10-08 | 三菱重工业株式会社 | Vane for turbine |
| CN205743994U (en) * | 2016-05-18 | 2016-11-30 | 中航商用航空发动机有限责任公司 | Flow-disturbing rib structure and turbo blade for turbo blade |
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN108150224A (en) * | 2017-12-21 | 2018-06-12 | 西安交通大学 | A kind of eddy flow is the same as impacting cooling structure inside the turbine blade being combined |
| CN112867844A (en) * | 2019-01-17 | 2021-05-28 | 三菱动力株式会社 | Turbine rotor blade and gas turbine |
| CN112867844B (en) * | 2019-01-17 | 2023-12-08 | 三菱重工业株式会社 | Turbine bucket and gas turbine |
| WO2021104002A1 (en) * | 2019-11-29 | 2021-06-03 | 大连理工大学 | Curvilinear exhaust slit structure for trailing edge of turbine blade |
| CN110773501A (en) * | 2019-12-02 | 2020-02-11 | 南京航空航天大学 | Device and method for removing blocking holes of thermal barrier coating of turbine blade by pulsating porous water jet |
| CN111485956A (en) * | 2020-04-24 | 2020-08-04 | 中国航发湖南动力机械研究所 | Turbine guide cooling blade |
| CN112178692A (en) * | 2020-10-27 | 2021-01-05 | 西北工业大学 | Longitudinal corrugated cooling structure with L-shaped impact orifice plate |
| CN112302727A (en) * | 2020-11-23 | 2021-02-02 | 华能国际电力股份有限公司 | Turbine blade leading edge cooling structure |
| CN112610284A (en) * | 2020-12-17 | 2021-04-06 | 东北电力大学 | Gas turbine blade with spiral band |
| CN114046180A (en) * | 2021-11-02 | 2022-02-15 | 西北工业大学 | Combined hole air film cooling structure utilizing rotational flow |
| CN114046180B (en) * | 2021-11-02 | 2024-01-30 | 西北工业大学 | Combined hole air film cooling structure utilizing rotational flow |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
| CN115247575A (en) * | 2022-05-12 | 2022-10-28 | 中国航发四川燃气涡轮研究院 | Spiral turbine blade cooling unit and cooling structure |
| CN115247575B (en) * | 2022-05-12 | 2024-05-03 | 中国航发四川燃气涡轮研究院 | Helical turbine blade cooling unit and cooling structure |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106761951A (en) | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade | |
| JP5306613B2 (en) | Tip inclined surface turbine blade | |
| US9175569B2 (en) | Turbine airfoil trailing edge cooling slots | |
| EP2428683B1 (en) | Centrifugal fan and air conditioner | |
| JP4184323B2 (en) | Hollow rotor blades for gas turbine engine turbines | |
| JP6001696B2 (en) | Turbine blade with swirling cooling channel and cooling method thereof | |
| US9334745B2 (en) | Gas turbine stator vane | |
| US10012087B2 (en) | Gas turbine including a contoured end wall section of a rotor blade | |
| JP2001214707A (en) | Turbine nozzle equipped with film cooling with gradient | |
| JP2007002843A (en) | Cooling circuit for movable blade of turbo machine | |
| JPH1089006A (en) | Cooling type blade | |
| CN106870015A (en) | A kind of labyrinth type internal cooling structure for high-temperature turbine movable vane trailing edge | |
| CN110043328A (en) | A kind of cooled change geometry low-pressure turbine guide vane | |
| CN107131006A (en) | Turbo blade | |
| JPH09195705A (en) | Axial-flow turbine blade | |
| CN113153447B (en) | Prerotation structure for strengthening cooling of leakage flow of end wall of turbine stationary blade | |
| JP2012202335A (en) | Impingement cooling structure and gas turbine stator blade using the same | |
| CN106703899A (en) | High-pressure turbine rotor blade front edge impingement cooling structure and engine with same | |
| CN106930977A (en) | A kind of direct-cooled noise reduction axial flow blower | |
| CN106837431A (en) | A kind of method for improving aero gas turbine engine transient process turbine blade air film cooling air-flow quality | |
| US11401820B1 (en) | Cooling structure and method of trailing-edge cutback of turbine blade, and turbine blade | |
| JP2003322003A (en) | Turbine airfoil part having single three-passage zigzag cooling circuit flowing rearward | |
| CN205445688U (en) | Gas turbine turbine does not have hat film cooling rotor blade | |
| CN206845538U (en) | A kind of direct-cooled noise reduction axial flow blower | |
| JPH09195703A (en) | Cooled moving blade for 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: 20170531 |