CN107013253A - A kind of efficient cooling blade of gas-turbine unit - Google Patents
A kind of efficient cooling blade of gas-turbine unit Download PDFInfo
- Publication number
- CN107013253A CN107013253A CN201710356212.7A CN201710356212A CN107013253A CN 107013253 A CN107013253 A CN 107013253A CN 201710356212 A CN201710356212 A CN 201710356212A CN 107013253 A CN107013253 A CN 107013253A
- Authority
- CN
- China
- Prior art keywords
- blade
- special
- channel
- film hole
- shaped air
- 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
- 238000001816 cooling Methods 0.000 title claims abstract description 38
- 239000007921 spray Substances 0.000 claims abstract description 6
- 238000000034 method Methods 0.000 claims description 3
- 230000000694 effects Effects 0.000 abstract description 7
- 230000009471 action Effects 0.000 abstract description 2
- 239000007789 gas Substances 0.000 description 5
- 238000012913 prioritisation Methods 0.000 description 4
- 239000000112 cooling gas Substances 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- 238000012546 transfer Methods 0.000 description 3
- 238000002485 combustion reaction Methods 0.000 description 2
- 208000002925 dental caries Diseases 0.000 description 2
- 239000002737 fuel gas Substances 0.000 description 2
- 239000012720 thermal barrier coating Substances 0.000 description 2
- 206010020843 Hyperthermia Diseases 0.000 description 1
- 241000237509 Patinopecten sp. Species 0.000 description 1
- 238000005452 bending Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000006835 compression Effects 0.000 description 1
- 238000007906 compression Methods 0.000 description 1
- 239000012809 cooling fluid Substances 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 230000002708 enhancing effect Effects 0.000 description 1
- 230000004907 flux Effects 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 230000036031 hyperthermia Effects 0.000 description 1
- 239000007769 metal material Substances 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 235000020637 scallop Nutrition 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 238000003786 synthesis reaction Methods 0.000 description 1
- 230000008646 thermal stress Effects 0.000 description 1
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
- 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
-
- 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
Abstract
The invention discloses a kind of efficient cooling blade of gas-turbine unit, comprising blade, end wall and tenon, blade is in hollow cylinder, includes pressure face and suction surface；The first to the 3rd bridge joint beam is sequentially provided with from blade inlet edge to blade trailing edge in blade, is air cavity, first to third channel by blade inside division；The first special-shaped air film hole that suction surface is provided with and air cavity is communicated；The 4th special-shaped air film hole that the 3rd special-shaped air film hole and third channel that the second special-shaped air film hole and second channel that pressure face is provided with and air cavity is communicated is communicated is communicated；First bridge joint beam is provided with impact jet flow hole；Blade inlet edge is provided with spray apertures；Air cavity, the interface channel of second channel and the 3rd special-shaped air film hole, it is equipped with heat exchange structure in the interface channel of third channel and the 4th special-shaped air film hole.The present invention has played the perturbation action that heat exchange structure flows to cooling air-flow, has enhanced the protection effect of gaseous film control while efficient internal cooling is implemented to blade.
Description
Technical field
The present invention relates to the Local Heat Transfer field of the high heat fluxs such as Aeronautics and Astronautics, dynamic power machine, more particularly to a kind of combustion
The efficient cooling blade of gas eddy turbine.
Background technology
In modern high performance gas-turbine unit, it is increase thrust-weight ratio to improve turbine inlet temperature (TIT), reduces rate of fuel consumption
Effective measures.But with the raising of fuel gas temperature at turbine inlet, the thermic load of turbo blade will constantly increase.Modern times combustion
Temperature is up to 1900K at the turbine inlet of gas eddy turbine, oneself beyond existing metal material permission temperature in use, so
Effective cooling is only carried out to turbo blade, Turbine Blade Temperature Field under working condition is reduced, can just make it normal safely
Work, and meet certain life requirements.In addition, when turbo blade (working-blade) works at high speed, in very high
Centrifugal field among.Under the collective effect of aerodynamic force, thermal stress and huge centrifugal force, blade had both needed enough cold
But, it is also desirable to retain on higher strength level.
On modern turbo blade, basically form by internal cooling, outside cooling and three part groups of thermal barrier coating
Into the synthesis type of cooling.Outside cooling is more combined using gaseous film control and thermal barrier coating by the way of, internal cooling is then used pair
Stream, impinging cooling, many bending strip rib passages, with cooling structures such as turbulence columns passages.In recent years, for the structure of gaseous film control
Optimization carried out a series of exploratory development both at home and abroad, developed as slit-like hole, dust-pan shaped hole, scallop hole, crescent hole with
And the diversified structure type such as bellmouth.In addition, substantial amounts of, there are some researches show internally increase coarse in cooling duct
Rib and turbulence columns largely improve the heat transfer effect of internal cooling channel.
Blade inlet edge carries out thermal protection frequently with the mode that impact aerating film is cooled down, and trailing edge then adds to split using turbulence columns to be stitched out
The technologies such as stream.In middle part of blade, due to aerodynamic loss in terms of consideration, often using less Film Cooling, form one piece of temperature
The of a relatively high region of degree.
The content of the invention
The technical problems to be solved by the invention are for turbine blade surface thermic load is too high, threaten in background technology
There is provided a kind of efficient cooling blade of gas-turbine unit for the defect of the safe and stable operation of gas-turbine unit.
The present invention uses following technical scheme to solve above-mentioned technical problem：
A kind of efficient cooling blade of gas-turbine unit, includes the blade, end wall and tenon being sequentially connected, the LJ use
Connected in the turbine disk；
The blade is in hollow cylinder, includes pressure face and suction surface；
The pressure face and suction surface be in cambered surface, wherein, suction surface is arranged on outside pressure face, and suction surface both sides respectively and
The both sides connection of pressure face, forms blade inlet edge and blade trailing edge；
In the blade the first to the 3rd bridge joint beam, and the described first to the 3rd bridge are sequentially provided with from blade inlet edge to blade trailing edge
Beam is connect to set along blade direction of extension；
The first bridge joint beam is set in cambered surface, along the blade inlet edge, and long and narrow air cavity is formed between blade inlet edge；
Space beyond air cavity in blade is divided into first passage, second channel and threeway by the described first to the 3rd bridge joint beam
Road；
Be sequentially provided with the end wall, LJ along blade direction of extension lead to first passage, second channel, third channel it is logical
Hole；
The suction surface is in the first special-shaped air film hole being provided with blade inlet edge and the air cavity is communicated；
The pressure face is at close blade inlet edge provided with the second special-shaped air film hole communicated with the air cavity, close to blade tail
Be provided with edge and the 4th special-shaped air film hole that the third channel is communicated, the second special-shaped air film hole and the 4th special-shaped air film hole it
Between be provided with and the 3rd special-shaped air film hole that the second channel is communicated；
The impact jet flow hole that some and described air cavity is communicated is uniformly provided with the first bridge joint beam；
The spray apertures that some and described air cavity is communicated are uniformly provided with the blade inlet edge；
In the air cavity, in the interface channel of second channel and the 3rd special-shaped air film hole, third channel and the 4th special-shaped air film hole
Interface channel in be equipped with heat exchange structure.
It is used as a kind of efficient further prioritization scheme of cooling blade of gas-turbine unit of the invention, the suction surface
On be additionally provided with the 6th special-shaped air film hole that the 5th special-shaped air film hole and third channel communicated with the second channel is communicated, and
The junction of the junction of the second channel and the 5th special-shaped air film hole, third channel and the 6th special-shaped air film hole is all provided with
There is heat exchange structure.
It is used as a kind of efficient further prioritization scheme of cooling blade of gas-turbine unit of the invention, the heat exchange knot
Structure includes some fins being arranged on two apparent surfaces, and the fin on same surface is parallel, and each fin and
Angle between the direction flowed by gas between two surfaces is between 30 ° to 70 °.
It is used as a kind of efficient further prioritization scheme of cooling blade of gas-turbine unit of the invention, described two phases
90 ° are less than or equal to the angle between the fin of surface.
It is used as a kind of efficient further prioritization scheme of cooling blade of gas-turbine unit of the invention, described two phases
The waterpower for being 0.25~1, fin cross section to the hydraulic diameter of fin cross section on surface and the scope of the ratio between rib spacing is straight
The scope of the ratio between footpath and rib height is 0.4~1.2, and the high scope the ratio between the distance between with two apparent surfaces of rib is 0.5
~1, the fin on one of surface is embedded in corresponding fin on another surface.
The present invention uses above technical scheme compared with prior art, with following technique effect：
1. heat exchange structure to shuttle above and below the cooling fluid flowed wherein, upper and lower surface and fin to passage are rushed
Hit, substantially increase the heat transfer effect of this part；
2. the special-shaped air film hole in heat exchange structure downstream is favorably improved the adherence quality and ductility of downstream air film, with realization pair
The more preferable thermal protection effect of blade surface；
3. the present invention has played heat exchange structure and cooling air-flow flowing has been disturbed while efficient internal cooling is implemented to blade
Action is used so that the air film of blowout is uniform, localized hyperthermia's phenomenon caused by air film hole downstream air film temperature can be avoided uneven, enhancing
The protection effect of gaseous film control.
Brief description of the drawings
Fig. 1 is the structural representation of the present invention；
Fig. 2 is the structural representation of Leaf of the present invention；
Fig. 3 is the diagrammatic cross-section of Leaf of the present invention；
Fig. 4 is a kind of structural representation of heat exchange structure in the present invention；
Fig. 5 is another structural representation of heat exchange structure in the present invention.
In figure, 1- blades, 2- end walls, 3- tenons, 4- leaf tails, 5- turbine disk center lines, 6- blade tips, 7- pressure faces, 8- inhales
Power face, 9- blade inlet edges, 10- blade trailing edges, the bridge joint beams of 11- first, the bridge joint beams of 12- second, the bridge joint beams of 13- the 3rd, 14- the
Fin on heat exchange structure at one passage, 15- second channels, 16-- third channels, 17- air cavitys, 18- blade inlet edge inwalls,
The company of fin on the bridge joint beams of 19- first, 20- spray apertures, 21- impact jet flows hole, 22- second channels and the 3rd special-shaped air film hole
Connect the heat exchange structure at place, the heat exchange structure of the junction of 23- third channels and the 4th special-shaped air film hole, 24- air cavitys, 25- second
Special-shaped air film hole, the special-shaped air film holes of 26- first, the special-shaped air film holes of 27- the 3rd, the special-shaped air film holes of 28- the 4th.
Embodiment
Technical scheme is described in further detail below in conjunction with the accompanying drawings：
As depicted in figs. 1 and 2, the invention discloses a kind of efficient cooling blade of gas-turbine unit, comprising being sequentially connected
Blade, end wall and tenon, described LJ be used for and the turbine disk connection；
The blade is in hollow cylinder, includes pressure face and suction surface；
The pressure face and suction surface be in cambered surface, wherein, suction surface is arranged on outside pressure face, and suction surface both sides respectively and
The both sides connection of pressure face, forms blade inlet edge and blade trailing edge；
In the blade the first to the 3rd bridge joint beam, and the described first to the 3rd bridge are sequentially provided with from blade inlet edge to blade trailing edge
Beam is connect to set along blade direction of extension；
The first bridge joint beam is set in cambered surface, along the blade inlet edge, and long and narrow air cavity is formed between blade inlet edge；
Space beyond air cavity in blade is divided into first passage, second channel and threeway by the described first to the 3rd bridge joint beam
Road；
Be sequentially provided with the end wall, LJ along blade direction of extension lead to first passage, second channel, third channel it is logical
Hole；
The suction surface is in the first special-shaped air film hole being provided with blade inlet edge and the air cavity is communicated；
The pressure face is at close blade inlet edge provided with the second special-shaped air film hole communicated with the air cavity, close to blade tail
Be provided with edge and the 4th special-shaped air film hole that the third channel is communicated, the second special-shaped air film hole and the 4th special-shaped air film hole it
Between be provided with and the 3rd special-shaped air film hole that the second channel is communicated；
The impact jet flow hole that some and described air cavity is communicated is uniformly provided with the first bridge joint beam；
The spray apertures that some and described air cavity is communicated are uniformly provided with the blade inlet edge；
In the air cavity, in the interface channel of second channel and the 3rd special-shaped air film hole, third channel and the 4th special-shaped air film hole
Interface channel in be equipped with heat exchange structure.
Turbine disk center line is radially inserted as axial reference line, tenon, and one is circumferentially installed along the turbine disk and encloses turbine
Blade.Blade is stretched to blade tip vertically since the leaf tail on end wall.During operation, high-temperature fuel gas impact blade, from compression
The air that power traction enters enters from tenon bottom, and flow into blade interior first provides cooling gas to third channel, to blade.
The 5th special-shaped air film hole and third channel communicated with the second channel can also be provided with the suction surface again
The 6th special-shaped air film hole communicated, and the junction of the second channel and the 5th special-shaped air film hole, third channel and the
The junction of six special-shaped air film holes is equipped with heat exchange structure.
The heat exchange structure includes some fins being arranged on two apparent surfaces, and the fin on same surface is equal
It is parallel, and each fin and by the angle between gas flows between two surfaces direction between 30 ° to 70 °.
Angle between described two apparent surface's fins is less than or equal to 90 °.
Hydraulic diameter of the upper fin cross section of described two apparent surfaces and the scope of the ratio between rib spacing be 0.25~1,
The scope of the hydraulic diameter of fin cross section and the ratio between rib height is 0.4~1.2, and rib is high the distance between with two apparent surfaces
The ratio between scope be 0.5~1, the fin on one of surface is embedded in corresponding fin on another surface.
Cooling air from compressor is introduced via first passage is entered, and is blown into air cavity by impact jet flow hole.Right
Blade inlet edge is implemented after impinging cooling, and this fraction is allocated in air cavity, and a part flows to the spray at blade inlet edge
Hole is drenched, it is remaining, spread respectively to both sides, wash away the heat exchange structure of air cavity both sides.By air cavity both sides heat exchange structure it is cold
But air-flow is respectively from the first special-shaped air film hole blowout on the second special-shaped air film hole and suction surface on pressure face, to downstream wall
Carry out gaseous film control.It can be seen that, at the heat exchange structure of air cavity both sides the distance between two apparent surfaces along cold air flow direction by
Decrescence small, this causes the rise of cooling gas flow velocity to reach the purpose of enhanced heat exchange.
Fig. 3 is the sectional stretch-out view along Fig. 2 sight a-a, as can be seen from the figure：The cooling sprayed from impact jet flow hole
Air shuttles up and down between the heat exchange structure of air cavity both sides, is finally sprayed from first, second special-shaped air film hole, real to downstream wall
Apply gaseous film control.It should be noted that the special-shaped air film hole shown in accompanying drawing is only used as representative using slit air film hole.
As shown in Figure 4, Figure 5, The present invention gives the structural representation of two kinds of heat exchange structures, the fin section shown in figure
For square-section.
First provides cooling air-flow to blade inlet edge, blade middle part, blade trailing edge near zone respectively to third channel.The
Two passages, third channel correspond to the 3rd special-shaped air film hole, the 4th special-shaped air film hole respectively, and cooling gas is respectively from the 3rd special-shaped gas
Fenestra, the 4th special-shaped air film hole blowout implement gaseous film control to downstream wall.The connection of second channel and the 3rd special-shaped air film hole
Heat exchange structure is equipped with passage, in the interface channel of third channel and the 4th special-shaped air film hole.
Interface channel between third channel and the 4th special-shaped air film hole can be designed to pressure face and suction surface it
Between, extend to the gallery of blade trailing edge, its pressure face and suction surface to the section is carried out internal cooling, then from the
Four air film holes are sprayed, and trailing edge region is formed and protected.
The efficient cooling knot of the present invention combined by the special-shaped air film hole of latticed internal cooling channel and downstream
Structure, can be used for the thermal protection to turbine engine bucket piece, it can also be used to stator blade or any other need the heat of thermal protection
End pieces, with the service life for improving engine efficiency and extending hot-end component.
Those skilled in the art of the present technique are it is understood that unless otherwise defined, all terms used herein（Including skill
Art term and scientific terminology）With the general understanding identical meaning with the those of ordinary skill in art of the present invention.Also
It should be understood that those terms defined in such as general dictionary should be understood that with the context of prior art
The consistent meaning of meaning, and unless defined as here, will not be explained with idealization or excessively formal implication.
Above-described embodiment, has been carried out further to the purpose of the present invention, technical scheme and beneficial effect
Describe in detail, should be understood that the embodiment that the foregoing is only the present invention, be not limited to this hair
Bright, within the spirit and principles of the invention, any modification, equivalent substitution and improvements done etc. should be included in the present invention
Protection domain within.
Claims (5)
1. the efficient cooling blade of a kind of gas-turbine unit, it is characterised in that include the blade, end wall and tenon being sequentially connected
Head, described LJ is used for and turbine disk connection；
The blade is in hollow cylinder, includes pressure face and suction surface；
The pressure face and suction surface be in cambered surface, wherein, suction surface is arranged on outside pressure face, and suction surface both sides respectively and
The both sides connection of pressure face, forms blade inlet edge and blade trailing edge；
In the blade the first to the 3rd bridge joint beam, and the described first to the 3rd bridge are sequentially provided with from blade inlet edge to blade trailing edge
Beam is connect to set along blade direction of extension；
The first bridge joint beam is set in cambered surface, along the blade inlet edge, and long and narrow air cavity is formed between blade inlet edge；
Space beyond air cavity in blade is divided into first passage, second channel and threeway by the described first to the 3rd bridge joint beam
Road；
Be sequentially provided with the end wall, LJ along blade direction of extension lead to first passage, second channel, third channel it is logical
Hole；
The suction surface is in the first special-shaped air film hole being provided with blade inlet edge and the air cavity is communicated；
The pressure face is at close blade inlet edge provided with the second special-shaped air film hole communicated with the air cavity, close to blade tail
Be provided with edge and the 4th special-shaped air film hole that the third channel is communicated, the second special-shaped air film hole and the 4th special-shaped air film hole it
Between be provided with and the 3rd special-shaped air film hole that the second channel is communicated；
The impact jet flow hole that some and described air cavity is communicated is uniformly provided with the first bridge joint beam；
The spray apertures that some and described air cavity is communicated are uniformly provided with the blade inlet edge；
In the air cavity, in the interface channel of second channel and the 3rd special-shaped air film hole, third channel and the 4th special-shaped air film hole
Interface channel in be equipped with heat exchange structure.
2. the efficient cooling blade of gas-turbine unit according to claim 1, it is characterised in that on the suction surface
It is additionally provided with the 6th special-shaped air film hole that the 5th special-shaped air film hole and third channel communicated with the second channel is communicated, and institute
The junction for stating junction, third channel and the 6th special-shaped air film hole of the second channel and the 5th special-shaped air film hole is equipped with
Heat exchange structure.
3. the efficient cooling blade of gas-turbine unit according to claim 1 or 2, it is characterised in that the heat exchange
Structure includes some fins being uniformly arranged on two apparent surfaces, and the fin on same surface is parallel, and each
Angle between the direction that gas flows between two surfaces of fin and process is between 30 ° to 70 °.
4. the efficient cooling blade of gas-turbine unit according to claim 3, it is characterised in that described two relative
Angle between the fin of surface is less than or equal to 90 °.
5. the efficient cooling blade of gas-turbine unit according to claim 4, it is characterised in that described two relative
The scope of the ratio between the hydraulic diameter of fin cross section and rib spacing is 0.25~1, hydraulic diameter of fin cross section on surface
Scope with the ratio between rib height be the high scope the ratio between the distance between with two apparent surfaces of 0.4~1.2, rib be 0.5~
1, the fin on one of surface is embedded in corresponding fin on another surface.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710356212.7A CN107013253A (en) | 2017-05-19 | 2017-05-19 | A kind of efficient cooling blade of gas-turbine unit |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710356212.7A CN107013253A (en) | 2017-05-19 | 2017-05-19 | A kind of efficient cooling blade of gas-turbine unit |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN107013253A true CN107013253A (en) | 2017-08-04 |
Family
ID=59449834
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710356212.7A Pending CN107013253A (en) | 2017-05-19 | 2017-05-19 | A kind of efficient cooling blade of gas-turbine unit |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN107013253A (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109736899A (en) * | 2019-01-13 | 2019-05-10 | 中国航发四川燃气涡轮研究院 | A kind of turbine blade tail with microchannel partly splits seam cooling structure |
| CN110617114A (en) * | 2019-09-02 | 2019-12-27 | 上海大学 | Ceramic-coated high-temperature alloy stator blade |
| CN112855285A (en) * | 2019-11-28 | 2021-05-28 | 中国航发商用航空发动机有限责任公司 | Turbine blade and aircraft engine |
| CN113236372A (en) * | 2021-06-07 | 2021-08-10 | 南京航空航天大学 | Gas turbine guide vane blade with jet oscillator and working method |
| CN114810216A (en) * | 2021-01-27 | 2022-07-29 | 中国航发商用航空发动机有限责任公司 | Aeroengine blade and aeroengine |
| CN115875084A (en) * | 2023-03-02 | 2023-03-31 | 中国航发四川燃气涡轮研究院 | Laminate cooling structure applied to pressure surface of turbine blade |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4859147A (en) * | 1988-01-25 | 1989-08-22 | United Technologies Corporation | Cooled gas turbine blade |
| US8585365B1 (en) * | 2010-04-13 | 2013-11-19 | Florida Turbine Technologies, Inc. | Turbine blade with triple pass serpentine cooling |
| CN103806951A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Turbine blade combining cooling seam gas films with turbulence columns |
| CN207526530U (en) * | 2017-05-19 | 2018-06-22 | 南京航空航天大学 | A kind of efficient cooling blade of gas-turbine unit |
-
2017
- 2017-05-19 CN CN201710356212.7A patent/CN107013253A/en active Pending
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4859147A (en) * | 1988-01-25 | 1989-08-22 | United Technologies Corporation | Cooled gas turbine blade |
| US8585365B1 (en) * | 2010-04-13 | 2013-11-19 | Florida Turbine Technologies, Inc. | Turbine blade with triple pass serpentine cooling |
| CN103806951A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Turbine blade combining cooling seam gas films with turbulence columns |
| CN207526530U (en) * | 2017-05-19 | 2018-06-22 | 南京航空航天大学 | A kind of efficient cooling blade of gas-turbine unit |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109736899A (en) * | 2019-01-13 | 2019-05-10 | 中国航发四川燃气涡轮研究院 | A kind of turbine blade tail with microchannel partly splits seam cooling structure |
| CN110617114A (en) * | 2019-09-02 | 2019-12-27 | 上海大学 | Ceramic-coated high-temperature alloy stator blade |
| CN110617114B (en) * | 2019-09-02 | 2021-12-03 | 上海大学 | Ceramic-coated high-temperature alloy stator blade |
| CN112855285A (en) * | 2019-11-28 | 2021-05-28 | 中国航发商用航空发动机有限责任公司 | Turbine blade and aircraft engine |
| CN112855285B (en) * | 2019-11-28 | 2023-03-24 | 中国航发商用航空发动机有限责任公司 | Turbine blade and aircraft engine |
| CN114810216A (en) * | 2021-01-27 | 2022-07-29 | 中国航发商用航空发动机有限责任公司 | Aeroengine blade and aeroengine |
| CN113236372A (en) * | 2021-06-07 | 2021-08-10 | 南京航空航天大学 | Gas turbine guide vane blade with jet oscillator and working method |
| CN113236372B (en) * | 2021-06-07 | 2022-06-10 | 南京航空航天大学 | Gas turbine guide vane blade with jet oscillator and working method |
| CN115875084A (en) * | 2023-03-02 | 2023-03-31 | 中国航发四川燃气涡轮研究院 | Laminate cooling structure applied to pressure surface of turbine blade |
| CN115875084B (en) * | 2023-03-02 | 2023-06-30 | 中国航发四川燃气涡轮研究院 | Laminate cooling structure applied to turbine blade pressure surface |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN107013253A (en) | A kind of efficient cooling blade of gas-turbine unit | |
| CN207526530U (en) | A kind of efficient cooling blade of gas-turbine unit | |
| US10513932B2 (en) | Cooling pedestal array | |
| US7887294B1 (en) | Turbine airfoil with continuous curved diffusion film holes | |
| US4515526A (en) | Coolable airfoil for a rotary machine | |
| JP3367697B2 (en) | Blades for turbines | |
| US9039371B2 (en) | Trailing edge cooling using angled impingement on surface enhanced with cast chevron arrangements | |
| US20130115103A1 (en) | Film hole trench | |
| CN108884716B (en) | Turbine airfoil with internal cooling passage having flow splitter feature | |
| US20090304494A1 (en) | Counter-vortex paired film cooling hole design | |
| JPH10274002A (en) | Turbulence unit structure of cooling passage of moving blade for gas turbine engine | |
| IT8224878A1 (en) | COOLING STRUCTURE FOR AERODYNAMIC ELEMENTS OF ROTARY MACHINES | |
| JP2011247261A (en) | System and method for enhanced turbine wake mixing via fluidic-generated vortices | |
| JP2015083916A (en) | Transition duct assembly with modified trailing edge in turbine system | |
| CN103452595A (en) | Novel air film hole with improved cooling efficiency | |
| US20140116666A1 (en) | Film Cooling Channel Array Having Anti-Vortex Properties | |
| CN108442985A (en) | It is a kind of that there is the line of rabbet joint cooling structure for improving stator blade passage end wall cooling efficiency | |
| CN109931114A (en) | A kind of novel impinging cooling turbulence structure | |
| CN108757047A (en) | Turbine blade of gas turbine with cooling structure inside the droplet-shaped rib of column | |
| CN107143382A (en) | The CMC article with small complex characteristic cooled down for advanced film | |
| CN111706409B (en) | Corrugated air film hole with branch hole | |
| CN112780354B (en) | Tail edge crack-splitting cooling structure and method suitable for turbine blade and turbine blade | |
| JPS5951104A (en) | Internal structure of turbine stage | |
| CN106014487A (en) | Jet flow impact control structure with confined space internally provided with cross flow | |
| CN115898957A (en) | Compressor rotor cooling structure suitable for high-Mach-number incoming flow working condition |
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 | ||
| WD01 | Invention patent application deemed withdrawn after publication | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20170804 |