CN106703899B - High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it - Google Patents
High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it Download PDFInfo
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- CN106703899B CN106703899B CN201710048920.4A CN201710048920A CN106703899B CN 106703899 B CN106703899 B CN 106703899B CN 201710048920 A CN201710048920 A CN 201710048920A CN 106703899 B CN106703899 B CN 106703899B
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- Prior art keywords
- impact
- chamber
- inlet edge
- high pressure
- blade
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Classifications
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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
- 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
Abstract
The invention discloses a kind of High Pressure Turbine Rotor blade inlet edge impinging cooling structure and with its engine, it is related to technical field of engines.The blade inlet edge (1) of the High Pressure Turbine Rotor blade inlet edge impinging cooling structure is provided with impact chamber, and the impact chamber is divided into multiple mutually independent unit impacts chamber (2) in the short transverse of blade.Rotor blade on the engine includes High Pressure Turbine Rotor blade inlet edge impinging cooling structure as described above.The present invention has the advantages that original integrated impact chamber is divided into multiple units by High Pressure Turbine Rotor blade inlet edge impinging cooling structure of the invention impacts chamber, individual impact opening and air film hole are provided on each unit impact chamber, realize the independent control of blade inlet edge different zones cold air, cold air utilization rate is improved, the performance for improving engine is conducive to；Blade safety in utilization is improved simultaneously, reduces the risk of blade global failure.
Description
Technical field
The present invention relates to technical field of engines, and in particular to a kind of High Pressure Turbine Rotor blade inlet edge impinging cooling structure
And the engine with it.
Background technique
High Pressure Turbine Rotor blade work long hours high temperature, high pressure environment in, bad environments, load is big, especially
Blade inlet edge position is the entire highest region of blade thermic load.Independent impact would generally be designed by edge in front of the blade when design
Chamber carries out individual impinging cooling to blade inlet edge by the high-speed jet of impact opening, and then cold air in leading edge by being arranged
Air film hole flows to blade surface and forms air film protection to blade.Even if using individual impact aerating film compound cooling structure,
Blade inlet edge position is still the highest region of blade thermic load, once ablation, which occurs, forms notch, a large amount of cold air will be from notch
Position leakage upsets the cold air distribution of entire blade, causes blade global failure.
In prior art designs, blade inlet edge impact chamber is typically designed as an entirety.From cooling technology angle analysis, this is set
It is insufficient to there is both sides in meter.First is that the unfavorable fining adjustment with cold air when being a whole chamber of impact chamber.Due to combustion gas temperature
Degree has temperature gradient along the high direction of leaf, and the intensity of cooling that each section needs is different, and the high region of temperature needs cold air more, temperature
Cold air is suitably reduced in low region, and whole chamber cannot achieve the independent control of cold air.Second is that once ablation shape occurs for blade inlet edge
At notch, a large amount of cold air will be revealed from gap position, be upset the cold air distribution of entire blade, be caused blade global failure.
Summary of the invention
Start the object of the present invention is to provide a kind of High Pressure Turbine Rotor blade inlet edge impinging cooling structure and with it
Machine realizes the independent control of blade inlet edge different zones cold air, improves cold air utilization rate, improves blade safety in utilization, reduces
Blade global failure risk, with solve the problems, such as or at least mitigate background technique in the presence of at least one at.
The technical solution adopted by the present invention is that: a kind of High Pressure Turbine Rotor blade inlet edge impinging cooling structure, blade are provided
Leading edge is provided with impact chamber, and the impact chamber is divided into multiple mutually independent unit impact chambers in the short transverse of blade.
Preferably, the side wall of the unit impact chamber includes close to the first side wall in trailing edge direction and far from trailing edge direction
Second sidewall is provided with impact opening on the first side wall, and the axis of the impact opening is perpendicular to place the first side wall；Described
Air film hole is provided on two side walls, the axis of the air film hole forms a folder in the short transverse and place second sidewall of blade
Angle.
Preferably, multiple impact openings, height of multiple impact openings in blade are provided on each unit impact chamber
Degree direction is set gradually；Multiple air film holes, height of multiple air film holes in blade are provided on each unit impact chamber
Degree direction is set gradually.
Preferably, each unit impact chamber is provided with multiple air film holes in the same section in blade height direction,
The axis shape of any 2 air film holes has angle in the multiple air film hole, and any two air film hole is impacting chamber far from unit
One end distance be greater than close to unit impact chamber one end distance.
Preferably, the unit impact chamber is set as unidirectional contracted section, institute in the section perpendicular to blade height direction
State unidirectional contracted section to refer to: second side where first side to the air film hole where the section from the impact opening is gradually
It shrinks.
Preferably, the impact chamber impacts chamber by being divided into multiple units every rib, described every rib and the impact chamber
Side wall is integrally formed.
Preferably, the impact chamber impacts chamber by being divided into multiple units every rib, described removably to pacify every rib
It is intracavitary mounted in the impact.
Preferably, be provided on the inner sidewall of the impact chamber it is multiple every rib mounting portion, it is multiple described to exist every rib mounting portion
The short transverse of blade is uniformly arranged.
Preferably, the impact opening is set as bellmouth, and the small end opening of the bellmouth is towards the inside of blade inlet edge
Wall.
The present invention also provides a kind of engine, the rotor blade on the engine includes high-pressure turbine as described above
Rotor blade leading edge impinging cooling structure.
The beneficial effects of the present invention are: High Pressure Turbine Rotor blade inlet edge impinging cooling structure of the invention will be original
Integrated impact chamber is divided into multiple unit impact chambers, is provided with individual impact opening and air film on each unit impact chamber
Hole realizes the independent control of blade inlet edge different zones cold air, improves cold air utilization rate, is conducive to the property for improving engine
Energy；Blade safety in utilization is improved simultaneously, reduces the risk of blade global failure.
Detailed description of the invention
Fig. 1 is the schematic diagram of the High Pressure Turbine Rotor blade inlet edge impinging cooling structure of one embodiment of the invention.
Fig. 2 is that High Pressure Turbine Rotor blade inlet edge impinging cooling structure shown in FIG. 1 is shown in the section in blade height direction
It is intended to.
Wherein, 1- blade inlet edge, 2- unit impact chamber, 21- the first side wall, 22- second sidewall, 23- impact opening, 24- gas
Fenestra, 3- is every rib.
Specific embodiment
To keep the purposes, technical schemes and advantages of the invention implemented clearer, below in conjunction in the embodiment of the present invention
Attached drawing, technical solution in the embodiment of the present invention is further described in more detail.In the accompanying drawings, identical from beginning to end or class
As label indicate same or similar element or element with the same or similar functions.Described embodiment is the present invention
A part of the embodiment, instead of all the embodiments.The embodiments described below with reference to the accompanying drawings are exemplary, it is intended to use
It is of the invention in explaining, and be not considered as limiting the invention.Based on the embodiments of the present invention, ordinary skill people
Member's every other embodiment obtained without creative efforts, shall fall within the protection scope of the present invention.Under
Face is described in detail the embodiment of the present invention in conjunction with attached drawing.
In the description of the present invention, it is to be understood that, term " center ", " longitudinal direction ", " transverse direction ", "front", "rear",
The orientation or positional relationship of the instructions such as "left", "right", "vertical", "horizontal", "top", "bottom" "inner", "outside" is based on attached drawing institute
The orientation or positional relationship shown, is merely for convenience of description of the present invention and simplification of the description, rather than the dress of indication or suggestion meaning
It sets or element must have a particular orientation, be constructed and operated in a specific orientation, therefore should not be understood as protecting the present invention
The limitation of range.
As shown in Figure 1 and Figure 2, a kind of High Pressure Turbine Rotor blade inlet edge impinging cooling structure, blade inlet edge 1 are provided with punching
Chamber is hit, the impact chamber is divided into multiple mutually independent unit impact chambers 2 in the short transverse of blade.In the present embodiment
In, the impact chamber is separated into 3 independent unit impact chambers, integrated impact chamber compared to the prior art, the present invention
Integrated impact chamber is divided into 3 unit impact chambers 2, each unit impacts 2 independent gas supply of chamber.By leading edge in practical engineering application
Originally independent one-piece type core is divided into three sections of multi-cavity structures that can be realized in this example, and technology difficulty is low.Before realizing blade
The independent control of 1 different zones cold air of edge, improves cold air utilization rate, is conducive to the performance for improving engine；Realizing multi-cavity
Also the application risk of blade is directly reduced while independent gas supply.Blade generation ablation is formed scarce under the original structure of the prior art
When mouth, the cold air of entire ante-chamber can be largely lost from gap position, caused other region cold air of leading edge insufficient, finally sent out blade
Raw global failure；Integration impact chamber is divided into multiple independent units and impacts chamber 2 by the present invention, when blade inlet edge chamber is burnt
When erosion, the leak area of cold air is limited to the individual unit impact corresponding position of chamber 2.Maximum cold air amount of leakage depends on unit
The area of chamber 2 is impacted, this reduces influence of the notch to other units impact chamber cool-air feed to a certain extent, improves leaf
Piece safety in utilization reduces the risk of blade global failure.
In the present embodiment, the side wall of unit impact chamber 2 includes close to the first side wall 21 in trailing edge direction and far from trailing edge
The second sidewall 22 in direction is provided with impact opening 23 on the first side wall 21, and the axis of impact opening 23 is perpendicular to place the first side wall
21；It is provided with air film hole 24 in second sidewall 22, the short transverse and place second sidewall 22 of the axis of air film hole 24 in blade
Shape has angle.Cold air can directly be impacted on the inner surface of second sidewall 22 by impact opening 23, and cold air enters unit punching
It after hitting 2 inside of chamber, is impacted inside chamber 2 and is flowed out from unit by air film hole 24, flowed to blade surface and air film protection is formed to blade.
Individual impact opening 23 and air film hole 24, the more temperature of blade in the height direction are provided on each unit impact chamber 2
Gradient, can individually control unit impact chamber 2 on impact opening 23 cold air flow, realize blade different zones fining control
System, improves the utilization rate of cold air, preferably controls the surface temperature of blade, improves the performance of engine.
In the present embodiment, multiple impact openings 23 are provided on each unit impact chamber 2, multiple impact openings 23 are in blade
Short transverse is set gradually；Multiple air film holes 24, height of multiple air film holes 24 in blade are provided on each unit impact chamber 2
Direction is set gradually.It is understood that the diameter of multiple impact openings 23 can be set according to actual needs, multiple impacts
The pore size in hole 23 may be the same or different；Equally, the pore size of multiple air film holes 24 can be identical, can also be with
It is different.Impact opening 23 and air film hole 24 also may be set according to actual conditions in the relative position in blade height direction not in this reality
It applies in example, the axis of impact opening 23 is fallen between two air film holes 24, the advantage is that, cold air is rushed through 23 air admission unit of impact opening
It after hitting chamber 2, will not directly be flowed out from air film hole 24, and air-flow flowing can be formed in the inner surface that unit impacts chamber 2, be conducive to tie up
The gas flow temperature in unit impact chamber 2 is held, dead space is avoided the formation of.
In the present embodiment, each unit impact chamber 2 is provided with multiple air films in the same section in blade height direction
Hole 24, the axis shape of any 2 air film holes has angle in multiple air film holes, and any two air film hole is impacting chamber far from unit
The distance of 2 one end is greater than the distance close to one end of unit impact chamber 2.It the advantage is that, be conducive to the diffusion of cold air, in leaf
The up-front outer surface of piece forms larger range of cold air film.
In the present embodiment, unit impact chamber 2 is set as unidirectional contracted section in the section perpendicular to blade height direction,
The unidirectional contracted section refers to: gradually receiving from 23 place first side of impact opening to 24 place second side of air film hole in the section
Contracting, the advantage is that, is conducive to the flow velocity for accelerating cold air, is formed and impacted to second side, cooling effect is more preferable.
In the present embodiment, the impact chamber impacts chamber 2 by being divided into multiple units every rib 3, every rib 3 and impacts chamber
Side wall is integrally formed, and structure is simple, easy to process.It is understood that the structure every rib 3 and impact chamber can also be according to reality
Situation setting.For example, in an alternative embodiment, impact chamber impacts chambers 2 by being divided into multiple units every rib 3, every rib 3 with
It is intracavitary to be removably mounted impact, the advantage is that, the replacement every rib and the control of position may be implemented, realizes unit punching
Hit the control of the size of chamber 2；It is understood that can be set on the inner sidewall of the impact chamber in above-mentioned alternative embodiment
Be equipped with it is multiple every rib mounting portion, it is multiple it is described be uniformly arranged every rib mounting portion in the short transverse of blade, the advantage is that, pass through
Be selected in it is different may be mounted at different positions every rib every rib mounting portion, thus realize unit impact 2 size of chamber adjusting.
In the present embodiment, impact opening 23 is set as cylindrical hole, it is to be understood that the impact opening 23 can also be set
It is set to bellmouth, and the small end opening of bellmouth the advantage is that towards the inner sidewall of blade inlet edge, cold airflow can be improved
Speed improves cooling effect.
The present invention also provides a kind of engine, the rotor blade on the engine includes high-pressure turbine as described above
Rotor blade leading edge impinging cooling structure.The surface temperature of blade inlet edge can be effectively reduced in the structure, according in blade height
The different zones in direction realize Precise control, reduce the loss of cold airflow, improve cooling efficiency, before reducing blade
The risk of global failure occurs for chamber, improves the safety of blade, and then improve the performance of engine.
Finally it is noted that the above embodiments are merely illustrative of the technical solutions of the present invention, rather than its limitations.To the greatest extent
Present invention has been described in detail with reference to the aforementioned embodiments for pipe, those skilled in the art should understand that: it is still
It is possible to modify the technical solutions described in the foregoing embodiments, or part of technical characteristic is equally replaced
It changes；And these are modified or replaceed, the essence for technical solution of various embodiments of the present invention that it does not separate the essence of the corresponding technical solution
Mind and range.
Claims (9)
1. High Pressure Turbine Rotor blade inlet edge impinging cooling structure, blade inlet edge (1) are provided with impact chamber, it is characterised in that:
The impact chamber is divided into multiple mutually independent unit impacts chamber (2), the unit punching in the short transverse of blade
The side wall for hitting chamber (2) includes the second sidewall (22) of the first side wall (21) and separate trailing edge direction close to trailing edge direction, described
It is provided on the first side wall (21) impact opening (23), the axis of the impact opening (23) is perpendicular to place the first side wall (21)；Institute
It states and is provided on second sidewall (22) air film hole (24), the axis of the air film hole (24) is at the short transverse of blade and place the
Two side walls (22) shape has angle.
2. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as described in claim 1, it is characterised in that: each list
It is provided with multiple impact openings (23) in member impact chamber (2), multiple impact openings (23) set gradually in the short transverse of blade；
It is provided with multiple air film holes (24) on each unit impact chamber (2), multiple air film holes (24) are in the height side of blade
To setting gradually.
3. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as claimed in claim 2, it is characterised in that: each list
Member impact chamber (2) is provided with multiple air film holes (24) in the same section in blade height direction, the multiple air film hole (24)
In the axis shapes of any 2 air film holes have angle, any two air film hole (24) is in one end far from unit impact chamber (2)
Distance is greater than the distance close to one end of unit impact chamber (2).
4. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as claimed in claim 3, which is characterized in that the unit punching
It hits chamber (2) and is set as unidirectional contracted section in the section perpendicular to blade height direction, the unidirectional contracted section refers to: this section
Second side where first side to the air film hole where face from the impact opening gradually tapers up.
5. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as claimed in claim 4, it is characterised in that: the impact chamber
It is described to be integrally formed every rib (3) with the side wall of the impact chamber by being divided into multiple unit impacts chamber (2) every rib (3).
6. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as claimed in claim 4, it is characterised in that: the impact chamber
Chambers are impacted by being divided into multiple units every rib, and described every rib to be removably mounted in the impact intracavitary.
7. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as claimed in claim 6, it is characterised in that: the impact chamber
Inner sidewall on be provided with it is multiple every rib mounting portion, it is multiple described to be uniformly arranged every rib mounting portion in the short transverse of blade.
8. High Pressure Turbine Rotor blade inlet edge impinging cooling structure as described in claim 5 or 7, it is characterised in that: the punching
It hits hole and is set as bellmouth, the inner sidewall of the small end opening of the bellmouth towards blade inlet edge.
9. a kind of engine, it is characterised in that: the rotor blade on the engine includes high pressure whirlpool as claimed in claim 8
Rotor blades leading edge impinging cooling structure.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710048920.4A CN106703899B (en) | 2017-01-23 | 2017-01-23 | High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201710048920.4A CN106703899B (en) | 2017-01-23 | 2017-01-23 | High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it |
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| Publication Number | Publication Date |
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| CN106703899A CN106703899A (en) | 2017-05-24 |
| CN106703899B true CN106703899B (en) | 2019-08-23 |
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| CN201710048920.4A Active CN106703899B (en) | 2017-01-23 | 2017-01-23 | High Pressure Turbine Rotor blade inlet edge impinging cooling structure and the engine with it |
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Families Citing this family (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110593961B (en) * | 2019-09-29 | 2020-09-15 | 华北电力大学 | Divided cabin type turbine blade |
| CN113513371A (en) * | 2021-08-19 | 2021-10-19 | 北京全四维动力科技有限公司 | Double-wall cooling blade, turbine blade using same and gas turbine |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0887515A1 (en) * | 1997-06-26 | 1998-12-30 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" | Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall |
| JP2005069236A (en) * | 2004-12-10 | 2005-03-17 | Toshiba Corp | Turbine cooling blade |
| JP2015067902A (en) * | 2013-09-26 | 2015-04-13 | ゼネラル・エレクトリック・カンパニイ | Manufacturing method and heat management method of component |
| CN106401667A (en) * | 2015-07-29 | 2017-02-15 | 通用电气公司 | Article, airfoil component and method for forming article |
-
2017
- 2017-01-23 CN CN201710048920.4A patent/CN106703899B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0887515A1 (en) * | 1997-06-26 | 1998-12-30 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" | Blading with a helical ramp having a serial impingement cooling through a system of ribs in a double shell wall |
| JP2005069236A (en) * | 2004-12-10 | 2005-03-17 | Toshiba Corp | Turbine cooling blade |
| JP2015067902A (en) * | 2013-09-26 | 2015-04-13 | ゼネラル・エレクトリック・カンパニイ | Manufacturing method and heat management method of component |
| CN106401667A (en) * | 2015-07-29 | 2017-02-15 | 通用电气公司 | Article, airfoil component and method for forming article |
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| CN106703899A (en) | 2017-05-24 |
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