CN111120008A - Novel turbine blade rotational flow cooling structure - Google Patents
Novel turbine blade rotational flow cooling structure Download PDFInfo
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- CN111120008A CN111120008A CN201911259529.4A CN201911259529A CN111120008A CN 111120008 A CN111120008 A CN 111120008A CN 201911259529 A CN201911259529 A CN 201911259529A CN 111120008 A CN111120008 A CN 111120008A
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- Prior art keywords
- target surface
- jet
- holes
- rotational flow
- cooling
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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
- F01D5/187—Convection cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- 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
- F05D2260/2212—Improvement of heat transfer by creating turbulence
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a novel turbine blade rotational flow cooling structure, which comprises a rotational flow cooling cavity, a jet hole and the like; the top of the rotational flow cooling cavity is of a plane structure, and the target surface is of a curved surface structure; the jet holes are arranged on one side of the top surface of the cyclone cooling cavity at equal intervals, the outflow grooves are arranged on the other side of the top surface of the cyclone cooling cavity opposite to the jet holes at equal intervals, and the outflow grooves and the jet holes are arranged in a staggered manner; the ball socket structures and the air film holes are arranged on the target surface of the cyclone cooling cavity; when the device works, a cooling working medium is injected from the jet hole to enter the rotational flow cooling cavity, flows along one side of a target surface where the jet hole is located, forms a transverse large-scale vortex under the action of the bent target surface, carries out strong heat exchange with the target surface, generates flow separation and reattachment when flowing through the ball socket structure, enhances the local heat transfer of the ball socket, obtains higher fluid turbulence energy under the suction action of the air film hole, eliminates the local heat transfer deterioration phenomenon, and discharges the cooling working medium after heat exchange through the outflow groove adjacent to the jet hole.
Description
Technical Field
The invention belongs to the technical field of turbine blade cooling, and particularly relates to a novel turbine blade rotational flow cooling structure.
Background
As one of the core components of the gas turbine, the turbine is close to the downstream of the combustion chamber and is continuously flushed by high-temperature and high-pressure gas in the operation process, the working environment is extremely severe, and when the turbine is designed, besides the consideration of pneumatic factors, an efficient cooling technology is also needed to reduce the temperature of the turbine to be within a material allowable range so as to ensure the safe and stable operation of the gas turbine. The high temperature combustion gases output from the combustor directly impinge on the turbine blade leading edge as they enter the cascade channels, and therefore this region, particularly the first stage guide vane leading edge region, has an extremely high thermal load and is a significant concern in blade cooling design.
In order to improve the cooling performance of the leading edge of the turbine blade, an impingement cooling mode with the highest heat transfer enhancement effect is often adopted, a plurality of jet holes are formed in the surface of an inner liner plate of a leading edge area, cooling air supplied from a blade root impacts a curved target surface of the leading edge through the jet holes, strong heat and mass exchange is generated in the area near an impact point, and the cooling air absorbs heat from the target surface, so that the temperature of the inner surface and the outer surface of the leading edge is effectively reduced.
With the development of gas turbines, according to the intrinsic law of the brayton cycle, the increase in power and efficiency requires the turbine inlet temperature to rise continuously, which brings great challenges to the cooling design of the turbine components, especially the cooling design of the leading edge of the first stage guide vane of the turbine. Because the traditional impingement cooling can not deal with the rapid rise of the thermal load of the front edge of the blade under many conditions, related experts and designers propose a rotational flow cooling structure, namely, jet holes are arranged on one side area of the top surface, and a transverse large-scale vortex system is formed along an arc target surface after a cooling working medium enters a channel, so that the turbulent kinetic energy of the fluid is enhanced, and the cooling performance obviously superior to that of the traditional impingement structure is obtained.
However, the adoption of the rotational flow cooling structure can cause local heat transfer deterioration among the jet holes, aggravate the nonuniformity of the temperature distribution of the front edge of the blade, and meanwhile, the cross flow formed by the aggregation of the upstream rotational flow working medium can obviously weaken the enhanced heat transfer effect of the downstream jet flow, so that the potential cooling performance of the rotational flow structure cannot be fully exerted. Therefore, there is a need for a new turbine blade cyclone cooling structure that achieves better cyclone cooling performance by overcoming the above problems, thereby achieving efficient thermal protection of the leading edge of the turbine blade.
Disclosure of Invention
In order to solve the problems, the invention provides a novel turbine blade rotational flow cooling structure, which eliminates the influence of cross flow through a distributed outflow groove, avoids the rapid attenuation of rotational flow heat transfer performance in the flow direction, improves the overall heat transfer level and reduces the pressure loss; the ball socket and the air film hole combined structure arranged on the target surface strengthens local heat transfer, obviously reduces the heat transfer deterioration area of the traditional rotational flow target surface, improves the heat transfer performance and simultaneously reduces the temperature difference of the target surface, thereby improving the overall cooling performance.
The invention is realized by adopting the following technical scheme:
a novel turbine blade rotational flow cooling structure comprises a rotational flow cooling cavity, a jet hole, a flow outlet groove, a ball socket structure and a gas film hole; the top of the rotational flow cooling cavity is of a plane structure, and the target surface is of a curved surface structure; the jet holes are arranged on one side of the top surface of the cyclone cooling cavity at equal intervals, the outflow grooves are arranged on the other side of the top surface of the cyclone cooling cavity opposite to the jet holes at equal intervals, and the outflow grooves and the jet holes are arranged in a staggered manner; the ball socket structures and the air film holes are arranged on the target surface of the cyclone cooling cavity;
when the device works, a cooling working medium is injected from the jet hole to enter the rotational flow cooling cavity, flows along one side of a target surface where the jet hole is located, forms a transverse large-scale vortex under the action of the bent target surface, carries out strong heat exchange with the target surface, generates flow separation and reattachment when flowing through the ball socket structure, enhances the local heat transfer of the ball socket, obtains higher fluid turbulence energy under the suction action of the air film hole, eliminates the local heat transfer deterioration phenomenon, and discharges the cooling working medium after heat exchange through the outflow groove adjacent to the jet hole.
The invention is further improved in that the target surface of the rotational flow cooling cavity is in a shape of a circular arc, an elliptic arc, a parabola or a hyperbolic curve.
The invention has the further improvement that the jet holes are rectangular, straight notch-shaped, circular and oval, the S/w ratio of the arrangement space of the jet holes to the width of the jet holes is within the range of 2-5, and the w/t ratio of the width to the thickness of the jet holes is within the range of 1-6.
The invention is further improved in that the outflow groove is arranged on the other side of the area between the adjacent jet holes, the shape of the outflow groove is rectangular, straight notch-shaped, circular or elliptical, and the structural size and the arrangement mode of the outflow groove are the same as those of the jet holes.
The invention is further improved in that a ball-and-socket structure is arranged in the target surface area between adjacent jet holes, and the depth-to-diameter ratio e/D of the ball-and-socket structure (4)dIn the range of 0.1-0.3, the target surface is uniformly arranged in the circumferential direction, and the number of the circumferential arrangement is 1-6.
The invention is further improved in that the air film holes are arranged in the target surface area between the adjacent jet holes, the cross section of the air film holes is circular or elliptical, the arrangement angle α of the air film holes is changed within the range of 0-180 degrees, and the circumferential arrangement number of the air film holes is 1-4.
The invention is further improved in that the air film hole is arranged on the surface of the ball-and-socket structure or in the target surface area near the ball-and-socket structure.
The invention has at least the following beneficial technical effects:
according to the turbine blade rotational flow cooling structure provided by the invention, the plurality of distributed outflow grooves are introduced, the cooling working medium injected into the channel by the jet holes can be discharged through the outflow grooves after heat exchange is completed, and the cross flow which commonly exists in the traditional rotational flow channel can not be formed, so that the weakening effect of the cross flow on the downstream heat transfer can be eliminated, the impact rotational flow can keep a better enhanced heat transfer effect in the radial direction, and the overall heat transfer performance of a target surface is obviously improved.
Furthermore, the flexibly variable target surface shape can adapt to different blade leading edge molded lines, the height adaptation with the turbine blade leading edge structure is realized, the shapes and the structural sizes of the jet hole and the flow outlet groove can be selected according to the heat load on the leading edge when the turbine operates, and the adaptability and the cooling performance of the novel rotational flow structure are improved to the maximum extent;
furthermore, the ball socket structure has the excellent characteristics of high heat transfer and low flow resistance, only generates small resistance loss while enhancing heat transfer, can increase the heat transfer area, and can realize local flow heat transfer control by arranging the ball socket structure on the target surface of the rotational flow cooling cavity, thereby eliminating heat transfer deterioration areas and realizing more uniform temperature distribution while improving the heat transfer level;
furthermore, the pumping action of the gas film holes can enhance the turbulent kinetic energy of the fluid in the area nearby the gas film holes, destroy the flow boundary layer and promote the heat and mass exchange nearby the wall surface, in addition, the outflow of the gas film holes can also form a protective gas film on the outer surface of the blade, the heat exchange between high-temperature gas and the blade is weakened, and efficient internal and external coupling cooling is realized.
According to the invention, the novel turbine blade rotational flow cooling structure is established, the negative influence of transverse flow in the traditional rotational flow channel on heat transfer is eliminated through the distributed outflow grooves, so that the high heat transfer level is maintained in the whole channel, meanwhile, the arrangement of the ball socket structure and the air film holes realizes local flow control, the heat transfer performance is improved, and more uniform temperature distribution can be obtained, so that the better comprehensive cooling performance is realized.
Drawings
FIG. 1 is a three-dimensional view of a turbine blade cyclonic cooling structure;
FIG. 2 is a schematic view of the arrangement of fluidic orifices and outflow channels on the top surface of a channel;
FIG. 3 is a schematic view of the orifice and the outflow groove shape, wherein FIG. 3(a) is a rectangular orifice/outflow groove, FIG. 3(b) is a straight notch-shaped orifice/outflow groove, FIG. 3(c) is a circular orifice/outflow groove, and FIG. 3(d) is an elliptical orifice/outflow groove;
FIG. 4 is a schematic view of a ball and socket structure and an arrangement of air film holes in a localized area of the target surface;
FIG. 5 is a cross-sectional schematic view of a ball and socket arrangement;
FIG. 6 is a schematic cross-sectional view of the structure and arrangement of the gas film holes.
Description of reference numerals:
1 is the whirl cooling chamber, 2 is the jet orifice, 3 is the play chute, 4 is the ball and socket structure, 5 is the air film hole, 231 is rectangle jet orifice/play chute, 232 is straight notch shape jet orifice/play chute, 233 is circular jet orifice/play chute, 234 is oval jet orifice/play chute.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings.
Referring to fig. 1 and 2, the novel turbine blade rotational flow cooling structure provided by the invention comprises a rotational flow cooling cavity 1, a plurality of jet holes 2, a plurality of outflow grooves 3, a plurality of ball socket structures 4 and a plurality of film holes 5. The top of the cyclone cooling cavity 1 is of a plane structure, and the target surface is of a curved surface structure; the jet holes 2 are arranged on one side of the top surface of the cyclone cooling cavity 1 at equal intervals, the outflow grooves 3 are arranged on the other side of the top surface of the cyclone cooling cavity 1 opposite to the jet holes 2 at equal intervals, and the outflow grooves 3 and the jet holes 2 are arranged in a staggered manner; a plurality of ball and socket structures 4 and air film holes 5 are arranged on the target surface of the cyclone cooling cavity 1. The whole flow heat transfer process comprises the following steps: the cooling working medium is injected from the jet hole 2 to enter the rotational flow cooling cavity 1, flows along one side of a target surface where the jet hole 2 is located, forms a transverse large-scale vortex under the action of a bent target surface, carries out strong heat exchange with the target surface, generates flow separation and reattachment when flowing through the ball socket structure 4, enhances the local heat transfer of the ball socket, obtains higher fluid turbulence kinetic energy under the suction action of the air film hole 5, eliminates the local heat transfer deterioration phenomenon, and discharges the cooling working medium after heat exchange through the outflow groove 3 adjacent to the jet hole 2.
The target surface shape may take the form of a circular arc, elliptical arc, parabolic curve, hyperbolic curve, or other curve that approximates the configuration of the leading edge of the turbine blade.
Referring to fig. 2 and 3, the shape of the jet hole 2 may be rectangular, straight notch, circular, oval, etc., the S/w ratio of the arrangement pitch of the jet holes 2 to the width of the jet hole 2 is preferably in the range of 2 to 5, and the w/t ratio of the width to the thickness of the jet hole 2 is preferably in the range of 1 to 6. The outflow groove 3 is arranged on the other side of the area between the adjacent jet holes 2, the shape of the outflow groove can be rectangular, straight notch-shaped, circular, oval and the like, and the structural size and the arrangement mode of the outflow groove are the same as those of the jet holes 2. Accordingly, the present invention has a rectangular orifice/outflow groove 231, a straight notch-shaped orifice/outflow groove 232, a circular orifice/outflow groove 233, and an oval orifice/outflow groove 234.
Referring to fig. 1, 4, 5 and 6, a ball and socket structure 4 is disposed in a target surface region between adjacent jet holes 2, the depth to diameter ratio e/D of the ball and socket structure 4dThe optimal range is 0.1-0.3, the target surface is uniformly arranged in the circumferential direction, and the circumferential arrangement number is 1-6. The film holes 5 are arranged in the target surface area between the adjacent jet holes 2, and the cross section of the film holes isThe shape can be circular, oval and the like, the arrangement angle α of the air film holes 5 can be changed within the range of 0-180 degrees, the circumferential arrangement number of the air film holes 5 is preferably 1-4, and the air film holes 5 can be arranged on the surface of the ball socket structure 4 or in the target surface area near the ball socket.
Claims (7)
1. A novel turbine blade rotational flow cooling structure is characterized by comprising a rotational flow cooling cavity (1), a jet hole (2), an outflow groove (3), a ball socket structure (4) and an air film hole (5); wherein the content of the first and second substances,
the top of the cyclone cooling cavity (1) is of a plane structure, and the target surface is of a curved surface structure; the jet holes (2) are arranged on one side of the top surface of the cyclone cooling cavity (1) at equal intervals, the outflow grooves (3) are arranged on the other side of the top surface of the cyclone cooling cavity (1) opposite to the jet holes (2) at equal intervals, and the outflow grooves (3) and the jet holes (2) are arranged in a staggered manner; a plurality of ball socket structures (4) and air film holes (5) are arranged on the target surface of the cyclone cooling cavity (1);
during operation, a cooling working medium is injected from the jet hole (2) to enter the cyclone cooling cavity (1), flows along one side of a target surface where the jet hole (2) is located, forms a transverse large-scale vortex under the action of a bent target surface, and carries out strong heat exchange with the target surface, flow separation and reattachment are generated when the cooling working medium flows through the ball socket structure (4), so that local heat transfer of the ball socket is enhanced, higher fluid turbulence energy is obtained under the suction action of the air film hole (5), the local heat transfer deterioration phenomenon is eliminated, and the cooling working medium after heat exchange is discharged through the outflow groove (3) adjacent to the jet hole (2).
2. The novel turbine blade cyclone cooling structure according to claim 1, wherein the target surface shape of the cyclone cooling cavity (1) is a circular arc, an elliptical arc, a parabola or a hyperbola.
3. The novel turbine blade rotational flow cooling structure as claimed in claim 1, wherein the jet holes (2) are rectangular, straight notch, circular and oval in shape, the S/w ratio of the arrangement pitch of the jet holes (2) to the width of the jet holes is within the range of 2 to 5, and the w/t ratio of the width of the jet holes to the thickness is within the range of 1 to 6.
4. The new turbine blade cyclone cooling structure according to claim 3, characterized in that the outflow slots (3) are arranged on the other side of the area between the adjacent jet holes (2), and have the shape of rectangle, straight slot, circle and ellipse, and the structure size and arrangement are the same as the jet holes (2).
5. The new turbine blade cyclone cooling structure according to claim 1, characterized in that the ball and socket structure (4) is arranged in the target surface area between adjacent jet holes (2), the depth to diameter ratio e/D of the ball and socket structure (4)dIn the range of 0.1-0.3, the target surface is uniformly arranged in the circumferential direction, and the number of the circumferential arrangement is 1-6.
6. The novel turbine blade rotational flow cooling structure as claimed in claim 1, wherein the gas film holes (5) are arranged in the target surface area between the adjacent jet holes (2), the cross section of the gas film holes is circular or elliptical, the arrangement angle α of the gas film holes (5) is changed within the range of 0-180 degrees, and the circumferential arrangement number of the gas film holes (5) is 1-4.
7. The novel turbine blade cyclone cooling structure according to claim 1, characterized in that the film holes (5) are arranged on the surface of the ball-and-socket structure (4) or in the target surface area near the ball-and-socket structure (4).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911259529.4A CN111120008A (en) | 2019-12-10 | 2019-12-10 | Novel turbine blade rotational flow cooling structure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911259529.4A CN111120008A (en) | 2019-12-10 | 2019-12-10 | Novel turbine blade rotational flow cooling structure |
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| CN111120008A true CN111120008A (en) | 2020-05-08 |
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| CN201911259529.4A Pending CN111120008A (en) | 2019-12-10 | 2019-12-10 | Novel turbine blade rotational flow cooling structure |
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| CN113266427A (en) * | 2021-04-28 | 2021-08-17 | 西安交通大学 | Inside compound cooling structure of turbine movable vane |
| CN113404546A (en) * | 2021-07-09 | 2021-09-17 | 中国联合重型燃气轮机技术有限公司 | Blade, turbine and gas turbine |
| CN113404545A (en) * | 2021-07-09 | 2021-09-17 | 中国联合重型燃气轮机技术有限公司 | Gas turbine and turbine blade thereof |
| 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 |
| CN115045721A (en) * | 2022-08-17 | 2022-09-13 | 中国航发四川燃气涡轮研究院 | Series-type rotational flow impact turbine blade cooling unit and turbine blade |
-
2019
- 2019-12-10 CN CN201911259529.4A patent/CN111120008A/en active Pending
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| CN113266427A (en) * | 2021-04-28 | 2021-08-17 | 西安交通大学 | Inside compound cooling structure of turbine movable vane |
| CN113266427B (en) * | 2021-04-28 | 2022-07-12 | 西安交通大学 | Inside compound cooling structure of turbine movable vane |
| CN113404546A (en) * | 2021-07-09 | 2021-09-17 | 中国联合重型燃气轮机技术有限公司 | Blade, turbine and gas turbine |
| CN113404545A (en) * | 2021-07-09 | 2021-09-17 | 中国联合重型燃气轮机技术有限公司 | Gas turbine and turbine blade thereof |
| 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 |
| CN115045721A (en) * | 2022-08-17 | 2022-09-13 | 中国航发四川燃气涡轮研究院 | Series-type rotational flow impact turbine blade cooling unit and turbine blade |
| CN115045721B (en) * | 2022-08-17 | 2022-12-06 | 中国航发四川燃气涡轮研究院 | Series-type rotational flow impact turbine blade cooling unit and turbine blade |
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