CN108167026B - Baffle plate with depressions and turbine blade internal cooling channel - Google Patents
Baffle plate with depressions and turbine blade internal cooling channel Download PDFInfo
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- CN108167026B CN108167026B CN201711433488.7A CN201711433488A CN108167026B CN 108167026 B CN108167026 B CN 108167026B CN 201711433488 A CN201711433488 A CN 201711433488A CN 108167026 B CN108167026 B CN 108167026B
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- side wall
- turbine blade
- partition plate
- recess
- 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
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- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A spacer for partitioning cooling passages within a turbine blade includes a spacer body and a plurality of dimples disposed on the spacer body. A cooling passage and a turbine blade including the above-described partition and cooling passage are also disclosed. The depressions arranged on the partition plate can generate near-wall secondary flow, so that the turbulence intensity of the near-wall surface is enhanced, and the heat exchange capacity of the wall surface of the partition plate is improved, thereby improving the overall heat exchange capacity of the cooling channel.
Description
Technical Field
The invention relates to the technical field of cooling of high-temperature components of gas turbines or aircraft engines, in particular to a partition plate with a recess and a cooling channel inside a turbine blade.
Background
Increasing the pre-turbine gas temperature is a major trend in the development of modern aircraft engines and gas turbines. More efficient, advanced cooling techniques are therefore urgently needed to ensure reliable operation of gas turbine blades. Since the cooling air is extracted from the compressor, the enhanced cooling technique means that the amount of cooling fluid used can be reduced, which is beneficial for improving the overall efficiency of the aircraft engine and gas turbine.
Existing turbine blades are generally of hollow construction and have a plurality of internal cooling flow passages defined by the pressure and suction side wall faces 1, 3 of the turbine blade (see fig. 1). Turbulence ribs are arranged on the inner sides of the pressure side and the suction side wall surface in the cooling channel, and cooling fluid flows in the cooling channel with the turbulence ribs to cool the turbine blade.
A search of prior art documents found that Daigo Fujimura et al (U.S. patent No.8556583B2), Ching-pang Lee (U.S. patent No.5797726) proposed a ribbed cooling structure with multiple channels inside the turbine blade. And the inner wall surface of the middle chord area of the blade is provided with a U-shaped cooling channel of a ribbed spoiler. Daigo Fujimura et al (U.S. patent No.8556583b2) propose to arrange a recess on a wall surface with flow disturbing ribs in order to keep the heat transfer capacity constant but reduce the flow resistance. However, as described in the patent (U.S. patent No.8556583B2), the depressed vortices and the turbulator ribs are arranged on the inner wall surfaces of the pressure side wall surface and the suction side wall surface, and although this can improve the heat transfer performance, the aeroengine turbine blade wall surfaces are generally thin, and the arrangement of the depressed vortices on the wall surfaces will cause the wall surfaces to be weakened; in addition, because the flow separation area is also arranged in the recess, the local heat transfer is deteriorated, and local high-temperature hot spots are easy to appear. However, it is only feasible to provide a recess in the wall of the turbine blade of a gas turbine for power generation, if the wall is thick. In addition, since the turbulator ribs have been disposed on the inner wall surfaces of the pressure side wall surface and the suction side wall surface of the turbine blade, the surface area on which the recesses are disposed on these wall surfaces is limited. This will also affect the exertion of the enhanced heat transfer capability of the depressed vortex.
Disclosure of Invention
In view of the above-mentioned defects of the prior art, the technical problem to be solved by the present invention is to provide a cooling channel capable of improving convective heat transfer performance while maintaining low flow resistance, and a turbine blade having the cooling channel.
To achieve the above object, a first aspect of the present invention provides a partition plate for partitioning a cooling passage inside a turbine blade, the partition plate including a partition plate body and a plurality of recesses provided on the partition plate body.
Further, the plurality of recesses are staggered on the separator body, but not limited thereto. Compared with other arrangement forms, the plurality of the recesses are staggered on the separator body, so that a better heat transfer enhancement effect can be obtained.
Further, the depth to diameter ratio of the depression is less than 0.5.
Further, a plurality of recesses are provided on both side wall surfaces or one of the side wall surfaces of the bulkhead body.
Further, the plurality of recesses include a first recess having a depth less than or half of a thickness of the separator body and a second recess having a depth equal to or half of the thickness of the separator body.
Further, the second recess is provided at an end portion of the separator body, which is a portion of the separator near the cooling passage turning region, and the recesses at both sides of the separator may penetrate.
A second aspect of the present invention provides a cooling passage provided inside a turbine blade, the cooling passage being defined by an inner sidewall of the turbine blade and any one of the above-described partitions.
Further, the partition plate is provided in plurality so that the cooling passage has a plurality.
Further, be provided with one or more vortex rib on the inside wall of turbine blade for cooling channel is injectd by vortex rib, turbine blade's inside wall, baffle body and sunken and forms.
Further, one or more rib-depression composite structures are disposed on the inner sidewall of the turbine blade.
A third aspect of the present invention provides a turbine blade comprising any one of the above-described partitions and any one of the above-described cooling passages, the inner sidewall including a pressure-side wall surface and a suction-side wall surface, both end portions of the partition being connected to the pressure-side wall surface and the suction-side wall surface, respectively, the plurality of partitions dividing the cooling passage into a plurality of cooling passages.
In a preferred embodiment of the present invention, a turbine blade comprises an inner sidewall, a diaphragm, and a cooling channel, the inner sidewall comprising a pressure sidewall surface and a suction sidewall surface, both ends of the diaphragm being connected to the pressure sidewall surface and the suction sidewall surface, respectively, the cooling channel being defined by the diaphragm, the pressure sidewall surface, and the suction sidewall surface; the plurality of partitions divide the cooling passage into a plurality of cooling passages. Wherein, the pressure side wall surface and/or the suction side wall surface are/is provided with one or more turbulence ribs. The partition board comprises a partition board body and a plurality of recesses, wherein the plurality of recesses are arranged on two side wall surfaces or one side wall surface of the partition board body.
In another preferred embodiment of the present invention, a turbine blade includes an inner sidewall, a partition, and a cooling channel, the inner sidewall including a pressure sidewall surface and a suction sidewall surface, both end portions of the partition being connected to the pressure sidewall surface and the suction sidewall surface, respectively, the cooling channel being defined by the partition, the pressure sidewall surface, and the suction sidewall surface; the plurality of partitions divide the cooling passage into a plurality of cooling passages. Wherein, the pressure side wall surface and/or the suction side wall surface are/is provided with one or more turbulence ribs. The partition board comprises a partition board body and a plurality of recesses, wherein the plurality of recesses are arranged on two side wall surfaces or one side wall surface of the partition board body. The plurality of recesses include a first recess having a depth less than the thickness of the separator body or less than a half of the thickness of the separator body, and a second recess having a depth equal to the thickness of the separator body or equal to a half of the thickness of the separator body. The second recesses are provided at the end of the separator body, which is the portion of the separator near the turn region of the cooling passage, and the second recesses on both sides of the separator are communicated with each other. In the existing turbine blade, because the side wall surface of the partition is a smooth side wall surface, the heat exchange capability is low, and because the partition is connected with the pressure side wall surface and the suction side wall surface of the turbine blade, heat can be transferred to cooling fluid through the side wall surface of the partition.
Compared with the prior art, the invention has the following advantages that the plurality of the depressions are arranged on the partition plate:
1. the depressions on the side wall surface of the partition plate can generate secondary flow near the wall surface, the turbulence intensity near the wall surface is enhanced, and the heat exchange capacity of the side wall surface of the partition plate is improved, so that the overall heat exchange capacity of the cooling channel is improved, and the pressure loss of the cooling channel is not increased. Or the number of the turbulence ribs can be reduced under the condition of keeping the same heat transfer capacity of the cooling channel, so that the pressure loss of the cooling channel is obviously reduced. Reducing the number of turbulator ribs also facilitates reducing the weight of the turbine blade.
2. The provision of the recess in the side wall surface of the diaphragm contributes to the reduction in weight of the turbine blade.
3. The concave part arranged on the partition board is beneficial to inhibiting the flow separation at the flow turning part of the internal cooling channel, the flow loss is reduced, and the heat exchange uniformity of the turning area of the internal cooling channel is improved.
4. One or more depressions having a depth equal to the thickness of the partition or half of the thickness of the partition, that is, depressions punched through the partition, are provided on the partition, particularly at the end of the partition (near the flow separation region), so that flow separation of the flow passage on the other side (that is, the flow passage after flow turning) can be suppressed, pressure loss of the cooling passage inside the turbine blade is reduced, and uniformity of heat transfer of the wall surface having the turbulator ribs is improved.
5. By providing shallow recesses (the depth to diameter ratio of the recesses is less than 0.5) in the opposite sidewall faces of the partition, it is advantageous to reduce the pressure loss of the cooling passages and maintain the heat transfer and cooling capacity of the cooling passages.
6. The arrangement of the depressions on the partition plate and the arrangement of the depressions on the inner wall surfaces of the pressure side and the suction side of the turbine blade form a rib-depression composite structure, so that the heat transfer capacity of the cooling channel can be further improved, and the increase of the flow pressure drop is not obvious.
Compared with the traditional turbulent flow rib cooling channel, the turbulent flow rib cooling channel with the sunken vortexes has the advantage that the overall heat transfer performance is improved by 15-20%.
The conception, the specific structure and the technical effects of the present invention will be further described with reference to the accompanying drawings to fully understand the objects, the features and the effects of the present invention.
Drawings
FIG. 1 is a schematic view of a prior art turbine blade;
FIG. 2 is a schematic view of a turbine blade in a preferred embodiment of the invention;
FIG. 3 is a partial schematic view of a turbine blade in accordance with a preferred embodiment of the present invention;
FIG. 4 is a partial schematic view of a turbine blade in accordance with another preferred embodiment of the present invention;
FIG. 5 is a graph comparing overall heat exchange performance data for two cooling passages.
Detailed Description
As shown in fig. 3, a preferred embodiment of the present invention provides a turbine blade 10, which includes a pressure side wall surface 1, a suction side wall surface 2, a partition plate 3 and a cooling channel, wherein a plurality of spoiler ribs 4 are disposed on the pressure side wall surface 1 and the suction side wall surface 2, the spoiler ribs 4 are in a long strip shape, and the spoiler ribs 4 are disposed in parallel with each other. Both ends of the diaphragm 3 are respectively connected to the pressure side wall face 1 and the suction side wall face 2. The cooling channel is defined by a partition 3, a pressure side wall face 1 and a suction side wall face 2. If a plurality of separators 3 are provided, the cooling passage will be divided into a plurality of cooling passages.
Wherein the partition 3 is provided with a plurality of recesses 32. The plurality of recesses 32 are staggered on the separator body 31, which is a preferred embodiment, and in other embodiments, other arrangements are possible. A plurality of recesses 32 are provided on both side wall surfaces or one of the side wall surfaces of the bulkhead body 31. The depth of the recess 32 is less than or equal to the thickness of the separator body 31, or the depth of the recess 32 is less than or equal to half the thickness of the separator body 31.
In the present embodiment, the ratio of the depth to the diameter of the recess 32 is less than 0.5, and preferably the ratio of the depth to the diameter of the recess 32 is 0.2.
In one embodiment, the plurality of recesses 32 include a first recess and a second recess, which are provided at one of the side wall surfaces of the separator body 31, the depth of the first recess being smaller than the thickness of the separator body 31, and the depth of the second recess being equal to the thickness of the separator body 31, i.e., the bottom of the second recess penetrates the separator body 31 in the thickness direction.
In one embodiment, the plurality of recesses 32 include a first recess and a second recess, which are provided on both side wall surfaces of the separator body 31, the depth of the first recess is less than half of the thickness of the separator body 31, and the depth of the second recess is equal to half of the thickness of the separator body 31, that is, the bottoms of the second recesses on both sides each penetrate half of the thickness of the separator body 31 in the thickness direction, so that the separator body 31 is perforated.
In one embodiment, the second recesses are provided at the upper end and the lower end of the bulkhead body 31, and the end of the bulkhead body 31 refers to a portion of the bulkhead near the turning region of the cooling passage, and more specifically, refers to a connection or a connection near the turbine blade at the upper end and the lower end of the bulkhead body 31, respectively, when the second recesses at both sides of the bulkhead body 31 are penetrated.
The cooling passage of the present embodiment is defined by the turbulator ribs 4, the pressure side wall surfaces 1 and the suction side wall surfaces 2 of the turbine blades, the partition body 31, and the dimples 32. The depressions 32 on the side wall surface of the partition plate 3 will generate secondary flow near the wall surface, thereby enhancing the turbulence intensity near the wall surface, and improving the heat exchange capability of the side wall surface of the partition plate 3, so that the whole heat exchange capability of the cooling channel is improved, and the pressure loss of the cooling channel is not increased. Or the number of the turbulence ribs 4 can be reduced under the condition of keeping the same heat transfer capacity of the cooling channel, so that the pressure loss of the cooling channel is obviously reduced. Reducing the number of turbulator ribs 4 also contributes to reducing the weight of the turbine blade. The provision of the recess 32 in the partition plate 3 also contributes to a reduction in the weight of the turbine blade.
The provision of the recess 32 in the partition plate 3, a portion of which penetrates the partition plate 3, is advantageous in suppressing flow separation on the other side, reducing pressure loss in the cooling passage inside the turbine blade, and improving uniformity of heat transfer in the wall surface having the turbulator ribs 4.
As shown in fig. 4, another embodiment of the present invention provides a turbine blade, which includes a pressure side wall surface 1, a suction side wall surface 2, a partition plate 3, and a cooling channel, wherein a plurality of spoiler ribs 4 are provided on the pressure side wall surface 1 and the suction side wall surface 2, the spoiler ribs 4 are in a long strip shape, and the spoiler ribs 4 are arranged in parallel with each other. Both ends of the diaphragm 3 are respectively connected to the pressure side wall face 1 and the suction side wall face 2. The cooling channel is defined by a partition 3, a pressure side wall face 1 and a suction side wall face 2. If a plurality of separators 3 are provided, the cooling passage will be divided into a plurality of cooling passages. Wherein the partition 3 is provided with a plurality of recesses 32. The plurality of recesses 32 are arranged in a staggered manner on the separator body 31. A plurality of recesses 32 are provided on both side wall surfaces or one of the side wall surfaces of the bulkhead body 31. The depth of the recess 32 is less than or equal to the thickness of the separator body 31, or the depth of the recess 32 is less than or equal to half the thickness of the separator body 31.
In this embodiment, the density of the plurality of recesses 32 on the separator body 31 is greater. Because the high-density depressions 32 are formed in the separator body 31, the heat exchange capacity of the surface of the separator 3 is further improved, and the heat transfer capacity of the whole cooling channel is further improved.
Fig. 5 shows a graph comparing the overall heat exchange performance data of two cooling channels, i.e. a turbulator rib cooling channel and a turbulator rib cooling channel with a trapped vortex. The turbulator rib cooling channels are prior art here, referring to the case: the pressure side wall surface and the suction side wall surface are provided with turbulence ribs, but the partition plate is not provided with a recess, so that the turbulence rib cooling channel is limited by the pressure side wall surface, the suction side wall surface, the turbulence ribs and the partition plate. Turbulator rib cooling channels with recessed vortices refer to the case of the present invention: turbulence ribs are arranged on the pressure side wall surface and the suction side wall surface, and a recess is arranged on the partition plate, so that the turbulence rib cooling channel with the recessed vortex is formed by limiting the pressure side wall surface, the suction side wall surface, the turbulence ribs, the recess and the partition plate. As can be seen from fig. 5, the overall heat transfer performance of the cooling channel with the concave vortex of the embodiment of the invention is improved by 15% -20% compared with that of the conventional cooling channel with the turbulent rib.
The foregoing detailed description of the preferred embodiments of the invention has been presented. It should be understood that numerous modifications and variations could be devised by those skilled in the art in light of the present teachings without departing from the inventive concepts. Therefore, the technical solutions available to those skilled in the art through logic analysis, reasoning and limited experiments based on the prior art according to the concept of the present invention should be within the scope of protection defined by the claims.
Claims (7)
1. A partition plate for partitioning a turbine blade internal cooling passage, characterized by comprising a partition plate body and a plurality of recesses provided on the partition plate body, the plurality of recesses being provided on one or both side wall surfaces of the partition plate body, the plurality of recesses including a first recess having a depth less than half of a thickness of the partition plate body and a second recess having a depth equal to half of the thickness of the partition plate body, when the plurality of recesses are provided on both side wall surfaces of the partition plate body, the second recess being symmetrically arranged in a thickness direction of the partition plate body; when the plurality of recesses are provided in one of the side wall surfaces of the separator body, the depth of the first recess is smaller than the thickness of the separator body, and the depth of the second recess is equal to the thickness of the separator body; the second recess is provided at an end portion of the separator body, the end portion of the separator body being a portion of the separator near a turning region of the cooling passage, and the second recess is through-shaped for suppressing flow separation at the turning region of the flow of the internal cooling passage.
2. The separator plate of claim 1 wherein said plurality of depressions are staggered on said separator plate body.
3. The separator according to claim 1, wherein the depth to diameter ratio of said depression is less than 0.5.
4. A cooling passage provided inside a turbine blade, characterized in that the cooling passage is defined by an inner side wall of the turbine blade and the partition plate of any one of claims 1 to 3.
5. The cooling passage according to claim 4, wherein the partition plate is provided in plurality so that the cooling passage has a plurality.
6. The cooling channel of claim 4, wherein one or more turbulator ribs are provided on an inner sidewall of the turbine blade.
7. A turbine blade comprising the cooling passage of any one of claims 4-6, a plurality of partitions dividing the cooling passage into a plurality of cooling passages.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711433488.7A CN108167026B (en) | 2017-12-26 | 2017-12-26 | Baffle plate with depressions and turbine blade internal cooling channel |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201711433488.7A CN108167026B (en) | 2017-12-26 | 2017-12-26 | Baffle plate with depressions and turbine blade internal cooling channel |
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| Publication Number | Publication Date |
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| CN108167026A CN108167026A (en) | 2018-06-15 |
| CN108167026B true CN108167026B (en) | 2020-02-07 |
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| CN201711433488.7A Active CN108167026B (en) | 2017-12-26 | 2017-12-26 | Baffle plate with depressions and turbine blade internal cooling channel |
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| CN114215609B (en) * | 2021-12-30 | 2023-07-04 | 华中科技大学 | Blade internal cooling channel capable of enhancing cooling and application thereof |
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| US5472316A (en) * | 1994-09-19 | 1995-12-05 | General Electric Company | Enhanced cooling apparatus for gas turbine engine airfoils |
| GB2382383B (en) * | 2001-11-27 | 2005-09-21 | Rolls Royce Plc | Gas turbine engine aerofoil |
| US7334992B2 (en) * | 2005-05-31 | 2008-02-26 | United Technologies Corporation | Turbine blade cooling system |
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