CN117449916A - Cooling channel applied to middle chord area of turbine blade - Google Patents
Cooling channel applied to middle chord area of turbine blade Download PDFInfo
- Publication number
- CN117449916A CN117449916A CN202311522601.4A CN202311522601A CN117449916A CN 117449916 A CN117449916 A CN 117449916A CN 202311522601 A CN202311522601 A CN 202311522601A CN 117449916 A CN117449916 A CN 117449916A
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- CN
- China
- Prior art keywords
- cooling
- turbine blade
- channel
- passage
- cooling passage
- 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 63
- 239000000112 cooling gas Substances 0.000 claims abstract description 6
- 239000007789 gas Substances 0.000 claims description 9
- 230000008646 thermal stress Effects 0.000 claims description 4
- 239000002826 coolant Substances 0.000 claims 1
- 230000001419 dependent effect Effects 0.000 claims 1
- 230000000694 effects Effects 0.000 abstract description 10
- 238000012546 transfer Methods 0.000 abstract description 7
- 238000004519 manufacturing process Methods 0.000 abstract description 2
- 238000011161 development Methods 0.000 description 3
- 239000000463 material Substances 0.000 description 3
- 238000013461 design Methods 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000009825 accumulation Methods 0.000 description 1
- 230000006978 adaptation Effects 0.000 description 1
- 238000004458 analytical method Methods 0.000 description 1
- 238000003491 array Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000000903 blocking effect Effects 0.000 description 1
- 239000012809 cooling fluid Substances 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000000428 dust Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The present invention discloses a turbine cooling passage with a plurality of curves, wherein the blade is provided with 6 chordwise turning passages, and 5 radial shorter cooling passages and 2 radial longer cooling passages. The cooling gas enters from the lower part of the channel and is discharged from the outlets at the upper parts of the blades through a plurality of turns. The blades with the cooling channels in the form can better weaken the influence of the rotation force on the heat transfer of the blades under the rotation working condition of the blades, meanwhile, the local heat exchange strength is improved without depending on the streaming structures such as ribs and the like, and the balance of the heat transfer strength of the front edge surface and the rear edge surface is realized. The manufacturing difficulty of blade is lower, and the cooling effect is more even, and life is longer.
Description
Technical Field
The invention relates to the field of turbomachinery, in particular to a cooling channel applied to a chord area in a turbine blade.
Background
Under the high-temperature working condition, the blades in the turbine machinery are easily affected by thermal stress, so that the problems of thermal fatigue, deformation and the like of the blade materials are caused. To reduce the impact of these problems on turbine mechanical performance and life, cooling channels are often used for active blade cooling. However, the conventional cooling channels have problems of uneven flow, poor cooling effect and the like, and seriously affect the working efficiency and reliability of the turbine machinery.
Under the rotating state of the turbine blade, due to the action of the rotating type force, the front edge surface and the rear edge surface of the traditional blade cooling channel have the problem of uneven heat transfer and too depend on the turbulent flow structure to improve the local heat exchange strength, and meanwhile, the existence of the turbulent flow structure increases the dust accumulation rate and shortens the service life of the blade cooling channel.
Disclosure of Invention
The present invention aims to provide a cooling passage for a chord zone in a turbine blade, which aims to solve the problems of the conventional passage.
The invention is realized by adopting the following technical scheme:
a cooling channel for the chord region of turbine blade is composed of several S-shaped channels with radially shorter vertical channels, the outlets of former S-shaped channels are connected to the inlets of next S-shaped channels, the inlets of first S-shaped channels are at blade root, the outlets of last S-shaped channels are at blade top, and cooling gas is introduced from the inlets of first S-shaped channels and discharged from the outlets of last S-shaped channels.
The invention is further improved in that the cooling channels are provided with 2 180-degree turning channels in each S-shaped channel, the cooling working medium is rapidly turned in the turning channels, and the heat exchange strength of the nearby area can be enhanced through impact and rotational flow cooling.
A further development of the invention is that the cooling channels are uniformly arranged with ribs in two shorter vertical channel areas, wherein the shorter vertical channel areas are 40% of the cooling area of the blade.
A further development of the invention is that the ribs extend obliquely or horizontally in the width direction of the turbine blade.
A further development of the invention is that the cooling channel can also be left-side inflow or right-side inflow, while the opposite side outflow is chosen for adaptation to the turbine design.
A further improvement of the invention is that there are at least three S-channels in the cooling channel.
The invention is further improved in that a plurality of air film holes are arranged in the cooling channel in order to reduce the temperature difference between the inside and outside of the blade and reduce the thermal stress.
The invention is further improved in that the number of the gas film holes is related to the temperature of the outside main flow gas.
The invention has at least the following beneficial technical effects:
the novel turbine cooling channel is provided with a plurality of 180-degree bends, so that the disturbance of working medium flow in a turning area is large, and the heat exchange capacity is extremely high.
The vertical channel of the novel channel is relatively short, and the longest section is only 40% of the height of the cooling area of the blade, so that the channel is relatively less influenced by the rotating force, the heat transfer of the rotating front edge surface and the rear edge surface of the channel is more even, and if ribs are required to be arranged, the novel channel can be arranged in the short vertical channel area, so that the manufacturing difficulty of the blade is reduced.
In summary, the novel turbine cooling channel of the invention can improve the cooling effect of the turbine blade and ensure the performance and the service life of the turbine machinery.
Drawings
FIG. 1 is a schematic view of the turbine cooling passage structure of the present invention.
FIG. 2 is a simplified schematic of a turbine cooling passage of the present invention.
Fig. 3 is a diagram showing the heat transfer effect of a conventional cooling passage in a rotating state.
FIG. 4 is a graph comparing heat transfer effects of the novel cooling channels in a rotated state.
Detailed Description
Reference will now be made in detail to embodiments of the present invention, examples of which are illustrated in the accompanying drawings. The embodiments of the present invention are not limited to the above description, but may be appropriately adjusted and modified according to specific needs. And the present invention may be used in combination with other known techniques.
As shown in FIG. 1, the present invention is implemented with a turbine mid-chord zone cooling channel having a plurality of S-shaped channels, into which cooling gas flows from below, thereby passing through 6 180 turns. The flow channel also includes two relatively short vertically long channels, and finally the cooling fluid flows out from the outlet of the 6 # turn.
For ease of analysis and understanding, fig. 2 shows a simplified cooling channel. Compared with the traditional S-shaped cooling channel, the overall channel length of the channel is not changed, the positions of the inlet and the outlet are completely consistent, and the combined heat exchange effect is improved without completely depending on a bypass structure.
It should be noted that the cross-sectional area of the cooling channel may be different at different locations, and in particular, may be determined according to the curvature of the surface where the channel is located and the different cooling requirements of the region.
Fig. 3-4 show a comparison of the heat transfer effects of the novel channels and the conventional cooling channels. It can be seen intuitively that the heat exchange strength and uniformity of the novel vane are far better than those of the traditional channel, and even no additional bypass structure in any form is needed.
Optionally, as shown in fig. 2, to further enhance the heat exchange effect of the channels 1 and the long channels 2, it is conceivable to add a flow-around structure (rib, socket, bulb etc.) in the vicinity of both channels. It can be appreciated that the space size of the turbulence structure is smaller than the cross-sectional area of the cooling channel, so as to avoid blocking the cooling channel. The specific arrangement may be selected based on the actual cooling requirements and on-site process conditions.
Alternatively, as the radial length of the blade is greater, more S-bend regions may be optionally added.
Alternatively, the inlets of the channels may be provided on the left and right sides of the whole in order to facilitate the flow of cooling gas into the cooling channels without significantly changing the cooling effect.
Alternatively, the novel channel may also be applied to the leading edge region remote from the mid-chord region, and impingement cooling or film cooling structures for the leading edge may also be provided at the channel.
Optionally, the novel channel can also be applied to a tail edge area far away from the mid-chord area, and a spoiler column array aiming at the tail edge or a tail edge split joint structure can also be arranged at the channel.
Optionally, when the temperature of the outside main stream gas is higher, in order to reduce the temperature difference between the inside and outside of the blade and reduce the thermal stress, mid-chord film cooling holes can be added on the front edge surface and the rear edge surface, and the shape of the holes is not limited.
Optionally, the number of film holes arranged is related to the ambient mainstream gas temperature. The number of the gas film holes is related to the size of the gas film holes and the temperature difference between the inside and the outside, and when the temperature difference is relatively large, more gas film hole arrays with larger sizes are required to be arranged, so that the gas film can be well covered on the surface of the blade.
Specifically, 1-3 exhaust film holes are uniformly arranged at the position of the turbine blade suction surface close to the front edge (namely, the position of a long channel 2 and the positions of turns 1, 4 and 5 in the picture), and the size and the form of the film holes are variable.
The specific design steps of the invention are as follows:
(1) And determining the position, the S-bend number and the inlet and outlet positions of the novel channel according to the three-dimensional shape data of the blade.
(2) And determining the flow of cooling gas of the blade according to the thermal load distribution of the surface of the blade, and checking the pressure drop of the cooling channel.
(3) The flow-around configuration of the vertical channels is selected according to the thermal load of the blades.
(4) And selecting additional air film hole forms and arrangement modes according to the internal and external temperature difference and the attribute of the blade material.
The above is only for illustrating the technical idea of the present invention, and the protection scope of the present invention is not limited by this, and any modification made on the basis of the technical scheme according to the technical idea of the present invention falls within the protection scope of the claims of the present invention.
In the present invention, the width between the outer surface of the blade and the channel may vary depending on the strength of the blade material. In particular, when the blade strength is poor, the passage sectional area can be appropriately reduced.
Claims (8)
1. A cooling channel for a chord zone in a turbine blade, characterized in that the cooling channel consists of a plurality of radially shorter S-shaped channels of vertical channels, the outlet of the former S-shaped channel being connected to the inlet of the next S-shaped channel, the inlet of the first S-shaped channel being located at the blade root, the outlet of the last S-shaped channel being located at the blade tip, cooling gas entering from the inlet of the first S-shaped channel and exiting from the outlet of the last S-shaped channel.
2. A cooling passage for a mid-chord region of a turbine blade according to claim 1, wherein the cooling passage has 2 180 degree turn passages in each S-shaped passage, the cooling medium turns rapidly at the turn, and heat exchange strength in the vicinity can be enhanced by impingement and swirl cooling.
3. A cooling passage for a chord zone in a turbine blade according to claim 1 wherein the cooling passage is evenly arranged with fins in two shorter vertical passage zones, wherein the shorter vertical passage zone is 40% of the blade cooling zone.
4. A cooling channel for a chord zone in a turbine blade according to claim 3, wherein the rib extends obliquely or horizontally in the width direction of the turbine blade.
5. A cooling passage for a mid-chord region of a turbine blade according to claim 1, wherein the cooling passage is further adapted to accommodate changes in turbine configuration by either a left side inflow or a right side inflow while a side outflow.
6. A cooling passage for a chord zone in a turbine blade according to claim 1 wherein the number of S-shaped passages in the cooling passage is at least three.
7. A cooling passage for a mid-chord region of a turbine blade according to claim 1, wherein a plurality of film holes are disposed in the cooling passage for reducing the temperature differential between the inner and outer surfaces of the blade to reduce thermal stresses.
8. A cooling passage for a mid-chord region of a turbine blade according to claim 1 wherein the number of film holes is dependent on the ambient mainstream gas temperature.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311522601.4A CN117449916A (en) | 2023-11-15 | 2023-11-15 | Cooling channel applied to middle chord area of turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311522601.4A CN117449916A (en) | 2023-11-15 | 2023-11-15 | Cooling channel applied to middle chord area of turbine blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN117449916A true CN117449916A (en) | 2024-01-26 |
Family
ID=89583507
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202311522601.4A Pending CN117449916A (en) | 2023-11-15 | 2023-11-15 | Cooling channel applied to middle chord area of turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN117449916A (en) |
-
2023
- 2023-11-15 CN CN202311522601.4A patent/CN117449916A/en active Pending
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