CN211448781U - System for cooling turbine blade with participation of fuel - Google Patents
System for cooling turbine blade with participation of fuel Download PDFInfo
- Publication number
- CN211448781U CN211448781U CN202020048567.7U CN202020048567U CN211448781U CN 211448781 U CN211448781 U CN 211448781U CN 202020048567 U CN202020048567 U CN 202020048567U CN 211448781 U CN211448781 U CN 211448781U
- Authority
- CN
- China
- Prior art keywords
- fuel
- turbine
- cooling
- blade
- blades
- 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.)
- Expired - Fee Related
Links
- 239000000446 fuel Substances 0.000 title claims abstract description 110
- 238000001816 cooling Methods 0.000 title claims abstract description 80
- 238000001704 evaporation Methods 0.000 claims abstract description 18
- 230000008020 evaporation Effects 0.000 claims abstract description 18
- 239000002828 fuel tank Substances 0.000 claims abstract description 9
- -1 turbine vanes Substances 0.000 claims 1
- 239000000567 combustion gas Substances 0.000 description 11
- 239000007789 gas Substances 0.000 description 10
- 238000000034 method Methods 0.000 description 6
- 230000001681 protective effect Effects 0.000 description 6
- 239000003921 oil Substances 0.000 description 5
- 238000010521 absorption reaction Methods 0.000 description 3
- 230000000694 effects Effects 0.000 description 3
- 238000006243 chemical reaction Methods 0.000 description 2
- 238000002485 combustion reaction Methods 0.000 description 2
- 238000002309 gasification Methods 0.000 description 2
- 239000007788 liquid Substances 0.000 description 2
- 239000002826 coolant Substances 0.000 description 1
- 238000000354 decomposition reaction Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000000295 fuel oil Substances 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000003303 reheating Methods 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
Images
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A system for cooling turbine blades by fuel belongs to the field of gas turbine structures. The turbine comprises a fuel tank, turbine vanes, an oil pipe, a combustor, a rotor shaft and a turbine disc. The fuel tank is connected with the turbine stator blade through an oil pipe, the turbine stator blade is connected with the combustor through an oil pipe, and the fuel tank is connected with the rotor shaft through an oil pipe. The turbine stationary blade is provided with an inlet radial fuel channel, a circumferential fuel channel and an outlet radial fuel channel which are sequentially connected, and the surface of the turbine stationary blade is provided with a plurality of fuel evaporation holes connected with the outlet radial fuel channel. The turbine disk and the rotor shaft are provided with fuel holes, the turbine moving blades are provided with fuel channels connected with the fuel holes of the turbine disk and the rotor shaft, and the surfaces of the turbine moving blades are provided with a plurality of fuel evaporation holes connected with the fuel channels. The turbine vanes and blades are provided with cooling air passages spaced from the fuel passages. The invention reduces the consumption of cooling air, reduces the consumption of fuel and improves the unit output work of the gas turbine.
Description
Technical Field
The utility model relates to a fuel participates in turbine cooling's system belongs to gas turbine structure field.
Background
A gas turbine includes a compressor for compressing air, a combustor for combusting the compressed air together with fuel, a turbine for converting thermal energy of combustion gas into rotational energy, and the like. For a gas turbine, an effective method for improving the cycle thermal efficiency is to increase the temperature of combustion gas at the inlet of the turbine, but because the turbine blade material can bear the temperature limitation, the surface temperature of the turbine blade is at least 300-500K lower than the temperature of the combustion gas, and the turbine blade needs to be effectively cooled.
The turbine blades of the gas turbine are mostly cooled by using cooling air, and air of a certain stage of a compressor is generally led to the blades needing cooling. Some land based gas turbines are cooled with cold air or water prior to directing the cooling air to the turbine blades, and the cooling medium for the turbine is still a single air. The blades are cooled by introducing steam, and the turbine cooling adopts two media of steam and air, but the cooling mode must have enough steam. The turbine blade is cooled completely by cooling air, and the consumption of the cooling air is large and even accounts for 30 percent of the total air at the inlet of the gas turbine. The large cooling air usage reduces the air mass flow participating in work, interferes with the flow of the main air flow, and reduces the power and efficiency of the gas turbine. The large amount of cooling air forms a film of air over the cooled turbine blade surface, causing non-uniform temperatures in the main air stream as it flows through the blade.
SUMMERY OF THE UTILITY MODEL
The utility model discloses a for solving among the prior art turbine cooling and consuming a large amount of cooling air's not enough, provide a fuel and participate in turbine cooling's system, this system saves considerable cooling air under the same cooling effect.
In order to achieve the above purpose, the utility model adopts the following technical scheme:
a system for assisting in cooling turbine blades with fuel flow comprising a liquid, a gas, or a mixture of liquid and gas, wherein the fuel flow through the turbine blades absorbs heat, reducing the amount of cooling air used by the turbine blades. Fuel from the fuel tank is routed through the turbine vanes and blades, respectively. The fuel flowing through the turbine stationary blade absorbs heat of the turbine stationary blade, and most of the fuel is sent to the combustor to be combusted, and a small amount of the fuel is discharged through the fuel evaporation hole in the turbine stationary blade, so that a cooling protective film is formed on the surface of the turbine stationary blade. The fuel flowing through the turbine rotor blades absorbs heat of the turbine rotor blades, and is discharged from all the fuel evaporation holes in the turbine rotor blades, thereby forming a cooling protective film on the surfaces of the turbine rotor blades. Systems where fuel participates in turbine cooling employ cooling air to cool the turbine blades simultaneously.
The utility model has the advantages that: the fuel participates in the system for cooling the turbine blade, so that the consumption of cooling air is reduced, the fuel can be heated at first in the heat absorption process, the fuel can be evaporated and gasified after the temperature reaches a certain value, the fuel can absorb a large amount of heat in the physical process of heating and gasification, the fuel can be subjected to a continuous heat absorption process after the gasification is finished, the temperature is continuously raised, the fuel can be subjected to a chemical reaction of molecular decomposition when the temperature is raised to a certain value, the chemical reaction is also a heat absorption process, and therefore flowing fuel can replace part of the cooling air. At the same time, the temperature of the fuel is already raised to a high value before it reaches the burner, and since the fuel is heated for the purpose of raising the temperature, the fuel temperature is raised so as to reduce the amount of fuel used. The fuel participates in the turbine blade cooling system, except that the use amount of cooling air is reduced, part of the fuel is combusted in the turbine interstage to release heat, interstage reheating in the thermal process is formed, and the output work of the gas turbine is improved.
Drawings
FIG. 1 is a system diagram of fuel participating in turbine blade cooling.
FIG. 2 is a turbine vane cooling configuration view.
FIG. 3 is a turbine bucket cooling configuration view.
The utility model discloses reference number explains in the figure:
1-compressor 10 rotor shaft 11-turbine disk 2-air 21-compressed air 22-cooling air 23-cooling air 3-combustor 411-main fuel 412-fuel 41-vane cooling fuel 42-blade cooling fuel 5-combustion gas 6-turbine 61-turbine vane 62-turbine blade 611-blade tip fuel inlet 612-radial fuel inlet channel 613-circumferential fuel channel 614-radial fuel outlet channel 615-blade tip fuel outlet 616-fuel evaporation hole 617-cooling air channel 618-cooling air evaporation hole 621-fuel inlet 622-fuel channel 624-fuel evaporation hole 626-cooling air channel 628-cooling air evaporation hole 7-fuel tank 81 Regulator 82-regulator 91-oil pipe 92-oil pipe
Detailed Description
The present invention will be further explained with reference to the drawings and examples.
Referring to fig. 1, in the cooling structure of the present embodiment in which fuel participates in cooling of turbine stationary blades and turbine moving blades, a compressor 1 sucks air 2 from the outside and compresses the air, the compressed air 21 enters a combustor 3 and is mixed with main fuel 411 and then is combusted to form high-temperature combustion gas 5, and the combustion gas 5 flows through a turbine 6 to perform work.
The fuel of the fuel tank 7 is divided into two paths: the vane cooling fuel 41 and the blade cooling fuel 42. The vane cooling fuel 41 is adjusted in pressure and flow rate by the adjuster 81, and then sent to the inside of the turbine vane 61 through the oil pipe 91, and after the vane cooling fuel 41 absorbs heat in the inside of the turbine vane 61, most of the main fuel 411 is sent to the combustor 3, and the remaining small fuel flows out of the turbine vane 61 and is merged into the high-temperature combustion gas 5.
The pressure and flow rate of the blade cooling fuel 42 are adjusted by the regulator 82, and then the blade cooling fuel 42 enters the rotor shaft 10 from the compressor 1 end with a low temperature through the oil pipe 92, flows into the turbine blades 62 along the rotor shaft 10 and the fuel holes in the turbine disk 11, absorbs heat, and then flows out of the turbine blades 62 to join the high-temperature combustion gas 5.
In order to sufficiently cool the turbine blades, the cooling air 22 and the cooling air 23 introduced from a certain stage of the compressor 1 cool the turbine vanes 61 and the turbine blades 62, respectively.
As shown in FIG. 2, is an example of fuel participating in cooling a turbine vane.
Vane cooling fuel 41 delivered by an oil pipe enters a radial fuel inlet passage 612 through a tip fuel inlet 611 of the turbine vane 61, and the radial fuel inlet passage 612 is connected to a radial fuel outlet passage 614 through a segment of a circumferential fuel passage 613. The main fuel oil 411 in the vane cooling fuel 41 after two turns inside the blade is delivered from the tip fuel outlet 615 to the combustor 3.
The surface of the turbine stationary blade 61 is provided with a plurality of fuel evaporation holes 616 connected to the radial fuel outlet passage 614, and a small portion of the fuel 412 of the stationary blade cooling fuel 41 evaporates from the fuel evaporation holes 616 and is discharged from the turbine stationary blade 61, and a cooling protective film is formed on the blade surface and then high-temperature combustion gas 5 is merged. In order to ensure the cooling effect, a cooling air channel 617 is provided in the hollow turbine stationary blade 61, a plurality of cooling air evaporation holes 618 are provided on the blade surface, and all the cooling air is discharged from the cooling air evaporation holes 618 out of the blade to form a cooling protective film on the blade surface.
The fuel and the cooling air are separated in the passage of the blade, so that the fuel and the cooling air are prevented from being mixed in the blade to generate a combustion reaction, and the heat generated by combustion can cause the blade to be over-heated. The fuel can leave the blade surface certain distance behind the fuel discharge blade, and the fuel burns outside the blade and can not cause the super temperature damage to the blade, alleviates the temperature inhomogeneity that cooling air caused in the combustion gas of high temperature simultaneously.
As shown in FIG. 3, is an example of fuel participating in cooling turbine blades.
The blade cooling fuel 42 enters the fuel passage 622 from a fuel inlet 621 on the blade root. The turbine blades 62 have a plurality of fuel evaporation holes 624 formed on the surface thereof and connected to the fuel passage 622, and the blade cooling fuel 42 absorbs heat to cool the turbine blades 62, and then all of the fuel is discharged from the fuel evaporation holes 624 to the turbine blades 62, and a cooling protective film is formed on the blade surface, and high-temperature combustion gas 5 is introduced. In order to ensure the cooling effect, a cooling air channel 626 is arranged in the hollow turbine moving blade 62, a plurality of cooling air evaporation holes 628 are arranged on the surface of the blade, and after the cooling air cools the blade, all the cooling air is discharged out of the blade from the cooling air evaporation holes 628, so that a protective film is formed on the surface of the blade. The two passages for fuel and cooling air in the blade are separate.
Claims (3)
1. A system for fuel to participate in cooling of turbine blades includes a fuel tank, turbine vanes, fuel lines, a combustor, a rotor shaft, and a turbine disk; the fuel-efficient turbine is characterized in that the fuel tank is connected with the turbine stationary blade through an oil pipe, the turbine stationary blade is connected with the combustor through an oil pipe, and the fuel tank is connected with the rotor shaft through an oil pipe.
2. The system of claim 1, wherein the turbine vane is provided with an inlet radial fuel channel, a circumferential fuel channel and an outlet radial fuel channel connected in series, the surface of the turbine vane is provided with a plurality of fuel evaporation holes connected to the outlet radial fuel channel, and the turbine vane is provided with a cooling air channel spaced from the fuel channels.
3. The system of claim 1, wherein the turbine disk and the rotor shaft are provided with fuel holes, the turbine blades are provided with fuel passages connected with the fuel holes of the turbine disk and the rotor shaft, the surfaces of the turbine blades are provided with a plurality of fuel evaporation holes connected with the fuel passages, and the turbine blades are provided with cooling air passages separated from the fuel passages.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202020048567.7U CN211448781U (en) | 2020-01-10 | 2020-01-10 | System for cooling turbine blade with participation of fuel |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202020048567.7U CN211448781U (en) | 2020-01-10 | 2020-01-10 | System for cooling turbine blade with participation of fuel |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN211448781U true CN211448781U (en) | 2020-09-08 |
Family
ID=72316958
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202020048567.7U Expired - Fee Related CN211448781U (en) | 2020-01-10 | 2020-01-10 | System for cooling turbine blade with participation of fuel |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN211448781U (en) |
-
2020
- 2020-01-10 CN CN202020048567.7U patent/CN211448781U/en not_active Expired - Fee Related
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| GR01 | Patent grant | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20200908 |
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| CF01 | Termination of patent right due to non-payment of annual fee |