CN113202566B - Turbine guide vane and gas turbine engine - Google Patents
Turbine guide vane and gas turbine engine Download PDFInfo
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- CN113202566B CN113202566B CN202110417270.2A CN202110417270A CN113202566B CN 113202566 B CN113202566 B CN 113202566B CN 202110417270 A CN202110417270 A CN 202110417270A CN 113202566 B CN113202566 B CN 113202566B
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- turbine
- fuel
- guide vane
- cooling cavity
- gas
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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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/22—Fuel supply systems
- F02C7/224—Heating fuel before feeding to the burner
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a turbine guide vane and a gas turbine engine, wherein a cooling cavity is arranged in the turbine guide vane, the turbine guide vane comprises a front edge, a tail edge, a pressure surface and a suction surface, a premixed gas injection hole communicated with the cooling cavity is arranged on the turbine guide vane, fuel is injected to the inner wall surface of the cooling cavity from the cooling cavity to absorb heat and gasify the fuel on the inner wall surface of the cooling cavity, and then the premixed gas is mixed with cold gas in the cooling cavity to form premixed gas and is injected to high-temperature high-pressure gas around the turbine guide vane from the premixed gas injection hole to be combusted. According to the turbine guide vane, the energy added by fuel combustion is improved by supplementing combustion in the turbine guide vane, so that the temperature margin of a subsequent turbine is fully utilized, the output power of the whole turbine is greatly increased, and the power-to-weight ratio of a gas turbine engine is improved.
Description
Technical Field
The invention relates to the technical field of gas turbine engines, in particular to a turbine guide vane and a gas turbine engine.
Background
An aircraft gas turbine engine is typically constructed from components such as a compressor, combustor, and turbine. The working principle of the air-cooled combustor is that after air is compressed by an air compressor, the pressure is increased, the temperature is increased, and the air flows into a combustion chamber. The high-pressure air is mixed with the sprayed fuel in the combustion chamber to be combusted to form high-temperature and high-pressure fuel gas, the temperature is greatly increased, and the heat energy contained in the fuel gas is increased. The high-temperature and high-pressure gas flows into the turbine, power extraction is realized in the turbine to drive the compressor, and the residual energy is used for outputting shaft power or generating thrust. The gas is finally discharged into the atmosphere.
The power-to-weight ratio (power-to-weight ratio) and the fuel consumption rate (fuel amount consumed per unit power) of an aircraft engine are important indicators for measuring the performance of the engine. The energy added by fuel combustion is a main way for improving the power-weight ratio, but in the scheme of the conventional aircraft engine, the fuel combustion is completely carried out in a combustion chamber, so that the outlet temperature of the combustion chamber, namely the inlet temperature of a turbine, is greatly increased. The mainstream level of the current turbine inlet temperature is far beyond the material limit, so that the energy cannot be increased by adding fuel to a combustion chamber for combustion to improve the power-to-weight ratio.
Disclosure of Invention
The invention provides a turbine guide vane and a gas turbine engine, and aims to solve the technical problem that the existing gas turbine engine still needs to improve the energy added by fuel combustion to improve the power-to-weight ratio.
According to one aspect of the invention, a turbine guide vane is provided, a cooling cavity is arranged in the turbine guide vane, the turbine guide vane comprises a front edge, a tail edge, a pressure surface and a suction surface, a premixed gas injection hole communicated with the cooling cavity is arranged on the turbine guide vane, fuel is injected to the inner wall surface of the cooling cavity from the cooling cavity, so that the fuel absorbs heat and is gasified on the inner wall surface of the cooling cavity, and then is mixed with cold gas in the cooling cavity to form premixed gas, and the premixed gas is injected to high-temperature high-pressure gas around the turbine guide vane from the premixed gas injection hole to be combusted.
Furthermore, the cooling cavity is divided into a cooling front cavity and a cooling rear cavity along the direction from the front edge to the tail edge, and the premixed gas injection hole is formed in the wall body of the cooling front cavity.
Furthermore, a cooling film hole is formed in the wall body of the cooling rear cavity and is positioned on the pressure surface of the turbine guide vane, so that the cold air in the cooling rear cavity is sprayed out from the cooling film hole and forms a cooling film on the pressure surface.
Further, the premix gas injection orifices are located at the leading edge and near the pressure surface.
Furthermore, the premixed gas injection holes are arranged along the direction from the blade root to the blade top of the turbine guide blade, and the premixed gas injection holes are arranged in a plurality of rows in the direction from the front edge to the pressure surface.
Further, the turbine guide vane further includes a fuel injection member extending into the cooling cavity and connected to the fuel supply pipe for injecting fuel to an inner wall surface of the cooling cavity.
Furthermore, the fuel injection member is a fuel injection rod of a hollow rod structure, and a fuel injection hole for injecting fuel in the hollow cavity is arranged on an inner extending section of the fuel injection rod extending into the cooling cavity.
Furthermore, the fuel injection holes and the premixed gas injection holes are oppositely arranged; and/or the fuel injection hole is located in a peripheral region of the premixed gas injection hole to inject the fuel onto an inner wall surface around the premixed gas injection hole.
Further, a plurality of fuel injection holes are arranged in the axial direction of the fuel injection rod, and a plurality of rows of fuel injection holes are arranged in the circumferential direction of the fuel injection rod.
According to another aspect of the invention, there is also provided a gas turbine engine comprising the above turbine guide vane.
The invention has the following beneficial effects:
according to the turbine guide vane, in a multi-stage turbine (the turbine guide vane and a turbine rotor vane form a turbine grade), because the temperature is gradually reduced from the first stage to the last stage, the temperature born by the first stage turbine is highest, the temperature of the subsequent turbine stage is relatively low, the first stage turbine becomes a main factor for limiting the temperature increase of a combustion chamber outlet, and the subsequent turbine stage has the characteristic of large temperature margin.
In addition to the objects, features and advantages described above, other objects, features and advantages of the present invention are also provided. The present invention will be described in further detail below with reference to the drawings.
Drawings
The accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and not to limit the invention. In the drawings:
FIG. 1 is a cross-sectional structural schematic view of a turbine guide vane of a preferred embodiment of the present invention;
FIG. 2 is a schematic structural view of a turbine guide vane of a preferred embodiment of the present invention;
fig. 3 is a schematic structural view of a fuel injection rod according to a preferred embodiment of the present invention.
Illustration of the drawings:
1. a turbine guide vane; 2. cooling the cavity; 21. cooling the front cavity; 22. cooling the back cavity; 3. a premixed gas injection hole; 4. cooling the film holes; 5. a fuel injection rod; 51. a fuel injection hole.
Detailed Description
The embodiments of the invention will be described in detail below with reference to the accompanying drawings, but the invention can be embodied in many different forms, which are defined and covered by the following description.
FIG. 1 is a cross-sectional structural schematic view of a turbine guide vane of a preferred embodiment of the present invention; FIG. 2 is a schematic structural view of a turbine guide vane of a preferred embodiment of the present invention; fig. 3 is a schematic structural view of a fuel injection rod according to a preferred embodiment of the present invention.
As shown in fig. 1 and 2, in the turbine guide vane 1 of the present embodiment, a cooling cavity 2 is arranged in the turbine guide vane 1, the turbine guide vane 1 includes a front edge, a rear edge, a pressure surface and a suction surface, a premixed gas injection hole 3 communicated with the cooling cavity 2 is arranged on the turbine guide vane 1, fuel is injected from the cooling cavity 2 to an inner wall surface of the cooling cavity 2, so that the fuel absorbs heat and is gasified in the inner wall surface of the cooling cavity 2, and is further mixed with cold air in the cooling cavity 2 to form premixed gas, and the premixed gas is injected from the premixed gas injection hole 3 to high-temperature and high-pressure gas around the turbine guide vane 1 to be combusted. Known from the above structure, when the turbine guide vane 1 of this embodiment realizes supplementary combustion, the turbine guide vane 1 is not added, and the blade profile of the turbine guide vane 1 is also not changed, so as to avoid the blade profile from changing and increase the start loss and influence the turbine efficiency, but the existing cooling cavity 2 is modified, and the fuel is injected into the inner wall surface of the cooling cavity 2 for heat absorption gasification, so as to avoid the fuel from being directly injected onto the outer surface of the turbine guide vane 1 with higher temperature for combustion and causing the ablation of the turbine guide vane 1. Optionally, the fuel is fuel oil. Optionally, the fuel is liquid hydrogen.
According to the turbine guide vane 1, in a multistage turbine (the turbine guide vane 1 and a turbine rotor blade form a turbine grade), because the temperature is gradually reduced from the first stage to the last stage, the temperature born by the first stage turbine is highest, the temperature of the subsequent turbine stage is relatively low, the first stage turbine becomes a main factor for limiting the temperature increase of a combustion chamber outlet, and the subsequent turbine stages have the characteristic of larger temperature margin, the fuel is subjected to heat absorption and gasification on the inner surface of a cooling cavity 2 to form fuel steam by arranging a premixed gas injection hole 3 communicated with the cooling cavity 2 on the turbine guide vane 1 and injecting the fuel from the cooling cavity 2 to the inner wall surface of the cooling cavity 2, so that the fuel steam is mixed with cold air in the cooling cavity 2 to form premixed gas and is injected into high-temperature and high-pressure gas around the turbine guide vane 1 from the premixed gas injection hole 3 to be combusted, the energy added by fuel combustion is increased in the turbine guide vane 1, the temperature margin of the subsequent turbine is fully utilized, the output power is greatly increased, and the power-weight ratio of a gas turbine engine is increased. Further, the fuel injected onto the inner wall surface of the cooling cavity 2 absorbs the heat of the turbine guide vane 1, reducing the metal temperature of the turbine guide vane 1 itself, thereby protecting the turbine guide vane 1 from burning.
The cooling cavity 2 is divided into a cooling front cavity 21 and a cooling rear cavity 22 along the direction from the front edge to the rear edge, and the premixed air injection holes 3 are arranged on the wall body of the cooling front cavity 21. The cooling forechamber 21 is closer to the incoming flow direction of the high-temperature and high-pressure gas, the temperature of the wall body of the cooling forechamber 21 is higher, the premixed gas injection holes 3 are arranged on the wall body of the cooling forechamber 21, so that the fuel is injected onto the inner wall surface of the cooling forechamber 21 to be quickly gasified in a heat absorption mode to form fuel steam, meanwhile, the wall body of the cooling forechamber 21 is cooled, and then the premixed gas is mixed with the cold air in the cooling forechamber 21 to form premixed gas, and then the premixed gas is injected into the high-temperature and high-pressure gas flowing around the turbine guide vane 1 from the premixed gas injection holes 3 to be combusted. Alternatively, the cooling cavity 2 is divided into a plurality of cooling compartments in the leading edge to trailing edge direction, and the premix spray holes 3 are located on the wall of the cooling compartment closest to the leading edge.
The wall body of the cooling back cavity 22 is provided with a cooling film hole 4 and is positioned on the pressure surface of the turbine guide vane 1, so that the cold air in the cooling back cavity 22 is sprayed out from the cooling film hole 4 and forms a cooling film on the pressure surface. The included angle between the cold air flow direction flowing out from the cooling air film hole 4 and the high-temperature high-pressure gas flow direction is smaller than 90 degrees, so that the cold air flow sprayed out from the cooling air film hole 4 forms a cooling air film covered on the pressure surface under the action of the high-temperature high-pressure gas flow. The outer surface of the turbine guide vane 1 is burnt with the premixed gas in the combustion gas by the cooling gas film.
In the present embodiment, the premix gas injection holes 3 are located at the leading edge, close to the pressure surface. The leading edge of the turbine guide vane 1 is firstly contacted with high-temperature and high-pressure combustion gas discharged from a combustion chamber, the temperature is highest, and the fuel sprayed from the cooling cavity 2 absorbs the heat at the leading edge to be quickly gasified into fuel steam, so that the temperature at the leading edge of the turbine guide vane 1 is reduced. The high-temperature high-pressure gas flows into one side of the pressure surface, and the premixed gas injection holes 3 are close to the pressure surface, so that the sprayed premixed gas quickly flows into the high-temperature high-pressure gas for combustion.
The premixed gas injection holes 3 are arranged in the direction from the root to the tip of the turbine guide vane 1. The premixed fuel gas in the cooling cavity 2 is sprayed from the plurality of premixed gas spraying holes 3, and is distributed more uniformly, the combustion is more sufficient, and the combustion efficiency is higher. The multiple rows of premixed gas injection holes 3 are arranged in the direction extending from the front edge to the pressure surface. The sprayed premixed gas is uniformly distributed in the high-temperature high-pressure gas flowing into the pressure surface, so that the combustion is more sufficient, and the combustion efficiency is higher. In this embodiment, two rows of premix gas injection holes 3 are provided at the leading edge of the turbine guide vane. One row of premixed gas injection holes 3 is located in the middle of the front edge, and one row of premixed gas injection holes 3 is located on the side of the front edge connected with the pressure surface.
The turbine guide vane 1 further includes a fuel injection member extending into the cooling cavity 2 and connected to the fuel supply line for injecting fuel to the inner wall surface of the cooling cavity 2. As shown in fig. 3, in the present embodiment, the fuel injection member is a hollow rod-structured fuel injection rod 5, and a fuel injection hole 51 for injecting fuel in the hollow chamber is provided on an inward extension of the fuel injection rod 5 extending into the cooling cavity 2. The fuel injection rod 5 is simple in structure, facilitates machining of the fuel injection hole 51, and is easy to install. Optionally, the fuel injection member employs a fuel nozzle.
The fuel injection holes 51 are arranged opposite to the premixed gas injection holes 3; and/or the fuel injection hole 51 is located in the peripheral region of the premixed gas injection hole 3 to inject fuel onto the inner wall surface around the premixed gas injection hole 3. A plurality of fuel injection holes 51 are arranged in the axial direction of the fuel injection rod 5, and a plurality of rows of fuel injection holes 51 are arranged in the circumferential direction of the fuel injection rod 5. In the present embodiment, three rows of fuel injection holes 51 are provided on the inward extending section of the fuel injection rod 5, and the three rows of fuel injection holes 51 are respectively directed toward the middle of the leading edge and both sides connected to the pressure surface and the suction surface, respectively.
The gas turbine engine of the embodiment comprises a gas compressor, a combustion chamber and a turbine part, wherein the turbine part comprises a plurality of stages of turbines which are arranged along the direction of gas flow, one turbine stage comprises a guide turbine and a rotor turbine which are arranged along the direction of gas flow, and the guide turbine is provided with the turbine guide blade 1. After the air is compressed by the compressor, the pressure is raised, the temperature is raised, and the air flows into the combustion chamber. The high-pressure air is mixed with the sprayed fuel in the combustion chamber to be combusted to form high-temperature and high-pressure fuel gas, the temperature is greatly increased, and the heat energy contained in the fuel gas is increased. The high-temperature and high-pressure gas discharged from the combustion chamber flows into the multi-stage turbine in sequence. The fuel is added into the turbine guide vane 1 in the multi-stage turbine for supplementary combustion, and the multi-stage turbine is close to the temperature margin, so that the output power of the gas turbine engine is improved, and the power-to-weight ratio is improved. Alternatively, the amount of fuel injected into the cooling cavity 2 in the turbine guide vane 1 in each stage of the turbine is controlled according to the difference between the temperature of the multistage turbine and the temperature margin, respectively.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (7)
1. A gas turbine engine comprises a gas compressor, a combustion chamber and a turbine part, and is characterized in that the turbine part comprises a plurality of stages of turbines which are arranged along the airflow direction, one turbine stage comprises a guide turbine and a rotor turbine which are arranged along the airflow direction, a turbine guide vane (1) is arranged on the guide turbine, a cooling cavity (2) is arranged in the turbine guide vane (1), the turbine guide vane (1) comprises a front edge, a tail edge, a pressure surface and a suction surface,
the turbine guide vane (1) is provided with a premixed gas injection hole (3) communicated with the cooling cavity (2),
injecting fuel from the cooling cavity (2) to the inner wall surface of the cooling cavity (2) to ensure that the fuel absorbs heat and is gasified on the inner wall surface of the cooling cavity (2), further mixing the fuel with cold air in the cooling cavity (2) to form premixed gas, and injecting the premixed gas from the premixed gas injection hole (3) to high-temperature and high-pressure gas around the turbine guide vane (1) for combustion;
controlling the quantity of fuel injected into a cooling cavity (2) in a turbine guide vane (1) in each stage of turbine according to the temperature of the multistage turbine and the difference between the temperature margins;
the premixed gas injection hole (3) is positioned at the front edge and is close to the pressure surface;
the premixed gas injection holes (3) are arranged along the direction from the blade root to the blade top of the turbine guide blade (1), and the premixed gas injection holes (3) in multiple rows are arranged in the direction extending from the front edge to the pressure surface.
2. The gas turbine engine of claim 1,
the cooling cavity (2) is divided into a front cooling cavity (21) and a rear cooling cavity (22) along the direction from the front edge to the tail edge, and the premixed gas injection holes (3) are formed in the wall body of the front cooling cavity (21).
3. The gas turbine engine of claim 2,
and a cooling air film hole is formed in the wall body of the cooling rear cavity (22) and is positioned on the pressure surface of the turbine guide blade (1), so that the cold air in the cooling rear cavity (22) is sprayed out from the cooling air film hole (4) and forms a cooling air film on the pressure surface.
4. The gas turbine engine of claim 1,
the turbine guide vane (1) further comprises a fuel injection member extending into the cooling cavity (2) and connected to the fuel supply line for injecting fuel to the inner wall surface of the cooling cavity (2).
5. The gas turbine engine of claim 4,
the fuel injection piece is a fuel injection rod (5) with a hollow rod structure, and a fuel injection hole (51) used for injecting fuel in the hollow cavity is arranged on an inward extending section of the fuel injection rod (5) extending into the cooling cavity (2).
6. The gas turbine engine of claim 5,
the fuel injection hole (51) and the premixed gas injection hole (3) are oppositely arranged; and/or the fuel injection hole (51) is located in a peripheral region of the premixed gas injection hole (3) to inject fuel onto an inner wall surface around the premixed gas injection hole (3).
7. The gas turbine engine of claim 6,
a plurality of fuel injection holes (51) are arranged in the axial direction of the fuel injection rod (5), and a plurality of rows of fuel injection holes (51) are arranged in the circumferential direction of the fuel injection rod (5).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110417270.2A CN113202566B (en) | 2021-04-19 | 2021-04-19 | Turbine guide vane and gas turbine engine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110417270.2A CN113202566B (en) | 2021-04-19 | 2021-04-19 | Turbine guide vane and gas turbine engine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113202566A CN113202566A (en) | 2021-08-03 |
| CN113202566B true CN113202566B (en) | 2022-12-02 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202110417270.2A Active CN113202566B (en) | 2021-04-19 | 2021-04-19 | Turbine guide vane and gas turbine engine |
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Family Cites Families (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| ES2288687T3 (en) * | 2003-07-04 | 2008-01-16 | Siemens Aktiengesellschaft | OPEN REFRIGERATION COMPONENT FOR A GAS TURBINE, COMBUSTION CHAMBER AND GAS TURBINE. |
| JP4412081B2 (en) * | 2004-07-07 | 2010-02-10 | 株式会社日立製作所 | Gas turbine and gas turbine cooling method |
| DE102009009129B4 (en) * | 2009-02-17 | 2022-11-03 | BMTS Technology GmbH & Co. KG | Turbocharger with variable turbine geometry |
| US8387245B2 (en) * | 2010-11-10 | 2013-03-05 | General Electric Company | Components with re-entrant shaped cooling channels and methods of manufacture |
| CN103266922B (en) * | 2013-06-15 | 2014-11-12 | 厦门大学 | Turbine stator blade with interstage combustor |
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