CN113932251B - Gas turbine annular combustion chamber based on detonation combustion - Google Patents
Gas turbine annular combustion chamber based on detonation combustion Download PDFInfo
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- CN113932251B CN113932251B CN202111402024.6A CN202111402024A CN113932251B CN 113932251 B CN113932251 B CN 113932251B CN 202111402024 A CN202111402024 A CN 202111402024A CN 113932251 B CN113932251 B CN 113932251B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/52—Toroidal combustion chambers
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Abstract
The invention discloses a gas turbine annular combustion chamber based on detonation combustion, which mainly comprises a combustion chamber air inlet cavity, a rotary detonation combustion chamber, an annular combustion cavity, a cooling air central channel and an air film cooling hole. Because the detonation combustion has the advantages of high thermal cycle efficiency and high heat release rate, the rotary detonation combustor and the gas turbine annular combustor are combined, on one hand, the thermal cycle efficiency of the combustor is increased, and the volume and the mass of the combustor are reduced; on the other hand, the plurality of RDC combustion chambers are arranged at the head part of the combustion chamber, so that the working stability of the combustion chamber is increased, and the uniformity of the pressure and the temperature distribution of the burnt gas at the outlet of the combustion chamber is improved. On the premise of not obviously increasing the structural complexity of the combustion chamber, the invention optimizes the structural form of the combustion chamber of the gas turbine, has reasonable design, improves the working stability of the combustion chamber of the gas turbine and simultaneously reduces the volume of the combustion chamber.
Description
Technical Field
The invention belongs to the technical field of gas turbine combustors, and particularly relates to a gas turbine annular combustor based on detonation combustion.
Background
The gas turbine is widely applied to a plurality of fields such as ship power, power generation and oil exploitation, and plays a very important role in a modern industrial system. The compressor, the combustor and the turbine are three basic parts of the gas turbine, wherein the combustor of the gas turbine usually adopts a slow combustion mode based on isobaric cycle, the technical maturity is high, and the improvement of the thermal cycle efficiency is very difficult by further optimizing the structure of the combustor.
Disclosure of Invention
The invention aims to provide a gas turbine annular combustion chamber based on detonation combustion, which aims to improve the thermal cycle efficiency of the gas turbine combustion chamber, reduce the structural size of the combustion chamber, lighten the mass of the combustion chamber and improve the flexibility of the combustion chamber as a power device. The invention can be applied to the fields of ship gas turbines, industrial gas turbines and the like.
In order to achieve the purpose, the invention is realized by adopting the following technical scheme:
the annular combustion chamber of the gas turbine based on the detonation combustion comprises a combustion chamber air inlet cavity, an RDC combustion chamber, an annular combustion cavity, a cooling air central channel and an air film cooling hole;
the combustion chamber air inlet cavity is positioned at the head part of the combustion chamber and consists of a combustion chamber air inlet cavity outer ring and a combustion chamber air inlet cavity inner ring, the combustion chamber air inlet cavity outer ring is of a hollow round platform structure, the left side, namely the inlet of the combustion chamber air inlet cavity, is connected with the outlet of a gas turbine compressor, and air pressurized by the gas compressor enters the combustion chamber from the combustion chamber air inlet cavity; the inner ring of the combustion chamber air inlet cavity is of a hollow round table structure and is distributed along the same axis as the outer ring of the combustion chamber air inlet cavity; an airflow channel between the outer ring and the inner ring of the combustion chamber air inlet cavity is a combustion chamber air inlet ring cavity, and the outer diameter of the combustion chamber air inlet cavity is continuously increased from an inlet to an outlet;
the RDC combustion chamber is positioned on one side of an outlet of an air inlet cavity of the combustion chamber and consists of an RDC cover plate, swirler vanes, a swirler center cone and an RDC outer ring; the RDC cover plate is positioned at the head part of the RDC combustion chamber and is of a cylindrical thin plate structure; the outer edge of the RDC cover plate is connected with the inner ring of the combustion chamber air inlet cavity; compared with the smaller radius position of the mounting hole, 12 swirler air channels are reserved and are uniformly distributed along the circumferential direction of the combustion chamber, and swirler vanes and a swirler central cone are positioned in the swirler air channels; the RDC outer ring is positioned on the right side of the RDC cover plate, is fixedly connected with the RDC cover plate and is simultaneously installed in a rotary detonation combustion cavity installation hole reserved in an air inlet cavity inner ring of the combustion chamber, the RDC outer ring is of a circular ring structure, and a central cylindrical cavity is an RDC combustion cavity;
the annular combustion cavity is positioned on one side of the RDC combustion cavity outlet and consists of an annular combustion cavity outer ring and an annular combustion cavity inner ring, the combustion chamber air inlet cavity outer ring and the combustion chamber air inlet cavity inner ring are distributed along the same axis, an air flow channel between the combustion chamber air inlet cavity outer ring and the combustion chamber inner ring is an annular combustion cavity, and the outer diameter of the combustion cavity is continuously reduced from the annular combustion cavity inlet to the annular combustion cavity outlet; one side of the outlet of the combustion cavity is connected with a first-stage nozzle ring of a turbine of the gas turbine, and high-temperature gas enters the turbine blade through the outlet of the annular combustion cavity;
the cooling gas central channel is of a hollow cylindrical structure, is positioned in the center of the combustion chamber and close to one side of the annular combustion chamber, is provided with a cooling gas central supply hole, and supplies cooling gas required by the combustion chamber through the cooling gas central supply hole;
the film cooling holes are of straight circular hole structures and are respectively distributed on the wall surfaces of the RDC outer ring, the annular combustion cavity outer ring and the annular combustion cavity inner ring.
The invention has the further improvement that 12 screw mounting holes are reserved at the outer edge of the RDC cover plate, are uniformly distributed along the circumferential direction of the combustion chamber and are connected with the inner ring of the air inlet cavity of the combustion chamber through screws.
The invention has the further improvement that the combustion chamber air inlet cavity is an annular cavity channel with a variable cross section, the head part of the combustion chamber air inlet cavity is connected with the outlet of the gas compressor, and the outer diameter of the head part of the combustion chamber air inlet cavity is smaller; the outlet of the combustion chamber air inlet cavity is connected with the RDC combustion cavity which is distributed along the circumferential direction and has a larger radial position, and the outer diameter of the outlet of the combustion chamber air inlet cavity is larger; the inner wall surface of the combustion chamber air inlet cavity outer ring and the outer wall surface of the combustion chamber air inlet cavity inner ring are in smooth transition through curved surfaces, and flow loss of gas is reduced.
The invention has the further improvement that the RDC combustion chambers are uniformly distributed along the circumferential direction of the combustion chamber, the number of the RDC combustion chambers is 20, and the structures of all the RDC combustion chambers are completely the same; as for a single RDC combustion chamber, the single RDC combustion chamber is of an empty barrel-shaped combustion chamber structure consisting of an RDC cover plate and an RDC outer ring, 12 swirler air channels are located on the RDC cover plate and close to one side of the outer edge and are evenly distributed along the circumferential direction of the RDC combustion chamber, the swirler air channels are of circular through hole structures, the direction of the central axis is consistent with the axial direction of the combustion chamber, and a tapered profile structure is adopted on one side of an inlet, so that air can smoothly flow in the channel along the axial direction.
A further development of the invention is that the angle of the tapered profile to the axis of the combustion chamber is 15 °.
The invention has the further improvement that the swirler vanes are positioned in the swirler air passage, the number of the swirler vanes is 12, and the swirl vanes are uniformly distributed along the circumferential direction of the swirler air passage; at the inlet of the swirler vane, the direction of the swirler vane is consistent with the axial direction of the combustion chamber; the angle of the blades is changed continuously from the inlet to the outlet of the swirler vane, the included angle between the direction of the swirler vane and the axis of the combustion chamber is 45 degrees at the outlet of the swirler vane, a swirler fuel cavity is arranged on the inner side of the swirler vane, the external profile is in a football shape, the flow resistance of air flowing through the swirler fuel cavity can be reduced, and the internal part of the swirler vane is in a cylindrical cavity structure; the swirler fuel spray holes are positioned on one side of the swirler fuel cavity close to the outlet, 12 swirler fuel spray holes are uniformly distributed along the circumference of the swirler fuel cavity, the swirler fuel spray holes are of a circular through hole structure, and the included angle between the axis direction of the through holes and the axis of the combustion chamber is 60 degrees.
The invention has the further improvement that the annular combustion chamber is divided into three parts, namely an inlet straight circular ring section, a transition section and an outlet straight circular ring section; the central radial position of the inlet straight circular ring section is the same as the central radial position of the RDC combustion chamber; the outlet straight circular ring section is used for being connected with a first-stage stationary blade ring of the turbine; because the diameter of the first stage stationary blade ring of the turbine is small, a transition section structure is adopted between the inlet straight circular ring section and the outlet straight circular ring section, and the wall surface of the transition section is a smooth circular arc profile.
The invention has the further improvement that an RDC outer ring cooling cavity and an annular combustion chamber cooling cavity are respectively arranged inside the RDC outer ring, the annular combustion cavity outer ring and the annular combustion cavity inner ring, a gas film cooling hole is arranged between the cooling cavity and the combustion cavity, cooling gas is supplied to the combustion cavity for cooling through the gas film cooling hole, the cooling gas film hole is of a circular through hole structure, the included angle between the axis direction of the cooling gas film hole and the axis of the combustion cavity is 45 degrees, and the jet flow direction of the gas film hole is along the counter flow direction of the burnt gas.
The invention has at least the following beneficial technical effects:
the invention provides a gas turbine annular combustion chamber based on detonation combustion, which combines an RDC (radial fuel cell) with the gas turbine annular combustion chamber, wherein fuel and air are mainly combusted in the RDC, and combusted gas enters the annular combustion chamber. Firstly, due to the self-pressurization effect of the detonation fuel, the number of stages of the gas compressor can be reduced under the condition of the same output power; the thermal cycle efficiency of the gas turbine can be improved under the condition of the same pressure ratio. And secondly, because the heat release rate of the detonation combustion is high, the volume of the combustion chamber of the gas turbine can be reduced, the quality of the combustion chamber is reduced, and the flexibility of the power device is improved. Thirdly, as the head of the combustion chamber adopts 20 RDC combustion chambers, the working stability of the combustion chamber can be improved, and the gas turbine can still work after the single RDC combustion chamber is flamed out. Finally, because the RDC outlet is of an annular combustion chamber structure, the pressure and temperature distribution of the outlet of burned gas after detonation combustion is more uniform after the burned gas flows through the annular combustion chamber. Compared with the traditional annular combustion chamber of the gas turbine, the annular combustion chamber of the gas turbine optimizes the structural form of the combustion chamber of the gas turbine, has reasonable design, can improve the working stability of the combustion chamber of the gas turbine, and simultaneously reduces the volume of the combustion chamber.
Drawings
FIG. 1 is a schematic diagram of a gas turbine annular combustor based on detonation combustion;
FIG. 2 isbase:Sub>A schematic cross-sectional view of an annular combustion chamber A-A ofbase:Sub>A gas turbine based on detonation combustion;
FIG. 3 is a schematic view of an RDC combustion chamber configuration, wherein FIG. 3 (b) is a schematic view of a section C-C of FIG. 3 (a);
FIG. 4 is a schematic view of a swirler vane and fuel supply cavity configuration, wherein FIG. 4 (a) is a perspective view and FIG. 4 (b) is a front view;
FIG. 5 is a schematic cross-sectional view of swirler vanes and fuel supply cavity D-D;
FIG. 6 is a schematic view of a film cooling hole configuration;
FIG. 7 is a schematic view of a gas turbine annular combustor inner ring structure, wherein FIG. 7 (B) is a schematic view of a section B-B of FIG. 7 (a).
Description of reference numerals:
1 is an outer ring of a combustion chamber air inlet cavity, 2 is an inner ring of the combustion chamber air inlet cavity, 3 is the combustion chamber air inlet cavity, 4 is an RDC combustion chamber, 5 is an annular combustion cavity, 6 is an outer ring of the annular combustion cavity, 7 is an inner ring of the annular combustion cavity, 8 is a flange plate, 9 is a mounting hole of a rotary detonation combustion cavity, 10 is a cooling air supply hole, 11 is a cooling air supply cavity, 12 is a cooling air film hole, 13 is a cooling air central channel cover, 14 is a cooling air central channel, 15 is a central cooling gas supply hole, 4-1 is an RDC combustion cavity, 4-2 is an RDC outer ring, 4-3 is a swirler vane, 4-4 is a swirler central cone, 4-5 is a swirler fuel spray hole, 4-6 is a swirler fuel cavity, 4-7 is a swirler air channel, 4-8 is an RDC cover plate, 4-9 is a bolt mounting hole, 4-10 is an RDC outer ring cooling cavity, 4-11 is an RDC cooling gas film hole, and 4-12 is an RDC igniter mounting hole.
Detailed Description
In the description of the present invention, it should be noted that, unless otherwise explicitly specified or limited, the terms "mounted," "connected," and "connected" are to be construed broadly, e.g., as meaning either a fixed connection, a removable connection, or an integral connection; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.
Referring to fig. 1 and fig. 2, the annular combustion chamber of the gas turbine based on Detonation combustion provided by the invention is composed of a combustion chamber air inlet cavity 3, a Rotary Detonation Combustor (RDC) 4, an annular combustion cavity 5, a cooling gas central channel housing 13 and a cooling gas film hole 12.
The combustion chamber air inlet cavity 3 is positioned at the leftmost end of the annular combustion chamber of the gas turbine, is an annular cavity-shaped channel consisting of an outer combustion chamber air inlet cavity ring 1 and an inner combustion chamber air inlet cavity ring 2, and is used for supplying high-pressure air at the outlet of the compressor into the RDC combustion chamber 4. The inlet of the combustion cavity air inlet cavity 3 is connected with the outlet of the compressor, and the outlet of the combustion cavity air inlet cavity is connected with the RDC combustion chamber 4; because the diameter of the compressor outlet is smaller than the diameter of the RDC combustion chamber 4, the outer diameter of the annular cavity-shaped channel is continuously increased from the inlet to the outlet of the combustion chamber air inlet annular cavity 3. The outer ring 1 of the combustion chamber air inlet cavity is of a hollow round table structure, the left end of the outer ring is connected with an outlet of a gas compressor of the gas turbine, and the right end of the outer ring is connected with the outer ring 6 of the annular combustion cavity; the combustion chamber air inlet cavity inner ring 2 is also of a hollow circular truncated cone structure and is positioned on the inner side of the combustion chamber air inlet cavity outer ring 1, the central axes of the two parts are the same, a cooling air central channel housing 13 is nested in the center of the combustion chamber air inlet cavity inner ring 2, the outer profile is a smooth curved surface, and the flow loss of gas can be effectively reduced.
Referring to fig. 3-6, the RDC combustion chamber 4 is located between the combustion chamber inlet chamber 3 and the annular combustion chamber 5, and is composed of an RDC combustion chamber 4-1, an RDC outer ring 4-2, swirler vanes 4-3, a swirler center cone 4-4, an RDC cover plate 4-8, and RDC cooling air film holes 4-11, and functions to organize air and fuel combustion and convert chemical energy into heat energy. RDC combustion chamber 4-1 is the empty bucket structure of constituteing by RDC apron 4-8 and RDC outer ring 4-2, and RDC apron 4-8 is located the leftmost end of RDC combustion chamber 4-1, is cylindrical plate structure. The RDC outer ring 4-2 is positioned outside the RDC combustion chamber and is of a circular ring structure.
Close to one side of the outer edge of the RDC cover plate 4-8, 12 swirler air channels 4-7 and 1 RDC igniter mounting hole 4-12 are reserved, the swirler air channels 4-7 are used for supplying air at the outlet of the combustion chamber air inlet cavity 3 to the RDC combustion chamber 4, the swirler air channels 4-7 are all in a through hole structure, a contraction molded surface is adopted at an inlet, smooth flowing of the air in the channels along the axial direction is facilitated, and an included angle between the contraction molded surface and the axis of the combustion chamber is 15 degrees. The interior of each swirler air channel 4-7 is provided with swirler vanes 4-3 and a swirler central cone 4-4; the swirler vanes 4-3 are positioned between a swirler air channel 4-7 and a swirler central cone 4-4, the number of the swirler vanes 4-3 is 12, the swirler vanes are uniformly distributed along the circumferential direction of the swirler air channel 4-7, the swirler vanes are used for guiding air entering the swirler air channel 4-7 to generate swirling flow, the direction of the vanes at an inlet of the swirler vanes 4-3 is the same as the axial direction of the combustion chamber, and the included angle between the direction of the vanes at an outlet and the axial direction of the combustion chamber is 45 degrees; the swirler central cone 4-4 is positioned at the center of a swirler air channel 4-7 and is of a rugby-ball structure, so that the flow loss of gas can be effectively reduced, the swirler fuel cavity 4-6 is positioned inside the swirler central cone 4-4 and is of a cylindrical cavity structure, 12 swirler fuel spray holes 4-5 distributed along the circumferential direction are reserved on the right side close to the swirler fuel cavity 4-6 and are of a circular through hole structure, and the included angle between the axial direction of the through hole and the axial direction of a combustion chamber is 60 degrees.
The RDC outer ring 4-2 is nested in the mounting hole 9 of the rotary detonation combustion chamber, and an outer ring cooling chamber 4-10 is designed inside the RDC outer ring and is of an annular cavity structure and used for supplying cooling air required by cooling of the RDC combustion chamber 4. On one side of an outer ring cooling cavity 4-10 close to the inner wall surface of an RDC outer ring 4-2, RDC cooling air film holes 4-11 are designed, 18 rows are uniformly distributed along the circumferential direction of a combustion cavity, 11 rows are uniformly distributed along the axial direction of the combustion cavity, namely the total number is 198, the RDC cooling air film holes 4-11 are of a circular through hole structure, the included angle between the axial direction of the RDC cooling air film holes and the axial direction of the combustion cavity is 45 degrees, and the RDC cooling air film holes can achieve a better cooling effect along the counter-flow direction of burnt gas.
Referring to fig. 1, 2 and 7, the annular combustion chamber 5 is located at the rightmost end of the annular combustion chamber of the gas turbine, and is an annular chamber structure composed of an annular combustion chamber outer ring 6 and an annular combustion chamber inner ring 7, and is used for conveying high-temperature and high-pressure burned gas generated in the RDC combustion chamber 4 to the turbine; the inlet is connected with the outlet of the RDC combustion chamber 4-1, the outlet is connected with the first stage stationary blade ring of the turbine, and the diameter of the first stage stationary blade ring of the turbine is obviously smaller than that of the position of the RDC combustion chamber 4-1, so that a smooth transition section is designed from the inlet to the outlet of the annular combustion chamber, and the flow loss of the combusted gas can be effectively reduced.
The leftmost end of the annular combustion cavity inner ring 7 is provided with a flange plate 8, 20 rotary detonation combustion cavity mounting holes 9 are uniformly distributed on the flange plate 8 along the circumferential direction, and 20 RDC outer rings 4-2 are respectively nested in the corresponding rotary detonation combustion cavity mounting holes 9 and are connected with each other through bolt mounting holes 4-9.
A cooling air supply cavity 11 with a ring cavity-shaped structure is designed in each of the annular combustion cavity outer ring 6 and the annular combustion cavity inner ring 7, and is used for supplying cooling air to a cooling air film hole 12; the cooling gas is supplied from the cooling gas supply hole 10; on the inboard of annular combustion chamber outer loop 6 and the outside wall of annular combustion chamber inner ring 7, all along 22 rows of axial equipartitions, along 30 rows of circumference equipartitions, the total is the gas film cooling hole of 660 circular through-hole structures, and its axis direction is 45 with the axial contained angle of combustion chamber, and along the counterflow direction of burnt gas, can reach better cooling effect.
Referring to fig. 1 and 2, a cooling gas central passage housing 13 is located at the center of the annular combustion chamber of the gas turbine, and has a hollow circular tube structure, and a cooling gas central passage 14 is formed inside the housing, and is used for supplying cooling gas to the cooling gas supply chamber 11 through a central cooling gas supply hole 15 and a cooling gas supply hole 10, respectively.
1 is an outer ring of a combustion chamber air inlet cavity, 2 is an inner ring of the combustion chamber air inlet cavity, 3 is the combustion chamber air inlet cavity, 4 is an RDC combustion chamber, 5 is an annular combustion cavity, 6 is an outer ring of the annular combustion cavity, 7 is an inner ring of the annular combustion cavity, 8 is a flange plate, 9 is a mounting hole of a rotary detonation combustion cavity, 10 is a cooling air supply hole, 11 is a cooling air supply cavity, 12 is a cooling air film hole, 13 is a cooling air central channel cover, 14 is a cooling air central channel, 15 is a central cooling gas supply hole, 4-1 is an RDC combustion cavity, 4-2 is an RDC outer ring, 4-3 is a swirler vane, 4-4 is a swirler central cone, 4-5 is a swirler fuel spray hole, 4-6 is a swirler fuel cavity, 4-7 is a swirler air channel, 4-8 is an RDC cover plate, 4-9 is a bolt mounting hole, 4-10 is an RDC outer ring cooling cavity, 4-11 is an RDC cooling gas film hole, and 4-12 is an RDC igniter mounting hole.
The working cycle process of the invention is as follows:
the invention relates to a gas turbine annular combustion chamber based on detonation combustion. When the combustion chamber starts to work, firstly, the air compressor is driven to work through a jigger, high-pressure air generated by the air compressor is supplied to the RDC combustion chamber 4 through the combustion chamber air inlet cavity 3, a fuel supply valve is opened, and fuel is supplied to the RDC combustion chamber 4 through a swirler fuel cavity 4-6; meanwhile, cooling gas is supplied to the RDC cooling film holes 4-11 and the cooling film holes 12 of the RDC combustion chamber 4-1 and the annular combustion chamber 5 respectively through the RDC outer ring cooling chamber 4-10 and the cooling gas supply chamber 11. When sufficient fuel and air reactants are filled in the RDC combustion cavity 4-1, an igniter in an RDC igniter mounting hole 4-12 is started, a stably propagated slow combustion wave is formed in the RDC combustion cavity 4-1 after ignition, and the slow combustion wave gradually develops into a stably propagated rotary detonation wave under the turbulent flow action of the wall surface of the combustion cavity and the RDC cooling air film hole 4-11 in the circumferential propagation process of the RDC combustion cavity 4-1; high-temperature and high-pressure burnt gas generated by slow combustion or detonation combustion is conveyed to a turbine for doing work through an annular combustion cavity 5, further, shaft work is generated to drive a gas compressor and a generator to rotate, and an annular combustion chamber of the gas turbine enters a stable working stage.
When the work of the combustion chamber is finished, the fuel supply valve is closed, the fuel cannot be continuously supplied to the RDC combustion chamber 4, the continuous propagation of the rotary detonation wave is maintained due to the lack of combustible mixtures in the RDC combustion chamber 4-1, the rotary detonation wave is decoupled and changed into a slow combustion wave, then the flame is gradually extinguished, and meanwhile, the supply of cooling gas in the cooling gas central channel housing 13 is closed, and the work of the combustion chamber is finished.
It will be evident to those skilled in the art that the invention is not limited to the details of the foregoing illustrative embodiments, and that the present invention may be embodied in other specific forms without departing from the spirit or essential attributes thereof. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive, the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be embraced therein. Any reference sign in a claim should not be construed as limiting the claim concerned.
Claims (4)
1. The annular combustion chamber of the gas turbine based on the detonation combustion is characterized by comprising a combustion chamber air inlet cavity, an RDC combustion chamber, an annular combustion cavity, a cooling air central channel and an air film cooling hole;
the combustion chamber air inlet cavity is positioned at the head part of the combustion chamber and consists of a combustion chamber air inlet cavity outer ring and a combustion chamber air inlet cavity inner ring, the combustion chamber air inlet cavity outer ring is of a hollow round platform structure, the left side, namely the inlet of the combustion chamber air inlet cavity, is connected with the outlet of a gas turbine compressor, and air pressurized by the gas compressor enters the combustion chamber from the combustion chamber air inlet cavity; the inner ring of the combustion chamber air inlet cavity is of a hollow round table structure and is distributed along the same axis as the outer ring of the combustion chamber air inlet cavity; an airflow channel between the outer ring and the inner ring of the combustion chamber air inlet cavity is a combustion chamber air inlet ring cavity, and the outer diameter of the combustion chamber air inlet cavity is continuously increased from an inlet to an outlet;
the RDC combustion chamber is positioned on one side of an outlet of an air inlet cavity of the combustion chamber and consists of an RDC cover plate, swirler vanes, a swirler center cone and an RDC outer ring; the RDC cover plate is positioned at the head part of the RDC combustion chamber and is of a cylindrical thin plate structure; the outer edge of the RDC cover plate is connected with the inner ring of the combustion chamber air inlet cavity; compared with the position with a smaller radius of the mounting hole, 12 swirler air channels are reserved and are uniformly distributed along the circumferential direction of the combustion chamber, and swirler vanes and a swirler center cone are positioned in the swirler air channels; the RDC outer ring is positioned on the right side of the RDC cover plate, is fixedly connected with the RDC cover plate and is simultaneously installed in a rotary detonation combustion cavity installation hole reserved in an air inlet cavity inner ring of the combustion chamber, the RDC outer ring is of a circular ring structure, and a central cylindrical cavity is an RDC combustion cavity;
the annular combustion cavity is positioned on one side of the RDC combustion cavity outlet and consists of an annular combustion cavity outer ring and an annular combustion cavity inner ring, the combustion chamber air inlet cavity outer ring and the combustion chamber air inlet cavity inner ring are distributed along the same axis, an air flow channel between the combustion chamber air inlet cavity outer ring and the combustion chamber inner ring is an annular combustion cavity, and the outer diameter of the combustion cavity is continuously reduced from the annular combustion cavity inlet to the annular combustion cavity outlet; one side of the outlet of the combustion cavity is connected with a first-stage nozzle ring of a turbine of the gas turbine, and high-temperature gas enters the turbine blade through the outlet of the annular combustion cavity;
the cooling gas central channel is of a hollow cylindrical structure, is positioned in the center of the combustion chamber and close to one side of the annular combustion chamber, is provided with a cooling gas central supply hole, and supplies cooling gas required by the combustion chamber through the cooling gas central supply hole;
the film cooling holes are of straight circular hole structures and are respectively distributed on the wall surfaces of the RDC outer ring, the annular combustion cavity outer ring and the annular combustion cavity inner ring;
the combustion chamber air inlet cavity is an annular cavity channel with a variable cross section, the head of the combustion chamber air inlet cavity is connected with the outlet of the compressor, and the outer diameter of the head of the combustion chamber air inlet cavity is smaller; the outlet of the combustion chamber air inlet cavity is connected with the RDC combustion cavity which is distributed along the circumferential direction and has a larger radial position, and the outer diameter of the outlet of the combustion chamber air inlet cavity is larger; the inner wall surface of the outer ring of the combustion chamber air inlet cavity and the outer wall surface of the inner ring of the combustion chamber air inlet cavity are in smooth transition by adopting curved surfaces, so that the flow loss of gas is reduced;
the RDC combustion chambers are uniformly distributed along the circumferential direction of the combustion chamber, the number of the RDC combustion chambers is 20, and the structures of the RDC combustion chambers are completely the same; for a single RDC combustion chamber, the single RDC combustion chamber is of an empty barrel-shaped combustion chamber structure consisting of an RDC cover plate and an RDC outer ring, 12 swirler air channels are located on the RDC cover plate and close to one side of the outer edge and are uniformly distributed along the circumferential direction of the RDC combustion chamber, the swirler air channels are of circular through hole structures, the direction of the central axis is consistent with the axial direction of the combustion chamber, and a tapered profile structure is adopted on one side of an inlet, so that air can smoothly flow in the channels along the axial direction;
the swirler vanes are positioned in the swirler air channel, the number of the swirler vanes is 12, and the swirler vanes are uniformly distributed along the circumferential direction of the swirler air channel; at the inlet of the swirler vane, the direction of the swirler vane is consistent with the axial direction of the combustion chamber; the angle of the blades is constantly changed from the inlet to the outlet of the swirler, the included angle between the direction of the swirl blades and the axis of the combustion chamber is 45 degrees at the outlet of the swirler blades, a swirler fuel cavity is arranged on the inner side of the swirler blades, the outer profile is in the shape of a football, the flow resistance of air flowing through the swirler fuel cavity can be reduced, and the inner part of the swirler fuel cavity is in a cylindrical cavity structure; the swirler fuel spray holes are positioned at one side of the swirler fuel cavity close to the outlet, 12 swirler fuel spray holes are uniformly distributed along the circumference of the swirler fuel cavity and are of a circular through hole structure, and the included angle between the axis direction of the through holes and the axis of the combustion chamber is 60 degrees;
the RDC outer ring cooling cavity and the annular combustion chamber cooling cavity are respectively arranged in the RDC outer ring, the annular combustion cavity outer ring and the annular combustion cavity inner ring, air film cooling holes are arranged between the cooling cavity and the combustion cavity, cooling gas is supplied to the combustion cavity through the air film cooling holes to be cooled, the cooling air film holes are of a circular through hole structure, the included angle between the axis direction of the cooling air film holes and the axis of the combustion cavity is 45 degrees, and the jet flow direction of the air film holes is along the counter flow direction of the burnt gas.
2. The annular combustion chamber of the gas turbine based on the detonation combustion as claimed in claim 1, wherein 12 bolt mounting holes are reserved at the outer edge of the RDC cover plate, are uniformly distributed along the circumferential direction of the combustion chamber, and are connected with the inner ring of the combustion chamber air inlet cavity through bolts.
3. The annular combustor of a gas turbine engine based on detonation combustion of claim 1, wherein the tapered profile is angled 15 ° from the combustor axis.
4. The annular combustion chamber of the gas turbine based on the detonation combustion as claimed in claim 1, characterized in that the annular combustion chamber is divided into three parts, namely an inlet straight circular ring section, a transition section and an outlet straight circular ring section; the central radial position of the inlet straight circular ring section is the same as the central radial position of the RDC combustion chamber; the outlet straight circular ring section is used for being connected with a first-stage stationary blade ring of the turbine; because the diameter of the first stage stationary blade ring of the turbine is small, a transition section structure is adopted between the inlet straight circular ring section and the outlet straight circular ring section, and the wall surface of the transition section is a smooth circular arc profile.
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| CN115143490B (en) * | 2022-06-15 | 2023-08-01 | 南京航空航天大学 | Combustion chamber with circumferential staggered opposite flushing jet flow and full-ring large-scale rotational flow coupling |
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| CN104153884B (en) * | 2014-08-06 | 2015-10-28 | 西安热工研究院有限公司 | A kind of rotation pinking gas turbine |
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