CN113834095B - Gas turbine annular tube type combustion chamber based on detonation combustion - Google Patents
Gas turbine annular tube type combustion chamber based on detonation combustion Download PDFInfo
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- CN113834095B CN113834095B CN202111290887.9A CN202111290887A CN113834095B CN 113834095 B CN113834095 B CN 113834095B CN 202111290887 A CN202111290887 A CN 202111290887A CN 113834095 B CN113834095 B CN 113834095B
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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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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/02—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof the jet being intermittent, i.e. pulse-jet
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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
- F23R7/00—Intermittent or explosive combustion chambers
Abstract
The invention discloses a gas turbine annular pipe type combustion chamber based on detonation combustion, which applies a rotary detonation combustion cavity to a gas turbine to replace the conventional annular pipe type isobaric combustion chamber. Due to the self-pressurization effect of detonation combustion, compared with a combustion chamber based on isobaric combustion, the thermal cycle efficiency of the gas turbine can be improved, the volume of the combustion chamber of the gas turbine is reduced, and the flexibility of a power device is improved; the invention adopts a structure of 20 annular combustion chambers which are uniformly distributed along the circumferential direction, on one hand, the gas pressure and the temperature distribution at the outlet of the combustion chamber can be ensured to be more uniform, and on the other hand, the invention is also beneficial to the maintenance and the replacement of a single combustion chamber.
Description
Technical Field
The invention belongs to the technical field of gas turbine combustors, and particularly relates to a gas turbine ring-pipe type combustor based on detonation combustion.
Background
The gas turbine is widely applied to the fields of aeroengines, ship power, power generation, oil exploitation and the like, plays a very important role in modern industrial systems, and is also an important mark for measuring the development level of national industry. The compressor, the combustor and the turbine are three basic parts of the gas turbine, wherein the combustor usually adopts a slow combustion mode based on isobaric cycle, the technical maturity is high, and it is very difficult to further improve the thermal cycle efficiency. According to the difference of the intensity of combustion, another combustion mode, namely detonation combustion, exists in nature, and the propagation characteristic of the detonation combustion is that a shock wave structure exists in front of a combustion zone, and the shock wave and the combustion zone are coupled together for propagation. Because the shock wave compresses the unburned gas, the pressure of detonation combustion is higher, and the detonation combustion is a combustion mode with the self-pressurization characteristic. Accordingly, the knocking combustion has advantages of low entropy increase and fast heat release rate.
Disclosure of Invention
The invention aims to provide a gas turbine annular tube type combustion chamber based on detonation combustion, which is intended to replace a conventional gas turbine combustion chamber based on an isobaric combustion mode so as to improve the heat cycle efficiency of a gas turbine, reduce the volume and weight 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 aircraft engines, industrial gas turbines, ship gas turbines and the like.
In order to achieve the purpose, the invention adopts the technical scheme that:
a gas turbine circular pipe type combustion chamber based on detonation combustion comprises an air supply cavity, a fuel supply cavity, a swirler, a rotary detonation combustion cavity, a burnt gas exhaust cavity and a circular pipe type combustion chamber annular channel;
the air supply cavity is positioned at the head part of the combustion chamber and is a hollow cylindrical channel, and the outer side of the air supply cavity is an outer ring of an air supply cavity with a circular ring structure; the left side of the rotary detonation combustor is provided with an installation flange, the rotary detonation combustor is connected with an outlet of a gas compressor through a bolt installation hole reserved on the installation flange, and the right side of the rotary detonation combustor is connected with a rotary detonation combustion cavity;
the fuel supply cavity is positioned in the center of the air supply cavity and is a conical channel, the outer side of the fuel supply cavity is a wall surface of the fuel supply cavity with a thin-wall conical structure, the outlet of the fuel supply cavity is a fuel supply hole with a circular hole structure, and the direction of the central axis of the circular hole is the same as the radial direction of the combustion chamber; the outlet of the fuel supply hole is a fuel supply pipeline with a circular tube-shaped structure, the fuel supply pipeline is close to one end of the fuel supply hole, the axial direction of the fuel supply pipeline is the same as the radial direction of the combustion chamber, the other end of the fuel supply pipeline is connected with the swirler, and the axial direction of the fuel supply pipeline is the same as the axial direction of the combustion chamber;
the cyclone is positioned between the air supply cavity and the rotary detonation combustion cavity, is arranged in a cyclone mounting hole at the head of the rotary detonation combustion cavity, has one inlet end connected with a fuel supply pipeline and one outlet end provided with the rotary detonation combustion cavity, and consists of a cyclone fuel cavity, a cyclone fuel cavity wall surface, a cyclone fuel supply hole and a cyclone blade, and the axial direction of the cyclone is the same as the axial direction of the combustion chamber;
the rotary detonation combustion cavity is an annular cavity-shaped channel consisting of an outer ring and an inner column of the rotary detonation combustion chamber, and the axial direction of the rotary detonation combustion cavity is the same as the axial direction of the combustion chamber; the outer ring of the rotary detonation combustion chamber is of a circular structure, a cooling gas supply annular cavity of an annular cavity structure is reserved in one side close to the rotary detonation combustion cavity, and meanwhile, two radial circular holes are reserved on one side close to the head of the outer ring of the rotary detonation combustion chamber and are respectively a flame connecting pipe mounting hole and a combustion chamber pressure monitoring hole, and the two circular holes are spaced by 90 degrees; the inner column of the rotary detonation combustion chamber is of a cylindrical structure, and a cooling gas supply annular cavity of an annular cavity structure is reserved in one side close to the rotary detonation combustion chamber; the axial lengths of the outer ring and the inner column of the rotary detonation combustion chamber are the same, the axial positions along the combustion chamber are also completely the same, and the outlet of the rotary detonation combustion chamber is a combustion chamber outlet guide cone which is of a solid conical structure;
the burned gas exhaust cavity is positioned on one side of an outlet of the rotary detonation combustion cavity, the structure of the burned gas exhaust cavity is an irregular round-to-square runner, and the outer side of the burned gas exhaust cavity is the outer wall surface of the exhaust section of the combustion chamber;
the annular channel of the annular tube type combustion chamber is of an annular structure, consists of an outer ring of the annular channel of the annular tube type combustion chamber and an inner ring of the annular channel of the annular tube type combustion chamber, is an installation channel of 20 annular tube type combustion chambers which are uniformly distributed along the circumferential direction, and is connected with the gas compressor at one end of an inlet and connected with a first-stage stationary blade ring of the turbine at one end of an outlet.
The invention has the further improvement that the outer wall surface of the fuel supply cavity is of a conical structure, the left end of the fuel supply cavity is of a pointed structure, the right end of the fuel supply cavity is of a circular structure, and the transition is smooth from the left end to the right end through an arc, so that the air can be effectively prevented from forming a backflow area at the head of the fuel supply cavity, and the flow loss of the air is reduced.
The invention is further improved in that the fuel supply pipeline is a circular tube structure with variable flow direction, one end of the inlet is connected with the fuel supply hole, the axial direction of the circular tube is the same as the radial direction of the combustion chamber, one end of the outlet is connected with the swirler, the axial direction of the circular tube is the same as the axial direction of the combustion chamber, and circular arc smooth transition is adopted from the inlet to the outlet of the circular tube.
The invention has the further improvement that 12 cyclones are respectively arranged in 12 cyclone mounting holes which are uniformly distributed along the circumferential direction at the head part of the rotary detonation combustion chamber, the cyclone mounting holes are of a circular through hole structure, the wall surfaces of outlets of inlets of circular holes are all subjected to chamfering treatment, air is ensured to flow into and out of the cyclones, and the inner diameter of the circular holes is the same as the outer diameter of the cyclone blades.
The invention is further improved in that the swirler vanes of a single swirler are positioned between the swirler fuel cavity and the swirler mounting holes, the number of the swirler vanes of each swirler is 10, and the swirler vanes are uniformly distributed along the circumferential direction of the swirler mounting holes; the inlet of the swirler vane is consistent with the axial direction of the combustion chamber, the included angle between the outlet direction of the swirler vane and the axis of the combustion chamber is 60 degrees, and the angle of the swirler vane is continuously changed from the inlet to the outlet, so that the mixture of air and fuel forms rotational flow movement deflected along a certain angle at the outlet of the swirler vane.
The invention has the further improvement that the fuel cavity of the swirler is positioned in the center of the swirler, and one side of the fuel cavity, which is close to the outlet, of the swirler is of a round platform structure, so that the air at the outlet of the swirler blade flows to the center along an arc to be quickly mixed with the fuel; and a swirler fuel supply hole with a circular through hole structure is arranged at the center of the swirler fuel cavity and close to one side of the outlet, and the axial direction of the circular hole is the same as the axial direction of the combustion chamber.
The invention has the further improvement that 2 flame tube mounting holes with circular through hole structures are reserved on one side of the outer ring of the rotary detonation combustor, which is close to the head of the combustion chamber, and are spaced by 180 degrees, and the central axis direction of the flame tube mounting holes is the same as the radial direction of the combustion chamber.
The invention has the further improvement that 1 combustion chamber pressure monitoring hole is reserved in the outer ring of the rotary detonation combustion chamber and is of a circular through hole structure, the axial positions of the combustion chamber pressure monitoring hole and the axial position of the flame connecting pipe mounting hole are the same, the interval between the circumferential direction of the combustion chamber pressure monitoring hole and the flame connecting pipe mounting hole is 90 degrees, a pressure sensor can be mounted in the combustion chamber pressure monitoring hole, and the propagation state of rotary detonation waves can be monitored in real time in the working process of the combustion chamber.
The further improvement of the invention is that cooling gas supply ring cavities with circular cavity structures are respectively designed inside one side of the outer ring and one side of the inner column of the rotary detonation combustion chamber, 160 gas film cooling holes are arranged between the gas film cooling holes and the rotary detonation combustion chamber, the gas film cooling holes of the outer ring and the inner column of the combustion chamber are distributed in the same way, 20 rows are distributed along the circumferential direction, and 8 rows are distributed along the axial direction; the included angle between the flowing direction of the gas in the gas film cooling hole and the axis of the combustion cavity is 30 degrees, and the cooling air flowing out of the gas film hole flows along the downstream direction of the rotary detonation combustion cavity.
The invention has the further improvement that the inlet of the burnt gas exhaust cavity is connected with the outlet of the outer ring of the rotary detonation combustor, and the section of the burnt gas exhaust cavity is circular; the outlet is connected with a first-stage stationary blade ring of the turbine, and the section of the outlet is a sector ring of 18 degrees; the cross section of the burnt gas exhaust cavity is smoothly transited from a circle to a fan-shaped structure from the inlet to the outlet of the burnt gas exhaust cavity; the material needs to be high-temperature resistant alloy material, and the inner wall surface needs to be sprayed with a heat insulation coating.
The invention has at least the following beneficial technical effects:
the invention relates to a gas turbine annular pipe type combustion chamber based on detonation combustion, which applies a rotary detonation combustion cavity to a gas turbine to replace the conventional annular pipe type isobaric combustion chamber. Due to the self-pressurization effect of detonation combustion, compared with a combustion chamber based on isobaric combustion, the burnt gas pressure at the outlet of the combustion chamber is higher under the condition of the same compressor pressure ratio, and the heat cycle efficiency of the gas turbine can be improved; the exhaust pressure of the compressor can be reduced under the condition of the same exhaust pressure of the combustion chamber, and the reduction of the stage number of the compressor is facilitated. Meanwhile, because the flame propagation rate of the detonation combustion is higher, the combustion is quickly completed in a smaller rotary detonation combustion cavity, the size of a combustion chamber of the gas turbine is favorably reduced, and the flexibility of the power device is improved. Finally, when a ring-pipe type combustion chamber structure is adopted, on one hand, the 20 combustion chambers uniformly distributed along the circumferential direction can ensure that the gas pressure and the temperature at the outlet of the combustion chamber are more uniformly distributed; on the other hand, the individual combustion chambers are easier to repair and replace after damage. Compared with the traditional annular combustion chamber of the gas turbine, the annular combustion chamber of the gas turbine has the advantages that the structural form of the combustion chamber of the gas turbine is optimized, the design is reasonable, the heat cycle efficiency of the gas turbine can be improved, and the volume of the combustion chamber is reduced.
Drawings
FIGS. 1 and 2 are schematic views of a can-annular combustor of a gas turbine;
FIG. 3 is a front view of a gas turbine can-annular combustor;
FIG. 4 is a schematic view of a single can-annular combustor;
FIG. 5 is a sectional view of a single can-annular combustion chamber;
FIGS. 6 and 7 are schematic views of a fuel supply configuration, respectively;
FIGS. 8 and 9 are schematic structural views of a rotary detonation combustor, respectively;
fig. 10 and 11 are schematic views of the cyclone structures respectively.
Description of reference numerals:
the combustion chamber comprises a combustion chamber 1, a fuel supply chamber 2, a swirler 3, a rotary detonation combustion chamber 4, a burnt gas exhaust chamber 5, a circular-tube-type combustion chamber annular channel 6, a mounting flange 7, an air supply chamber outer ring 8, a fuel supply chamber wall 9, a fuel supply hole 10, a fuel supply pipeline 11, a swirler fuel chamber 12, a swirler fuel chamber wall 13, a swirler fuel chamber wall 14, a swirler fuel supply hole 15, swirler vanes 15, a rotary detonation combustion chamber outer ring 16, a cooling gas supply annular cavity 17, a rotary detonation combustion chamber inner column 18, a swirler mounting hole 19, a gas film cooling hole 20, a flame connecting tube mounting hole 21, a combustion chamber pressure monitoring hole 22, a combustion chamber outlet guide cone 23, a combustion chamber exhaust section outer wall 24, a circular-tube-type combustion chamber annular channel outer ring 25 and a circular-tube-type combustion chamber annular channel inner ring 26.
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 and may be, for example, fixedly connected, detachably connected, or integrally connected; 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 to 5, the gas turbine annular combustion chamber based on detonation combustion of the present invention is composed of an air supply chamber 1, a fuel supply chamber 2, a swirler 3, a rotary detonation combustion chamber 4, a burned gas exhaust chamber 5 and an annular channel 6 of the annular combustion chamber.
The air supply cavity 1 is positioned at the leftmost end of the ring pipe type combustion chamber, is a cylindrical channel formed by an outer ring 8 of the air supply cavity, and is used for supplying high-pressure air generated by the compressor into the swirler 3, and is connected with the compressor by taking the leftmost end of the air supply cavity 1 as a mounting flange 2.
Referring to fig. 5, 6 and 7, the fuel supply system is composed of a fuel supply chamber 2, a fuel supply chamber wall surface 9, a fuel supply hole 10 and a fuel supply line 11 1. The fuel supply cavity 2 is located at the center of the air supply cavity 1, and is a conical structure formed by wall surfaces 9 of the fuel supply cavity, and the function of the conical structure is mainly to store sufficient fuel for a fuel supply system. The outlet of the fuel supply cavity 2 is provided with 12 fuel supply holes 10 which are uniformly distributed along the circumferential direction, and the fuel supply holes are in a circular through hole structure, and the axial direction of the fuel supply holes is the same as the radial direction of the ring pipe type combustion chamber. The fuel supply holes 10 are connected to corresponding 12 fuel supply pipes 11, the fuel in the fuel supply chamber 2 is supplied to the swirler 3, the axial direction of the inlet of the fuel supply pipe 11 is the same as the radial direction of the can-annular combustion chamber, the axial direction of the outlet is the same as the axial direction of the can-annular combustion chamber, and the transition from the inlet to the outlet of the fuel supply pipe 11 is a 1/4 circular arc.
Referring to fig. 5, 8, 9, 10 and 11, each swirler 3 is composed of a swirler fuel cavity 12, a swirler fuel cavity wall 13, swirler fuel supply holes 14 and swirler vanes 15, each rotary detonation combustion cavity 4 has 12 swirlers 3 which are uniformly distributed along the circumferential direction of the combustion chamber, the interval between adjacent swirlers 3 is 30 °, and the 12 swirlers 3 are mounted in corresponding 12 swirler mounting holes 19. The swirler fuel cavity 12 is a cylindrical channel formed by swirler fuel cavity wall surfaces 13 in a circular structure, one inlet end of the cylindrical channel is connected with the fuel supply pipeline 11, one outlet end of the cylindrical channel is provided with 1 swirler fuel supply hole 14, and the supply flow rate of swirler fuel is adjusted by changing the diameter of the swirler fuel supply hole 14. The outer side of the wall surface 13 of the fuel cavity of the swirler is provided with 10 swirler vanes 15, each swirler 3 is provided with 10 swirler vanes 15 which are uniformly distributed along the circumferential direction of the swirler 3, the direction of the inlet of each swirler vane 15 is consistent with the axial direction of the combustion chamber, the included angle between the direction of the outlet of each swirler vane 15 and the axial direction of the combustion chamber is 60 degrees, smooth circular arc transition is adopted from the inlet to the outlet of each swirler vane 15, and the air generates rotational flow movement deflected along a certain angle at the outlet of each swirler vane 15 through the change of the flowing direction of the air in each swirler vane 15.
The rotary detonation combustion chamber 4 is positioned at the outlet of the swirler 3 and is an annular cavity-shaped passage consisting of an outer ring 16 of the rotary detonation combustion chamber and an inner column 18 of the rotary detonation combustion chamber. Wherein, the export of rotatory detonation combustion chamber inner prop 18 is connected with the combustor export water conservancy diversion awl 23 of hemispherical structure, and its aim at avoids forming the backward flow district at the export of rotatory detonation combustion chamber 4 through the circular arc transition structure, and then causes the flow loss of burnt gas. 2 flame tube mounting holes 21 are reserved on one side, close to the head part of the combustion chamber, of the outer ring 16 of the rotary detonation combustion chamber, are spaced by 180 degrees, and the direction of the central axis of each flame tube mounting hole is the same as the radial direction of the annular-tube type combustion chamber, so that the flame tube is mounted between the adjacent rotary detonation combustion chambers; the purpose of flame tube is in order to be in the combustion chamber stage of detonating, and the rotatory detonation wave of the stable propagation that the rotatory detonation combustion chamber of part produced diffuses fast to other combustion chambers for quick detonating is all realized to all rotatory detonation combustion chambers, and through the flame tube of the installation of flame tube mounting hole 21, can make the rotatory detonation wave that forms in the rotatory detonation combustion chamber of part 4 diffuse fast to other rotatory detonation combustion chambers, and then realize quick detonating. The axial line position is the same as that of the flame tube mounting hole 21, 1 combustion chamber pressure monitoring hole 22 is reserved at the position of 90 degrees in the circumferential direction, a pressure sensor can be mounted, and the propagation state of the rotation detonation wave in each rotation detonation combustion chamber 4 can be monitored in real time.
The fuel supply pipeline 11 is a circular tube structure with variable flow direction, one end of the inlet is connected with the fuel supply hole, the axial direction of the circular tube is the same as the radial direction of the combustion chamber, one end of the outlet is connected with the swirler, the axial direction of the circular tube is the same as the axial direction of the combustion chamber, and circular arc smooth transition is adopted from the inlet to the outlet of the circular tube. The fuel supply pipeline adopting the variable flow direction aims to ensure that the radius position of the fuel supply cavity in the combustion chamber is smaller than that of the swirler, so that the fuel supply cavity is prevented from blocking air entering the swirler, and the air can smoothly enter the swirler. The fuel supply pipe is made of a pressure-resistant material.
The material of the swirler 3 blade needs to adopt high-strength pressure-resistant alloy material.
Inside the outer ring 16 and the inner column 18 of the rotary detonation combustor, a cooling air supply ring cavity 17 with a ring cavity-shaped structure is designed at one side close to the rotary detonation combustion cavity 4, and the cooling air required in the rotary detonation combustion cavity 4 is stored. The cooling air in the film cooling holes 17 is supplied to the rotary detonation combustion chamber 4 through the film cooling holes 20, 160 film cooling holes 20 are arranged on two sides of the rotary detonation combustion chamber outer ring 16 and two sides of the rotary detonation combustion chamber inner column 18, 20 rows are distributed in the circumferential direction, and 8 rows are distributed in the axial direction. The axial direction of each film cooling hole 20 and the axial direction of the rotary detonation combustion cavity 4 form an included angle of 30 degrees. The cooling air flowing out of the air film hole is along the downstream direction of the rotary detonation combustion cavity, so that the cooling air is guaranteed to be discharged from the rotary detonation combustion cavity along with the burnt gas, and the flow loss of the cooling air is reduced. The materials of the outer ring and the inner column of the rotary detonation combustor need to be high-temperature-resistant alloy materials, and the inner wall surface of the outer ring of the rotary detonation combustor and the outer wall surface of the inner column of the rotary detonation combustor need to be sprayed with heat-insulating coatings respectively.
Referring to fig. 1 to 5, the burned gas exhaust cavity 5 is located at the outlet of the rotary detonation combustion cavity 4, and is a variable cross-section channel formed by the outer wall surface 24 of the combustion chamber exhaust section, and one end of the inlet of the variable cross-section channel is connected with the rotary detonation combustion chamber outer ring 16, and the cross section of the inlet of the burned gas exhaust cavity 5 is circular because the cross section of the outlet of the rotary detonation combustion chamber outer ring 16 is circular; one end of the outlet of the combustion gas exhaust cavity is connected with a first-stage stationary blade ring of the turbine, and the section of the combustion gas exhaust cavity is an 18-degree fan-shaped ring structure, so that the section of the inlet of the combustion gas exhaust cavity 5 is also an 18-degree fan-shaped ring; in order to reduce the flow loss of the burned gas in the burned gas exhaust chamber 5, a smooth curved surface transition is adopted from the inlet to the outlet of the burned gas exhaust chamber 5.
Referring to fig. 1 and 3, the annular channel 6 of the can-annular combustor is an annular channel composed of an outer ring 25 of the annular channel of the can-annular combustor and an inner ring 26 of the annular channel of the can-annular combustor, 20 can-annular combustors uniformly distributed along the circumferential direction are installed in the annular channel 6 of the can-annular combustor, one end of the inlet of the can-annular combustor is connected with the annular cavity of the compressor outlet, and one end of the outlet of the can-annular combustor is connected with the first-stage stationary blade ring of the turbine. Because the cross sections of the air supply cavity 1 and the rotary detonation combustion cavity 4 at the head part of the annular cavity combustion chamber are circular, the section of the outlet of the burnt gas exhaust cavity 5 at the tail part of the annular cavity combustion chamber is a fan-shaped ring, and the radial width of the circular cross section is greater than that of the fan-shaped ring, the width of the head part of the annular channel 6 of the annular tube type combustion chamber is greater than that of the tail part, namely, one side of the annular channel 6 of the annular tube type combustion chamber, which is close to the tail part, is a contraction type channel.
The working cycle process of the invention is as follows:
the invention relates to a gas turbine annular pipe type combustion chamber based on detonation combustion. At the initial stage of the combustion chamber, firstly, the air compressor is driven to work by the jigger, and high-pressure gas compressed by the air compressor enters the rotary detonation combustion chamber 4 through the air supply chamber 1 and the swirler 3. The fuel supply valve is opened and fuel enters the rotary detonation combustion chamber 4 through the fuel supply chamber 2 and the swirler 3 to be mixed with air. Meanwhile, high-pressure air compressed by a part of the compressor is supplied to the cooling air supply ring cavity 17, and then the cooling air enters the rotary detonation combustion cavity 4 through the air film cooling holes 20 to cool the wall surface of the combustion cavity. In the stage of initiating a detonation, select a ring pipe type combustion chamber, install the firearm in its 1 continuous flame pipe mounting hole, open the ignition ware and ignite, form the slow combustion ripples in this rotatory detonation combustion chamber 4, the slow combustion ripples is followed the circumference propagation process in rotatory detonation combustion chamber 4, receives the vortex effect of combustion chamber wall face and gas film cooling hole 20, evolves gradually the rotatory detonation ripples of stable propagation, rotatory detonation ripples propagates to adjacent ring pipe type combustion chamber through continuous flame pipe afterwards. And repeating the processes in sequence, and quickly diffusing the rotary detonation wave to the rotary detonation combustion cavity 4 corresponding to each ring pipe type combustion chamber to realize the detonation of each combustion cavity. High-temperature and high-pressure combusted gas generated by slow combustion or detonation combustion flows to each stage of static and dynamic blades of the turbine through the combusted gas exhaust cavity 5 to do work, further shaft work is generated to drive the compressor and the generator to rotate, and the ring-pipe type 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 rotary detonation combustion cavity 4, the rotary detonation wave is decoupled into a slow combustion wave due to the fact that combustible mixture is lacked to maintain continuous propagation of the rotary detonation wave, then flame is gradually extinguished, 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 (10)
1. A gas turbine annular tube type combustion chamber based on detonation combustion is characterized by comprising an air supply cavity, a fuel supply cavity, a swirler, a rotary detonation combustion cavity, a burnt gas exhaust cavity and an annular tube type combustion chamber annular channel;
the air supply cavity is positioned at the head part of the combustion chamber and is a hollow cylindrical channel, and the outer side of the air supply cavity is an outer ring of an air supply cavity with a circular ring structure; the left side of the rotary detonation combustor is provided with an installation flange, the rotary detonation combustor is connected with an outlet of a gas compressor through a bolt installation hole reserved on the installation flange, and the right side of the rotary detonation combustor is connected with a rotary detonation combustion cavity;
the fuel supply cavity is positioned in the center of the air supply cavity and is a conical channel, the outer side of the fuel supply cavity is a wall surface of the fuel supply cavity with a thin-wall conical structure, the outlet of the fuel supply cavity is a fuel supply hole with a circular hole structure, and the direction of the central axis of the circular hole is the same as the radial direction of the combustion chamber; the outlet of the fuel supply hole is a fuel supply pipeline with a circular tube-shaped structure, the fuel supply pipeline is close to one end of the fuel supply hole, the axial direction of the fuel supply pipeline is the same as the radial direction of the combustion chamber, the other end of the fuel supply pipeline is connected with the swirler, and the axial direction of the fuel supply pipeline is the same as the axial direction of the combustion chamber;
the cyclone is positioned between the air supply cavity and the rotary detonation combustion cavity, is arranged in a cyclone mounting hole at the head of the rotary detonation combustion cavity, has one inlet end connected with a fuel supply pipeline and one outlet end provided with the rotary detonation combustion cavity, and consists of a cyclone fuel cavity, a cyclone fuel cavity wall surface, a cyclone fuel supply hole and a cyclone blade, and the axial direction of the cyclone is the same as the axial direction of the combustion chamber;
the rotary detonation combustion cavity is an annular cavity-shaped channel consisting of an outer ring and an inner column of the rotary detonation combustion chamber, and the axial direction of the rotary detonation combustion cavity is the same as the axial direction of the combustion chamber; the outer ring of the rotary detonation combustion chamber is of a circular structure, a cooling gas supply circular cavity of a circular cavity structure is reserved in one side close to the rotary detonation combustion cavity, and meanwhile, two circular holes in the radial direction are reserved on one side close to the head of the outer ring of the rotary detonation combustion chamber and are respectively a flame connecting pipe mounting hole and a combustion chamber pressure monitoring hole, and the circular holes are spaced by 90 degrees; the inner column of the rotary detonation combustion chamber is of a cylindrical structure, and a cooling gas supply annular cavity of an annular cavity structure is reserved in one side close to the rotary detonation combustion chamber; the axial lengths of the outer ring and the inner column of the rotary detonation combustion chamber are the same, the axial positions along the combustion chamber are also completely the same, and the outlet of the rotary detonation combustion chamber is a combustion chamber outlet guide cone which is of a solid conical structure;
the burned gas exhaust cavity is positioned on one side of the outlet of the rotary detonation combustion cavity, the structure of the burned gas exhaust cavity is an irregular round-to-square flow channel, and the outer side of the burned gas exhaust cavity is the outer wall surface of the exhaust section of the combustion chamber;
the annular channel of the annular tube type combustion chamber is of an annular structure, consists of an outer ring of the annular channel of the annular tube type combustion chamber and an inner ring of the annular channel of the annular tube type combustion chamber, is an installation channel of 20 annular tube type combustion chambers which are uniformly distributed along the circumferential direction, and is connected with the gas compressor at one end of an inlet and connected with a first-stage stationary blade ring of the turbine at one end of an outlet.
2. The annular combustion chamber of a gas turbine based on knocking combustion as set forth in claim 1, wherein the outer wall surface of the fuel supply chamber has a conical structure, the left end of the chamber has a pointed structure, the right end of the chamber has a circular structure, and the transition from the left end to the right end is smooth by an arc, so that a backflow area of air at the head of the fuel supply chamber can be effectively avoided, and the flow loss of gas can be reduced.
3. The gas turbine can-annular combustion chamber based on detonation combustion of claim 1, characterized in that the fuel supply pipeline is a circular tube structure with variable flow directions, one inlet end of the fuel supply pipeline is connected with the fuel supply hole, the axial direction of the circular tube is the same as the radial direction of the combustion chamber, one outlet end of the circular tube is connected with the swirler, the axial direction of the circular tube is the same as the axial direction of the combustion chamber, and a circular arc smooth transition is adopted from the inlet to the outlet end of the circular tube.
4. The gas turbine annular-tube type combustion chamber based on detonation combustion as claimed in claim 1, wherein 12 swirlers are respectively installed in 12 swirler installation holes uniformly distributed along the circumferential direction at the head of the rotary detonation combustion chamber, the swirler installation holes are of a circular through hole structure, the wall surfaces of outlets of circular hole inlets are all processed by rounding off, so that air flows into and out of the swirlers, and the inner diameter of the circular hole is the same as the outer diameter of the swirl vanes.
5. The gas turbine can-annular combustor based on detonation combustion of claim 1, wherein swirler vanes of a single swirler are located between a swirler fuel cavity and swirler mounting holes, the number of the swirler vanes of each swirler is 10, and the swirler vanes are uniformly distributed along the circumferential direction of the swirler mounting holes; the inlet of the swirler vane is consistent with the axial direction of the combustion chamber, the included angle between the outlet direction of the swirler vane and the axis of the combustion chamber is 60 degrees, and the angle of the swirler vane is continuously changed from the inlet to the outlet, so that the mixture of air and fuel forms rotational flow movement deflected along a certain angle at the outlet of the swirler vane.
6. The annular combustion chamber of the gas turbine based on the detonation combustion as claimed in claim 1, wherein a fuel cavity of the swirler is positioned in the center of the swirler, and one side of the fuel cavity close to the outlet is of a circular truncated cone structure, so that air at the outlet of the swirler vanes flows to the center along an arc to be rapidly mixed with the fuel; and a swirler fuel supply hole with a circular through hole structure is arranged at the center of the swirler fuel cavity and close to one side of the outlet, and the axial direction of the circular hole is the same as the axial direction of the combustion chamber.
7. The gas turbine annular tube type combustion chamber based on the detonation combustion as claimed in claim 1, wherein 2 flame tube mounting holes with circular through hole structures are reserved on one side of the outer ring of the rotary detonation combustion chamber close to the head of the combustion chamber, the flame tube mounting holes are spaced by 180 degrees, and the central axis direction of the flame tube mounting holes is the same as the radial direction of the combustion chamber.
8. The gas turbine annular-tube type combustion chamber based on detonation combustion as claimed in claim 1, wherein 1 combustion chamber pressure monitoring hole is reserved in the outer ring of the rotary detonation combustion chamber, the combustion chamber pressure monitoring hole is of a circular through hole structure, the axial positions of the combustion chamber pressure monitoring hole and the axial position of the flame connecting tube mounting hole are the same, the interval between the circumferential direction and the flame connecting tube mounting hole is 90 degrees, a pressure sensor can be mounted in the combustion chamber pressure monitoring hole, and the propagation state of the rotary detonation wave can be monitored in real time in the working process of the combustion chamber.
9. The gas turbine annular-tube type combustion chamber based on the detonation combustion as claimed in claim 1, wherein cooling gas supply annular cavities with circular cavity-shaped structures are respectively designed inside one side of the outer ring and one side of the inner column of the rotary detonation combustion chamber, 160 film cooling holes are formed between the film cooling holes and the rotary detonation combustion chamber, the film cooling holes of the outer ring and the inner column of the combustion chamber are distributed in the same way, 20 rows are distributed in the circumferential direction, and 8 rows are distributed in the axial direction; the included angle between the flowing direction of the gas in the gas film cooling hole and the axis of the combustion cavity is 30 degrees, and the cooling air flowing out of the gas film hole flows along the downstream direction of the rotary detonation combustion cavity.
10. The gas turbine can-annular combustor based on knocking combustion as claimed in claim 1, wherein the inlet of burned gas exhaust cavity is connected with the outlet of the outer ring of the rotary knocking combustor and has a circular cross section; the outlet is connected with a first-stage stationary blade ring of the turbine, and the section of the outlet is a sector ring of 18 degrees; the cross section of the burnt gas exhaust cavity is smoothly transited from a circle to a fan-shaped structure from the inlet to the outlet of the burnt gas exhaust cavity; the material needs to be high-temperature resistant alloy material, and the inner wall surface needs to be sprayed with a heat insulation coating.
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