A turbine interstage combustion chamber

Abstract

The present invention discloses a turbine interstage combustion chamber, comprising: an outer casing having a U-shaped recirculation cavity on its rear side and an air bleed interface on the rear wall of the U-shaped recirculation cavity; an arc-shaped guide body located inside the outer casing, a diverter plate at the front end of the arc-shaped guide body, a secondary flow channel connected to the U-shaped recirculation cavity between the top surface of the diverter plate, the top surface of the arc-shaped guide body, and the bottom surface of the outer casing; a concave area of ​​the arc-shaped guide body forming a trapped vortex cavity connected to the U-shaped recirculation cavity; an inner casing located below the arc-shaped guide body, a primary flow channel between the top surface of the inner casing and the bottom surface of the diverter plate, the bottom surface of the arc-shaped guide body, and the bottom surface of the tail wall; an igniter extending into the trapped vortex cavity; an air bleed pipe having one end connected to the air bleed interface and the other end connected to the oil supply pipe; an oil-gas mixing pipe connected to the oil supply pipe; and the oil supply pipe communicating with both the U-shaped recirculation cavity and the trapped vortex cavity. This turbine interstage combustion chamber can achieve low-resistance gas flow and efficient and stable combustion.

Description

Technical Field

[0001] The present invention relates to the technical field of aviation engines, and more particularly to a turbine interstage combustion chamber. Background Art

[0002] The ideal cycle for modern aircraft engines is the Brayton cycle. Based on the thermodynamic cycle, it is known that the cycle temperature increase ratio is one of the determining factors of thermal efficiency. Increasing the turbine inlet temperature has become the primary method for increasing thermal efficiency. Consequently, numerous researchers have conducted extensive research on high-temperature combustion chambers. However, higher combustion chamber outlet temperatures place severe demands on turbine blade materials and cooling technologies, limiting the approach of simply increasing the main combustion chamber outlet temperature to improve engine performance. In response, Srignano, Liu, and others proposed the concept of a turbine-burner (TB), which implements an isothermal thermodynamic cycle within the turbine, thereby increasing the work done by the entire thermodynamic cycle and boosting engine power. Through computational studies of thermodynamic cycle parameters, they found that compared to conventional engines, TB engines offer significantly higher specific thrust with little or no increase in thrust-specific fuel consumption, and also extend the operating range of flight Mach numbers and compressor pressure ratios. Later, the U.S. Air Force Research Laboratory (AFRL) proposed the concept of an ultra-compact combustion chamber (UCC), bringing the turbine-burner concept to engineering application. Based on the UCC and TB structures, researchers have proposed an inter-stage turbine burner (ITB) scheme, which adds a circumferential cavity to the casing wall at the passage between the high-pressure turbine and the low-pressure turbine. There are two main combustion schemes based on circumferential cavities. One is the high-g combustor (HGC), which injects air into the circumferential cavity at a certain angle, causing the airflow to rotate around the engine axis. Fuel is simultaneously supplied to the cavity, utilizing the high-speed rotation of the tangential airflow to enhance mixed combustion. However, the magnitude of the g load is proportional to the engine diameter, which to some extent limits the application of this scheme. The other is the trapped-vortex combustor (TVC), which provides intake channels on the front and rear walls of the circumferential cavity. Under the action of the intake jet, a stable trapped vortex zone is formed within the cavity, increasing the residence time of the oil-air mixture, resulting in more complete combustion and higher combustion efficiency. In interstage combustion chamber designs with circumferentially recessed cavities, most research utilizes bleed air devices to draw high-pressure air from behind the compressor, introducing high-energy unburned gases directly into the turbine. This has a negative impact on the thermodynamic cycle. Although this increases the total energy of the gas within the turbine, it does not solely benefit the overall engine cycle.

[0003] Publication number CN112524641 A describes a turbine interstage combustor structure that utilizes a V-shaped flame stabilizer and other devices downstream of the swirl guide vanes, achieving increased engine specific thrust at a low fuel consumption ratio. However, most currently disclosed interstage combustor solutions utilize a forced jet design, forming a single or dual vortex within the cavity to increase gas residence time and enhance primary and secondary flow mixing, but pay little attention to gas flow losses.

[0004] Therefore, how to provide a turbine interstage combustion chamber that can achieve low-resistance flow and efficient and stable combustion under high-speed inlet conditions is an urgent problem that needs to be solved by those skilled in the art. Summary of the Invention

[0005] Existing interstage combustor designs, based on UCC or TVC, mostly employ separate bleed air and forced jets, forming single or dual vortices within the concave cavity to increase gas residence time, enhance trapped vortex combustion, and improve primary and secondary flow mixing. However, these designs result in significant overall flow losses, making it difficult to improve combustion efficiency. In light of this, embodiments of the present invention disclose a turbine interstage combustor that achieves low-resistance flow and efficient, stable combustion under high-speed, strong swirl inlet conditions.

[0006] In order to achieve the above object, the present invention adopts the following technical solutions:

[0007] A turbine interstage combustion chamber, comprising:

[0008] An outer shell, wherein the rear side of the outer shell has a U-shaped reflux cavity, the lower side horizontal plate of the U-shaped reflux cavity is integrally connected to a rear wall plate extending downward, the lower end of the rear wall plate is integrally connected to a tail wall plate extending rearward, and an air bleed interface is provided on the rear wall surface of the U-shaped reflux cavity;

[0009] An arc-shaped guide body, the arc-shaped guide body is located on the inner side of the outer shell, and the two are fixed by a connecting column. A diverter plate is provided at the front end of the arc-shaped guide body near the front end of the outer shell. The rear end of the arc-shaped guide body extends rearward to the cavity opening of the U-shaped return cavity. The area between the top surface of the diverter plate, the top surface of the arc-shaped guide body, and the bottom end surface of the outer shell is a secondary flow channel connected to the U-shaped return cavity. The concave area of ​​the arc-shaped guide body is a trapped vortex cavity connected to the U-shaped return cavity. The rear wall plate is located on the rear side of the trapped vortex cavity.

[0010] An inner casing, the inner casing being located below the arc-shaped guide body; the area between the front ends of the inner casing and the outer casing being a high-temperature, high-speed gas inlet communicating with the outlet of the high-pressure turbine of the aircraft gas turbine; the area between the top end surface of the inner casing and the bottom end surface of the manifold, the bottom end surface of the arc-shaped guide body, and the bottom end surface of the tail wall plate being a main flow channel; the first channel opening of the main flow channel and the front end port of the secondary flow channel are both in communication with the high-temperature, high-speed gas inlet; and the second channel opening of the main flow channel is in communication with the inlet of the low-pressure turbine of the aircraft gas turbine;

[0011] an igniter, the igniter being mounted on the outer shell, and an ignition end of the igniter extending through the arc-shaped guide body into the trapped vortex cavity;

[0012] An air bleed tube, the air bleed tube being located outside the U-shaped reflux cavity, one end of the air bleed tube being in communication with the air bleed interface;

[0013] An oil supply pipe, one end of which is an oil supply port, and the other end of the air bleed pipe is connected to the oil supply pipe near the oil supply port;

[0014] an oil-gas mixing pipe, the oil-gas mixing pipe being connected to and in communication with the other end of the oil supply pipe;

[0015] Among them, the lower horizontal plate is provided with a first oil-gas mixing inlet for connecting the U-shaped reflux cavity and the oil-gas mixing tube, and the rear wall plate is provided with a second oil-gas mixing inlet for connecting the trapped vortex cavity and the oil-gas mixing tube.

[0016] Through the above technical solution, it can be known that compared with the prior art, the present invention discloses a turbine interstage combustion chamber, in which the high-temperature and high-speed combustion gas at the high-pressure turbine outlet enters the turbine interstage combustion chamber through the high-temperature and high-speed combustion gas inlet, and is divided into a secondary flow and a mainstream flow under the action of the diverter plate, and enters the secondary flow channel and the mainstream channel respectively. Most of the secondary flow passes through the U-shaped recirculation cavity, and a small part of the secondary flow passes through the U-shaped recirculation cavity on the outer shell and enters the oil supply pipe under the guidance of the bleed pipe. At this time, the oil supply port of the oil supply pipe supplies liquid fuel. Under the action of the high-temperature airflow in the bleed pipe, the liquid fuel is evaporated to form an oil-gas mixture that enters the oil-gas mixing pipe. The oil-gas mixture in the oil-gas mixing pipe enters the U-shaped recirculation cavity through the first oil-gas mixing inlet, and is quickly driven by most of the high-speed secondary flow refluxed in the U-shaped recirculation cavity into the trapped vortex cavity. At this time, the igniter ignites to ignite the secondary flow in the trapped vortex cavity. At this time, the combustion area is the trapped vortex main combustion chamber of the interstage combustion chamber. The second inlet is used to connect the second and third inlets to the oil and gas mixture, and the oil and gas mixture in the oil and gas mixing tube is used to connect the second and third inlets to the oil and gas mixture.

[0017] Therefore, the turbine interstage combustor proposed in the present invention differs from previously known interstage combustors. To increase the residence time of the fuel-gas mixture and reduce flow losses, the proposed interstage combustor structure primarily uses a diverter plate and an arc-shaped guide body to guide the airflow, forming a primary and secondary flow. The secondary flow is used to form a trapped vortex zone within the cavity. Furthermore, when the secondary flow enters the trapped vortex cavity to form the trapped vortex zone, it does not use a converging channel structure to force the gas into a jet. Instead, it is guided by a U-shaped recirculation cavity, which reverses the flow direction and forms a large, counterclockwise rotating trapped vortex with minimal flow losses. Therefore, the U-shaped recirculation cavity achieves low-resistance flow of the fuel gas while increasing its residence time and improving combustion efficiency. Furthermore, the use of an evaporator-type fuel supply utilizes a portion of the high-temperature fuel gas to preheat and evaporate the liquid fuel, enabling the fuel-gas mixture to burn more quickly within the trapped vortex cavity. Furthermore, the high-pressure turbine outlet airflow has a high swirl velocity, and the airflow maintains a circumferential rotational velocity upon entering the circumferential trapped vortex cavity, which helps enhance combustion.

[0018] Furthermore, it also includes a low-pressure turbine inlet guide vane arranged in the main flow channel, the bottom end of the low-pressure turbine inlet guide vane is fixed to the top end of the inner casing, and a radial cavity is opened on the pressure side of the low-pressure turbine inlet guide vane, and the radial cavity is located directly below the trapped vortex cavity.

[0019] The beneficial effects of adopting the above technical solution are: the low-pressure turbine inlet guide vanes are integrated with the mainstream channel, and the low-pressure turbine inlet guide vanes and the trapped vortex cavity are integrated into the combustion chamber, which shortens the axial length of the interstage combustion chamber, improves space utilization, and increases the compactness of the interstage combustion chamber. In addition, when the mainstream high-temperature and high-speed gas flows through the low-pressure turbine inlet guide vanes, a guide vane cavity trapped vortex vortex is also formed in the radial cavity, and the trapped vortex cavity vortex forms a double trapped vortex combustion of the gas. The radial cavity and the trapped vortex cavity are connected, which strengthens the mixing of the primary and secondary flows and facilitates the energy exchange between the secondary flow of the secondary combustion and the mainstream, thereby increasing the total energy of the airflow, enhancing the working capacity of the low-pressure turbine, and improving the engine's cyclic work capacity under the condition of low fuel consumption ratio.

[0020] Furthermore, the arc-shaped guide body is provided with a plurality of exhaust cooling holes for cooling the inner concave wall surface of the arc-shaped guide body corresponding to the trapped vortex cavity.

[0021] The beneficial effect of adopting the above technical solution is: since the static pressure of the secondary flow channel is higher than the internal pressure of the trapped vortex cavity, under the action of the pressure difference, the air flow passes through the exhaust cooling holes to form an air film on the concave wall of the trapped vortex cavity, thereby avoiding the trapped vortex combustion from burning at high temperature for a long time and ablating the concave wall surface of the arc-shaped guide body, thereby reducing the service life of the arc-shaped guide body.

[0022] Furthermore, the oil-gas mixing pipe is located in the interlayer formed between the lower side transverse plate and the tail wall plate.

[0023] The beneficial effect of adopting the above technical solution is that the oil-gas mixing tube is cleverly arranged in the interlayer, which not only makes the combustion chamber structure more compact, but also enables the oil-gas mixing tube to simultaneously provide oil-gas mixture to the U-shaped reflow cavity and the trapped vortex cavity, thereby realizing ultra-compact oil supply to the combustion chamber.

[0024] In summary, in the present invention, high-speed combustion gas is divided into two streams, the mainstream and the secondary stream, after passing through the arc-shaped guide body. A small part of the secondary stream evaporates the fuel, while the majority of the secondary stream passes through the U-shaped recirculation cavity and is mixed with the gaseous fuel before being injected back into the trapped vortex cavity, forming a trapped vortex main combustion zone in the trapped vortex cavity. At this time, the residence time of the oil-gas mixture is increased, and stable combustion is formed in the trapped vortex cavity under low oil-gas ratio, with less flow loss and increased temperature of the secondary stream. At the same time, the low-pressure turbine inlet guide vane is coupled with the trapped vortex cavity to improve the compactness of the interstage combustion chamber, and radial cavities are opened on the guide vanes. On the one hand, the oil-gas mixture circulates in the trapped vortex in the cavity, forming a combustion zone in the radial cavity, accelerating the combustion of the mainstream gas. On the other hand, the transport of components of the mainstream and secondary streams in the radial direction is faster, which strengthens the mixing of high and low temperature airflows in a short distance, thereby strengthening the energy exchange between the main and secondary streams, thereby increasing the overall gas temperature and increasing the Brayton cycle work of the gas turbine. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are merely embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying any creative work.

[0026] Figure 1 An overall schematic diagram of a turbine interstage combustion chamber provided by the present invention;

[0027] Figure 2 A front view of a turbine interstage combustion chamber provided by the present invention;

[0028] Figure 3 A schematic diagram of the internal flow of gas in a turbine interstage combustion chamber provided by the present invention;

[0029] Figure 3 The meanings of the arrows are as follows:

[0030] DETAILED DESCRIPTION

[0031] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0032] It should be noted that the existing turbine interstage combustion chamber is generally annular, that is, the outer shell 1, the inner shell 11, the arc-shaped guide body 6, the diverter plate 8, the U-shaped recirculation cavity, the lower side transverse plate 3, the rear wall plate 4, the tail wall plate 5, the oil-gas mixing pipe 16, and the trapped vortex cavity 10 are all annular structures, and there are multiple low-pressure turbine inlet guide vanes 17, which are uniformly distributed on the top surface of the inner shell 11 in the circumferential direction. In order to facilitate the display of the structure of the combustion chamber, the present invention only illustrates a partial structural schematic diagram of the annular combustion chamber. Specifically, as shown in FIG. Figure 1-Figure 3 As shown, an embodiment of the present invention discloses a turbine interstage combustion chamber, comprising:

[0033] The outer shell 1 has a U-shaped reflux chamber 2 on its rear side. The lower horizontal plate 3 of the U-shaped reflux chamber 2 is integrally connected to a rear wall plate 4 extending downward. The lower end of the rear wall plate 4 is integrally connected to a tail wall plate 5 extending rearward. An air bleed interface 201 is provided on the rear wall surface of the U-shaped reflux chamber 2.

[0034] The arc-shaped guide body 6 is located on the inner side of the outer shell 1, and the two are fixed by a connecting column 7. A diverter plate 8 is provided at the front end of the arc-shaped guide body 6 near the front end of the outer shell 1. The rear end of the arc-shaped guide body 6 extends backward to the cavity opening of the U-shaped return flow cavity 2. The area between the top surface of the diverter plate 8, the top surface of the arc-shaped guide body 6 and the bottom surface of the outer shell 1 is a secondary flow channel 9 connected to the U-shaped return flow cavity 2. The concave area of ​​the arc-shaped guide body 6 is a trapped vortex cavity 10 connected to the U-shaped return flow cavity 2. The rear wall plate 4 is located on the rear side of the trapped vortex cavity 10. A plurality of exhaust cooling holes 601 are opened on the arc-shaped guide body 6 for cooling the concave wall surface of the arc-shaped guide body 6 corresponding to the trapped vortex cavity 10;

[0035] The inner casing 11 is located below the arcuate guide body 6. The area between the front ends of the inner casing 11 and the outer casing 1 is a high-temperature, high-speed gas inlet 19 that communicates with the outlet of the high-pressure turbine of the aircraft gas turbine. The area between the top end surface of the inner casing 11 and the bottom end surfaces of the diverter plate 8, the arcuate guide body 6, and the tail wall plate 5 is the main flow channel 12. The first channel opening 121 of the main flow channel 12 and the front end 901 of the secondary flow channel 9 are both in communication with the high-temperature, high-speed gas inlet 19. The second channel opening 122 of the main flow channel 12 is in communication with the first-stage rotor inlet of the low-pressure turbine of the aircraft gas turbine.

[0036] An igniter 13 is mounted on the outer shell 1, and an ignition end of the igniter 13 passes through the arc-shaped guide body 6 and extends into the trapped vortex cavity 10;

[0037] The air bleed pipe 14 is located outside the U-shaped reflux cavity 2, and one end of the air bleed pipe 14 is connected to the air bleed interface 201;

[0038] The fuel supply pipe 15 has a fuel supply port 151 at one end, and the other end of the air bleed pipe 14 is connected to the fuel supply pipe 15 near the fuel supply port 151;

[0039] The oil-gas mixing pipe 16 is located in the interlayer 18 formed between the lower transverse plate 3 and the tail wall plate 5. The oil-gas mixing pipe 16 is connected to and communicates with the other end of the oil supply pipe 15;

[0040] Among them, a first oil-gas mixing inlet 301 for connecting the U-shaped reflux cavity 2 and the oil-gas mixing pipe 16 is opened on the lower horizontal plate 3, and a second oil-gas mixing inlet 401 for connecting the trapped vortex cavity 10 and the oil-gas mixing pipe 16 is opened on the rear wall plate 4.

[0041] The turbine interstage combustion chamber of the present invention also includes a low-pressure turbine inlet guide vane 17 arranged in the mainstream channel 12. The bottom end of the low-pressure turbine inlet guide vane 17 is fixed to the top end of the inner casing 11. A radial cavity 171 is opened on the pressure side of the low-pressure turbine inlet guide vane 17. The radial cavity 171 is located directly below the trapped vortex cavity 10.

[0042] like Figure 3As shown in the figure, the high-temperature and high-speed combustion gas at the outlet of the high-pressure turbine is divided into two streams, the main stream and the secondary stream, under the action of the diverter plate 7. When the secondary stream enters the U-shaped reflux cavity 2 on the outer shell 1 through the secondary stream channel, a small part of the secondary stream enters the oil supply pipe 15 under the guidance of the bleed pipe 14. At this time, the oil supply port 151 of the oil supply pipe 15 is filled with liquid fuel. Under the action of the high-temperature airflow in the bleed pipe 14, the evaporated oil-gas mixture enters the oil-gas mixing pipe 16. The oil-gas mixture in the oil-gas mixing pipe 16 passes through the first oil-gas mixing pipe 16. The inlet 301 enters the U-shaped recirculation chamber 2, and driven by most of the high-speed secondary flow recirculating in the U-shaped recirculation chamber 2, it quickly enters the trapped vortex cavity 10, and forms a large-scale counterclockwise rotating trapped vortex in the trapped vortex cavity 10. At this time, the igniter ignites and ignites the secondary flow in the trapped vortex cavity 10. At this time, the combustion area is the trapped vortex main combustion area of ​​the combustion chamber; in addition, the oil and gas mixture entering the U-shaped recirculation chamber 2 through the first oil and gas mixture inlet 301 will rush to the front of the trapped vortex cavity 10 driven by the high-speed secondary flow. The oil and gas content at the rear of the trapped vortex cavity 10 is less. At this time, the oil and gas mixture in the oil-gas mixing pipe 16 enters the rear of the trapped vortex cavity 10 through the second oil-gas mixing inlet 401, thereby filling the oil and gas content at the rear of the trapped vortex cavity 10. Therefore, under the cooperation of the first oil-gas mixing inlet 301 and the second oil-gas mixing inlet 401, the oil and gas in the trapped vortex cavity 10 are uniformly distributed in the radial and span directions, thereby achieving efficient and stable combustion. At this time, the temperature of the secondary flow increases; while the high temperature and high pressure of the mainstream flow When the high-speed gas flows through the low-pressure turbine inlet guide vane 17 in the mainstream channel 12, a trapped vortex vortex is also formed in the radial cavity 171 of the guide vane, and cooperates with the vortex of the trapped vortex cavity 10 to form a double trapped vortex combustion of the gas. In addition, the radial cavity 171 and the trapped vortex cavity 10 are connected, which strengthens the mixing of the primary and secondary flows and facilitates the energy exchange between the secondary flow of the secondary combustion and the mainstream, thereby increasing the total energy of the airflow, enhancing the working ability of the low-pressure turbine, and improving the cyclic work capacity of the engine under the condition of low fuel consumption ratio.

[0043] A turbine interstage combustion chamber of the present invention, first, does not draw air from the compressor, thereby reducing flow losses in the main combustion chamber, and adopts a splitter plate and an arc-shaped guide body structure to form two airflows, a primary flow and a secondary flow, and the secondary flow is used to form a trapped vortex zone in the concave cavity; secondly, when the secondary flow enters the concave cavity to form the trapped vortex zone, the structure of the contraction channel is not adopted to force the gas to form a jet, but the airflow direction is reversed through the guidance of the U-shaped recirculation cavity, forming a large-scale counterclockwise rotating trapped vortex with less flow loss; thirdly, the fuel supply of the combustion chamber uses the secondary flow high-temperature gas to achieve pre-evaporation of the liquid fuel, which is very important for the interstage combustion chamber with an extremely short oil-gas residence time. While ensuring stable combustion at a low oil-gas ratio, it also improves combustion efficiency; finally, the integrated design of the trapped vortex cavity and the low-pressure turbine inlet guide vane improves the space utilization of the interstage combustion chamber, strengthens the mixing of the primary and secondary flows, and is beneficial to energy exchange, which can form a more uniform temperature distribution at the outlet of the interstage combustion chamber, and extend the working life of the downstream low-pressure turbine blades.

[0044] In another embodiment of the present invention, the diverter plate and the arc-shaped guide body are connected by a hinge, so that the diverter plate can rotate around the hinge to change the area ratio of the secondary flow inlet and the primary flow inlet, thereby adjusting the flow ratio of the two air flows and further adjusting the combustion effect of the combustion chamber.

[0045] The various embodiments in this specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Reference can be made to the common and similar parts between the various embodiments. For the devices disclosed in the embodiments, since they correspond to the methods disclosed in the embodiments, the description is relatively simple, and the relevant parts can be referred to the method description.

[0046] The above description of the disclosed embodiments is intended to enable one skilled in the art to implement or use the present invention. Various modifications to these embodiments will be readily apparent to one skilled in the art, and the general principles defined herein may be implemented in other embodiments without departing from the spirit or scope of the present invention. Therefore, the present invention is not limited to the embodiments shown herein but is intended to conform to the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A turbine interstage combustion chamber, characterized in that: include: An outer shell (1), the rear side of the outer shell (1) having a U-shaped reflux chamber (2), a lower side transverse plate (3) of the U-shaped reflux chamber (2) being integrally connected to a rear wall plate (4) extending downward, a lower end of the rear wall plate (4) being integrally connected to a tail wall plate (5) extending rearward, and an air bleed interface (201) being provided on the rear wall surface of the U-shaped reflux chamber (2); An arc-shaped guide body (6), the arc-shaped guide body (6) is located on the inner side of the outer shell (1), and the two are fixed by a connecting column (7), a diverter plate (8) is provided at a position near the front end of the outer shell (1) at the front end of the arc-shaped guide body (6), and the rear end of the arc-shaped guide body (6) extends backward to the cavity opening of the U-shaped return flow cavity (2), the area between the top end surface of the diverter plate (8), the top end surface of the arc-shaped guide body (6) and the bottom end surface of the outer shell (1) is a secondary flow channel (9) connected to the U-shaped return flow cavity (2), the concave area of ​​the arc-shaped guide body (6) is a trapped vortex cavity (10) connected to the U-shaped return flow cavity (2), and the rear wall plate (4) is located on the rear side of the trapped vortex cavity (10); An inner casing (11), the inner casing (11) is located below the arc-shaped guide body (6), the area between the front end of the inner casing (11) and the front end of the outer casing (1) is a high-temperature and high-speed gas inlet (19) connected to the outlet of the high-pressure turbine of the aircraft gas turbine, and the area between the top end surface of the inner casing (11) and the bottom end surface of the diverter plate (8), the bottom end surface of the arc-shaped guide body (6), and the bottom end surface of the tail wall plate (5) is a main flow channel (12); the first channel opening (121) of the main flow channel (12) and the front end (901) of the secondary flow channel (9) are both connected to the high-temperature and high-speed gas inlet (19), and the second channel opening (122) of the main flow channel (12) is connected to the inlet of the low-pressure turbine of the aircraft gas turbine; an igniter (13), the igniter (13) being mounted on the outer shell (1), and an ignition end of the igniter (13) passing through the arc-shaped guide body (6) and extending into the trapped vortex cavity (10); An air bleed pipe (14), the air bleed pipe (14) being located outside the U-shaped reflux cavity (2), one end of the air bleed pipe (14) being in communication with the air bleed interface (201); An oil supply pipe (15), one end of the oil supply pipe (15) is an oil supply port (151), and the other end of the air bleed pipe (14) is connected to the oil supply pipe (15) near the oil supply port (151); an oil-gas mixing pipe (16), the oil-gas mixing pipe (16) being connected to and in communication with the other end of the oil supply pipe (15); The lower horizontal plate (3) is provided with a first oil-gas mixing inlet (301) for connecting the U-shaped reflux cavity (2) and the oil-gas mixing pipe (16), and the rear wall plate (4) is provided with a second oil-gas mixing inlet (401) for connecting the trapped vortex cavity (10) and the oil-gas mixing pipe (16).

2. A turbine interstage combustion chamber according to claim 1, characterized in that: The invention also includes a low-pressure turbine inlet guide vane (17) arranged in the main flow channel (12), the bottom end of the low-pressure turbine inlet guide vane (17) is fixed to the top end of the inner casing (11), and a radial cavity (171) is opened on the pressure side of the low-pressure turbine inlet guide vane (17), and the radial cavity (171) is located directly below the trapped vortex cavity (10).

3. A turbine interstage combustion chamber according to claim 1 or 2, characterized in that: The arc-shaped guide body (6) is provided with a plurality of exhaust cooling holes (601) for cooling the inner concave wall surface of the arc-shaped guide body (6) corresponding to the trapped vortex cavity (10).

4. A turbine interstage combustion chamber according to claim 1 or 2, characterized in that: The oil-gas mixing pipe (16) is located in an interlayer (18) formed between the lower transverse plate (3) and the tail wall plate (5).

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