Forcing type continuous detonation turbine engine

Abstract

The invention discloses a thrust augmentation type continuous detonation turbine engine. An afterburner is formed on the rear side of a rectification exhaust tail cone of the engine; the continuous detonation main combustion chamber consists of a shell and a central inner column which are coaxially arranged; a pre-explosion pipe, an air film cooling hole and a cold air mixing hole are formed in the shell; a fuel channel and a secondary air channel are arranged in the central inner column; a fuel swirler and a secondary air hole are arranged on the central inner column; the supporting plates are evenly distributed in the afterburner in the circumferential direction. A secondary inner column, a plane air ring and an annular protruding oil injection ring are arranged at the tail of the supporting plate. The annular protruding oil injection ring injects secondary fuel at high pressure in the axial direction, and the secondary fuel is mixed with tail gas of a combustion chamber for combustion. The plane air ring injects air in the axial direction, so that efficient mixing and combustion of fuel in the inner ring are achieved. According to the engine, a continuous detonation turbine engine and an afterburner are integrally designed, and the purpose of releasing a large amount of energy in a short time is achieved.

Description

Technical Field

[0001] The invention belongs to the technical field of aerospace engine structure design, and in particular relates to an afterburner continuous detonation turbine engine. Background Art

[0002] Conventional aircraft turbine engines generally employ an isobaric slow combustion method to organize combustion. The resulting high-temperature, high-pressure combustion gas propels the turbine, expands through the tail nozzle, and is then transmitted to the compressor via a drive shaft. However, due to the limitations of the isobaric slow combustion method and the thermal cycle, improving engine thermal and combustion efficiency faces bottlenecks. Detonation combustion, a new combustion method research hotspot in the aerospace propulsion field in recent years, approximates isochoric combustion and offers higher theoretical thermal efficiency. Using detonation to organize combustion holds promise for further optimizing and improving aircraft engine performance.

[0003] Continuous detonation engines utilize a self-sustaining detonation wave (rotating at supersonic speeds) formed by the fuel and oxidizer within the combustion chamber, achieving efficient energy release. These engines offer a range of advantages, including high thermal efficiency, self-pressurization, and single-shot detonation. Continuous detonation turbine engines, which utilize a continuous detonation main combustion chamber instead of the traditional aircraft turbine engine's main combustion chamber, are expected to significantly simplify the engine structure, improve thrust and thrust-to-weight ratio, and increase engine cruising time. However, because the detonation wave propagates within the combustion chamber head and is sensitive to incoming flow conditions and fuel distribution, it places high demands on rapid, short-range mixing of the fuel and oxidizer. Furthermore, cooling technology for continuous detonation turbine engines is a core challenge in their engineering application, primarily focusing on cooling the combustion chamber head walls and the high-temperature exhaust gases at the combustion chamber outlet. Furthermore, to meet the requirements of high maneuverability, supersonic cruise, or emergency acceleration for fighter aircraft, the addition of an afterburner at the turbine outlet of the continuous detonation turbine engine is necessary to further achieve high thrust for a short period of time.

[0004] Existing continuous detonation turbine engines mostly use a single row of fuel holes for counter-injection. However, this injection method does not allow for efficient mixing of fuel and oxidizer, resulting in uneven distribution of propellant in the flow field, which greatly affects the stable propagation of the detonation wave and reduces engine performance. In addition, existing continuous detonation turbine engines have not fully considered the problems of cooling the combustion chamber head wall and the tail high-temperature gas cooling. Finally, existing continuous detonation turbine engines lack an integrated design with the afterburner, which also makes it difficult to further increase engine thrust and achieve short-range acceleration.

[0005] Based on the above-mentioned prior art background, the inventors have developed an afterburner continuous detonation turbine engine. Summary of the Invention

[0006] The present invention provides an afterburner continuous detonation turbine engine, which integrates the continuous detonation turbine engine and the afterburner combustion chamber into an integrated design, thereby achieving the goal of releasing a large amount of energy in a short time, meeting the requirements of high maneuverability and emergency acceleration of the continuous detonation turbine engine, and fully considering the mixing and cooling of the continuous detonation main combustion chamber.

[0007] In order to achieve the above object, the present invention adopts the following specific technical solutions: An afterburner continuous detonation turbine engine, comprising a casing, a center body, compressor stator blades, compressor rotor blades, a continuous detonation main combustion chamber, turbine stator blades, turbine rotor blades, and a support plate; From the air inlet end to the air outlet end, the inner wall of the shell is provided with the air inlet lip, the compressor casing, the outer wall of the main combustion chamber, the turbine casing, the outer wall of the afterburner, and the tail nozzle in sequence. The center body is provided with the air inlet center cone, the compressor drum, the compressor wheel, the inner wall of the main combustion chamber, the turbine wheel, and the rectifying exhaust tail cone in sequence. The afterburner is formed at the rear side of the rectifying exhaust tail cone, and the combustion gas passing through the rectifying exhaust tail cone enters the afterburner chamber. The compressor stator blades are fixedly mounted on the inner wall of the compressor casing to form an annular flow channel; the compressor rotor blades are fixedly mounted on the compressor wheel; the turbine stator blades are fixedly mounted on the inner wall of the turbine casing; the turbine rotor blades are fixedly mounted on the turbine wheel; A continuous detonation main combustion chamber is arranged between the outer wall of the main combustion chamber and the outer wall of the main combustion chamber; the continuous detonation main combustion chamber consists of a shell fixedly connected to the inner wall of the outer shell and a central inner column coaxially arranged in the shell; an annular inner cavity is formed between the shell and the central inner column; a pre-detonation tube, air film cooling holes and cold air mixing holes are arranged on the shell; each air film cooling hole is connected to the low-pressure air collecting cavity; each cold air mixing hole is connected to the high-pressure air collecting cavity; the central inner column is fixedly connected to the turbine casing, and a fuel channel and a secondary air channel are arranged inside; a fuel swirler connected to the fuel channel and a secondary air hole connected to the secondary air channel are arranged on the central inner column; The support plates are evenly distributed in the afterburner combustion chamber along the circumferential direction; the support plates are provided with a secondary inner column, a plane air ring and an annular raised fuel injection ring coaxially arranged from the inside to the outside at one end facing the tail nozzle; the plane air ring and the annular raised fuel injection ring form a double-ring concave-convex structure; the annular raised fuel injection ring is used to axially inject secondary fuel into the afterburner combustion chamber at high pressure, and mix it with the combustion chamber exhaust for combustion; the plane air ring is used to axially inject air into the afterburner combustion chamber to achieve efficient mixing and combustion of the inner ring fuel.

[0008] Furthermore, the support plate adopts a triangular prism structure with a thickness gradually increasing from the air inlet end to the air outlet end, and the end of the support plate facing the rectification exhaust tail cone is the front end and the tip.

[0009] Furthermore, the air inlet lip is a pointed lip and is spaced apart on the outer peripheral side of the air inlet center cone; the air inlet center cone is a cone; the air inlet lip and the air inlet center cone cooperate to achieve a change in the flow area.

[0010] Furthermore, the compressor drum is connected in series with each compressor rotor blade; The low-pressure plenum is connected to the outlet of the low-pressure compressor; The high-pressure plenum is connected to the high-pressure compressor outlet.

[0011] Furthermore, each of the air film cooling holes and the cold air mixing holes is connected to a thin tube, and is connected to the corresponding low-pressure air collecting cavity and high-pressure air collecting cavity through the thin tube.

[0012] Furthermore, the rectifying exhaust tail cone is fixedly connected to the outer wall of the afterburner chamber.

[0013] Furthermore, the end of the central inner column is embedded in a mortise and tenon groove on the inner wall of the turbine casing; One end of the turbine stator blade is embedded in the inner wall of the turbine casing through a locking piece; The compressor stator blades are fixed to the inner wall of the compressor casing by bolts.

[0014] Furthermore, the tail nozzle adopts a Laval flow channel design.

[0015] Furthermore, the pre-explosion pipe and the two exhaust film cooling holes are arranged at the front end of the shell; the two rows of cold air mixing holes are arranged at the rear end of the shell; The shell is 180mm long and 120mm in outer diameter; the outer diameter of the central inner column is 60mm; There are 20-40 film cooling holes evenly distributed along the circumference of each exhaust film cooling hole, and the diameter of the film cooling hole is 1mm; Each row of cold air mixing holes has 20-40 holes evenly distributed along the circumference, and the diameter of the cold air mixing holes is 1.5mm; The front end of the central inner column is evenly distributed with 10-20 fuel swirlers along the circumference, and the rear end is distributed with two rows of secondary air holes, each row of secondary air holes is evenly distributed along the circumference, and the diameter of the secondary air holes is 1mm; The inner diameter of the pre-explosion tube is 80mm.

[0016] Furthermore, the outer diameter of the annular raised injection ring is 40 mm, the inner diameter is 30 mm, and the raised height is 3 mm; The outer diameter of the flat air ring is 30mm and the inner diameter is 15mm; The outer diameter of the secondary inner column is 15 mm and the length is 10 mm.

[0017] Compared with the prior art, the technical solution of the present invention has the following beneficial effects: 1. The afterburner continuous detonation turbine engine of the present invention is provided with a continuous detonation main combustion chamber arranged between the outer wall of the main combustion chamber and the inner wall of the main combustion chamber, an afterburner combustion chamber is formed on the rear side of the rectifying exhaust tail cone, and support plates are evenly distributed circumferentially in the afterburner combustion chamber. The afterburner combustion chamber and the continuous detonation turbine engine are integrated into a design to meet the requirements of high maneuverability and emergency acceleration of the continuous detonation turbine engine.

[0018] 2. The afterburner continuous detonation turbine engine of the present invention is provided with circumferentially uniformly distributed fuel swirlers on the central inner column, and the fuel swirlers and the incoming air flow are counter-mixed, which is beneficial to shortening the mixing distance, improving the mixing quality, and enhancing the stability of detonation propagation.

[0019] 3. The afterburner continuous detonation turbine engine of the present invention has secondary air holes uniformly distributed circumferentially at the axial rear end of the fuel swirler to inject air to adjust the premix equivalence ratio and further react with the fuel-rich premixed gas to form an explosive premixed gas.

[0020] 4. The afterburner continuous detonation turbine engine of the present invention has film cooling holes evenly distributed circumferentially at the front end of the casing of the continuous detonation main combustion chamber, and cold air mixing holes evenly distributed circumferentially at the rear end. The film cooling holes and cold air mixing holes are conducive to fully cooling the wall surface of the main combustion chamber, and to a certain extent reduce the ablation of the turbine blades, thereby improving the engine life and maximum operating time.

[0021] 5. The afterburner continuous detonation turbine engine of the present invention is provided with a coaxial secondary inner column, a planar air ring and an annular raised fuel injection ring at the tail of the support plate. The annular raised fuel injection ring and the planar air ring are arranged in a coaxial shear arrangement. The annular raised fuel injection ring and the planar air ring form a double-ring concave-convex structure. The secondary inner column is arranged at the center of the annular raised fuel injection ring and the planar air ring at the tail of the support plate. Through the above-mentioned structural design, a large number of vortices can be induced in the axial and lateral directions, which greatly enhances combustion and achieves the goal of releasing a large amount of energy in a short time. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] Figure 1 This is a simplified structural diagram of the afterburner continuous detonation turbine engine of the present invention; Figure 2 This is a schematic diagram of the structure of the main combustion chamber of the continuous detonation wave; Figure 3 Schematic diagram of the support plate structure.

[0023] Figure numerals: 1-inlet lip; 2-compressor stator blades; 3-compressor rotor blades; 4-continuous detonation main combustion chamber; 5-turbine stator blades; 6-turbine rotor blades; 7-support plate; 8-tail nozzle; 9-inlet center cone; 10-compressor drum; 11-compressor impeller; 12-turbine impeller; 13-rectifying exhaust tail cone; 14-afterburner; 41-casing; 42-center inner column; 43-pre-detonation pipe; 44-fuel swirler; 45-secondary air hole; 46-film cooling hole; 47-cold air mixing hole; 71-front end of support plate; 72-annular raised injection ring; 73-flat air ring; 74-secondary inner column. DETAILED DESCRIPTION

[0024] 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.

[0025] The embodiment of the present invention provides an afterburner continuous detonation turbine engine, Figure 1 It is a half-section view of a continuous detonation afterburning turbine engine, which includes a casing, a center body, a compressor stator blade 2, a compressor rotor blade 3, a continuous detonation main combustion chamber 4, a turbine stator blade 5, a turbine rotor blade 6 and a support plate 7; the continuous detonation afterburning turbine engine also includes a central shaft. Figure 1 The central axis is indicated by a dot-dash line. The central axis passes through the center of the central body and is used to support the central body. The central body is coaxially arranged with the outer shell and is installed in the outer shell.

[0026] like Figure 1As shown, the left side is the air intake end of the engine, and the right side is the air outlet end of the engine; along from the air intake end to the air outlet end, the inner wall of the outer shell is sequentially provided with an air intake lip 1, a compressor casing, an outer wall of the main combustion chamber, a turbine casing, an outer wall of the afterburner and a tail nozzle 8, and the center body is sequentially provided with an air intake center cone 9, a compressor drum 10, a compressor impeller 11, an inner wall of the main combustion chamber, a turbine impeller 12 and a straightening exhaust tail cone 13; along the radial direction of the central axis, the air intake lip 1 is opposite to the air intake center cone 9, the compressor casing is opposite to the compressor drum 10 and the compressor impeller 11, the outer wall of the main combustion chamber is opposite to the inner wall of the main combustion chamber, the turbine casing is opposite to the turbine impeller 12, the straightening exhaust tail cone 13 is opposite to the junction of the outer wall of the afterburner and the turbine casing, and the tail nozzle 8 is located at the tail end of the outer shell. The rectifying exhaust tail cone 13 is fixedly connected to the outer wall of the afterburner chamber, and an afterburner chamber 14 is formed at the rear side of the rectifying exhaust tail cone 13 . The combustion gas passing through the rectifying exhaust tail cone 13 enters the afterburner chamber 14 .

[0027] The air inlet lip 1 is a pointed lip and is spaced apart on the outer peripheral side of the air inlet center cone 9; part of the air inlet center cone 9 extends out of the outer shell; the air inlet center cone 9 is a cone-shaped body; the air inlet lip 1 and the air inlet center cone 9 cooperate to achieve a change in the flow area.

[0028] The compressor stator blades 2 are bolted to the inner wall of the compressor casing, forming an annular flow channel between the compressor stator blades 2. The compressor rotor blades 3 are fixedly mounted on the outer periphery of the compressor impeller 11. The compressor drum 10 connects the compressor rotor blades 3 in series. The compressor stator blades 2, the compressor rotor blades 3, the compressor impeller 11, and the compressor drum 10 constitute the compressor. The turbine stator blades 5 are fixedly mounted to the inner wall of the turbine casing. Specifically, one end of the turbine stator blade 5 is embedded in the inner wall of the turbine casing via a locking plate. The turbine rotor blades 6 are fixedly mounted to the turbine impeller 12. The turbine stator blades 5, the turbine rotor blades 6, and the turbine impeller 12 constitute the turbine.

[0029] like Figure 1 As shown, a continuous detonation main combustion chamber 4 is arranged in the annular cavity between the outer wall of the main combustion chamber and the inner wall of the main combustion chamber; Figure 2As shown, the continuous detonation main combustion chamber 4 consists of a shell 41 fixedly connected to the inner wall of the outer shell and a central inner column 42 coaxially arranged in the shell 41. An annular inner cavity is formed between the shell 41 and the central inner column 42. The shell 41 is provided with a pre-detonation tube 43, film cooling holes 46, and cold air mixing holes 47. The pre-detonation tube 43 and film cooling holes 46 are both arranged at the front end of the shell 41; the cold air mixing holes 47 are arranged at the rear end of the shell 41. The film cooling holes 46 and the cold air mixing holes 47 are arranged in two rows. Each film cooling hole 46 is connected to the low-pressure plenum via a capillary tube, and the low-pressure plenum is connected to the low-pressure compressor outlet. Each cold air mixing hole 47 is connected to the high-pressure plenum via a capillary tube, and the high-pressure plenum is connected to the high-pressure compressor outlet. The central inner column 42 is fixedly connected to the turbine casing, and the central inner column 42 can be fixedly connected to the turbine casing by embedding the rear end into the mortise and tenon of the inner wall of the turbine casing; a fuel channel and a secondary air channel are provided inside the central inner column 42; a fuel swirler 44 connected to the fuel channel and a secondary air hole 45 connected to the secondary air channel are arranged on the central inner column 42.

[0030] like Figure 1 and Figure 3 As shown, support plates 7 are evenly distributed circumferentially within the afterburner 14 and are used to inject secondary fuel and mixed air into the afterburner 14. At the end of the support plate 7 facing the tail nozzle 8, a secondary inner column 74, a flat air ring 73, and an annular raised injection ring 72 are coaxially arranged from the inside out. The flat air ring 73 and the annular raised injection ring 72 form a double-ring concave-convex structure. The annular raised injection ring 72 is used to axially inject secondary fuel into the afterburner 14 at high pressure, mixing it with the combustion chamber exhaust for combustion. The flat air ring 73 is used to axially inject air into the afterburner 14 to achieve efficient mixing and combustion of the inner ring fuel. The support plate 7 adopts a triangular prism structure with a thickness that gradually increases from the air inlet end to the air outlet end. The end of the support plate 7 facing the exhaust tail cone 13 is the support plate front end 71, which is the tip.

[0031] In the above embodiment, the tail nozzle 8 adopts a Laval flow channel design. The characteristic of a Laval flow channel is that the inner diameter of the front half gradually decreases toward the center to a narrow throat, and then gradually increases outward. The shell 41 of the continuous detonation main combustion chamber 4 can be 180 mm long and have an outer diameter of 120 mm. The outer diameter of the central inner column 42 is 60 mm. Each row of film cooling holes 46 has 20-40 holes evenly distributed along the circumference, and the diameter of the film cooling holes 46 is 1 mm. Each row of cold air mixing holes 47 has 20-40 holes evenly distributed along the circumference, and the diameter of the cold air mixing holes 47 is 1.5 mm. The front end of the central inner column 42 has 10-20 fuel swirlers 44 evenly distributed along the circumference, and the rear end has two rows of secondary air holes 45, each row has 20-40 holes evenly distributed along the circumference, and the diameter of the secondary air holes 45 is 1 mm. The inner diameter of the pre-detonation tube 43 is 80 mm. The outer diameter of the annular raised oil injection ring 72 is 40 mm, the inner diameter is 30 mm, and the raised height is 3 mm; the outer diameter of the plane air ring 73 is 30 mm, the inner diameter is 15 mm; the outer diameter of the secondary inner column 74 is 15 mm, and the length is 10 mm.

[0032] The above-mentioned afterburner continuous detonation turbine engine is provided with a continuous detonation main combustion chamber 4 between the outer wall of the main combustion chamber and the inner wall of the main combustion chamber, and an afterburner combustion chamber 14 is formed on the rear side of the rectifying exhaust tail cone 13, and support plates 7 are evenly distributed along the circumference in the afterburner combustion chamber 14. The afterburner combustion chamber 14 and the continuous detonation turbine engine are integrated into a design to meet the requirements of high maneuverability and emergency acceleration of the continuous detonation turbine engine.

[0033] The above-mentioned afterburner continuous detonation turbine engine is provided with circumferentially uniformly distributed fuel swirlers 44 on the central inner column 42 of the continuous detonation main combustion chamber 4, and the fuel swirlers 44 are mixed with the incoming air, which is beneficial to shortening the mixing distance, improving the mixing quality, and enhancing the detonation propagation stability.

[0034] The operating principle of the aforementioned afterburning continuous detonation turbine engine is as follows: compressed air flows through the intake duct and compressor into the continuous detonation main combustion chamber 4. The annularly distributed fuel swirlers 44 efficiently mix the ejected fuel jet with the incoming air, forming a fuel-rich premix. To further enhance mixing and reduce the equivalence ratio of the fuel-rich premix to an explosive range, high-pressure secondary air is injected through the secondary air holes 45 to further interact with the fuel-rich premix, forming a highly explosive premix. A detonation wave is initiated from the pre-detonation tube 43 and propagates into the annular cavity of the continuous detonation main combustion chamber 4, thereby initiating detonation in the continuous detonation main combustion chamber 4. Because the rotating detonation wave propagates at the front end of the combustion chamber, and the highest temperature is generally located in the combustion chamber casing 41, cold air is injected through the film cooling holes 46 to achieve full air film coverage of the outer wall surface of the continuous detonation main combustion chamber 4, thereby cooling the combustion chamber casing 41. Considering that the exhaust gas from the continuous detonation main combustion chamber 4 is hotter than the exhaust gas from the combustion chamber of a traditional aircraft engine, cold air is injected at high pressure from the cold air mixing holes 47 evenly distributed at the rear end of the combustion chamber to cool the exhaust gas from the combustion chamber and reduce damage to the turbine blades.

[0035] When the afterburner continuous detonation turbine engine is in the non-afterburner state, the above-mentioned detonation combustion exhaust gas is adjusted to the axial direction by the rectifying exhaust tail cone 13, and then expands through the turbine to do work, and further expands in the Laval nozzle to maximize the utilization of energy. When the engine is in the afterburner state, the annular protruding injection ring 72 at the tail of the support plate 7 axially injects secondary fuel at high pressure, which is mixed and burned with the exhaust gas in the combustion chamber. In order to further improve the fuel utilization rate and shorten the combustion mixing distance, the flat air ring 73 located at the center of the annular protruding injection ring 72 axially injects air to achieve further efficient mixing and combustion of the inner ring fuel. The double-ring concave-convex structure at the tail of the support plate 7 and the design of the secondary inner column 74 induce a large number of vortices in the axial and lateral directions, which greatly enhances combustion and achieves the goal of releasing a large amount of energy in a short time, meeting the high maneuverability and emergency acceleration requirements of the continuous detonation turbine engine.

[0036] Example 1 like Figure 2As shown, this embodiment provides a schematic diagram of the continuous detonation main combustion chamber 4 in the above-mentioned afterburning continuous detonation turbine engine. The total length of the combustion chamber is approximately 180 mm. The outer diameter of the shell 41 is 120 mm, and the outer diameter of the central inner column 42 is 60 mm. A pre-detonation tube 43 is mounted at the front end of the shell 41. A pair of exhaust film cooling holes 46 are evenly distributed along the circumference, with 30 holes per row and a diameter of 1 mm. At the rear end of the shell 41, two rows of cold air mixing holes 47 are evenly distributed along the circumference, with 30 holes per row and a diameter of 1.5 mm. Ten fuel swirlers 44 are evenly distributed along the circumference at the front end of the central inner column 42. At the rear end of the central inner column 42, two rows of secondary air holes 45 are evenly distributed along the circumference, with 30 holes per row and a diameter of 1 mm. The length of the pre-explosion tube 43 is 300 mm and the inner diameter is 80 mm.

[0037] like Figure 3 As shown, this embodiment provides a schematic structural diagram of a support plate 7. The front end 71 of the support plate 7 is 100 mm long, 50 mm high, and 8 mm thick. Five annular raised oil injection rings 72, five flat air rings 73, and five secondary inner pillars 74 are arranged on the rear end surface of the support plate 7. The annular raised oil injection rings 72 have an outer diameter of 40 mm, an inner diameter of 30 mm, and a raised height of 3 mm. The flat air rings 73 have an outer diameter of 30 mm and an inner diameter of 15 mm. The secondary inner pillars 74 have an outer diameter of 15 mm and a length of 10 mm.

[0038] Obviously, those skilled in the art may make various changes and modifications to the embodiments of the present invention without departing from the spirit and scope of the present invention. Thus, if such modifications and variations of the present invention fall within the scope of the claims and their equivalents, the present invention is intended to include such modifications and variations.

Claims

1. A continuous detonation afterburning turbine engine, characterized in that: It includes the outer casing, center body, compressor stator blades, compressor rotor blades, continuous detonation main combustion chamber, turbine stator blades, turbine rotor blades and support plates; From the air inlet end to the air outlet end, the inner wall of the shell is provided with the air inlet lip, the compressor casing, the outer wall of the main combustion chamber, the turbine casing, the outer wall of the afterburner, and the tail nozzle in sequence. The center body is provided with the air inlet center cone, the compressor drum, the compressor wheel, the inner wall of the main combustion chamber, the turbine wheel, and the rectifying exhaust tail cone in sequence. The afterburner is formed at the rear side of the rectifying exhaust tail cone, and the combustion gas passing through the rectifying exhaust tail cone enters the afterburner chamber. The compressor stator blades are fixedly mounted on the inner wall of the compressor casing to form an annular flow channel; the compressor rotor blades are fixedly mounted on the compressor wheel; the turbine stator blades are fixedly mounted on the inner wall of the turbine casing; the turbine rotor blades are fixedly mounted on the turbine wheel; A continuous detonation main combustion chamber is arranged between the outer wall of the main combustion chamber and the outer wall of the main combustion chamber; the continuous detonation main combustion chamber consists of a shell fixedly connected to the inner wall of the outer shell and a central inner column coaxially arranged in the shell; an annular inner cavity is formed between the shell and the central inner column; a pre-detonation tube, air film cooling holes and cold air mixing holes are arranged on the shell; each air film cooling hole is connected to the low-pressure air collecting cavity; each cold air mixing hole is connected to the high-pressure air collecting cavity; the central inner column is fixedly connected to the turbine casing, and a fuel channel and a secondary air channel are arranged inside; a fuel swirler connected to the fuel channel and a secondary air hole connected to the secondary air channel are arranged on the central inner column; The support plates are evenly distributed in the afterburner combustion chamber along the circumferential direction; the support plates are provided with a secondary inner column, a plane air ring and an annular raised fuel injection ring coaxially arranged from the inside to the outside at one end facing the tail nozzle; the plane air ring and the annular raised fuel injection ring form a double-ring concave-convex structure; the annular raised fuel injection ring is used to axially inject secondary fuel into the afterburner combustion chamber at high pressure, and mix it with the combustion chamber exhaust for combustion; the plane air ring is used to axially inject air into the afterburner combustion chamber to achieve efficient mixing and combustion of the inner ring fuel.

2. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The support plate adopts a triangular prism structure with a thickness gradually increasing from the air inlet end to the air outlet end, and the end of the support plate facing the rectification exhaust tail cone is the front end and the tip.

3. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The air inlet lip is a pointed lip and is spaced apart on the outer peripheral side of the air inlet center cone; the air inlet center cone is a cone; the air inlet lip and the air inlet center cone cooperate to achieve a change in the flow area.

4. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The compressor drum is connected in series with the compressor rotor blades; The low-pressure plenum is connected to the outlet of the low-pressure compressor; The high-pressure plenum is connected to the high-pressure compressor outlet.

5. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: Each air film cooling hole and cold air mixing hole is connected with a thin tube, and is connected to the corresponding low-pressure air collecting cavity and high-pressure air collecting cavity through the thin tube.

6. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The rectifying exhaust tail cone is fixedly connected to the outer wall of the afterburner chamber.

7. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The end of the central inner column fits into the tongue and groove of the inner wall of the turbine casing; One end of the turbine stator blade is embedded in the inner wall of the turbine casing through a locking piece; The compressor stator blades are fixed to the inner wall of the compressor casing by bolts.

8. The afterburner continuous detonation turbine engine according to claim 1, characterized in that: The tail nozzle adopts Laval flow channel design.

9. The afterburner continuous detonation turbine engine according to any one of claims 1 to 8, characterized in that: The pre-explosion pipe and two rows of film cooling holes are both arranged at the front end of the shell; the two rows of cold air mixing holes are arranged at the rear end of the shell; The shell is 180mm long and 120mm in outer diameter; the outer diameter of the central inner column is 60mm; There are 20-40 film cooling holes evenly distributed along the circumference of each exhaust film cooling hole, and the diameter of the film cooling hole is 1mm; Each row of cold air mixing holes has 20-40 holes evenly distributed along the circumference, and the diameter of the cold air mixing holes is 1.5mm; The front end of the central inner column is evenly distributed with 10-20 fuel swirlers along the circumference, and the rear end is distributed with two rows of secondary air holes, each row of secondary air holes is evenly distributed along the circumference, and the diameter of the secondary air holes is 1mm; The inner diameter of the pre-explosion tube is 80mm.

10. The afterburner continuous detonation turbine engine according to claim 9, characterized in that: The outer diameter of the annular raised injection ring is 40mm, the inner diameter is 30mm, and the raised height is 3mm; The outer diameter of the flat air ring is 30mm and the inner diameter is 15mm; The outer diameter of the secondary inner column is 15 mm and the length is 10 mm.

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