Multi-parallel rocket type rotary detonation engine and working method

Abstract

The multi-parallel rocket type rotary detonation engine comprises an oxidant cavity and a combustion chamber fixedly connected with the oxidant cavity, and the combustion chamber is internally provided with a plurality of fuel supply systems which communicate with one another, are fixedly connected with the oxidant cavity and are arranged around a rotation center in an array mode. And the ignition device is arranged on one fuel supply system. The multiple rocket-type rotary detonation engines are connected in parallel by sharing the oxidant cavity system, and the combustion chambers and the fuel supply system are coaxially arranged to form combustion chamber channels of the multiple rocket-type rotary detonation engines. The multiple parallel rocket type rotary detonation engines share the ignition device, ignition of the multiple parallel engines is achieved in the mode that windows are formed in the wall face, the system complexity brought by the ignition device can be obviously reduced, the reliability of the parallel rocket type rotary detonation engines can be improved, the structural design of the engines can be greatly simplified, and the cost is reduced. And the influence of the size effect is weakened on the premise of ensuring that larger thrust can be provided.

Description

Technical Field

[0001] The present invention relates to the technical field of engines, and in particular to a plurality of parallel rocket-type rotating detonation engines and a working method thereof. Background Art

[0002] At present, almost all aerospace propulsion systems use the deflagration mode to organize combustion. However, after decades of development, it is difficult for the propulsion system based on deflagration combustion to have a major breakthrough in performance, and it is difficult to significantly improve the thermal cycle efficiency and heat release rate of the engine.

[0003] Most existing rotating detonation rocket engines adopt a single-engine design. There are no public reports on the design of multiple parallel rotating detonation rocket engines. Rockets using a single engine have higher engine thrust requirements, which increases the difficulty of engine development.

[0004] The Chinese invention patent with the announcement number "CN117249019B" and the name "A rocket engine three-engine parallel structure and rocket" discloses a rocket engine three-engine parallel structure and a rocket, wherein the rocket engine three-engine parallel structure includes a frame arranged at the bottom of the rocket body and three engines fixed on the frame, and the frame includes an upper end face, a lower end face, and a plurality of main load-bearing connecting rods; the upper end face is set in a regular hexagon, and the lower end face is set in a regular triangle; the upper end face and the lower end face are connected by the main load-bearing connecting rod: the three top corners and bottom faces of the lower end face are fixed with positioning holes, and the mounting holes of the three engines are respectively connected and fixed with the positioning holes. The structural setting of this technical solution is reasonable, and it adopts a regular triangle lower end face, matched with a regular hexagon upper end face, with a simple structure, convenient disassembly and assembly, and can ensure the stability of installation; through the regular triangle lower end face, the three engines can be connected in parallel, which fills the gap of the parallel layout of three engines in the rocket field, reduces the installation cost and the engine development cost, and is conducive to improving the reliability of the overall structure. Although this technical solution can realize the parallel connection of multiple engines, its parallel structure is not completely suitable for the annular combustion chamber configuration of the rotating detonation engine, and the design of multiple connecting rods greatly increases the complexity of the system. In the above technical solution, multiple engines must use multiple ignition devices to ignite the engines one by one, which greatly increases the cost and reduces the reliability of the system, so its applicability and practicality are limited. Summary of the invention

[0005] In order to solve the problem that the parallel structure of multiple independent engines in the above-mentioned prior art is not suitable for the annular combustion chamber configuration of the rotating detonation engine and multiple independent engines in parallel require multiple ignition devices to ignite the engines one by one, which reduces the reliability of the engine system. The present invention proposes a multiple parallel rocket rotating detonation engine and a working method. The multiple parallel rocket rotating detonation engines of the present invention share a common ignition device and realize the ignition of multiple parallel engines in the form of wall windowing, which can significantly reduce the system complexity brought by the ignition device and improve the reliability of the parallel rocket rotating detonation engine.

[0006] The present invention is achieved through the following technical scheme: it includes an oxidant chamber and a combustion chamber fixedly connected to the oxidant chamber, wherein a plurality of fuel supply systems are arranged inside the combustion chamber, which are interconnected and fixedly connected to the oxidant chamber and arranged in an array around a rotation center, and also includes an ignition device arranged on one of the fuel supply systems.

[0007] Furthermore, the oxidant chamber as a whole is in the shape of a cylinder with a single-side opening and a hollow interior, including an oxidant chamber wall, a first connecting groove arranged on the opening side and fixedly connected to the combustion chamber, and a base arranged on the sealing side and fixedly connected to the fuel supply system in a one-to-one correspondence.

[0008] Furthermore, the combustion chamber is a hollow cylinder with openings on both sides, including a first connecting portion fixedly connected to the first connecting groove, a central column in the shape of an elongated ring that divides the combustion chamber into a plurality of cavities, and a window on the central column that is connected to a fuel supply system.

[0009] Furthermore, the fuel supply system is cylindrical as a whole, including a fixing rod fixedly connected to the base, a cylindrical fuel storage tank arranged on a side close to the fixing rod, spray holes arranged at equal intervals in the circumferential direction on a side away from the fixing rod and penetrating the wall of the fuel storage tank, and an annular inner cylinder with a certain wall thickness arranged away from the fixing rod.

[0010] Furthermore, a first roar duct outer wall surface is provided on the inner wall of the combustion chamber and on one end of the central column close to the first connecting portion, and a second roar duct outer wall surface is provided at a corresponding position on the fuel supply system.

[0011] Furthermore, the outer wall surface of the first roar duct and the outer wall surface of the second roar duct are both a section of contraction and expansion curved surface, consisting of a contraction section and an expansion section, with a contraction half-angle of 25° to 45°, an expansion half-angle of 15° to 30°, a length of the contraction section being smaller than a length of the expansion section, and a ratio of the length of the contraction section to the length of the expansion section being 2:3.

[0012] Furthermore, it also includes a conical center cone fixedly connected to the fuel supply system at one end away from the oxidizer chamber, a second connecting groove is provided on the inner tube at one end away from the fixed rod, the center cone includes a second connecting part fixedly connected to the second connecting groove and a center cone wall surface provided at one end close to the second connecting part, and the cone of the center cone is oriented in a direction away from the fixed rod.

[0013] Furthermore, the wall surface of the central cone expands first and then contracts, and is divided into an expansion section and a contraction section. The starting point of the expansion section is the vertex of the central cone, and the end point of the expansion section intersects with the starting point of the contraction section. The angle between the tangent at the end point of the contraction section and the centerline of the central cone is 25° to 45°. The cone angle of the central cone is 15° to 35°. The length of the contraction section is smaller than that of the expansion section, and the ratio of the length of the contraction section to the length of the expansion section is 1:8 to 12.

[0014] Furthermore, the number of the fuel supply systems is seven, with one arranged in the center and six arranged in the circumference, and the diameter of the spray hole is 0.3-1.5 mm.

[0015] The present invention also provides a working method applicable to the multiple parallel rocket rotating detonation engines of the present invention, comprising the following steps:

[0016] Step 1: The oxidant is injected into the combustion chamber through the roar channel;

[0017] Step 2: After the oxidant flows stably in the combustion chamber, the fuel supply system injects fuel through the nozzles to mix with the oxidant in the combustion chamber;

[0018] Step 3: The ignition device is started to ignite the combustible mixture;

[0019] Step 4: One or more detonation waves are formed in the combustion chamber, propagating in a circumferential direction and consuming fuel;

[0020] Step 5: The flame in the central detonation engine spreads to the detonation engines arranged around the periphery and ignites the combustible mixture therein to form a rotating detonation wave;

[0021] Step 6: After the rotating detonation wave in each detonation engine is generated, the detonation products rapidly expand and flow axially after passing through the central cone, and are discharged at the outlet of the combustion chamber to generate thrust.

[0022] Compared with the prior art, the present invention has the following beneficial effects:

[0023] 1. The rocket rotating detonation engine of the present invention adopts rotating detonation to organize combustion. Compared with traditional rocket engines, the engine has the advantages of simple structure, short length, large thrust-to-weight ratio, high thermal cycle efficiency, etc.

[0024] 2. A parallel form is proposed for rocket rotating detonation engines, filling the gap of multiple parallel layouts in the field of rocket rotating detonation engines.

[0025] 3. The present invention adopts a parallel form of a common supply system, which can simplify the structural design of the engine and reduce the influence of the size effect while ensuring that it can provide greater thrust.

[0026] 4. The present invention adopts a parallel form of a common supply system, and does not need to use a thrust connecting rod to transmit the engine thrust to the rocket body, thereby greatly reducing the weight of the system.

[0027] 5. The present invention adopts a common ignition device in combination with a wall window to realize the ignition of multiple parallel engines, which can significantly reduce the system complexity brought by the ignition device and improve the reliability of the parallel rocket rotating detonation engine.

[0028] 6. The fuel is injected into the combustion chamber through a nozzle, which can ensure the stability and reliability of fuel injection during the operation of the rocket rotating detonation engine.

[0029] 7. When the oxidant is injected through the annular contraction-expansion flow channel, it can flow into the combustion chamber at a supersonic speed, thereby enhancing the ability of the air supply system to resist detonation back pressure and improving the stability of detonation combustion.

[0030] 8. The ignition system adopts a common ignition device combined with a wall window, which can reduce the number of ignition devices, improve the reliability of the system, and significantly reduce costs and complexity.

[0031] 9. A tail center cone is also provided to help avoid or reduce the generation of shock waves, reduce the total pressure loss of the engine, and thus improve the efficiency and stability of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 It is a schematic diagram of the overall structure of the rotating detonation engine in the present invention.

[0033] Figure 2 It is a cross-sectional view of the overall structure of the rotating detonation engine in the present invention.

[0034] Figure 3 It is a partial cross-sectional view of the oxidant chamber in the present invention.

[0035] Figure 4 It is a partial cross-sectional view of the outer wall of the combustion chamber in the present invention.

[0036] Figure 5 It is a cross-sectional view of the fuel supply system for a single combustion chamber in the present invention.

[0037] Figure 6 It is a cross-sectional view of the central cone portion of the present invention.

[0038] Indicated in the figure:

[0039] 1. Oxidant chamber; 2. Combustion chamber; 3. Fuel supply system; 4. Center cone; 5. Ignition device; 101. Oxidant chamber wall; 102. Base; 103. First connecting groove; 201. First connecting part; 202. First roar channel outer wall; 203. Center column; 204. Window; 301. Fixing rod; 302. Fuel storage tank; 303. Second roar channel outer wall; 304. Nozzle; 305. Inner tube; 306. Second connecting groove; 401. Second connecting part; 402. Center cone wall. DETAILED DESCRIPTION

[0040] Advantages and features of the present invention will be illustrated and explained by the following non-limiting description of preferred embodiments thereof, which are given by way of example only with reference to the accompanying drawings.

[0041] like Figure 1 and Figure 2 As shown, the present invention provides a plurality of parallel rocket-type rotating detonation engines, including an oxidizer chamber 1 and a combustion chamber 2 fixedly connected to the oxidizer chamber 1, wherein a plurality of fuel supply systems 3 connected to each other are arranged inside the combustion chamber 2. The fuel supply system 3 is fixedly connected to the oxidizer chamber 1, and there is a certain distance between the fuel supply system 3 and the combustion chamber 2. The fuel supply systems 3 are arranged in an array around the center of rotation, preferably seven in number, with six arranged in a circle in the center, and an ignition device 5 is arranged on one fuel supply system 3, preferably on a fuel supply system 3 at the center. The rotating detonation engine has an inlet at one end close to the oxidizer chamber 1 and an outlet at the end away from the oxidizer chamber 1. Each fuel supply system 3 and the remaining components can form a detonation engine.

[0042] like Figure 3 As shown, the oxidant chamber 1 is in the shape of a cylinder with a single-side opening and a hollow interior, and includes an oxidant chamber wall 101, a base 102, and a first connecting groove 103. The first connecting groove 103 is arranged on one side of the opening, and the oxidant chamber 1 is fixedly connected to the combustion chamber 2 through the first connecting groove 103. The number of the bases 102 is the same as the number of the fuel supply system 3, and is arranged on the sealed side of the oxidant chamber 1 for fixed connection with the fuel supply system 3. Preferably, the first connecting groove 103 and the base 102 are both provided with internal threads, the first connecting groove 103 is fixedly connected to the combustion chamber 2 by threads, and the base 102 is fixedly connected to the fuel supply system 3 by threads.

[0043] The oxidant chamber 1 is mainly used to store and supply the oxidant necessary for the rocket-type rotating detonation engine, preferably oxygen which has more practical value.

[0044] The base 102 is located at the bottom end of the inner wall of the oxidant chamber 1 , that is, the sealed side. The inner wall of the base 102 is arranged with an internal thread. The number of the bases 102 is consistent with the number of the fuel supply systems 3 .

[0045] In this embodiment, a base 102 is arranged at the center of the bottom end of the inner wall surface of the oxidant chamber 1, and the other bases 102 are arranged in an array along the circumferential direction.

[0046] like Figure 4 As shown, the combustion chamber 2 is a hollow cylinder with openings on both sides. The combustion chamber 2 includes a first connection portion 201, a first roar channel outer wall 202, a central column 203 and a window 204. The first connection portion 201 is fixedly connected to the first connection groove 103. Preferably, an external thread is provided at the first connection portion 201, which is fixed with the internal thread at the first connection groove 103. To ensure the strength of the thread, the first connection portion 201 can appropriately protrude from the outer wall of the combustion chamber 2. The central column 203 is in an oblong shape as a whole, which divides the combustion chamber 2 into a plurality of cavities. The number of the cavities is consistent with the number of the fuel supply system 3, and is used to place the fuel supply system 3. A plurality of windows 204 are provided on the central column 203, and the number of the windows 204 is one less than the number of the fuel supply system 3, which are used to connect the interior of the central column 203 with the separated cavities. The inner wall of the combustion chamber 2 and the central column 203 are both provided with a first roar channel outer wall 202. The first roar channel outer wall surface 202 is disposed at an end close to the first connecting portion 201 .

[0047] An external thread is arranged on the outer surface of the first connecting portion 201, and the first connecting portion 201 matches the thread of the first connecting groove 103, so that the combustion chamber 2 and the oxidant chamber 1 are coaxially fixedly connected through the thread.

[0048] like Figure 4 As shown, the first roar channel outer wall surface 202 is a section of contraction and expansion curved surface. The first roar channel outer wall surface 202 is a C-shaped curved surface as a whole. The starting point and the end point of the first roar channel outer wall surface 202 are both the main body, and the opening faces the main body. The main body is the inner wall surface of the combustion chamber 2 or the outer wall surface of the center column. In order to avoid flow loss, the first roar channel outer wall surface 202 is composed of two sections of curved surfaces. The starting point of one section of the curved surface is the main body, and the end point of the other section of the curved surface is the main body. The other ends of the two sections of the curved surface intersect, that is, the end point of one section of the curved surface is the starting point of the other section of the curved surface. The curved surface with the starting point as the main body is the expansion section, which is located far away from the oxidant chamber 1; the curved surface with the end point as the main body is the contraction section, which is located close to the oxidant chamber 1. The contraction half-angle β1 is the angle between the tangent at the end point of the contraction section and the center line of the main body, preferably 25° to 45°. The expansion half-angle β2 is the angle between the tangent at the starting point of the expansion section and the center line of the main body, preferably 15° to 30°. The contraction section length D1 is smaller than the expansion section length D2, and the ratio of the contraction section length D1 to the expansion section length D2 is preferably 2:3.

[0049] The first roar channel outer wall surface 202 is coaxially arranged with the base 102 .

[0050] Through the window 204, the flame of the centrally arranged fuel supply system 3 can propagate to the surrounding fuel supply systems 3 and ignite the combustible mixture.

[0051] like Figure 5 As shown, the fuel supply system 3 includes a fixing rod 301, a fuel storage tank 302, a second roar channel outer wall 303, a spray hole 304, an inner tube 305 and a second connection groove 306. The fuel supply system 3 is cylindrical as a whole, and a fixing rod 301 is arranged at one end, and the fixing rod 301 is fixedly connected to the base 102. Preferably, the fixing rod 301 is provided with a thread, and the fixing rod 301 is fixedly connected to the base 102 through the thread. The fuel storage tank 302 is used to store the fuel required by the engine, and the fuel is preferably liquid kerosene with more practical value. The fuel storage tank 302 is a cylindrical cavity structure, which is arranged on the side close to the fixing rod 301.

[0052] The second roar duct outer wall surface 303 is coaxially arranged on the fuel storage tank 302 , and its cross-sectional area increases first and then decreases, and its design angle and length ratio are consistent with those of the first roar duct outer wall surface 202 .

[0053] The outer wall surface 303 of the second roar channel is coaxial with the outer wall surface 202 of the first roar channel and matches the same straight line, forming an annular flow channel that first contracts and then expands, namely, a roar channel, so that the oxidant forms a supersonic flow after passing through the annular flow channel, thereby enhancing the ability of the air supply system to resist the detonation back pressure and improving the stability of the detonation combustion.

[0054] Spray holes 304 are arranged at equal intervals along the circumferential direction away from the fixing rod 301 and the outer wall surface 303 of the second roar duct. The spray holes 304 penetrate the wall surface of the fuel storage tank 302. The diameter of the spray holes 304 is preferably 0.3-1.5 mm.

[0055] The inner cylinder 305 is a circular ring structure with a certain wall thickness and is arranged away from the fixing rod 301. A second connecting groove 306 is arranged at one end of the inner cylinder 305 away from the fixing rod 301. Preferably, a thread is arranged on the second connecting groove 306.

[0056] like Figure 6As shown, preferably, the multiple parallel rocket rotating detonation engines are further provided with a plurality of center cones 4 at one end away from the oxidizer chamber 1. The number of the center cones 4 is consistent with the number of the fuel supply system 3. The center cone 4 is a centrally symmetrical cone structure, including a second connecting portion 401 and a center cone wall 402. Preferably, a thread is provided at the second connecting portion 401. The cone of the center cone 4 faces away from the fixing rod 301, and the center cone 4 is fixedly connected to the fuel supply system 3, preferably fixedly connected to the second connecting groove 306 through the second connecting portion 401, and the connection method is a threaded connection, and each of the center cones 4 is coaxial with the corresponding fuel supply system 3.

[0057] like Figure 6 As shown, the central cone wall 402 is arranged at one end close to the second connecting portion 401, and the central cone wall 402 is in a shape that expands first and then contracts, and is divided into an expansion section and a contraction section. The starting point of the expansion section is the vertex of the central cone 4, and the end point of the expansion section intersects with the starting point of the contraction section. The contraction half angle φ is the angle between the tangent at the end point of the contraction section and the center line of the central cone 4, which is preferably 25° to 45°. The cone angle α of the central cone 4 is preferably 15° to 35°. The length L1 of the contraction section is smaller than the length L2 of the expansion section, and the ratio of the length L1 of the contraction section to the length L2 of the expansion section is preferably 1:8 to 12.

[0058] The center cone wall 402 and the combustion chamber 2 form an annular channel that first contracts and then expands. The design of the center cone 4 helps to avoid or reduce the generation of shock waves, reduce the total pressure loss of the engine, and thus improve the efficiency and stability of the engine.

[0059] The ignition device 5 is preferably a solid powder ignition device.

[0060] The present invention also provides a working method of a parallel rocket rotating detonation engine, comprising the following steps.

[0061] Step 1: The oxidant and the fuel are stored in the oxidant chamber 1 and the fuel supply system 3 respectively. Multiple detonation engines share the same oxidant chamber 1, and the oxidant is injected into the combustion chamber 2 through the roar channel.

[0062] Step 2: After the high-speed oxidant flows stably in the combustion chamber 2 , the fuel supply system 3 starts to inject fuel through the nozzles 304 to mix with the oxidant in the combustion chamber 2 .

[0063] Step 3: An ignition device 5 is arranged on the fuel supply system 3 located in the center. After the fuel and the oxidant are mixed and stabilized, the ignition device 5 is started to ignite the combustible mixture.

[0064] Step 4: After ignition, one or more detonation waves are formed in the combustion chamber 2, propagating in a circumferential direction and consuming fuel, while the combustion chamber 2 itself is stationary in the circumferential direction.

[0065] Step 5: Since the combustion chamber 2 is provided with a window 204, the flame in the detonation engine located in the center will spread to the detonation engines arranged around it and ignite the combustible mixture therein to form a rotating detonation wave.

[0066] Step 6: After the rotating detonation wave in each detonation engine is generated, the detonation products rapidly expand and flow axially after passing through the central cone 4, and are discharged at a high speed at the outlet of the combustion chamber 2 to generate thrust.

[0067] In addition to the above embodiments, the present invention may also have other implementation modes. Any technical solutions formed by equivalent replacement or equivalent transformation shall fall within the protection scope required by the present invention.

Claims

1. A plurality of parallel rocket-type rotating detonation engines, comprising an oxidant chamber (1) and a combustion chamber (2) fixedly connected to the oxidant chamber (1), characterized in that: The combustion chamber (2) is provided with a plurality of fuel supply systems (3) which are interconnected, fixedly connected to the oxidant chamber (1) and arranged in an array around a rotation center, and also includes an ignition device (5) arranged on one of the fuel supply systems (3).

2. A plurality of parallel rocket rotating detonation engines according to claim 1, characterized in that: The oxidant chamber (1) is generally in the shape of a cylinder with a single-side opening and a hollow interior, and comprises an oxidant chamber wall (101), a first connecting groove (103) arranged on the opening side and fixedly connected to the combustion chamber (2), and a base (102) arranged on the sealing side and fixedly connected to the fuel supply system (3) in a one-to-one correspondence.

3. A plurality of parallel rocket rotating detonation engines according to claim 2, characterized in that: The combustion chamber (2) is a hollow cylinder with openings on both sides, comprising a first connection portion (201) fixedly connected to the first connection groove (103), a central column (203) in the shape of an oblong ring that divides the combustion chamber (2) into a plurality of cavities, and a window (204) provided on the central column (203) and connected to the fuel supply system (3).

4. A plurality of parallel rocket rotating detonation engines according to claim 3, characterized in that: The fuel supply system (3) is cylindrical in shape as a whole, and comprises a fixing rod (301) fixedly connected to the base (102), a cylindrical fuel storage tank (302) arranged on a side close to the fixing rod (301), spray holes (304) arranged at equal intervals in the circumferential direction on a side away from the fixing rod (301) and penetrating the wall of the fuel storage tank (302), and a circular inner cylinder (305) with a certain wall thickness arranged away from the fixing rod (301).

5. A plurality of parallel rocket rotating detonation engines according to claim 4, characterized in that: A first roar duct outer wall surface (202) is provided on the inner wall of the combustion chamber (2) and on one end of the central column (203) close to the first connecting portion (201), and a second roar duct outer wall surface (303) is provided at a corresponding position on the fuel supply system (3).

6. A plurality of parallel rocket rotating detonation engines according to claim 5, characterized in that: The first roar duct outer wall surface (202) and the second roar duct outer wall surface (303) are both a section of contraction and expansion curved surface, consisting of a contraction section and an expansion section, wherein the contraction half angle is 25° to 45°, the expansion half angle is 15° to 30°, the length of the contraction section is smaller than the length of the expansion section, and the ratio of the contraction section length to the expansion section length is 2:

3.

7. A plurality of parallel rocket rotating detonation engines according to claim 6, characterized in that: It also includes a conical central cone (4) arranged at one end away from the oxidant chamber (1) and fixedly connected to the fuel supply system (3); a second connecting groove (306) is arranged on the end of the inner tube (305) away from the fixed rod (301); the central cone (4) includes a second connecting portion (401) fixedly connected to the second connecting groove (306) and a central cone wall surface (402) arranged at one end close to the second connecting portion (401); the cone of the central cone (4) is oriented in a direction away from the fixed rod (301).

8. A plurality of parallel rocket rotating detonation engines according to claim 7, characterized in that: The central cone wall (402) is in a shape that expands first and then contracts, and is divided into an expansion section and a contraction section. The starting point of the expansion section is the vertex of the central cone (4), and the end point of the expansion section intersects with the starting point of the contraction section. The angle between the tangent at the end point of the contraction section and the center line of the central cone (4) is 25° to 45°. The cone angle of the central cone (4) is 15° to 35°. The length of the contraction section is shorter than that of the expansion section, and the ratio of the length of the contraction section to the length of the expansion section is 1:8 to 12.

9. A plurality of parallel rocket rotating detonation engines according to claim 4, characterized in that: The number of the fuel supply systems (3) is seven, with one arranged in the center and six arranged in a circle, and the diameter of the spray hole (304) is 0.3-1.5 mm.

10. A method for operating a plurality of parallel rocket rotating detonation engines according to any one of claims 7 to 9, characterized in that: The steps include: Step 1: The oxidant is injected into the combustion chamber (2) through the roar channel; Step 2: After the oxidant flows stably in the combustion chamber (2), the fuel supply system (3) injects fuel through the nozzles (304) to mix with the oxidant in the combustion chamber (2); Step 3: The ignition device (5) is started to ignite the combustible mixture; Step 4: one or more detonation waves are formed in the combustion chamber (2), propagating in a circumferential direction and consuming fuel; Step 5: The flame in the central detonation engine spreads to the detonation engines arranged around the periphery and ignites the combustible mixture therein to form a rotating detonation wave; Step 6: After the rotating detonation wave in each detonation engine is generated, the detonation products rapidly expand and flow in the axial direction after passing through the central cone (4), and are discharged at the outlet of the combustion chamber (2) to generate thrust.

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