Series-flow dual combustion chamber rotating detonation ramjet engine and method of operation

Abstract

The application provides a series type double combustion chamber rotating detonation ramjet engine and a working method, which comprises a combustion chamber outer shell, an air inlet channel, a separation section, a first stage combustion chamber, a first stage contraction section, a second stage combustion chamber, a second stage contraction section, an outlet nozzle, a combustion chamber center body, a first stage fuel injection cavity, a second stage fuel injection cavity, a first stage fuel injection hole, a second stage fuel injection hole, a center body front support and a center body rear support. The application is aimed at the rotating detonation combustion organization flow characteristics under high or low Mach number working conditions, realizes the compatibility of high or low Mach number rotating detonation combustion working conditions in the same flow channel, and thus widens the working Mach number range of the rotating detonation ramjet engine. Considering the large difference in the combustion back pressure transmission distance under the high / low incoming flow Mach number working conditions, the two combustion chambers are arranged in series, and the combustion organization working mode is reasonably arranged, so that the combustion back pressure transmission problem under the wide incoming flow Mach number condition can be solved.

Description

Technical Field

[0001] This invention belongs to the field of rotary detonation ramjet engine technology, specifically relating to a tandem dual-combustion chamber rotary detonation ramjet engine and its operating method. Background Technology

[0002] Rotating detonation combustion, as a novel combustion mode, offers advantages such as rapid heat release, continuous operation upon single ignition, and higher thermal efficiency. It also boasts unique advantages in shortening combustion chamber length, reducing structural weight, and improving propulsion performance. It can be applied in aero-engines, rocket engines, and ramjet engines to replace existing detonation combustion modes, showing broad application prospects in the aerospace propulsion field. The rotating detonation ramjet engine is a new type of ramjet engine that utilizes rotating detonation combustion. Its main operating process is as follows: a high-speed airflow is captured and decelerated through the inlet, then enters the annular combustion chamber through an isolator section. There, it mixes thoroughly with fuel injected through the nozzle to form a combustible mixture. Subsequently, ignition initiates a detonation wave within the annular combustion chamber, which propagates at high speed and frequency along the circumference of the combustion chamber, achieving rapid heat release from the fuel. Following the detonation wave, high-temperature combustion products propagate downstream and are accelerated and discharged through the nozzle, thus converting engine thrust into output.

[0003] Detonation waves exhibit good adaptability to incoming flow. Previous research has shown that rotating detonation in ramjet mode can achieve efficient and stable combustion at different incoming flow Mach numbers. However, the combustion chamber configuration for stable combustion under rotating detonation varies across different Mach number ranges, making it difficult to achieve stable and efficient combustion over a wide Mach range in a single combustion chamber. Furthermore, under different flight Mach numbers, the high-speed airflow conditions change drastically, resulting in significant differences in the isolator's ability to resist combustion back pressure. At higher Mach numbers, the shock wave propagation distance in the isolator is shorter, requiring a shorter isolator length; however, at lower Mach numbers, the shock wave propagation distance is significantly longer, necessitating a longer isolator to prevent combustion back pressure propagation. The required isolator length differs considerably between high and low incoming flow Mach number ranges. To achieve stable operation of a rotating detonation ramjet engine across a wide incoming flow Mach number range, the isolator design must be carefully considered.

[0004] Therefore, to address the aforementioned problems, it is crucial to design a novel rotating detonation combustor configuration that achieves controllable back pressure forward propagation over a wide Mach number range and high combustion performance within the same flow channel. This invention proposes a tandem dual-combustion-chamber rotating detonation ramjet engine combustor configuration that can solve the aforementioned problems. Summary of the Invention

[0005] To achieve the above-mentioned objectives, the technical solution of this invention is as follows:

[0006] A series dual-combustion chamber rotary detonation ramjet engine includes an outer shell 1, a central body 2 placed inside the outer shell 1, and the central body 2 connected to the outer shell 1 through a central body front support 3 and a central body rear support 4.

[0007] An annular flow channel is formed between the outer shell 1 and the central body 2. Along the direction of airflow propagation, the annular flow channel includes, in sequence, an intake duct 5, an isolation section 6, a first-stage combustion chamber 10, a first-stage contraction section 11, a second-stage combustion chamber 14, a second-stage contraction section 15, and an outlet nozzle 16.

[0008] The central body 2 is equipped with a fuel tank 9; the front support 3 of the central body is located in the isolation section 6 between the intake duct 5 and the first-stage combustion chamber 10, and the rear support 4 of the central body is located in the second-stage converging section 15 between the second-stage combustion chamber 14 and the outlet nozzle 16; in order to ensure airflow, the front support 3 and the rear support 4 of the central body are evenly distributed circumferentially.

[0009] Both the first-stage injection chamber 7 and the second-stage injection chamber 12 are dual-injection-chamber structures, with the two injection chambers respectively located on the outer shell 1 and the central body 2. Along the airflow propagation direction, the first-stage injection chamber 7 is located between the tail end of the isolation section 6 away from the incoming flow and the front end of the first-stage combustion chamber 10 near the incoming flow. Along the airflow propagation direction, the second-stage injection chamber 12 is located between the tail end of the first-stage contraction section 11 away from the incoming flow and the front end of the second-stage combustion chamber 14 near the incoming flow.

[0010] The first-stage injection chamber 7 has a first-stage injection hole 8 on both the outer shell 1 and the central body 2. The first-stage injection hole 8 is connected to the first-stage injection chamber 7 through a flow channel. The second-stage injection chamber 12 has a second-stage injection hole 13 on both the outer shell 1 and the central body 2. The second-stage injection hole 13 is connected to the second-stage injection chamber 12 through a flow channel. The first-stage injection hole 8 and the second-stage injection hole 13 are evenly distributed along the circumferential direction.

[0011] The outer shell 1 is used to support the central body 2;

[0012] An annular flow channel is formed between the central body 2 and the outer shell 1;

[0013] The front support 3 and the rear support 4 of the central body are used to support the central body;

[0014] Intake 5 is used to capture the incoming airflow and form an annular airflow;

[0015] Fuel tank 9 is used to store fuel;

[0016] The first-stage injection chamber 7 and the second-stage injection chamber 12 are used to uniformly fill the fuel along the circumference of the engine.

[0017] The first-stage injection hole 8 and the second-stage injection hole 13 are used for injecting fuel.

[0018] The first-stage combustion chamber 10 and the second-stage combustion chamber 14 are used for ignition and detonation, and to generate a rotating detonation wave within the combustion chamber.

[0019] The exit nozzle 16 is used to convert the high-temperature airflow into a high-speed airflow, thereby generating thrust;

[0020] Under low Mach number inflow conditions, the second-stage combustion chamber 14 is ignited and detonated. The isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11 are used to isolate the back pressure of the second-stage combustion chamber 14 from entering the intake duct 5.

[0021] Under high Mach number inflow conditions, the isolation section 6 is used to isolate the back pressure of the first-stage combustion chamber 10 from entering the intake duct 5.

[0022] As a preferred embodiment, the widths δ1 and δ2 of the first-stage combustion chamber satisfy the following formula: δ ≥ 0.5λ, and when liquid fuel is used, δ ≥ d; the length of the first-stage combustion chamber... L 1 and the length of the second-stage combustion chamber L 2 The following formula must be satisfied: L ≥2(12±5)λ, where λ is the size of the knock wave cell corresponding to the mixture under the current combustion chamber pressure, and d is the minimum diameter of the fuel droplet;

[0023] Based on meeting the above requirements, the surface structure design of the first-stage combustion chamber 10 takes the inflow condition of Mach number Ma=5 as the standard design point to achieve rotational detonation combustion in the inflow range of Mach number Ma=4~6; the surface structure design of the second-stage combustion chamber 14 takes the inflow condition of Mach number Ma=3 as the standard design point to achieve rotational detonation combustion in the inflow range of Mach number Ma=2~4.

[0024] As a preferred embodiment, the dual-cavity structure of the first-stage injection cavity 7 and the second-stage injection cavity 12 has a square or circular cross-section with a cross-sectional area of ​​4~25mm². 2 The fuel pressure in the injection chamber is 1~5MPa.

[0025] As a preferred embodiment, the first-stage injection holes 8 and the second-stage injection holes 13 are evenly distributed along the circumference. To ensure the mixing effect of fuel and air, the diameter of the injection holes is 0.2~0.6mm. For liquid fuel, the number of injection holes n satisfies the following formula: ,in m For fuel flow rate, C d The flow coefficient of the injection orifice is... D The diameter of the injection hole, p The fuel pressure in the injection chamber, ρ Let n be the fuel density; for gaseous fuels, the number of injection orifices n satisfies In the formula, m is the fuel flow rate. T t Total fuel temperature p The fuel pressure in the injection chamber, C d The flow coefficient of the injection orifice is... D Where is the diameter of the injection orifice, and K is a fuel-related constant, obtained by the following formula: In the formula, γ For specific heat ratio, R The constant is the fuel gas constant. To ensure the mixing effect of fuel and air, the injection orifice diameter D is 0.2~0.6mm.

[0026] As a preferred embodiment, the flow area of ​​the first-stage contraction section 11 is... A th Smaller than the flow area of ​​the first-stage combustion chamber 10 A c And the flow area of ​​the second-stage contraction section 15 A th Smaller than the flow area of ​​the second-stage combustion chamber 14 A c The area shrinkage ratio AR ranges from 1.2 to 2.5, and AR is the flow area of ​​the first-stage combustion chamber 10. A c and the flow area of ​​the first-stage contraction section 11 A th The ratio, or the flow area of ​​the second-stage combustion chamber 14. A c The flow area of ​​the second-stage contraction section 15 A th The ratio of .

[0027] As a preferred embodiment, the central body 2 is fixed to the center of the outer shell 1 by the front support 3 and the rear support 4.

[0028] As a preferred approach, the high Mach number inflow condition is 4 < Ma ≤ 6;

[0029] The low Mach number inflow condition is defined as 2 ≤ Ma ≤ 4;

[0030] High-temperature airflow refers to airflow with a total temperature greater than 2000K;

[0031] High-speed airflow refers to airflow with a speed greater than twice the speed of sound.

[0032] The present invention also provides a method for operating the aforementioned series dual-combustion chamber rotary detonation ramjet engine, which is as follows:

[0033] Under low Mach number incoming air conditions, the incoming air is captured by the intake duct 5 and then enters the second-stage combustion chamber 14 through the isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11. It then mixes with the fuel injected into the second-stage injection port 13 and ignites and detonates in the second-stage combustion chamber 14, forming a rotating detonation wave. The combustion products generated by the rotating detonation combustion are discharged into the atmosphere through the second-stage contraction section 15 and the outlet nozzle 16, generating thrust. Under low Mach number incoming air conditions, the forward shock wave induced by the rotating detonation wave is isolated by the combined channel formed by the isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11, so that the combustion back pressure does not propagate forward to the intake duct 5.

[0034] Under high Mach number incoming flow conditions, the incoming air is captured by the intake duct 5 and enters the first-stage combustion chamber 10 through the isolation section 6. After mixing with the fuel injected into the first-stage injection port 8, a combustible mixture is formed. Subsequently, in the first-stage combustion chamber 10, the engine ignition device ignites the flow field and forms a rotating detonation wave in the combustion chamber. The detonation wave propagates circumferentially and consumes fuel. The combustion products are discharged into the atmosphere after passing through the first-stage contraction section 11, the second-stage combustion chamber 14, the second-stage contraction section 15, and the outlet nozzle 16, thereby generating thrust. Under high Mach number incoming flow conditions, the forward shock wave induced by the rotating detonation wave is isolated by the isolation section 6, so that the combustion back pressure does not propagate forward to the intake duct 5.

[0035] The beneficial effects of this invention are as follows: First, this invention addresses the flow characteristics of rotating detonation combustion under high / low Mach number operating conditions. It allows for separate design of the first and second stage combustion chamber configurations, thereby reducing the difficulty of combustion chamber design and achieving compatibility of rotating detonation combustion under high / low Mach number inflow conditions within the same flow channel. This enables efficient rotating detonation combustion under wide Mach number inflow conditions, thus broadening the operating Mach number range of the rotating detonation ramjet engine and improving its performance. Second, considering the significant difference in combustion back pressure forward transmission distance under high / low inflow Mach number operating conditions, the problem of combustion back pressure forward transmission under wide inflow Mach number conditions can be solved by arranging the two combustion chambers in series and rationally arranging the combustion organization operating modes. Third, by arranging the two combustion chambers in the same flow channel, the flow channel can be fully utilized under high / low inflow Mach number conditions, helping to reduce engine size and structural weight, and lowering engine structural complexity. Attached Figure Description

[0036] Figure 1 This is a cross-sectional view of the flow channel configuration of the tandem dual-combustion chamber rotary detonation ramjet engine of the present invention.

[0037] Figure 2 Figure 1 Cross-sectional view of the flow channel configuration of a rotating detonation ramjet engine (AA section).

[0038] Figure 3This is a partial enlarged view of the fuel injection chamber and injection hole of the present invention.

[0039] Figure 4 This is a schematic diagram of the operating modes of the present invention at low incoming Mach numbers.

[0040] Figure 5 This is a schematic diagram of the operating modes of the present invention at high incoming Mach numbers.

[0041] Figure Labels

[0042] 1. Outer shell; 2. Central body; 3. Front support of central body; 4. Rear support of central body; 5. Intake duct; 6. Isolation section; 7. First-stage injection chamber; 8. First-stage injection orifice; 9. Fuel tank; 10. First-stage combustion chamber; 11. First-stage contraction section; 12. Second-stage injection chamber; 13. Second-stage injection orifice; 14. Second-stage combustion chamber; 15. Second-stage contraction section; 16. Exit nozzle. Implementation

[0043] The following specific examples illustrate the implementation of the present invention. Those skilled in the art can easily understand other advantages and effects of the present invention from the content disclosed in this specification. The present invention can also be implemented or applied through other different specific embodiments, and various details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of the present invention.

[0044] This embodiment provides a tandem dual-combustion chamber rotary detonation ramjet engine, the main structure of which is as follows: Figure 1 As shown, it includes an outer shell 1, a central body 2 is placed inside the outer shell 1, and the central body 2 is connected to the outer shell 1 through a central body front support 3 and a central body rear support 4.

[0045] An annular flow channel is formed between the outer shell 1 and the central body 2. Along the direction of airflow propagation, the annular flow channel includes an intake duct 5, an isolation section 6, a first-stage combustion chamber 10, a first-stage contraction section 11, a second-stage combustion chamber 14, a second-stage contraction section 15, and an outlet nozzle 16.

[0046] The fuel tank 9 is installed inside the central body 2; the front support 3 of the central body is located in the isolation section 6 between the intake duct 5 and the first-stage combustion chamber 10, and the rear support 4 of the central body is located in the second-stage converging section 15 between the second-stage combustion chamber 14 and the outlet nozzle 16; to ensure airflow, the front support 3 and the rear support 4 of the central body are evenly distributed circumferentially, and the specific arrangement scheme is shown in [reference needed]. Figure 2 ;

[0047] Figure 3A partially enlarged view of the structure of the first-stage injection chamber 7 or the second-stage injection chamber 12 of the present invention is provided. As shown in the figure, both the first-stage injection chamber 7 and the second-stage injection chamber 12 are dual-injection-chamber structures, with the two injection chambers respectively disposed on the outer shell 1 and the central body 2. Along the airflow propagation direction, the first-stage injection chamber 7 is located between the tail end of the isolation section 6 away from the incoming flow and the front end of the first-stage combustion chamber 10 near the incoming flow. Along the airflow propagation direction, the second-stage injection chamber 12 is located between the tail end of the first-stage contraction section 11 away from the incoming flow and the front end of the second-stage combustion chamber 14 near the incoming flow.

[0048] The first-stage injection chamber 7 has a first-stage injection hole 8 on both the outer shell 1 and the central body 2. The first-stage injection hole 8 is connected to the first-stage injection chamber 7 through a flow channel. The second-stage injection chamber 12 has a second-stage injection hole 13 on both the outer shell 1 and the central body 2. The second-stage injection hole 13 is connected to the second-stage injection chamber 12 through a flow channel. The first-stage injection hole 8 and the second-stage injection hole 13 are evenly distributed along the circumferential direction.

[0049] The outer shell 1 is used to support the central body 2;

[0050] An annular flow channel is formed between the central body 2 and the outer shell 1;

[0051] The front support 3 and the rear support 4 of the central body are used to support the central body;

[0052] Intake 5 is used to capture the incoming airflow and form an annular airflow;

[0053] Fuel tank 9 is used to store fuel;

[0054] The first-stage injection chamber 7 and the second-stage injection chamber 12 are used to uniformly fill the fuel along the circumference of the engine.

[0055] The first-stage injection hole 8 and the second-stage injection hole 13 are used for injecting fuel.

[0056] The first-stage combustion chamber 10 and the second-stage combustion chamber 14 are used for ignition and detonation, and to generate a rotating detonation wave within the combustion chamber.

[0057] The exit nozzle 16 is used to convert the high-temperature airflow into a high-speed airflow, thereby generating thrust;

[0058] Under low Mach number inflow conditions, the second-stage combustion chamber 14 is ignited and detonated. The isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11 are used to isolate the back pressure of the second-stage combustion chamber 14 from entering the intake duct 5.

[0059] Under high Mach number inflow conditions, the isolation section 6 is used to isolate the back pressure of the first-stage combustion chamber 10 from entering the intake duct 5.

[0060] In some embodiments, the widths δ1 and δ2 of the first-stage combustion chamber satisfy the following formula: δ ≥ 0.5λ, and when liquid fuel is used, δ ≥ d; the length of the first-stage combustion chamber... L 1 and the length of the second-stage combustion chamber L 2 The following formula must be satisfied: L ≥2(12±5)λ, where λ is the size of the knock wave cell corresponding to the mixture under the current combustion chamber pressure, and d is the minimum diameter of the fuel droplet;

[0061] Based on meeting the above requirements, the surface structure design of the first-stage combustion chamber 10 takes the inflow condition of Mach number Ma=5 as the standard design point to achieve rotational detonation combustion in the inflow range of Mach number Ma=4~6; the surface structure design of the second-stage combustion chamber 14 takes the inflow condition of Mach number Ma=3 as the standard design point to achieve rotational detonation combustion in the inflow range of Mach number Ma=2~4.

[0062] In some embodiments, the dual-cavity structure of the first-stage injection cavity 7 and the second-stage injection cavity 12, during processing, has a square or circular cross-section for the injection cavities, with a cross-sectional area of ​​4~25mm². 2 The fuel pressure in the injection chamber is 1~5MPa.

[0063] In some embodiments, the first-stage injection holes 8 and the second-stage injection holes 13 are evenly distributed along the circumference. To ensure the mixing effect of fuel and air, the diameter of the injection holes is 0.2~0.6mm. For liquid fuel, the number of injection holes n satisfies the following formula: ,in m For fuel flow rate, C d The flow coefficient of the injection orifice is... D The diameter of the injection hole, p The fuel pressure in the injection chamber, ρ Let n be the fuel density; for gaseous fuels, the number of injection orifices n satisfies In the formula m For fuel flow rate, T t Total fuel temperature p The fuel pressure in the injection chamber, C d The flow coefficient of the injection orifice is... D Let K be the nozzle diameter, and K be a fuel-related constant, obtained from the following formula: In the formula, γ For specific heat ratio, R The constant is the fuel gas constant. To ensure the mixing effect of fuel and air, the injection orifice diameter D is 0.2~0.6mm.

[0064] In some embodiments, the flow area of ​​the first-stage contraction section 11 A th Smaller than the flow area of ​​the first-stage combustion chamber 10 A c And the flow area of ​​the second-stage contraction section 15 A th Smaller than the flow area of ​​the second-stage combustion chamber 14 A c The area shrinkage ratio AR ranges from 1.2 to 2.5, and AR is the flow area of ​​the first-stage combustion chamber 10. A c and the flow area of ​​the first-stage contraction section 11 A th The ratio, or the flow area of ​​the second-stage combustion chamber 14. A c The flow area of ​​the second-stage contraction section 15 A th The ratio of .

[0065] In some embodiments, the central body 2 is fixed to the center inside the outer shell 1 by the central body front support 3 and the central body rear support 4.

[0066] In some embodiments, the high Mach number inflow condition means 4 < Ma ≤ 6;

[0067] The low Mach number inflow condition is defined as 2 ≤ Ma ≤ 4;

[0068] High-temperature airflow refers to airflow with a total temperature greater than 2000K;

[0069] High-speed airflow refers to airflow with a speed greater than twice the speed of sound.

[0070] This embodiment also provides a method for operating the aforementioned series dual-combustion chamber rotary detonation ramjet engine, which is as follows:

[0071] Its working diagram under low Mach number inflow conditions is as follows: Figure 4 As shown. After being captured by the intake duct 5, the incoming air passes through the isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11 into the second-stage combustion chamber 14. Thereafter, it mixes with the fuel injected into the second-stage injection port 13 and ignites and detonates in the second-stage combustion chamber 14, forming a rotating detonation wave. The combustion products generated by the rotating detonation combustion are discharged into the atmosphere through the second-stage contraction section 15 and the outlet nozzle 16, generating thrust. Under low Mach number incoming flow conditions, the forward shock wave induced by the rotating detonation wave is isolated by the combined channel formed by the isolation section 6, the first-stage combustion chamber 10, and the first-stage contraction section 11, so that the combustion back pressure does not propagate forward to the intake duct 5.

[0072] Under high Mach number inflow conditions, its working schematic diagram is as follows: Figure 5As shown, incoming air is captured by the intake duct 5 and enters the first-stage combustion chamber 10 through the isolation section 6. After mixing with the fuel injected through the first-stage injection orifice 8, a combustible mixture is formed. Subsequently, within the first-stage combustion chamber 10, the engine ignition device ignites the flow field and forms a rotating detonation wave. The detonation wave propagates circumferentially and consumes fuel. The combustion products are discharged into the atmosphere after passing through the first-stage contraction section 11, the second-stage combustion chamber 14, the second-stage contraction section 15, and the outlet nozzle 16, thereby generating thrust. Under high Mach number incoming flow conditions, the forward shock wave induced by the rotating detonation wave is isolated by the isolation section 6, preventing the combustion back pressure from propagating forward to the intake duct 5.

[0073] The above embodiments are merely illustrative of the principles and effects of the present invention and are not intended to limit the invention. Any person skilled in the art can modify or alter the above embodiments without departing from the spirit and scope of the present invention. Therefore, all equivalent modifications or alterations made by those skilled in the art without departing from the spirit and technical concept disclosed in the present invention should still be covered by the claims of the present invention.

Claims

1. A tandem dual-combustion chamber rotary detonation ramjet engine, characterized in that: It includes an outer shell (1), a central body (2) is placed inside the outer shell (1), and the central body (2) is connected to the outer shell (1) through a front support (3) and a rear support (4); An annular flow channel is formed between the outer shell (1) and the central body (2). Along the airflow propagation direction of the annular flow channel, the annular flow channel includes, in sequence, an intake duct (5), an isolation section (6), a first-stage combustion chamber (10), a first-stage contraction section (11), a second-stage combustion chamber (14), a second-stage contraction section (15), and an outlet nozzle (16). The central body (2) is equipped with a fuel tank (9); the front support (3) of the central body is located in the isolation section (6) between the intake duct (5) and the first-stage combustion chamber (10); the rear support (4) of the central body is located in the second-stage contraction section (15) between the second-stage combustion chamber (14) and the outlet nozzle (16); the front support (3) and the rear support (4) of the central body are evenly distributed in the circumference. Both the first-stage injection chamber (7) and the second-stage injection chamber (12) are dual-injection-chamber structures, with the two injection chambers respectively located on the outer shell (1) and the central body (2). Along the airflow propagation direction, the first-stage injection chamber (7) is located between the tail end of the isolation section (6) away from the incoming flow and the front end of the first-stage combustion chamber (10) close to the incoming flow. Along the airflow propagation direction, the second-stage injection chamber (12) is located between the tail end of the first-stage contraction section (11) away from the incoming flow and the front end of the second-stage combustion chamber (14) close to the incoming flow. The first-stage injection chamber (7) is provided with first-stage injection holes (8) on both the outer shell (1) and the central body (2). The first-stage injection holes (8) are connected to the first-stage injection chamber (7) through a flow channel. The second-stage injection chamber (12) is provided with second-stage injection holes (13) on both the outer shell (1) and the central body (2). The second-stage injection holes (13) are connected to the second-stage injection chamber (12) through a flow channel. The first-stage injection holes (8) and the second-stage injection holes (13) are evenly distributed along the circumferential direction. The widths δ1 and δ2 of the first-stage combustion chamber satisfy the following formula: δ≥0.5λ. When liquid fuel is used, δ≥d is also satisfied. The lengths L1 and L2 of the first-stage combustion chamber satisfy the following formula: L≥2(12±5)λ, where λ is the detonation cell size of the mixture under the current combustion chamber pressure, and d is the minimum diameter of the fuel droplets. Based on meeting the above requirements, the surface structure design of the first-stage combustion chamber (10) takes the inflow condition of Mach number Ma=5 as the standard design point to achieve rotating detonation combustion in the inflow range of Mach number Ma=4~6; the surface structure design of the second-stage combustion chamber (14) takes the inflow condition of Mach number Ma=3 as the standard design point to achieve rotating detonation combustion in the inflow range of Mach number Ma=2~4. The structure consists of a first-stage injection chamber (7) and a second-stage injection chamber (12), with the injection chamber cross-section being square or circular and the cross-sectional area being 4~25mm². 2 The fuel pressure in the injection chamber is 1~5MPa; The first-stage injection holes (8) and the second-stage injection holes (13) are evenly distributed along the circumference, and the diameter of the injection holes is 0.2~0.6mm; for liquid fuel, the number of injection holes n satisfies the following formula: Where m is the fuel flow rate, C d Let D be the orifice flow coefficient, p be the fuel pressure in the injection chamber, and ρ be the fuel density; for gaseous fuels, the number of orifices n satisfies In the formula, m is the fuel flow rate, and T is the fuel flow rate. t Where P is the total fuel temperature, C is the fuel pressure in the injection chamber, and P is the total fuel temperature. d Let D be the injection orifice flow coefficient, D be the injection orifice diameter, and K be a fuel-related constant, obtained from the following formula: In the formula, γ is the specific heat ratio, R is the fuel gas constant, and the injection orifice diameter D is 0.2~0.6 mm; The flow area A of the first-stage contraction section (11) th The flow area A is smaller than that of the first-stage combustion chamber (10). c And the flow area A of the second-stage contraction section (15) th The flow area A is smaller than that of the second-stage combustion chamber (14). c The area shrinkage ratio AR ranges from 1.2 to 2.5, and AR is the flow area A of the first-stage combustion chamber (10). c The flow area A of the first-stage contraction section (11) th The ratio, or the flow area A of the second-stage combustion chamber (14). c The flow area A of the second-stage contraction section (15) th The ratio of .

2. The tandem dual-combustion chamber rotary detonation ramjet engine according to claim 1, characterized in that: The central body (2) is fixed to the center inside the outer shell (1) by the front support (3) and the rear support (4).

3. The operating method of the tandem dual-combustion chamber rotary detonation ramjet engine according to any one of claims 1 to 2, characterized in that: Under low Mach number incoming flow conditions, the incoming air is captured by the intake duct (5) and then enters the second-stage combustion chamber (14) through the isolation section (6), the first-stage combustion chamber (10), and the first-stage contraction section (11). Subsequently, it mixes with the fuel injected through the second-stage injection hole (13) and ignites and detonates in the second-stage combustion chamber (14), forming a rotating detonation wave. The combustion products generated by the rotating detonation combustion are discharged into the atmosphere through the second-stage contraction section (15) and the outlet nozzle (16), generating thrust. Under low Mach number incoming flow conditions, the forward shock wave induced by the rotating detonation wave is isolated by the combined channel formed by the isolation section (6), the first-stage combustion chamber (10), and the first-stage contraction section (11), so that the combustion back pressure does not propagate forward to the intake duct (5). Under high Mach number incoming flow conditions, the incoming air is captured by the intake duct (5) and enters the first-stage combustion chamber (10) through the isolation section (6). After mixing with the fuel injected through the first-stage injection hole (8), a combustible mixture is formed. Subsequently, in the first-stage combustion chamber (10), the engine ignition device detonates the flow field and forms a rotating detonation wave in the combustion chamber. The detonation wave propagates circumferentially and consumes fuel. The combustion products generated are discharged into the atmosphere after passing through the first-stage contraction section (11), the second-stage combustion chamber (14), the second-stage contraction section (15), and the outlet nozzle (16), thereby generating thrust. Under high Mach number incoming flow conditions, the forward shock wave induced by the rotating detonation wave is isolated by the isolation section (6), so that the combustion back pressure does not propagate forward to the intake duct (5).

4. The operating method of the tandem dual-combustion chamber rotary detonation ramjet engine according to claim 3, characterized in that: The inflow condition for high Mach number is 4 < Ma ≤ 6; the inflow condition for low Mach number is 2 ≤ Ma ≤ 4.

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