A dual flowpath combustion chamber and rocket based combined power engine

Abstract

The application discloses a double-flow combustion chamber and a rocket-based combined power engine, and solves the problem that the combustion chamber cannot work stably with high performance in a wide range. The combustion chamber comprises a first flow channel and a first combustion chamber, the first combustion chamber is provided with a first flame stabilizer near one side of the first isolation section, a second flow channel and a second combustion chamber, the second combustion chamber is provided with a second flame stabilizer near one side of the second isolation section, the first flame stabilizer and the second flame stabilizer are used for injecting fuel, the flow direction of the second flow channel is parallel to that of the first flow channel, the expansion ratio of the second flow channel is greater than or equal to that of the first flow channel, the first combustion chamber and the second combustion chamber are communicated with a second combustion chamber, and the expansion ratios of the first flow channel and the second flow channel are both less than that of the second combustion chamber, a power device is communicated with the second combustion chamber, and the injection direction of the power device is inclined to the flow direction of the second flow channel, and the power device is used for injecting high-temperature exhaust into the second combustion chamber.

Description

Technical Field

[0001] This invention relates to the field of rocket ramjet engine technology, specifically to a dual-flow combustion chamber and rocket-based combined propulsion engine. Background Technology

[0002] A rocket-ramjet combined engine organically combines a high-specific-impulse, low-thrust-to-weight-ratio ramjet engine with a low-specific-impulse, high-thrust-to-weight-ratio rocket engine, greatly expanding the engine's flight airspace and speed range, enabling it to operate over an ultra-wide range. As the core component, the engine combustor's stable and efficient combustion across various speed ranges is crucial for the engine's high-efficiency operation; however, the combustor's operating modes differ across different speed and airspace ranges.

[0003] Currently, multi-stage combustion chamber designs adapt to a wide range of combustion by adjusting the heat release position under different operating conditions. However, since the corresponding intake duct contraction ratio cannot be changed, the intake duct cannot be started at low speeds, resulting in high intake duct drag and a limited range of flight speeds that the combustion chamber can adapt to. Summary of the Invention

[0004] The purpose of this invention is to provide a combined propulsion engine with a dual-flow combustion chamber and a rocket base, which ensures that the engine can operate stably and efficiently over a wide speed range.

[0005] To achieve the above objectives, the present invention provides the following technical solution: In a first aspect, the present invention provides a dual-flow combustion chamber, comprising: The first flow channel has a first isolation section and a first combustion chamber connected together, and a first flame stabilizer is provided on the side of the first combustion chamber near the first isolation section; The second flow channel has a connected second isolation section and a second combustion chamber. A second flame stabilizer is provided on the side of the second combustion chamber near the second isolation section. Both the first flame stabilizer and the second flame stabilizer are used for fuel injection. The flow direction of the second flow channel is parallel to the flow direction of the first flow channel, and the expansion ratio of the second flow channel is greater than or equal to the expansion ratio of the first flow channel. The second-stage combustion chamber is connected to both the end of the first combustion chamber away from the first isolation section and the end of the second combustion chamber away from the second isolation section. The expansion ratio of the first flow channel and the expansion ratio of the second flow channel are both smaller than the expansion ratio of the second-stage combustion chamber. The power unit is connected to the secondary combustion chamber, and the injection direction of the power unit is inclined toward the flow direction of the second flow channel. The power unit is used to inject high-temperature tail flow into the secondary combustion chamber.

[0006] Optionally, in the above-mentioned dual-flow-channel combustion chamber, the expansion ratio of the first flow channel is 1.45 to 1.55.

[0007] Optionally, in the above-mentioned dual-flow combustion chamber, the expansion ratio of the second flow channel is 1.55 to 1.65.

[0008] Optionally, in the above-mentioned dual-flow combustion chamber, the expansion ratio of the secondary combustion chamber is 2.6 to 3.6.

[0009] Optionally, in the above-mentioned dual-flow combustion chamber, the angle between the injection direction of the power unit and the flow direction of the second flow channel is 8°~10°.

[0010] Optionally, in the above-mentioned dual-channel combustion chamber, multiple second flame stabilizers are provided, and the multiple second flame stabilizers are arranged at intervals along the flow direction of the second channel.

[0011] Optionally, in the aforementioned dual-flow combustion chamber, the propulsion device is a high-ratio rocket with a thrust ratio of 1:5.

[0012] Compared with the prior art, when the above technical solution is adopted, since the expansion ratio of the second flow channel is greater than or equal to the expansion ratio of the first flow channel, the first flow channel acts as a high-speed channel and the second flow channel acts as a low-speed channel. Under static conditions, since no airflow enters the first and second flow channels, the combustion chamber cannot work. When the power unit is working at full capacity, the power unit starts the aircraft with maximum thrust. Under low speed conditions after startup, the first and second flow channels work simultaneously. The first flame stabilizer injects fuel into the first combustion chamber, the second flame stabilizer injects fuel into the second combustion chamber, and the power unit injects a high-temperature exhaust into the secondary combustion chamber. When the fuel injected into the first and second flow channels reaches the secondary combustion chamber, the fuel and the high-temperature exhaust are partially or completely mixed. At this time, the mixed gas is ignited under the action of the high-temperature exhaust. The fuel injected into the first flow channel and the fuel injected into the second flow channel both undergo subsonic combustion in the secondary combustion chamber to obtain better combustion performance at low speeds. As the incoming flow velocity increases, the second flow channel, which serves as a low-speed channel, closes and the second flame stabilizer stops injecting fuel. Only the first flow channel, which serves as a high-speed channel, opens, and the secondary combustion chamber continues to operate. At this time, because the incoming flow temperature is high enough, the fuel is simultaneously organized to burn in the first and secondary combustion chambers. The power plant dynamically adjusts according to the aircraft's thrust requirements and the combustion chamber conditions. When the incoming flow velocity increases to a certain level, both the first and second combustion chambers are in a superburning state. Due to the higher velocity in the second combustion chamber, the fuel has a shorter residence time and cannot effectively organize combustion. Therefore, the fuel mainly burns in the first combustion chamber, which serves as a high-speed channel. The second combustion chamber then functions as a tail nozzle to provide thrust. Compared to the traditional multi-stage combustion chamber scheme that uses a compressor installed in a low-speed flow channel without organizing combustion in the flow channel, this application eliminates the need for a compressor and enables combustion in the first, second, and second combustion chambers. The start-up and shutdown of the first and second flow channels are adjusted according to different incoming flow velocities, allowing the dual-flow combustion chamber to achieve better combustion performance at low speeds.

[0013] Secondly, the present invention also provides a rocket-based combined propulsion engine, comprising: an air intake, a first adjusting plate, a nozzle, a second adjusting plate, and a dual-flow combustion chamber as described above. The first and second flow channels are both connected to the air intake. The first adjusting plate is rotatably disposed within the air intake and is used to rotatably close or open the air intake channel corresponding to the second flow channel. The nozzle is connected to the outlet end of the secondary combustion chamber. The second adjusting plate is rotatably disposed within the nozzle and is used to adjust the attitude of the nozzle under different combustion modes in the dual-flow combustion chamber.

[0014] Compared with existing technologies, when the above technical solution is adopted, when the rocket-based combined propulsion engine is in an initial static state or its speed is below a certain threshold, the first and second channels of the dual-channel combustion chamber operate simultaneously. Airflow is simultaneously introduced into the first and second channels through the air intake. Fuel is simultaneously injected and organized for combustion in the first combustion chamber within the first channel and the second combustion chamber within the second channel. At this time, the rocket-based combined propulsion engine is in the on-state. While the power unit provides high thrust, its high-temperature exhaust can better assist in flame stabilization. By rotating the second adjusting plate, the nozzle is made to be in a contracted state, forming a geometric throat, so that subsonic combustion can be better organized within the dual-channel combustion chamber. When the speed exceeds a certain threshold, the first adjusting plate of the intake duct is rotated to close the intake duct corresponding to the second flow channel, which serves as a low-speed channel. Simultaneously, the second flame stabilizer within the second flow channel stops injecting fuel, and the second flow channel ceases operation. At this time, the first and second stage combustion chambers remain in subsonic combustion mode, and the rotation of the second adjusting plate keeps the nozzle in a contracted state. As the speed increases, when it exceeds a second threshold, the second adjusting plate of the nozzle is rotated according to the needs of the dual-flow combustion chamber, causing the second adjusting plate to rotate to a horizontal position. At this point, the nozzle expansion angle is at its maximum, the dual-flow combustion chamber operates in a high-performance state, and the nozzle obtains greater thrust. The beneficial effects of the rocket-based combined propulsion engine provided by this invention are the same as the beneficial effects of the dual-flow combustion chamber described above, and will not be elaborated upon here. Attached Figure Description

[0015] The accompanying drawings, which are included to provide a further understanding of the invention and form part of this invention, illustrate exemplary embodiments of the invention and are used to explain the invention, but do not constitute an undue limitation of the invention. In the drawings: Figure 1 This is a schematic diagram of a dual-flow combustion chamber provided in an embodiment of the present invention.

[0016] Figure label: 1-First flow channel; 11-First isolation section; 12-First combustion chamber; 13-First flame stabilizer; 2-Second flow channel; 21-Second isolation section; 22-Second combustion chamber; 23-Second flame stabilizer; 3-Second stage combustion chamber; 4-High-ratio rocket. Detailed Implementation

[0017] To make the technical problems to be solved, the technical solutions, and the beneficial effects of the present invention clearer, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the present invention and are not intended to limit the present invention.

[0018] It should be noted that when a component is referred to as being "fixed to" or "set on" another component, it can be directly on or indirectly on that other component. When a component is referred to as being "connected to" another component, it can be directly connected to or indirectly connected to that other component.

[0019] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified. "Several" means one or more, unless otherwise explicitly specified.

[0020] In the description of this invention, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0021] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0022] like Figure 1 As shown in the figure, an embodiment of the present invention provides a dual-flow-channel combustion chamber, including: a first flow channel 1, a second flow channel 2, a secondary combustion chamber 3, and a power unit.

[0023] The first flow channel 1 has a first isolation section 11 and a first combustion chamber 12 connected together. A first flame stabilizer 13 is provided on the side of the first combustion chamber 12 near the first isolation section 11. The second flow channel 2 has a second isolation section 21 and a second combustion chamber 22 connected together. A second flame stabilizer 23 is provided on the side of the second combustion chamber 22 near the second isolation section 21. Both the first flame stabilizer 13 and the second flame stabilizer 23 are used for fuel injection. The flow direction of the second flow channel 2 is parallel to the flow direction of the first flow channel 1, and the expansion ratio of the second flow channel 2 is greater than or equal to that of the first flow channel 1. The expansion ratio of the first flow channel 1; the end of the first combustion chamber 12 away from the first isolation section 11 and the end of the second combustion chamber 22 away from the second isolation section 21 are both connected to the second-stage combustion chamber 3, and the expansion ratio of the first flow channel 1 and the expansion ratio of the second flow channel 2 are both less than the expansion ratio of the second-stage combustion chamber 3; the power unit is connected to the second-stage combustion chamber 3, and the injection direction of the power unit is inclined towards the flow direction of the second flow channel 2. The power unit is used to inject a high-temperature tail stream into the second-stage combustion chamber 3 while starting from a standstill, so as to better organize combustion in the second-stage combustion chamber 3.

[0024] In specific implementation, such as Figure 1As shown, since the expansion ratio of the second channel 2 is greater than or equal to the expansion ratio of the first channel 1, the first channel 1 acts as a high-speed channel and the second channel 2 acts as a low-speed channel. Under static conditions, since no airflow enters the first channel 1 and the second channel 2, the combustion chamber cannot work. When the power unit is working at full capacity, it starts the aircraft with maximum thrust. Under low speed conditions after startup, the first channel 1 and the second channel 2 work simultaneously. The first flame stabilizer 13 injects fuel into the first combustion chamber 12, the second flame stabilizer 23 injects fuel into the second combustion chamber 22, and the power unit injects a high-temperature exhaust into the secondary combustion chamber 3. When the fuel injected into the first channel 1 and the second channel 2 reaches the secondary combustion chamber 3, the fuel and the high-temperature exhaust are partially or completely mixed. At this time, the mixed gas is ignited under the action of the high-temperature exhaust. The fuel injected into the first flow channel 1 and the fuel injected into the second flow channel 2 both undergo subsonic combustion in the secondary combustion chamber 3 to obtain better combustion performance at low speeds. As the incoming flow velocity increases, the second flow channel 2, which serves as a low-speed channel, closes and the second flame stabilizer 23 stops injecting fuel. Only the first flow channel 1, which serves as a high-speed channel, opens, and the secondary combustion chamber 3 continues to operate. At this time, because the incoming flow temperature is high enough, the fuel is simultaneously organized for combustion in the first combustion chamber 12 and the secondary combustion chamber 3. The power plant dynamically adjusts according to the aircraft's thrust requirements and the combustion chamber conditions. When the incoming flow velocity increases to a certain level, both the first combustion chamber 12 and the second-stage combustion chamber 3 are in a superburning state. Due to the higher velocity in the second-stage combustion chamber 3, the fuel has a shorter residence time in the second-stage combustion chamber 3 and cannot be well organized for combustion. Therefore, the fuel is mainly organized for combustion in the first combustion chamber 12, which serves as a high-speed channel. The second-stage combustion chamber 3 at this time acts as a tail nozzle to provide thrust performance. Compared with the traditional multi-stage combustion chamber scheme that uses a compressor installed in a low-speed flow channel but does not organize combustion in the flow channel, this application does not require a compressor and can organize combustion in the first flow channel 1, the second flow channel 2, and the second-stage combustion chamber 3. The start-stop of the first flow channel 1 and the second flow channel 2 is adjusted according to different incoming flow velocities, so that the dual-flow combustor can obtain better combustion performance at low speeds.

[0025] Specifically, in this embodiment, the expansion ratio of the first flow channel 1 is 1.45~1.55. For example, the expansion ratio of the first flow channel 1 can be 1.45, 1.47, 1.49, 1.5, 1.52, 1.55, etc. With this setting, the operator can select different expansion ratios for the first flow channel 1 according to actual working needs. If the expansion ratio of the first flow channel 1 is too small (e.g., the expansion ratio of the first flow channel 1 is less than 1.45), it is easy to form blockage in the first flow channel 1 when the incoming flow velocity is low, causing the back pressure generated by combustion to rise and preventing the upstream intake from starting, resulting in the engine not working properly. If the expansion ratio is too large (e.g., the expansion ratio of the first flow channel 1 is greater than 1.55), it is difficult to form blockage in the first combustion chamber 12 at higher Mach numbers, which will lead to reduced engine performance or even combustion chamber shutdown. Therefore, when the first flow channel 1 is used as a high-speed channel, the expansion ratio is 1.45~1.55. This can prevent the back pressure from preventing the intake from starting and also prevent the combustion chamber from being blocked, thus preventing performance reduction. This ensures the normal operation of the dual-flow combustion chamber.

[0026] Specifically, in this embodiment, since the second flow channel 2 is a low-speed flow channel, its expansion ratio is relatively large compared to the first flow channel 1. The expansion ratio of the second flow channel 2 is 1.55~1.65. For example, the expansion ratio of the second flow channel 2 can be 1.55, 1.57, 1.58, 1.6, 1.61, 1.63, 1.65, etc. This allows operators to select different expansion ratios for the second flow channel 2 according to actual work needs. However, if the expansion ratio of the second flow channel 2 is too small (e.g., less than 1.55), the flow channel expansion ratio of the second flow channel 2 is small, making it unable to accommodate much heat. This can easily cause back pressure generated by combustion to rise, leading to the intake manifold not starting. If the expansion ratio of the second flow channel 2 is too large... If the expansion ratio of the second flow channel 2 is too large (e.g., greater than 1.65), the space occupied by the second flow channel 2 inside the engine will increase. After the second flow channel 2 is closed, it will no longer work, resulting in a waste of space and load. Considering that both excessively large and excessively small expansion ratios of the second flow channel 2 will have adverse effects, the expansion ratio of the second flow channel 2 as a low-speed channel is set at 1.55~1.65. This can meet the transition requirements of the inner wall of the flow channel, control the structural dimensions and weight of the second flow channel 2, reduce the processing difficulty of the second flow channel 2, and enable the second flow channel 2 to work better.

[0027] Specifically, in this embodiment, the expansion ratio of the secondary combustion chamber 3 is 2.6 to 3.6. For example, the expansion ratio of the secondary combustion chamber 3 can be 2.6, 2.8, 3.0, 3.2, 3.4, 3.6, etc. With this setting, the expansion ratios of the first combustion chamber 12 and the second combustion chamber 22, which are the primary combustion chambers, are both smaller than the expansion ratio of the secondary combustion chamber 3. This allows the mainstream flow velocity to decrease within the secondary combustion chamber 3, which has a larger expansion ratio, thus prolonging the residence time of the mainstream flow within the secondary combustion chamber 3. If the expansion ratio of the secondary combustion chamber 3 is too large (e.g., greater than 3.6), it will increase the space occupied by the secondary combustion chamber 3, resulting in an increase in engine size. At the same time, the increase in its inner surface area will also increase the thermal protection area. If the expansion ratio of the secondary combustion chamber 3 is too small (e.g., less than 2.6), the mainstream airflow velocity will be high during combustion, which may lead to combustion chamber shutdown. Considering that both excessively large and excessively small expansion ratios of the secondary combustion chamber 3 will have adverse effects, the expansion ratio of the secondary combustion chamber 3 is set to 2.6 to 3.6. This can reduce the space required for the combustion chamber and reduce the difficulty of thermal protection while satisfying the combustion organization requirements.

[0028] like Figure 1 As shown, specifically, in this embodiment, the angle between the injection direction of the power unit and the flow direction of the second flow channel 2 is 8°~10°. The angle between the injection direction of the power unit and the flow direction of the second flow channel 2 is α, which can be 8°, 8.5°, 9°, 9.5°, 10°, etc. This setting, by tilting the injection direction towards the flow direction of the second flow channel 2 (which serves as a low-speed channel), and selecting an angle α of 8°~10°, allows the high-temperature tail jet injected by the power unit to deflect towards the second flow channel 2. This prevents the high-temperature, high-pressure tail jet from expanding at the outlet of the first flow channel 1 and causing blockage at the outlet of the first flow channel 1 near the secondary combustion chamber 3. The jet momentum of the high-temperature tail jet assists the mainstream flow entering from the first flow channel 1 and the second flow channel 2, reducing interference with the velocity field. Without disrupting the stability of the mainstream, the jet trajectory is kept obliquely cutting into the mainstream, causing a strong turbulent vortex to be generated by the velocity difference between the jet and the mainstream, ensuring sufficient mixing and combustion within the secondary combustion chamber 3.

[0029] Specifically, in this embodiment, multiple second flame stabilizers 23 are provided, and the multiple second flame stabilizers 23 are arranged at intervals along the flow direction of the second flow channel 2. For example, two, three, four, five, etc., of the second flame stabilizers 23 can be provided. When two second flame stabilizers 23 are provided, the two second flame stabilizers 23 are respectively disposed opposite each other on the sidewalls within the second flow channel 2, and the two second flame stabilizers 23 are arranged at intervals along the flow direction of the second flow channel 2; for example... Figure 1As shown, when three second flame stabilizers 23 are configured, two of them are arranged opposite each other on the sidewalls of the second flow channel 2, while the third second flame stabilizer 23 is spaced apart from the two oppositely arranged second flame stabilizers 23 along the flow direction of the second flow channel 2, so that the three second flame stabilizers 23 are triangularly distributed in the second flow channel 2. This configuration, with multiple second flame stabilizers 23 positioned at different locations, expands the contact area between the incoming flow and the fuel in the second channel, allowing the incoming flow to fully contact the fuel in the second channel with a large expansion ratio, thus ensuring the combustion performance of the dual-flow-channel combustion chamber.

[0030] Specifically, in this embodiment, the propulsion unit is a high-ratio rocket 4 with a thrust ratio of 1:5. This configuration, by setting the rocket 4 to a 1:5 thrust ratio, ensures that the rocket thrust meets the engine's requirements for high thrust during stationary start-up, medium thrust requirements at low speeds, and low thrust requirements at high speeds to accelerate the spacecraft. The high-ratio rocket 4 also assists in a dual-flow, multi-stage combustion system, enabling the dual-flow combustor to operate stably and efficiently over a wide velocity range.

[0031] In other embodiments, the power unit may also be a gas generator or other power unit, as long as it can inject high-temperature exhaust into the secondary combustion chamber 3.

[0032] Meanwhile, the present invention also provides a rocket-based combined propulsion engine, comprising: an air intake, a first adjusting plate, a nozzle, a second adjusting plate, and a dual-flow combustion chamber as described above. The first flow channel 1 and the second flow channel 2 are both connected to the air intake. The first adjusting plate is rotatably disposed within the air intake and is used to rotatably close or open the air intake passage corresponding to the second flow channel 2. The nozzle is connected to the outlet end of the secondary combustion chamber 3. The second adjusting plate is rotatably disposed within the nozzle and is used to adjust the attitude of the nozzle under different combustion modes in the dual-flow combustion chamber.

[0033] In specific implementation, such as Figure 1As shown, when the rocket-based combined propulsion engine is in its initial static state or its speed is below a certain threshold, the first channel 1 and the second channel 2 of the dual-channel combustion chamber operate simultaneously. Airflow is simultaneously introduced into the first channel 1 and the second channel 2 through the air intake. Fuel is simultaneously injected into the first combustion chamber 12 in the first channel 1 and the second combustion chamber 22 in the second channel 2, and combustion is organized. At this time, the rocket-based combined propulsion engine is in the on state. The power unit provides high thrust, and its high-temperature exhaust can better assist in flame stabilization. By rotating the second adjusting plate, the nozzle is made to be in a contracted state, forming a geometric throat, so that subsonic combustion can be better organized in the dual-channel combustion chamber. When the speed is higher than... At a certain threshold, by rotating the first adjusting plate of the intake duct, the intake duct corresponding to the second flow channel 2, which serves as a low-speed channel, is closed. At the same time, the second flame stabilizer 23 in the second flow channel 2 stops injecting fuel, and the second flow channel 2 stops working. At this time, the first combustion chamber 12 and the second-stage combustion chamber 3 are still in the subsonic combustion mode, and the rotation of the second adjusting plate keeps the nozzle in a contracted state. As the speed increases, when the speed exceeds the second threshold, the second adjusting plate of the nozzle is rotated according to the needs of the dual-flow combustion chamber, so that the second adjusting plate rotates to a horizontal position. At this time, the nozzle expansion angle is at its maximum, the dual-flow combustion chamber is in a high-performance working state, and the nozzle obtains greater thrust.

[0034] In the description of the above embodiments, specific features, structures, materials, or characteristics may be combined in any suitable manner in one or more embodiments or examples.

[0035] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be determined by the scope of the claims.

Claims

1. A dual flow combustion chamber characterized by, include: The first flow channel has a first isolation section and a first combustion chamber connected together, and a first flame stabilizer is provided on the side of the first combustion chamber near the first isolation section; The second flow channel has a connected second isolation section and a second combustion chamber. A second flame stabilizer is provided on the side of the second combustion chamber near the second isolation section. Both the first flame stabilizer and the second flame stabilizer are used for fuel injection. The flow direction of the second flow channel is parallel to the flow direction of the first flow channel, and the expansion ratio of the second flow channel is greater than or equal to the expansion ratio of the first flow channel. The second-stage combustion chamber is connected to both the end of the first combustion chamber away from the first isolation section and the end of the second combustion chamber away from the second isolation section, and the expansion ratio of the first flow channel and the expansion ratio of the second flow channel are both less than the expansion ratio of the second-stage combustion chamber. The power unit is connected to the secondary combustion chamber, and the injection direction of the power unit is inclined toward the flow direction of the second flow channel. The power unit is used to inject high-temperature tail flow into the secondary combustion chamber.

2. The dual flow combustion chamber of claim 1, wherein The expansion ratio of the first flow channel is 1.45~1.

55.

3. The dual flow combustion chamber of claim 1, wherein, The expansion ratio of the second flow channel is 1.55~1.

65.

4. The dual flow combustion chamber of claim 1, wherein, The expansion ratio of the secondary combustion chamber is 2.6 to 3.

6.

5. The dual-flow combustion chamber according to claim 1, characterized in that, The angle between the injection direction of the power device and the flow direction of the second channel is 8° to 10°.

6. The dual-flow combustion chamber according to claim 1, characterized in that, Multiple second flame stabilizers are provided, and the multiple second flame stabilizers are arranged at intervals along the flow direction of the second flow channel.

7. The dual-flow combustion chamber according to claim 1, characterized in that, The propulsion device is a high-ratio rocket, and the thrust ratio of the high-ratio rocket is 1:

5.

8. A rocket-based combined propulsion engine, characterized in that, include: The system comprises an air intake, a first adjusting plate, a nozzle, a second adjusting plate, and a dual-flow combustion chamber as described in any one of claims 1 to 7, wherein both the first flow channel and the second flow channel are connected to the air intake, the first adjusting plate is rotatably disposed within the air intake, and the first adjusting plate is used to rotatably close or open the air intake channel corresponding to the second flow channel; the nozzle is connected to the outlet end of the secondary combustion chamber, the second adjusting plate is rotatably disposed within the nozzle, and the rotation of the second adjusting plate is used to adjust the attitude of the nozzle under different combustion modes in the dual-flow combustion chamber.

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