Double-runner combustion chamber and rocket-based combined power engine

Abstract

The invention discloses a double-flow-channel combustion chamber and a rocket-based combined power engine. The problem that a combustion chamber cannot stably work in a wide range in a high-performance mode is solved. The first flow channel is provided with a first isolation section and a first combustion chamber, and a first flame stabilizer is arranged on the side, close to the first isolation section, of the first combustion chamber; the second flow channel is provided with a second isolation section and a second combustion chamber, a second flame stabilizer is arranged on the side, close to the second isolation section, of the second combustion chamber, the first flame stabilizer and the second flame stabilizer are both used for injecting fuel, the second flow channel is parallel to the flowing direction of the first flow channel, and the expansion ratio of the second flow channel is larger than or equal to that of the first flow channel; the first combustion chamber and the second combustion chamber communicate with the second-stage combustion chamber, and 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 combustion chamber. The power device communicates with the second-stage combustion chamber, and the jetting direction of the power device inclines towards the flowing direction of the second flow channel and is used for jetting high-temperature wake flow into the second-stage combustion chamber.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of rocket ramjet engine, in particular to a double-flow combustion chamber and rocket-based combined power engine. BACKGROUND

[0002] The rocket ramjet combined engine is combined with the ramjet engine with high specific impulse and low thrust-weight ratio and the rocket engine with low specific impulse and high thrust-weight ratio, which can greatly expand the flight air space and speed range of the engine, so that the engine can work in a super wide range. The combustion chamber as a core component is the key to realize stable and efficient combustion in various speed ranges, but the working mode of the combustion chamber is different in different speed ranges and air spaces.

[0003] At present, the multi-stage combustion chamber scheme adjusts the heat release position under different working conditions to adapt to wide range combustion. Since the corresponding inlet contraction ratio cannot be changed, the inlet cannot be started under low speed conditions, resulting in large inlet resistance and limited combustion chamber speed range. SUMMARY

[0004] The purpose of the present application is to provide a double-flow combustion chamber and rocket-based combined power engine for ensuring that the engine can work stably and efficiently in a wide speed range.

[0005] In order to achieve the above purpose, the present application provides the following technical scheme: In a first aspect, the present application provides a double-flow combustion chamber, comprising: A first flow passage has a first isolation section and a first combustion chamber connected thereto, and the first combustion chamber is provided with a first flame stabilizer on one side close to the first isolation section; A second flow passage has a second isolation section and a second combustion chamber connected thereto, and the second combustion chamber is provided with a second flame stabilizer on one side close to the second isolation section, the first flame stabilizer and the second flame stabilizer are both used for injecting fuel, the flow direction of the second flow passage is parallel to the flow direction of the first flow passage, and the expansion ratio of the second flow passage is greater than or equal to the expansion ratio of the first flow passage; A two-stage combustion chamber is in communication with one end of the first combustion chamber away from the first isolation section and one end of the second combustion chamber away from the second isolation section, and the expansion ratio of the first flow passage and the expansion ratio of the second flow passage are both less than the expansion ratio of the second combustion chamber; A power device is in communication with the two-stage combustion chamber, and the injection direction of the power device is inclined to the flow direction of the second flow passage, and the power device is used for injecting high-temperature exhaust into the two-stage combustion chamber.

[0006] Optionally, in the above double-flow combustion chamber, the expansion ratio of the first flow passage is 1.45-1.55.

[0007] Optionally, in the double-flow combustion chamber, the expansion ratio of the second flow passage is 1.55-1.65.

[0008] Optionally, in the double-flow combustion chamber, the expansion ratio of the secondary combustion chamber is 2.6-3.6.

[0009] Optionally, in the double-flow combustion chamber, the injection direction of the power device is inclined to the flow direction of the second flow passage by an angle of 8-10°.

[0010] Optionally, in the double-flow combustion chamber, a plurality of second flame stabilizers are arranged, and the plurality of second flame stabilizers are arranged at intervals along the flow direction of the second flow passage.

[0011] Optionally, in the double-flow combustion chamber, the power device is a large variable ratio rocket, and the thrust variable ratio of the large variable ratio rocket is 1:5.

[0012] Compared with the prior art, when the above technical scheme is adopted, because 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 serves as a high-speed channel and the second flow channel serves as a low-speed channel, in the static working condition, because no airflow enters the first flow channel and the second flow channel, the combustion chamber cannot work; when the power device works in the full working condition, the power device starts the aircraft at the maximum thrust, and in the low-speed condition after starting, the first flow channel and the second flow channel 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 device injects high-temperature exhaust into the secondary combustion chamber; when the fuel injected into the first flow channel and the fuel injected into the second flow channel reach the secondary combustion chamber, the fuel and the high-temperature exhaust are partially or fully mixed, and 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 are both subjected to subsonic combustion in the secondary combustion chamber, so as to obtain better combustion performance in the low-speed condition; with the increase of the incoming flow speed, the second flow channel as the low-speed channel is closed and the second flame stabilizer stops injecting fuel, only the first flow channel as the high-speed channel is opened, and the secondary combustion chamber continues to work; at this time, because the incoming flow temperature is high enough, the fuel is simultaneously combusted in the first combustion chamber and the secondary combustion chamber. The power device is dynamically adjusted according to the demand of the aircraft for thrust and the working condition of the combustion chamber; when the incoming flow speed increases to a certain degree, at this time, the first combustion chamber and the secondary combustion chamber are both in the supersonic working state, because the speed in the secondary combustion chamber is high, the fuel stays in the secondary combustion chamber for a short time and cannot be well combusted, therefore, the fuel is mainly combusted in the first combustion chamber as the high-speed channel, and the secondary combustion chamber at this time is equivalent to the function of the tail nozzle and is used to provide thrust performance. Compared with the traditional multi-stage combustion chamber scheme in which a compressor is arranged in the low-speed flow channel and no combustion is organized in the flow channel, in the present application, no compressor needs to be arranged, and combustion can be organized in the first flow channel, the second flow channel and the secondary combustion chamber, the first flow channel and the second flow channel are adjusted to be started or stopped according to different incoming flow speeds, so that the double-flow combustion chamber can obtain better combustion performance in the low-speed condition.

[0013] In a second aspect, the present application further provides a rocket-based combined power engine, which comprises an air inlet, a first adjusting plate, a nozzle, a second adjusting plate and a double-flow combustion chamber according to any one of the above, the first flow channel and the second flow channel both communicate with the air inlet, the first adjusting plate is rotationally arranged in the air inlet and is used to rotate to close or open the air inlet passage corresponding to the second flow channel; the nozzle communicates with the outlet end of the secondary combustion chamber, and the second adjusting plate is rotationally arranged in the nozzle and is used to adjust the posture of the nozzle in different combustion modes of the double-flow combustion chamber.

[0014] Compared with the prior art, when the rocket-based combined power engine is in an initial static state or the speed is lower than a certain threshold, the first flow channel and the second flow channel of the double-flow combustion chamber work simultaneously, the inlet channel simultaneously introduces airflow into the first flow channel and the second flow channel, fuel is simultaneously injected into the first combustion chamber in the first flow channel and the second combustion chamber in the second flow channel and combustion is organized, at this time, the rocket-based combined power engine is in an open state, the power device provides large thrust, and the high-temperature tail flow of the power device can better assist flame stabilization, by rotating the second adjusting plate, the nozzle is in a contraction state, a geometric throat is formed, and subsonic combustion in the double-flow combustion chamber is better organized; when the speed is higher than a certain threshold, by rotating the first adjusting plate of the inlet channel, the inlet channel corresponding to the second flow channel as a low-speed channel is closed, at the same time, the second flame stabilizer in the second flow channel stops injecting fuel, the second flow channel stops working, at this time, the first combustion chamber and the secondary combustion chamber are still in a subsonic combustion mode, and the second adjusting plate is rotated so that the nozzle is still in a contraction state; as the speed increases, when the speed exceeds a second threshold, the second adjusting plate of the nozzle is rotated according to the requirement of the double-flow combustion chamber, so that the second adjusting plate is rotated to a horizontal position, at this time, the expansion angle of the nozzle is maximum, the double-flow combustion chamber is in a high-performance working state, and the nozzle obtains large thrust. The rocket-based combined power engine provided by the present application has the same beneficial effects as the double-flow combustion chamber of the above technical scheme, and details are not repeated here. BRIEF DESCRIPTION OF DRAWINGS

[0015] The accompanying drawings, which are included to provide a further understanding of the present application and are incorporated in and constitute a part of this application, illustrate embodiments of the present application and together with the description serve to explain the present application. In the drawings: Figure 1 A structure schematic view of a double-flow combustion chamber provided in an embodiment of the present application.

[0016] Reference signs: 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 - secondary combustion chamber; 4 - large variable-ratio rocket. DETAILED DESCRIPTION

[0017] In order to make the technical problems to be solved by the present application, the technical solutions and the beneficial effects more clearly understood, the present application 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 only used to explain the present application and are not used to limit the present application.

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

[0019] In addition, the terms "first", "second", etc. are used only for descriptive purposes and should not be construed as indicating or implying relative importance or an indicated number of technical features. Therefore, the features defined as "first", "second" can explicitly or implicitly include one or more of the features. In the description of the present application, the meaning of "a plurality of" is two or more, unless otherwise explicitly specified and limited.

[0020] In the description of the present application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", etc. indicate the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application.

[0021] In the description of the present application, it should be noted that unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium; it can be the internal communication of two elements or the interaction relationship between two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0022] As shown in Figure 1 The embodiment of the present application provides a double-flow combustion chamber, which comprises a first flow channel 1, a second flow channel 2, a two-stage combustion chamber 3 and a power device.

[0023] The first flow channel 1 has a first isolation section 11 and a first combustion chamber 12 connected thereto, and the first combustion chamber 12 is provided with a first flame stabilizer 13 on the side close to the first isolation section 11; the second flow channel 2 has a second isolation section 21 and a second combustion chamber 22 connected thereto, and the second combustion chamber 22 is provided with a second flame stabilizer 23 on the side close to the second isolation section 21; the first flame stabilizer 13 and the second flame stabilizer 23 are both used for injecting fuel; 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 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 communicated with the second 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 combustion chamber 22; the power device is communicated with the second combustion chamber 3, and the injection direction of the power device is obliquely arranged towards the flow direction of the second flow channel 2; the power device is used for injecting high-temperature wake flow into the second combustion chamber 3 from the static start, so as to better organize the combustion in the second combustion chamber 3.

[0024] In specific implementation, as Figure 1As shown, since the expansion ratio of the second flow channel 2 is greater than or equal to the expansion ratio of the first flow channel 1, the first flow channel 1 serves as a high-speed channel and the second flow channel 2 serves as a low-speed channel. In the static working condition, since no airflow enters the first flow channel 1 and the second flow channel 2, the combustion chamber cannot work. When the power device works in the full working condition, the power device starts the aircraft at the maximum thrust. After starting, the first flow channel 1 and the second flow channel 2 work simultaneously at low speed. 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 device injects high-temperature exhaust into the secondary combustion chamber 3. When the fuel injected into the first flow channel 1 and the second flow channel 2 reaches the secondary combustion chamber 3, the fuel and the high-temperature exhaust are partially or fully mixed, and 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 are both subsonic combustion in the secondary combustion chamber 3, so as to obtain better combustion performance at low speed. With the increase of the incoming flow speed, the second flow channel 2 as a low-speed channel is closed and the second flame stabilizer 23 stops injecting fuel. Only the first flow channel 1 as a high-speed channel is opened, and the secondary combustion chamber 3 continues to work. At this time, since the incoming flow temperature is high enough, the fuel is combusted in the first combustion chamber 12 and the secondary combustion chamber 3 simultaneously. The power device dynamically adjusts according to the demand of the aircraft for thrust and the working condition of the combustion chamber. When the incoming flow speed increases to a certain extent, the first combustion chamber 12 and the secondary combustion chamber 3 are both in the super-combustion working state. Since the speed in the secondary combustion chamber 3 is high, the fuel stays in the secondary combustion chamber 3 for a short time and cannot be well combusted. Therefore, the fuel is mainly combusted in the first combustion chamber 12 as a high-speed channel, and the secondary combustion chamber 3 at this time is equivalent to the function of an exhaust nozzle for providing thrust performance. Compared with the traditional multi-stage combustion chamber scheme in which a compressor is installed in a low-speed flow channel and no combustion is organized in the flow channel, no compressor needs to be arranged in the present application, and combustion can be organized in the first flow channel 1, the second flow channel 2 and the secondary combustion chamber 3. According to different incoming flow speeds, the first flow channel 1 and the second flow channel 2 are adjusted to start and stop, so that the double-flow combustion chamber can obtain better combustion performance at low speed.

[0025] Specifically, in the 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. In this way, the operator can select the first flow channel 1 with different expansion ratios according to actual work requirements. If the expansion ratio of the first flow channel 1 is too small (for example, the expansion ratio of the first flow channel 1 is less than 1.45), when the incoming flow speed is low, it is easy to form a choke in the first flow channel 1, which causes the back pressure generated by combustion to be disturbed, resulting in the upstream intake port not starting, which causes the engine to malfunction. If the expansion ratio is too large (for example, the expansion ratio of the first flow channel 1 is greater than 1.55), it is difficult to form a choke in the first combustion chamber 12 at a high Mach number, which can cause the engine performance to decrease or even the combustion chamber to be extinguished. Therefore, when the first flow channel 1 is used as a high-speed channel, the expansion ratio is 1.45-1.55, which can prevent the back pressure from causing the intake port to not start and can prevent the combustion chamber from not being able to form a choke to cause performance to decrease, thereby ensuring the normal operation of the dual-flow combustion chamber.

[0026] Specifically, in the embodiment, since the second flow channel 2 is a low-speed flow channel, its expansion ratio is larger than that of 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. In this way, the operator can select the second flow channel 2 with different expansion ratios according to actual work requirements. If the expansion ratio of the second flow channel 2 is too small (for example, the expansion ratio of the second flow channel 2 is less than 1.55), the flow channel expansion ratio of the second flow channel 2 is small, which makes it difficult for the second flow channel 2 to accommodate more heat, which can easily cause the back pressure generated by combustion to be disturbed and cause the intake port to not start. If the expansion ratio of the second flow channel 2 is too large (for example, the expansion ratio of the second flow channel 2 is greater than 1.65), the space occupied by the second flow channel 2 in the engine increases, and after the second flow channel 2 is closed, it no longer works, which can cause a waste of space and load. Considering that too large or too small expansion ratio of the second flow channel 2 can have adverse effects, when the second flow channel 2 is used as a low-speed channel, the expansion ratio is 1.55-1.65, which can meet the transition requirements of the inner wall surface of the flow channel, control the structure size and weight of the second flow channel 2, reduce the processing difficulty of the second flow channel 2, and make the second flow channel 2 work better.

[0027] Specifically, in the present embodiment, the expansion ratio of the secondary combustion chamber 3 is 2.6-3.6. Exemplarily, the expansion ratio of the secondary combustion chamber 3 can be 2.6, 2.8, 3.0, 3.2, 3.4, 3.6, etc. In this way, the expansion ratio of the first combustion chamber 12 as the primary combustion chamber and the expansion ratio of the second combustion chamber 22 are both smaller than the expansion ratio of the secondary combustion chamber 3, so that the main flow can be reduced in the secondary combustion chamber 3 with a larger expansion ratio, and the time for the main flow to stay in the secondary combustion chamber 3 is prolonged. If the expansion ratio of the secondary combustion chamber 3 is too large (e.g., the expansion ratio of the secondary combustion chamber 3 is greater than 3.6), the space occupied by the secondary combustion chamber 3 will increase, which will increase the size of the engine, and the increase of the inner surface of the secondary combustion chamber 3 will also increase the heat protection area. If the expansion ratio of the secondary combustion chamber 3 is too small (e.g., the expansion ratio of the secondary combustion chamber 3 is less than 2.6), the airflow velocity of the main flow in the combustion state will be high, which may cause the combustion chamber to be extinguished. Considering that the expansion ratio of the secondary combustion chamber 3 being too large or too small will have adverse effects, therefore, the expansion ratio of the secondary combustion chamber 3 is 2.6-3.6, which can meet the requirements of combustion organization while reducing the space required by the combustion chamber and reducing the difficulty of heat protection.

[0028] As shown in Figure 1 Specifically, in the present embodiment, the angle between the injection direction of the power device and the flow direction of the second flow passage 2 is 8°-10°. The angle between the injection direction of the power device and the flow direction of the second flow passage 2 is α, and the angle α can be 8°, 8.5°, 9°, 9.5°, 10°, etc. In this way, by tilting the injection direction towards the flow direction of the second flow passage 2 as a low-speed passage, and selecting an inclination angle α of 8°-10°, the high-temperature wake of the power device injection can be deflected towards the second flow passage 2, preventing the high-temperature and high-pressure wake from expanding at the outlet position of the first flow passage 1, causing the first flow passage 1 to form a choke near the outlet of the secondary combustion chamber 3. The jet momentum of the high-temperature wake assists the flow of the main flow entering from the first flow passage 1 and the second flow passage 2, reducing the disturbance to the flow velocity field, while maintaining the jet trajectory to be obliquely cut into the main flow, so that the speed difference between the jet and the main flow generates strong turbulent vortexes, ensuring sufficient mixing and combustion in the secondary combustion chamber 3.

[0029] Specifically, in the present embodiment, the second flame stabilizer 23 is provided in plurality, and the plurality of second flame stabilizers 23 are arranged at intervals along the flow direction of the second flow passage 2. Exemplarily, the second flame stabilizer 23 can be provided in two, three, four, five, etc. When the second flame stabilizer 23 is provided in two, the two second flame stabilizers 23 are respectively arranged opposite to the side walls in the second flow passage 2, and the two second flame stabilizers 23 are arranged at intervals along the flow direction of the second flow passage 2. As shown in 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 power engine is in an initial static state or the speed is lower than a certain threshold, the first flow channel 1 and the second flow channel 2 of the double-flow combustion chamber work simultaneously, the inlet channel simultaneously introduces air flow into the first flow channel 1 and the second flow channel 2, and the first combustion chamber 12 in the first flow channel 1 and the second combustion chamber 22 in the second flow channel 2 simultaneously inject fuel and organize combustion, at this time, the rocket-based combined power engine is in an open state, the power device provides large thrust, and the high-temperature tail flow of the power device can better assist flame stabilization, by rotating the second adjusting plate, the nozzle is in a contraction state, a geometric throat is formed, and subsonic combustion in the double-flow combustion chamber is better organized; when the speed is higher than a certain threshold, by rotating the first adjusting plate of the inlet channel, the inlet channel corresponding to the second flow channel 2 as a low-speed channel is closed, and the second flame stabilizer 23 in the second flow channel 2 stops injecting fuel, the second flow channel 2 stops working, at this time, the first combustion chamber 12 and the secondary combustion chamber 3 are still in a subsonic combustion mode, and the second adjusting plate rotates so that the nozzle is still in a contraction state; as the speed increases, when the speed exceeds a second threshold, the second adjusting plate of the nozzle is rotated according to the needs of the double-flow combustion chamber, so that the second adjusting plate rotates to a horizontal position, at this time, the expansion angle of the nozzle is maximum, the double-flow combustion chamber is in a high-performance working state, and the nozzle obtains larger thrust.

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

[0035] The above description is merely specific implementation of the present application, but the protection scope of the present application is not limited to this, any person skilled in the art can easily think of changes or replacements within the technical range disclosed by the present application, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection scope of the claims.

Claims

1. A dual-flow-channel combustion chamber, characterized in that, 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 smaller than the expansion ratio of the second 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 according to claim 1, characterized in that, The expansion ratio of the first flow channel is 1.45~1.

55.

3. The dual-flow combustion chamber according to claim 1, characterized in that, The expansion ratio of the second flow channel is 1.55~1.

65.

4. The dual-flow combustion chamber according to claim 1, characterized in that, 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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