A two-stage rocket layout RBCC combustion chamber and combustion organization method

Abstract

The application discloses a two-stage rocket layout RBCC combustion chamber, which comprises a combustion chamber shell, at least one strut rocket installed on the combustion chamber shell through an injection hole, one strut rocket comprising a strut combustion chamber located outside the combustion chamber shell and used for generating high-temperature and high-pressure gas, a gas pipe, a jet pipe, a fuel pipe, which is obliquely arranged, and the upper end of which is connected with an external fuel supply system, and the lower end of which extends into the combustion chamber shell through the injection hole and sprays fuel into the combustion chamber shell, and at least one oblique cutting rocket installed on the combustion chamber shell through an extrusion hole, which is used for extruding and accelerating the subsonic gas in the injection mode at the rear end of the combustion chamber and providing most of the thrust for the normal work of the engine in the injection mode; the application can change the injection flow of the RBCC engine by adjusting the flow and oxygen-fuel ratio of the strut rocket, and then injects secondary fuel to organize secondary combustion heat release, so as to provide the required thrust of the engine.

Description

Technical Field

[0001] This invention belongs to the field of rocket ramjet engine technology, and particularly relates to an RBCC combustion chamber with a two-stage rocket layout and a combustion organization method. Background Technology

[0002] A rocket-based combined-cycle (RBCC) engine is a propulsion system that integrates a high thrust-to-weight ratio rocket engine with a high specific impulse air-breathing ramjet engine. It includes ejection, subsonic combustion, scramjet, and pure rocket modes. The rocket plays a crucial role throughout the entire RBCC's operation. In ejection mode, it primarily ejects the incoming atmospheric flow, injects fuel into the fuel struts, generates thrust, and provides ignition and flame stabilization. In subsonic and scramjet modes, it mainly provides ignition and flame stabilization. In pure rocket mode, it primarily provides the required thrust. Therefore, the rocket's configuration has a significant impact on the RBCC engine's operating range and performance.

[0003] Existing RBCC engines suffer from poor RBCC ejector mode performance, difficulty in starting the engine, and the tendency for thrust traps to occur during mode transitions. Summary of the Invention

[0004] The purpose of this invention is to provide a two-stage rocket layout RBCC combustion chamber and combustion organization method to solve the problems of poor RBCC ejector mode performance, difficult engine start-up, and easy occurrence of thrust traps during mode transition.

[0005] This invention adopts the following technical solution: an RBCC combustor with a two-stage rocket layout, comprising:

[0006] The combustion chamber shell is a cylindrical structure open at both ends, with its two ends connected to the engine's isolation section and the nozzle, respectively. At least one injection port is formed on the side wall of the combustion chamber shell near the isolation section, and at least one extrusion port is formed on the side wall of the combustion chamber shell near the nozzle.

[0007] At least one bolster rocket, mounted on the combustion chamber shell through an ejector port, is used to change the ejector flow rate of the RBCC engine by adjusting the flow rate and oxygen-fuel ratio of the bolster rocket, thereby injecting secondary fuel to organize secondary combustion and heat release, thus providing the engine with the required thrust.

[0008] A support plate rocket includes:

[0009] The support plate combustion chamber, located outside the combustion chamber shell, is used to generate high-temperature, high-pressure combustion gases.

[0010] The gas pipe is installed at an angle, with its upper end connected to the outlet of the combustion chamber of the support plate, and its lower end extending into the combustion chamber shell through an injection hole.

[0011] The injection pipe is horizontally positioned, closed at its left end and open at its right end. The inner diameter of the opening first narrows and then widens from left to right, forming a trumpet-shaped nozzle. Its axis is parallel to the axis of the combustion chamber shell. An air inlet is provided on its outer wall, which connects to the lower end of the gas pipe. This allows the gas in the gas pipe to flow from top to bottom through the air inlet into the injection pipe, where it is accelerated before being ejected to the right from the nozzle.

[0012] The fuel pipe is inclined and parallel to and close to the gas pipe. Its upper end is connected to the external fuel supply system, and its lower end extends into the combustion chamber housing through an injection port to inject fuel into the combustion chamber housing.

[0013] At least one oblique-cut rocket, mounted on the combustion chamber shell through a compression hole, is used to compress and accelerate the subsonic gas of the ejector mode at the rear end of the combustion chamber and to provide most of the thrust for normal operation of the engine ejector mode.

[0014] Furthermore, a slanted rocket includes:

[0015] The oblique-cut combustion chamber, located outside the combustion chamber shell, is used to generate high-temperature, high-pressure combustion gases.

[0016] The obliquely cut air guide pipe is set at an angle, with its upper end connected to the outlet of the obliquely cut combustion chamber and its lower end connected to the combustion chamber shell through a compression hole. The obliquely cut air guide pipe consists of a straight section, a contraction section, and an expansion section from top to bottom. The straight section, contraction section, and expansion section are connected in sequence to form the obliquely cut air guide pipe. The inner diameter of the straight section is equal, the inner diameter of the contraction section gradually decreases from top to bottom, and the inner diameter of the expansion section gradually increases from top to bottom and extends into the combustion chamber shell.

[0017] Furthermore, the angle between the central axis of the obliquely cut air guide pipe and the central axis of the combustion chamber shell is 1-30 degrees.

[0018] Furthermore, the angle between the central axis of the gas pipe and the central axis of the combustion chamber shell is 30-60 degrees.

[0019] Furthermore, multiple support rockets are installed, and they are arranged radially around the outer wall of the combustion chamber shell.

[0020] Furthermore, multiple oblique-cut rockets are installed and arranged radially around the outer wall of the combustion chamber shell.

[0021] A two-stage rocket RBCC combustion chamber and combustion organization method comprises the following steps:

[0022] The support combustion chamber of the support rocket and the oblique combustion chamber of the oblique rocket are opened, so that the support combustion chamber delivers high-temperature and high-pressure gas into the combustion chamber shell through the gas pipe, thereby ejecting the incoming airflow, and the oblique combustion chamber delivers high-temperature and high-pressure gas into the combustion chamber shell through the oblique gas guide pipe.

[0023] As the flight Mach number gradually increases to 2.5-3, the oblique combustion chamber of the oblique-cut rocket is shut down;

[0024] As the flight Mach number gradually increases to 3-3.5, the support plate combustion chamber of the support plate rocket is closed, and the incoming flow passes through the isolation section and directly enters the combustion chamber shell to participate in the combustion reaction. At this time, the support plate rocket only plays the role of fuel injection and flame stabilization.

[0025] As the flight Mach number gradually increases, and reaches 10, the oblique combustion chamber of the oblique rocket is opened, and the oblique rocket provides the required thrust to the engine;

[0026] Among them, the RBCC combustor in the two-stage rocket layout is any of the combustors mentioned above.

[0027] The beneficial effects of this invention are:

[0028] 1. This invention can change the ejector flow rate of the RBCC engine by adjusting the flow rate and oxygen-fuel ratio of the support rocket, thereby injecting secondary fuel to organize secondary combustion and heat release, thus providing the thrust required by the engine.

[0029] 2. The oblique rocket plume of the present invention can compress and accelerate the subsonic gas in the ejector mode to a certain extent at the rear end of the combustion chamber, and provide most of the thrust for the normal operation of the engine ejector mode;

[0030] 3. In actual engine operation, as the incoming Mach number gradually increases, the engine operating mode slowly transitions from ejection to sub-fuel combustion. During this process, the intake gradually starts up, and its back pressure resistance increases accordingly. Furthermore, the incoming flow rate into the engine combustion chamber casing gradually increases, and the combustion chamber gradually acquires the ability to sustain combustion and heat release. Therefore, the support rocket can be gradually shut down, serving only to inject secondary fuel and stabilize the flame. As the intensity of secondary combustion heat release increases, the specific impulse of the slant rocket itself is limited. Continuing to keep it open would only have a counter-effect on the total specific impulse of the RBCC engine. Therefore, the slant rocket is also gradually shut down. By combining the support rocket and the slant rocket, the performance of the RBCC engine in both ejection and sub-fuel combustion modes is improved. Attached Figure Description

[0031] Figure 1 This is a schematic diagram of the structure of the present invention;

[0032] Figure 2 This is a three-dimensional schematic diagram of the present invention;

[0033] Figure 3 This is a cross-sectional view of the present invention;

[0034] Figure 4 This is a cross-sectional view of the present invention.

[0035] The components are: 1. Combustion chamber shell; 2. Support plate rocket; 3. Support plate combustion chamber; 4. Gas pipe; 5. Injection pipe; 6. Slanted rocket; 7. Slanted combustion chamber; 8. Slanted air guide pipe; 9. Straight section; 10. Contraction section; 11. Expansion section; 12. Isolation section; 13. Nozzle. Detailed Implementation

[0036] The present invention will now be described in detail with reference to the accompanying drawings and specific embodiments.

[0037] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicating orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, are only for the convenience of describing the 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, and therefore should not be construed as a limitation of the invention. 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 with "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, unless otherwise stated, "multiple" means two or more.

[0038] This invention discloses an RBCC combustion chamber with a two-stage rocket layout, such as... Figures 1-4 As shown, it includes a combustion chamber shell 1, at least one support plate rocket 2, and at least one oblique-cut rocket 6.

[0039] The combustion chamber housing 1 is a columnar structure with openings at both ends. The two ends of the combustion chamber housing 1 are connected to the engine isolation section 12 and the nozzle 13, respectively. At least one ejector hole is opened on the side wall of the combustion chamber housing 1 near the isolation section 12, and at least one extrusion hole is opened on the side wall of the combustion chamber housing 1 near the nozzle 13.

[0040] At least one propeller rocket 2 is mounted on the combustion chamber shell 1 through an ejector port. The propeller rocket 2 is used to change the ejector flow rate of the RBCC engine by adjusting the flow rate and oxygen-fuel ratio of the propeller rocket 2, thereby injecting secondary fuel to organize secondary combustion and heat release, thus providing the engine with the required thrust. Multiple propeller rockets 2 are provided and arranged radially around the outer wall of the combustion chamber shell 1.

[0041] A support plate rocket 2 includes: a support plate combustion chamber 3, a gas pipe 4, an injection pipe 5, and a fuel pipe. The support plate combustion chamber 3 is located outside the combustion chamber shell 1, and is used to generate high-temperature and high-pressure gas inside the support plate combustion chamber 3.

[0042] The gas pipe 4 is inclined, with its upper end connected to the outlet of the combustion chamber 3 of the support plate, and its lower end extending into the combustion chamber housing 1 through an injection hole. The angle between the central axis of the gas pipe 4 and the central axis of the combustion chamber housing 1 is 30-60 degrees.

[0043] The injection pipe 5 is horizontally set, with its left end closed and the inner diameter of the opening at the right end of the injection pipe 5 first decreasing and then increasing from left to right to form a trumpet-shaped nozzle. The axis of the injection pipe 5 is parallel to the axis of the combustion chamber shell 1. An air inlet is provided on the outer wall of the injection pipe 5, which is used to connect with the lower end of the gas pipe 4, so that the gas in the gas pipe 4 enters the injection pipe 5 from top to bottom through the air inlet, is accelerated in the injection pipe 5, and is ejected to the right from the nozzle. The nozzle expands and accelerates the gas.

[0044] The fuel pipe is inclined and parallel to and close to the gas pipe 4. The upper end of the fuel pipe is connected to the external fuel supply system, and the lower end of the fuel pipe extends into the combustion chamber housing 1 through the injection hole and injects fuel into the combustion chamber housing 1.

[0045] At least one oblique-cut rocket 6 is mounted on the combustion chamber shell 1 through a compression hole. The oblique-cut rocket 6 is used to compress and accelerate the subsonic gas of the ejector mode to a certain extent at the rear end of the combustion chamber, and to provide most of the thrust for the engine to operate normally in the ejector mode. Multiple oblique-cut rockets 6 are provided and are arranged radially around the outer wall of the combustion chamber shell 1.

[0046] An oblique-cut rocket 6 includes: an oblique-cut combustion chamber 7 and an oblique-cut gas guide pipe 8. The oblique-cut combustion chamber 7 is located outside the combustion chamber shell 1 and is used to generate high-temperature and high-pressure gas.

[0047] The obliquely cut air guide pipe 8 is inclined, with its upper end connected to the outlet of the obliquely cut combustion chamber 7. The lower end of the obliquely cut air guide pipe 8 is connected to the combustion chamber shell 1 through a compression hole. The obliquely cut air guide pipe 8 consists of a straight section 9, a contraction section 10, and an expansion section 11, arranged sequentially from top to bottom. The straight section 9, contraction section 10, and expansion section 11 are connected in sequence to form the obliquely cut air guide pipe 8. The inner diameter of the straight section 9 is equal, the inner diameter of the contraction section 10 gradually decreases from top to bottom, and the inner diameter of the expansion section 11 gradually increases from top to bottom, extending into the combustion chamber shell 1. High-temperature, high-pressure combustion gas is generated in the obliquely cut combustion chamber 7, directly enters the straight section 9, and then expands and accelerates sequentially through the contraction section 10 and the expansion section 11 before finally entering the combustion chamber shell 1. The angle between the central axis of the obliquely cut air guide pipe 8 and the central axis of the combustion chamber shell 1 is 1-30 degrees.

[0048] This invention also discloses an RBCC combustor with a two-stage rocket layout and a combustion organization method, comprising the following steps:

[0049] The support plate combustion chamber 3 of the support plate rocket 2 and the oblique cutting combustion chamber 7 of the oblique cutting rocket 6 are opened, so that the support plate combustion chamber 3 delivers high-temperature and high-pressure gas into the combustion chamber shell 1 through the gas pipe 4, thereby ejecting the incoming airflow, and the oblique cutting combustion chamber 7 delivers high-temperature and high-pressure gas into the combustion chamber shell 1 through the oblique cutting gas guide pipe 8.

[0050] As the flight Mach number gradually increases to 2.5-3, the oblique combustion chamber 7 of the oblique-cut rocket 6 is shut down;

[0051] As the flight Mach number gradually increases to 3-3.5, the support plate combustion chamber 3 of the support plate rocket 2 is closed, and the incoming flow passes through the isolation section 12 and directly enters the combustion chamber shell 1 to participate in the combustion reaction. At this time, the support plate rocket 2 only plays the role of fuel injection and flame stabilization.

[0052] As the flight Mach number gradually increases and reaches 10, the oblique combustion chamber 7 of the oblique rocket 6 is opened, and the oblique rocket 6 provides the required thrust to the engine;

[0053] Among them, the RBCC combustor in the two-stage rocket layout is any of the combustors mentioned above.

[0054] As the incoming Mach number gradually increases, the present invention gradually closes the support rocket 2 and the oblique rocket 6, which is beneficial to the overall performance of the engine. The closed support rocket 2 can also be used as a way of injecting fuel. The support rocket 2 is located inside the flow channel. At this time, injecting fuel through the fuel pipe can greatly help the mixing of fuel with the incoming atmosphere.

[0055] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. An RBCC combustion chamber with a two-stage rocket layout, characterized in that, include: The combustion chamber housing (1) is a columnar structure with openings at both ends. Its two ends are connected to the engine isolation section (12) and the nozzle (13) respectively. At least one ejector hole is opened on the side wall of the combustion chamber housing (1) near the isolation section (12), and at least one extrusion hole is opened on the side wall of the combustion chamber housing (1) near the nozzle (13). At least one propeller rocket (2) is mounted on the combustion chamber shell (1) through an ejector port. It is used to change the ejector flow rate of the RBCC engine by adjusting the flow rate and oxygen-fuel ratio of the propeller rocket (2), thereby injecting secondary fuel to organize secondary combustion and heat release, and thus providing the engine with the required thrust. One of the said support plate rockets (2) includes: The support plate combustion chamber (3) is located outside the combustion chamber shell (1) and is used to generate high-temperature and high-pressure gas. The gas pipe (4) is inclined, with its upper end connected to the outlet of the support plate combustion chamber (3), and its lower end extending into the combustion chamber shell (1) through an injection hole. The injection pipe (5) is horizontally positioned, with its left end closed and its right end open. The inner diameter of the opening first decreases and then increases from left to right to form a trumpet-shaped nozzle. Its axis is parallel to the axis of the combustion chamber shell (1). An air inlet is provided on its outer wall. The air inlet is used to connect with the lower end of the gas pipe (4), so that the gas in the gas pipe (4) enters the injection pipe (5) from top to bottom through the air inlet, is accelerated in the injection pipe (5), and is then ejected to the right from the nozzle. The fuel pipe is inclined and parallel to and close to the gas pipe (4). Its upper end is connected to the external fuel supply system, and its lower end extends into the combustion chamber housing (1) through the injection hole and injects fuel into the combustion chamber housing (1). At least one oblique rocket (6) is mounted on the combustion chamber shell (1) through a compression hole for compressing and accelerating the subsonic gas of the ejector mode at the rear end of the combustion chamber and providing most of the thrust for normal operation of the engine ejector mode; The combustion organization method for the RBCC combustion chamber in the two-stage rocket configuration comprises the following steps: The support combustion chamber (3) of the RBCC combustion chamber support plate rocket (2) with a two-stage rocket layout and the oblique combustion chamber (7) of the oblique rocket (6) are opened, so that the support combustion chamber (3) delivers high-temperature and high-pressure gas into the combustion chamber shell (1) through the gas pipe (4), thereby ejecting the incoming airflow, and the oblique combustion chamber (7) delivers high-temperature and high-pressure gas into the combustion chamber shell (1) through the oblique gas guide pipe (8); As the flight Mach number gradually increases to 2.5-3, the oblique combustion chamber (7) of the oblique rocket (6) is shut down. As the flight Mach number gradually increases to 3-3.5, the support plate combustion chamber (3) of the support plate rocket (2) is closed, and the incoming flow passes through the isolation section (12) and directly enters the combustion chamber shell (1) to participate in the combustion reaction. At this time, the support plate rocket (2) only plays the role of fuel injection and flame stabilization. As the flight Mach number gradually increases and reaches 10, the oblique combustion chamber (7) of the oblique rocket (6) is opened, and the oblique rocket (6) provides the required thrust to the engine.

2. The RBCC combustion chamber of a two-stage rocket according to claim 1, characterized in that, One of the oblique-cut rockets (6) includes: An obliquely cut combustion chamber (7) is located outside the combustion chamber shell (1) and is used to generate high-temperature and high-pressure combustion gas. The obliquely cut air guide pipe (8) is set at an angle. Its upper end is connected to the outlet of the obliquely cut combustion chamber (7), and its lower end is connected to the combustion chamber shell (1) through the extrusion hole. The obliquely cut air guide pipe (8) is composed of a straight section (9), a contraction section (10), and an expansion section (11) from top to bottom. The straight section (9), the contraction section (10), and the expansion section (11) are connected in sequence to form the obliquely cut air guide pipe (8). The inner diameter of the straight section (9) is equal. The inner diameter of the contraction section (10) gradually decreases from top to bottom. The expansion section (11) gradually increases from top to bottom and extends into the combustion chamber shell (1).

3. The RBCC combustion chamber of a two-stage rocket according to claim 2, characterized in that, The angle between the central axis of the obliquely cut air guide pipe (8) and the central axis of the combustion chamber shell (1) is 1-30 degrees.

4. The RBCC combustion chamber of a two-stage rocket according to claim 1, characterized in that, The angle between the central axis of the gas pipe (4) and the central axis of the combustion chamber shell (1) is 30-60 degrees.

5. The RBCC combustion chamber of a two-stage rocket according to any one of claims 2-4, characterized in that, Multiple support rockets (2) are provided and are arranged radially around the outer wall of the combustion chamber shell (1).

6. The RBCC combustion chamber of a two-stage rocket according to claim 5, characterized in that, Multiple oblique-cut rockets (6) are provided and arranged radially around the outer wall of the combustion chamber shell (1).

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