CN116122989A - RBCC combustion chamber with two-stage rocket layout and combustion organization method - Google Patents
RBCC combustion chamber with two-stage rocket layout and combustion organization method Download PDFInfo
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- CN116122989A CN116122989A CN202310284660.6A CN202310284660A CN116122989A CN 116122989 A CN116122989 A CN 116122989A CN 202310284660 A CN202310284660 A CN 202310284660A CN 116122989 A CN116122989 A CN 116122989A
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- support plate
- pipe
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/18—Composite ram-jet/rocket engines
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- Engineering & Computer Science (AREA)
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- Combustion & Propulsion (AREA)
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- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Abstract
The invention discloses a RBCC combustion chamber with a two-stage rocket layout, which comprises: the combustion chamber casing, at least one extension board rocket installs on the combustion chamber casing through the ejector port, and an extension board rocket includes: the support plate combustion chamber is positioned outside the combustion chamber shell, is used for generating high-temperature and high-pressure gas, a gas pipe, an injection pipe and a fuel pipe in the combustion chamber shell, is obliquely arranged, the upper end of the support plate combustion chamber is communicated with an external fuel supply system, the lower end of the support plate combustion chamber extends into the combustion chamber shell through an injection hole and injects fuel into the combustion chamber shell, and at least one inclined cutting rocket is arranged on the combustion chamber shell through an extrusion hole and is used for extruding and accelerating subsonic gas in an injection mode at the rear end of the combustion chamber and providing most of thrust for normal operation of the injection mode of the engine; the jet flow of the RBCC engine can be changed by adjusting the flow and the oxygen-fuel ratio of the support plate rocket, so that secondary fuel is injected to organize secondary combustion to release heat, and the thrust required by the engine is provided.
Description
Technical Field
The invention belongs to the technical field of rocket ramjet engines, and particularly relates to a RBCC combustion chamber with a two-stage rocket layout and a combustion organization method.
Background
The Rocket-Based Combined-Cycle (RBCC) engine is a propulsion system integrating a Rocket engine with high thrust-weight ratio and an air-breathing ramjet engine with high specific impulse, and comprises injection, sub-combustion, super-combustion and pure Rocket modes. The rocket can play a very important role in the whole working process of the RBCC, and mainly plays roles of injecting the incoming flow atmosphere and the fuel support plate, generating thrust and igniting and stabilizing flame in an injection mode; the ignition and flame stabilization effects are mainly exerted in the sub-combustion mode and the super-combustion mode; in pure rocket mode, the required thrust is mainly provided. Therefore, the layout of the rocket has an extremely important influence on the operating range and performance of the RBCC engine.
The RBCC engine in the prior art has the problems that the RBCC ejection mode performance is poor, the engine is difficult to start, and a thrust trap is easy to appear in the mode conversion process.
Disclosure of Invention
The invention aims to provide a RBCC combustion chamber with a two-stage rocket layout and a combustion organization method, which are used for solving the problems that RBCC ejection mode performance is poor, engine starting is difficult, and thrust traps are easy to occur in the mode conversion process.
The invention adopts the following technical scheme: a two-stage rocket layout RBCC combustion chamber comprising:
the combustion chamber shell is a columnar structure with two open ends, the two ends of the combustion chamber shell are respectively communicated with an isolation section and a spray pipe of the engine, at least one injection hole is arranged on the side wall of the combustion chamber shell, which is close to the isolation section, at least one extrusion hole is arranged on the side wall of the combustion chamber shell, which is close to the spray pipe,
at least one support plate rocket is arranged on the combustion chamber shell through injection holes and used for changing the injection flow of the RBCC engine by adjusting the flow and the oxygen-fuel ratio of the support plate rocket so as to inject secondary fuel tissue for secondary combustion and heat release, thereby providing required thrust for the engine,
a stent rocket comprising:
the support plate combustion chamber is positioned outside the combustion chamber shell and used for generating high-temperature high-pressure gas,
the upper end of the gas pipe is communicated with the outlet of the supporting plate combustion chamber, the lower end of the gas pipe extends into the combustion chamber shell through the injection hole,
the jet pipe is horizontally arranged, the left end of the jet pipe is closed, the inner diameter of the opening at the right end of the jet pipe is firstly reduced from left to right and then increased to form a horn-shaped nozzle, the axis is parallel to the axial direction of the combustion chamber shell, the outer wall of the jet pipe is provided with an air inlet hole, the air inlet hole is communicated with the lower end of the gas pipe, so that the gas in the gas pipe enters the jet pipe from top to bottom along the gas pipe through the air inlet hole, and is accelerated in the jet pipe and then is ejected from the nozzle to right,
the fuel pipe is obliquely arranged, is parallel to the gas pipe and is close to the gas pipe, the upper end of the fuel pipe is communicated with an external fuel supply system, the lower end of the fuel pipe extends into the combustion chamber shell through the injection hole and injects fuel into the combustion chamber shell,
the at least one beveling rocket is arranged on the combustion chamber shell through the extrusion hole and is used for extruding and accelerating subsonic fuel gas in an injection mode at the rear end of the combustion chamber and providing most of thrust for normal operation of the injection mode of the engine.
Further, a chamfer rocket includes:
a chamfer combustion chamber located outside the combustion chamber housing for generating high temperature and high pressure gas therein,
the upper end of the inclined cutting air duct is communicated with the outlet of the inclined cutting combustion chamber, the lower end of the inclined cutting air duct is communicated with the combustion chamber shell through an extrusion hole, the inclined cutting air duct sequentially comprises an equal straight section, a contraction section and an expansion section from top to bottom, the equal straight section, the contraction section and the expansion section are sequentially communicated to form the inclined cutting air duct, the inner diameters of the equal straight section are equal, the inner diameters of the contraction section are gradually reduced from top to bottom, and the expansion section is gradually increased from top to bottom and stretches into the combustion chamber shell.
Further, the included angle between the central axis of the inclined cutting air duct and the central axis of the combustion chamber shell is 1-30 degrees.
Further, the included angle between the central axis of the gas pipe and the central axis of the combustion chamber shell is 30-60 degrees.
Further, the support plate rocket is provided with a plurality of support plate rockets and is radially distributed around the outer wall of the combustion chamber shell.
Further, the bevel rockets are provided with a plurality of bevel rockets and are arranged radially around the outer wall of the combustion chamber shell.
The RBCC combustion chamber with two-stage rocket layout and the combustion organization method thereof comprise the following steps:
the support plate combustion chamber of the support plate rocket and the beveling combustion chamber of the beveling rocket are opened, so that the support plate combustion chamber conveys high-temperature and high-pressure fuel gas into the combustion chamber shell through the fuel gas pipe, thereby injecting incoming air, and the beveling combustion chamber conveys the high-temperature and high-pressure fuel gas into the combustion chamber shell through the beveling gas guide pipe;
closing the chamfer combustion chamber of the chamfer rocket when the flight Mach number is gradually increased and is 2.5-3;
when the flight Mach number is gradually increased and is 3-3.5, the support plate combustion chamber of the support plate rocket is closed, and the incoming flow directly enters the combustion chamber shell through the isolation section to participate in the combustion reaction, and at the moment, the support plate rocket only plays roles of fuel injection and flame stabilization;
starting a chamfer combustion chamber of the chamfer rocket when the flight Mach number is gradually increased and is 10, and providing required thrust for the engine by the chamfer rocket;
wherein the RBCC combustion chamber of the two-stage rocket layout is any one of the combustion chambers.
The beneficial effects of the invention are as follows:
1. according to the invention, the injection flow of the RBCC engine can be changed by adjusting the flow and the oxygen-fuel ratio of the support plate rocket, so that secondary fuel is injected to organize secondary combustion for releasing heat, and the thrust required by the engine is provided;
2. the bevel rocket plume can carry out extrusion acceleration to a certain extent on subsonic fuel gas in an injection mode at the rear end of the combustion chamber, and provides most of thrust for normal operation of the injection mode of the engine;
3. when the engine is actually operated, the engine working mode is gradually changed from injection to sub-combustion along with the gradual increase of the Mach number of incoming flow, the air inlet channel is gradually started in the process, the back pressure resistance of the air inlet channel is increased along with the gradual increase of the incoming flow entering the combustion chamber shell of the engine, and the combustion chamber is slowly provided with the self-sustaining combustion heat release capacity, so that the support plate rocket can be gradually closed and only plays roles of injecting secondary fuel and stabilizing flame, and the beveling rocket is limited in specific impulse along with the increase of the intensity of secondary combustion heat release, and continuously opened to only react with the total specific impulse of the RBCC engine, so that the beveling rocket is also gradually closed; the performance of the RBCC engine in injection and sub-combustion modes is improved by combining the support plate rocket and the bevel rocket.
Drawings
FIG. 1 is a schematic diagram of the structure of the present invention;
FIG. 2 is a schematic perspective view of the present invention;
FIG. 3 is a cross-sectional view of the present invention;
fig. 4 is a cross-sectional view of the present invention.
Wherein: 1. a combustion chamber housing; 2. a support plate rocket; 3. a support plate combustion chamber; 4. a gas pipe; 5. a jet pipe; 6. a beveling rocket; 7. a chamfer combustor; 8. chamfering the airway tube; 9. a straight section; 10. a constriction section; 11. an expansion section; 12. an isolation section; 13. a spray pipe.
Detailed Description
The invention will be described in detail below with reference to the drawings and the detailed description.
In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, are merely for convenience in describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and the like, are used for descriptive purposes only and are not to be construed as indicating or implying a relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defining "a first" or "a second" may explicitly or implicitly include one or more such feature. In the description of the present invention, unless otherwise indicated, the meaning of "a plurality" is two or more.
The invention discloses a RBCC combustion chamber with a two-stage rocket layout, which is shown in figures 1-4 and comprises a combustion chamber shell 1, at least one support plate rocket 2 and at least one beveling rocket 6.
The combustion chamber shell 1 is of a columnar structure with two open ends, the two ends of the combustion chamber shell 1 are respectively communicated with the isolation section 12 and the spray pipe 13 of the engine, at least one injection hole is formed in the side wall, close to the isolation section 12, of the combustion chamber shell 1, and at least one extrusion hole is formed in the side wall, close to the spray pipe 13, of the combustion chamber shell 1.
At least one support plate rocket 2 is arranged on the combustion chamber shell 1 through injection holes, and the support plate rocket 2 is used for changing the injection flow of the RBCC engine by adjusting the flow and the oxygen-fuel ratio of the support plate rocket 2, so that secondary fuel is injected for organizing secondary combustion for releasing heat, and the required thrust is provided for the engine. The support plate rockets 2 are arranged in a plurality of ways and are radially distributed around the outer wall of the combustion chamber shell 1.
A launch pad 2 comprising: the support plate combustion chamber 3, the gas pipe 4, the injection pipe 5 and the fuel pipe are arranged outside the combustion chamber shell 1, and the support plate combustion chamber 3 is used for generating high-temperature and high-pressure fuel gas.
The gas pipe 4 is obliquely arranged, the upper end of the gas pipe 4 is communicated with the outlet of the supporting plate combustion chamber 3, and the lower end of the gas pipe 4 extends into the combustion chamber shell 1 through the injection hole. The included angle between the central axis of the gas pipe 4 and the central axis of the combustion chamber shell 1 is 30-60 degrees.
The injection pipe 5 is horizontally arranged, the left end of the injection pipe 5 is closed, the inner diameter of the opening at the right end of the injection pipe 5 is firstly reduced from left to right and then increased to form a horn-shaped nozzle, the axis of the injection pipe 5 is parallel to the axial direction of the combustion chamber shell 1, the outer wall of the injection pipe 5 is provided with an air inlet hole, and the air inlet hole is used for being communicated 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 along the gas pipe 4 through the air inlet hole, and is sprayed rightward from the nozzle after being accelerated in the injection pipe 5, and the nozzle expands and accelerates the gas.
The fuel pipe is obliquely arranged, is parallel to the gas pipe 4 and is arranged close to the gas pipe 4, the upper end of the fuel pipe is communicated with an external fuel supply system, and the lower end of the fuel pipe extends into the combustion chamber shell 1 through the injection hole and injects fuel into the combustion chamber shell 1.
At least one inclined cutting rocket 6 is arranged on the combustion chamber shell 1 through an extrusion hole, and the inclined cutting rocket 6 is used for carrying out extrusion acceleration to a certain degree on subsonic fuel gas in an ejection mode at the rear end of the combustion chamber and providing most of thrust for normal operation in the ejection mode for the engine. The bevel rockets 6 are provided in plurality and are arranged radially around the outer wall of the combustion chamber housing 1.
A chamfer rocket 6 comprising: a chamfer combustor 7 and a chamfer air duct 8, wherein the chamfer combustor 7 is positioned outside the combustor housing 1, and the chamfer combustor 7 is used for generating high-temperature and high-pressure fuel gas.
The inclined cutting air duct 8 is obliquely arranged, the upper end of the inclined cutting air duct 8 is communicated with the outlet of the inclined cutting combustion chamber 7, the lower end of the inclined cutting air duct 8 is communicated with the combustion chamber shell 1 through an extrusion hole, the inclined cutting air duct 8 sequentially comprises an equal straight section 9, a shrinkage section 10 and an expansion section 11 from top to bottom, the equal straight section 9, the shrinkage section 10 and the expansion section 11 sequentially communicate to form the inclined cutting air duct 8, the inner diameters of the equal straight section 9 are equal, the inner diameter of the shrinkage section 10 gradually reduces from top to bottom, the expansion section 11 gradually increases from top to bottom, and the inclined cutting air duct extends into the combustion chamber shell 1. The high-temperature high-pressure fuel gas is generated in the inclined cutting combustion chamber 7, directly enters the equal straight section 9, sequentially expands and accelerates through the contraction section 10 and the expansion section 11, and finally enters the combustion chamber shell 1. The included angle between the central axis of the inclined cutting air duct 8 and the central axis of the combustion chamber shell 1 is 1-30 degrees.
The invention also discloses a RBCC combustion chamber with a two-stage rocket layout and a combustion organization method, which comprise the following steps:
the support plate combustion chamber 3 of the support plate rocket 2 and the inclined cutting combustion chamber 7 of the inclined cutting rocket 6 are opened, so that the support plate combustion chamber 3 conveys high-temperature and high-pressure fuel gas into the combustion chamber shell 1 through the fuel gas pipe 4, thereby injecting the incoming air, and the inclined cutting combustion chamber 7 conveys high-temperature and high-pressure fuel gas into the combustion chamber shell 1 through the inclined cutting gas guide pipe 8;
closing the chamfer combustion chamber 7 of the chamfer rocket 6 when the flight Mach number is gradually increased and is 2.5-3;
when the flight Mach number is gradually increased and is 3-3.5, the support plate combustion chamber 3 of the support plate rocket 2 is closed, and the incoming flow directly enters the combustion chamber shell 1 through the isolation section 12 to participate in the combustion reaction, and at the moment, the support plate rocket 2 only plays roles of fuel injection and flame stabilization;
starting a chamfer combustion chamber 7 of the chamfer rocket 6 when the flight Mach number is gradually increased and is 10, and providing required thrust for an engine by the chamfer rocket 6;
wherein the RBCC combustion chamber of the two-stage rocket layout is any one of the combustion chambers.
According to the invention, along with the gradual increase of the Mach number of the incoming flow, the support plate rocket 2 and the beveling rocket 6 are gradually closed, so that the improvement of the overall performance of the engine is facilitated, the closed support plate rocket 2 can also be used as a mode for injecting fuel, the support plate rocket 2 is positioned in the flow channel, and the fuel injection through the fuel pipe can greatly facilitate the mixing of the fuel and the incoming flow atmosphere.
The foregoing description of the preferred embodiments of the invention is not intended to limit the invention to the precise form disclosed, and any such modifications, equivalents, and alternatives falling within the spirit and scope of the invention are intended to be included within the scope of the invention.
Claims (7)
1. A two-stage rocket-deployed RBCC combustion chamber, comprising:
the combustion chamber shell (1) is of a columnar structure with two open ends, the two ends of the combustion chamber shell are respectively communicated with an isolation section (12) and a spray pipe (13) of the engine, at least one injection hole is formed in the side wall, close to the isolation section (12), of the combustion chamber shell (1), at least one extrusion hole is formed in the side wall, close to the spray pipe (13), of the combustion chamber shell (1),
at least one support plate rocket (2) is arranged on the combustion chamber shell (1) through injection holes and is used for changing the injection flow of the RBCC engine by adjusting the flow and the oxygen-fuel ratio of the support plate rocket (2) so as to inject secondary fuel tissue secondary combustion heat release, thereby providing the engine with required thrust,
one of the plate rockets (2) comprises:
a support plate combustion chamber (3) which is positioned outside the combustion chamber shell (1) and is used for generating high-temperature high-pressure gas,
the gas pipe (4) is obliquely arranged, the upper end of the gas pipe is communicated with the outlet of the supporting plate combustion chamber (3), the lower end of the gas pipe extends into the combustion chamber shell (1) through the injection hole,
the jet pipe (5) is horizontally arranged, the left end of the jet pipe is closed, the inner diameter of the opening at the right end of the jet pipe is firstly reduced from left to right and then enlarged to form a horn-shaped nozzle, the axis is parallel to the axial direction of the combustion chamber shell (1), the outer wall of the jet pipe is provided with an air inlet hole, the air inlet hole is communicated with the lower end of the gas pipe (4), so that the gas in the gas pipe (4) enters the jet pipe (5) from top to bottom along the gas pipe (4) through the air inlet hole, and is accelerated in the jet pipe (5) and then is ejected from the nozzle to the right,
the fuel pipe is obliquely arranged and parallel to the gas pipe (4) and is arranged close to the gas pipe (4), the upper end of the fuel pipe is communicated with an external fuel supply system, the lower end of the fuel pipe extends into the combustion chamber shell (1) through the injection hole and injects fuel into the combustion chamber shell (1),
the at least one beveling rocket (6) is arranged on the combustion chamber shell (1) through the extrusion hole, is used for extruding and accelerating subsonic fuel gas in injection mode at the rear end of the combustion chamber, and provides most of thrust for normal operation of the injection mode of the engine.
2. A two-stage rocket-deployed RBCC combustion chamber according to claim 1, wherein one of said chamfer rockets (6) comprises:
a chamfer combustion chamber (7) located outside the combustion chamber housing (1) for generating high temperature and high pressure gas therein,
the inclined cutting air duct (8) is obliquely arranged, the upper end of the inclined cutting air duct is communicated with the outlet of the inclined cutting combustion chamber (7), the lower end of the inclined cutting air duct is communicated with the combustion chamber shell (1) through an extrusion hole, the inclined cutting air duct (8) is composed of an equal straight section (9), a shrinkage section (10) and an expansion section (11) from top to bottom, the equal straight section (9), the shrinkage section (10) and the expansion section (11) are sequentially communicated to form the inclined cutting air duct (8), the inner diameters of the equal straight section (9) are equal, the inner diameter of the shrinkage section (10) is gradually reduced from top to bottom, and the expansion section (11) is gradually increased from top to bottom and stretches into the combustion chamber shell (1).
3. A two-stage rocket layout RBCC combustion chamber according to claim 2, wherein the angle between the central axis of the bevelled air duct (8) and the central axis of the combustion chamber housing (1) is 1-30 degrees.
4. A two-stage rocket layout RBCC combustion chamber according to claim 1, wherein the central axis of the gas pipe (4) is at an angle of 30-60 degrees to the central axis of the combustion chamber housing (1).
5. A two-stage rocket layout RBCC combustion chamber according to any of claims 2-4, wherein the support plate rocket (2) is provided with a plurality of support plate rockets and is arranged radially around the outer wall of the combustion chamber housing (1).
6. A two-stage rocket layout RBCC combustion chamber according to claim 5, wherein the bevelled rockets (6) are arranged in a plurality and radially around the outer wall of the combustion chamber housing (1).
7. The RBCC combustion chamber with the two-stage rocket layout and the combustion organization method are characterized by comprising the following steps:
starting a support plate combustion chamber (3) of a RBCC (two-stage rocket) combustion chamber support plate rocket (2) and a beveling combustion chamber (7) of a beveling rocket (6), so that the support plate combustion chamber (3) conveys high-temperature and high-pressure fuel gas into a combustion chamber shell (1) through a fuel gas pipe (4), thereby injecting inflow air, and the beveling combustion chamber (7) conveys high-temperature and high-pressure fuel gas into the combustion chamber shell (1) through a beveling gas guide pipe (8);
closing a chamfer combustion chamber (7) of the chamfer rocket (6) when the flight Mach number is gradually increased and is 2.5-3;
when the flight Mach number is gradually increased and is 3-3.5, the support plate combustion chamber (3) of the support plate rocket (2) is closed, and incoming flow directly enters the combustion chamber shell (1) through the isolation section (12) to participate in combustion reaction, and at the moment, the support plate rocket (2) only plays roles of fuel injection and flame stabilization;
starting a bevel combustion chamber (7) of the bevel rocket (6) when the flight Mach number is gradually increased and is 10, wherein the bevel rocket (6) provides required thrust for the engine;
wherein the two-stage rocket layout RBCC combustion chamber is the combustion chamber of any of claims 1-6.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202310284660.6A CN116122989A (en) | 2023-03-22 | 2023-03-22 | RBCC combustion chamber with two-stage rocket layout and combustion organization method |
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| Application Number | Priority Date | Filing Date | Title |
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| CN202310284660.6A CN116122989A (en) | 2023-03-22 | 2023-03-22 | RBCC combustion chamber with two-stage rocket layout and combustion organization method |
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| CN116122989A true CN116122989A (en) | 2023-05-16 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN202310284660.6A Pending CN116122989A (en) | 2023-03-22 | 2023-03-22 | RBCC combustion chamber with two-stage rocket layout and combustion organization method |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116447042A (en) * | 2023-06-09 | 2023-07-18 | 西安航天动力研究所 | Rocket-based combined engine |
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2023
- 2023-03-22 CN CN202310284660.6A patent/CN116122989A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN116447042A (en) * | 2023-06-09 | 2023-07-18 | 西安航天动力研究所 | Rocket-based combined engine |
| CN116447042B (en) * | 2023-06-09 | 2023-10-20 | 西安航天动力研究所 | Rocket-based combined engine |
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