CN108757179B - Combined cycle engine and hypersonic aircraft - Google Patents
Combined cycle engine and hypersonic aircraft Download PDFInfo
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- CN108757179B CN108757179B CN201810529474.3A CN201810529474A CN108757179B CN 108757179 B CN108757179 B CN 108757179B CN 201810529474 A CN201810529474 A CN 201810529474A CN 108757179 B CN108757179 B CN 108757179B
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/32—Inducing air flow by fluid jet, e.g. ejector action
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
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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
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/78—Other construction of jet pipes
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fuel-Injection Apparatus (AREA)
Abstract
The invention discloses a combined cycle engine and a hypersonic aircraft, which comprise a rotary knocking ramjet engine and a rocket engine arranged in the rotary knocking ramjet engine. A rotary detonation ramjet engine comprising: the shell is in a hollow cylinder shape with two communicated ends. The rear body of the central cone extends into the shell from the air inlet end of the shell and is connected with the shell, a drainage channel for introducing air into the shell is formed in a gap between the shell and the rear body, a detonation chamber communicated with the drainage channel is formed by a cavity between the rear end face of the rear body and the inner wall of the shell, and a first spray pipe communicated with the detonation chamber. The rocket engine includes: the combustion chamber is arranged in the central cone, the second spray pipe is communicated with the combustion chamber, and a nozzle of the second spray pipe is communicated with the detonation chamber. The outer wall of the shell is provided with an outer nozzle, and two ends of the outer nozzle are respectively communicated with the fuel supply device and the drainage channel. And/or a plurality of inner nozzles are processed on the outer wall of the rear body, and two ends of each inner nozzle are respectively communicated with the fuel supply device and the drainage channel.
Description
Technical Field
the invention relates to the field of aircraft engines, in particular to a combined cycle engine. Furthermore, the invention also relates to a hypersonic aircraft comprising the combined cycle engine.
Background
The hypersonic aircraft has the flight speed of more than 5 times of sound velocity, the flight height of 18-35 kilometers generally, is a novel 'killer mace' weapon with remote, quick and accurate striking capability, is the focus of competition in the current international aviation aerospace field, and has the core of a hypersonic propulsion technology. The scramjet engine is a hypersonic propulsion device that has been extensively studied to date to organize combustion in an isobaric mode. The continuous rotation detonation ramjet engine is a novel ram propulsion scheme recently proposed, combustion is organized in a detonation mode, the combustion has higher thermal cycle efficiency compared with isobaric combustion, and a power device with higher performance is expected to be provided for a hypersonic aircraft. The documents [ Wang Chao, Liu Weidong, Liu Shijie, Jiang Luxin, Linzhiong.Experimental version of Air-breaking Continuous rotation testing Fuel by Hydrogen, International Journal of Hydrogen Energy,2015,40:9530-9538 ], and [ Shijie Liu, WeidingLiu, Yi Wang, Zhiying Lin, Free Jet Test of Continuous rotation testing Jet Engine, AIAA 2017-2282 ] respectively verify the feasibility of a Continuous rotation Detonation Ramjet Engine by direct connection and Free Jet Test.
Although the key technology of the prior scramjet engine is broken through and the feasibility of the continuous rotation detonation ramjet engine is verified, the engines can only work in a ramjet mode. If zero-speed starting and hypersonic cruise flight are realized, the ramjet engine and other power are combined to develop the technical research of the combined engine. Two common types of combination are rocket-based combined engines (RBCC) and turbine-based combined engines (TBCC), each of which has advantages and disadvantages.
the RBCC is an organic integration of rocket engine technology and ramjet engine technology, has the advantages of high specific thrust of the ramjet engine and high thrust-weight ratio of the rocket engine, and has the working capacity of full speed domain and full airspace. The RBCC is originally designed as a single-stage on-track power system, but can also be used for a zero-speed start and hypersonic cruise flight power system.
the traditional RBCC engine mainly comprises an injection rocket and a ramjet engine runner, and the rocket can be arranged at a proper position in the runner as required. For a hypersonic cruise RBCC engine, three working modes can be divided, and as shown in FIG. 1, when Ma is 0-3, the engine works in an injection mode. Under the dual functions of ejection and speed stamping, air enters the flow channel, is mixed with high-temperature fuel gas generated by the ejection rocket in the ejection mixing area, fuel is sprayed into the mixed gas for secondary afterburning, and the ejection rocket and the downstream afterburning can generate thrust. As shown in fig. 2, when Ma is 3-5, the RBCC operates in a sub-combustion mode, the ejector rocket is closed, the pressure of the high-total-pressure subsonic airflow captured by the air inlet is recovered at the upstream of the combustion chamber, fuel is injected at the downstream of the combustion chamber for subsonic combustion, and finally, combustion products are discharged through the tail nozzle. As shown in FIG. 3, when Ma5 is higher, the engine is shifted to a scramjet mode, the ejector rocket is still closed, supersonic airflow captured by an air inlet enters a combustion chamber, fuel is injected at the upstream position of the combustion chamber to realize supersonic combustion, and finally combustion products are discharged through a tail nozzle.
The traditional RBCC engine has more working modes and high mode conversion difficulty; the RBCC engine is organized to burn in an isobaric mode, the thermal cycle efficiency is low, and the thrust performance is poor; in addition, the injection mode needs to inject a mixing area, and the isobaric combustion heat release speed is low, so that a longer combustion chamber is needed, and the whole length of the engine is larger.
disclosure of Invention
the invention provides a combined cycle engine and a hypersonic aircraft, and aims to solve the technical problems of multiple working modes, high mode conversion difficulty, low thermal cycle efficiency, poor thrust performance and long combustion chamber length of the traditional RBCC engine.
The technical scheme adopted by the invention is as follows:
A combined cycle engine comprising a rotary detonation ramjet engine and a rocket engine disposed therein; a rotary detonation ramjet engine comprising: the shell is in a hollow cylinder shape with two communicated ends; the rear body of the central cone extends into the shell from the air inlet end of the shell and is connected with the shell, a drainage channel for introducing air into the shell is formed in a gap between the shell and the rear body, a detonation chamber communicated with the drainage channel and a first spray pipe communicated with the detonation chamber are formed in a cavity between the rear end surface of the rear body and the inner wall of the shell, and a spray nozzle of the first spray pipe is communicated with the atmosphere; the rocket engine includes: the combustion chamber is arranged in the central cone, and a second spray pipe is communicated with the combustion chamber; the outer wall of the shell is provided with a plurality of outer nozzles which are sequentially arranged at intervals along the circumferential direction of the shell, and two ends of each outer nozzle are respectively communicated with the fuel supply device and the drainage channel so that the fuel is sprayed into the drainage channel through the plurality of outer nozzles; and/or a plurality of inner nozzles are processed on the outer wall of the rear body at intervals along the circumferential direction of the rear body, and two ends of each inner nozzle are respectively communicated with the fuel supply device and the drainage channel, so that the fuel is sprayed into the drainage channel through the inner nozzles.
Furthermore, each outer nozzle is 10 mm-20 mm away from the outlet of the drainage channel; and/or the distance between each inner nozzle and the outlet of the drainage channel is 10 mm-20 mm.
further, the plurality of outer nozzles and the plurality of inner nozzles are arranged in a one-to-one correspondence manner, or the plurality of outer nozzles and the plurality of inner nozzles are arranged in a staggered manner.
furthermore, the outer walls of the first spray pipe and the detonation chamber are provided with first wall surface cooling channels, and the first wall surface cooling channels are connected with a fuel supply device; the outer nozzles are respectively communicated with the first wall surface cooling channel, so that the fuel which cools the first spray pipe and the detonation chamber is sprayed into the drainage channel through the outer nozzles.
Furthermore, a second wall surface cooling channel is arranged on the outer wall of the combustion chamber and the second spray pipe, and is connected with the fuel supply device; the inner nozzles and the head nozzles of the combustion chamber are respectively communicated with the second wall surface cooling channel, so that part of the fuel for cooling the second spray pipe and the combustion chamber is sprayed into the flow guide channel by the inner nozzles, and the other part of the fuel is sprayed into the combustion chamber by the head of the combustion chamber.
Further, the detonation chamber is in a straight cylinder shape; the length of the detonation chamber is 400 mm-800 mm.
Furthermore, rotatory detonation ramjet still includes and is used for making many the connection ribs that afterbody and shell link to each other, and many the connection ribs are arranged along drainage channel's circumference interval in proper order, and each connection rib's up end links to each other with the inner wall of shell, and each connection rib's lower terminal surface links to each other with the outer wall of afterbody.
Furthermore, the central cone also comprises a front body which is axially connected with the rear body, the front body is positioned outside the shell, and the outer wall surface of the front body forms an air inlet channel for compressing air; the combustion chamber is positioned in the front body; the second spray pipe is connected with an outlet of the combustion chamber, and extends to be communicated with the rear end face of the rear body along the axial direction of the central cone.
furthermore, an oxidant flow channel for guiding the oxidant and a fuel flow channel for guiding the fuel are arranged in the central cone; the inlet of the oxidant runner is connected with the oxidant supply device, and the outlet of the oxidant runner is connected with the head of the combustion chamber; the inlet of the fuel flow passage is connected with the fuel supply device, and the outlet of the fuel flow passage is connected with the second wall surface cooling channel.
According to another aspect of the invention there is also provided a hypersonic aircraft comprising a combined cycle engine as defined in any one of the preceding claims.
The invention has the following beneficial effects:
the combined cycle engine adopts the scheme of combining the conventional rocket engine and the rotary knocking ramjet engine, wherein the rocket engine is positioned in the central cone of the air inlet passage of the rotary knocking ramjet engine, and compared with the conventional RBCC engine, the combined cycle engine has a simpler and more compact integral structure. In the combined cycle engine, the rotary knocking ramjet engine organizes combustion in a rotary knocking mode, and the conventional RBCC engine organizes combustion in an isobaric mode, so that compared with the conventional RBCC engine, the combined cycle engine has higher cycle thermal efficiency and better thrust performance, is more suitable for remote hypersonic cruise flight and has better economic performance; in addition, for the conventional RBCC, an injection mixing area is required in an injection mode, and the isobaric combustion heat release speed is low, so that a longer combustion chamber is required, and the overall length of the engine is larger. In the invention, the detonation combustion has high heat release speed, so that the required length of the combustion chamber is short, and the whole length of the engine is shorter. The traditional RBCC engine has more working modes (including an injection mode, a sub-combustion stamping mode and a super-combustion stamping mode) and large mode conversion difficulty, and the engine has less working modes (including the injection mode and the stamping mode) and small mode conversion difficulty. In the prior art, the rotary detonation engine adopts the annular combustion chamber, and in the scheme of the application, the shell is in a hollow cylinder shape with two communicated ends, and only the rear body of the central cone extends into the shell, so that the detonation chamber is in a hollow cylinder shape, and compared with the annular combustion chamber, the cylindrical combustion chamber has stronger detonation combustion organization capacity and good combustion stability of flame;
The hypersonic aircraft has the advantages of small working mode, small mode conversion difficulty, high thermal cycle efficiency, good thrust performance and short combustion chamber length.
In addition to the objects, features and advantages described above, other objects, features and advantages of the present invention are also provided. The present invention will be described in further detail below with reference to the drawings.
Drawings
the accompanying drawings, which are incorporated in and constitute a part of this application, illustrate embodiments of the invention and, together with the description, serve to explain the invention and not to limit the invention. In the drawings:
FIG. 1 is a schematic diagram of injection mode operation of a conventional RBCC engine;
FIG. 2 is a schematic diagram of a prior art RBCC engine operating in a sub-combustion ramjet mode;
FIG. 3 is a schematic diagram of a scramjet mode operation of a conventional RBCC engine;
FIG. 4 is a schematic cross-sectional view of a combined cycle engine in accordance with a preferred embodiment of the present invention;
fig. 5 is a schematic diagram of the combined cycle engine propellant supply of fig. 4.
Description of the figures
10. a rotary detonation ramjet engine; 101. a drainage channel; 11. a housing; 1110. an outer spout; 112. a detonation chamber; 113. a first nozzle; 12. a central cone; 121. a rear body; 1210. an inner spout; 122. a precursor; 13. a connecting rib; 20. a rocket motor; 21. a combustion chamber; 22. a second nozzle; 30. an oxidant supply device; 40. a fuel supply device.
Detailed Description
The embodiments of the invention will be described in detail below with reference to the drawings, but the invention can be implemented in many different ways as defined and covered by the claims.
Referring to FIG. 4, a preferred embodiment of the present invention provides a combined cycle engine including a rotary detonation ramjet engine 10 and a rocket engine 20 disposed therein. The rotary detonation ramjet engine 10 comprises: the shell 11, shell 11 is the hollow tube-shape that both ends communicate. The detonation jet engine further comprises a central cone 12, a rear body 121 of the central cone 12 extends into the shell 11 from the air inlet end of the shell 11 and is connected with the shell 11, a drainage channel 101 for introducing air into the shell 11 is formed in a gap between the shell 11 and the rear body 121, a detonation chamber 112 communicated with the drainage channel 101 and a first jet pipe 113 communicated with the detonation chamber 112 are formed in a cavity between the rear end surface of the rear body 121 and the inner wall of the shell 11, and a jet port of the first jet pipe 113 is communicated with the atmosphere. The rocket engine 20 includes: a combustion chamber 21 arranged in the central cone 12, a second nozzle 22 communicated with the combustion chamber 21, the nozzle of the second nozzle 22 is communicated with the detonation chamber 112, and the combustion chamber 21 is communicated with an oxidant supply device 30 for supplying oxidant and a fuel supply device 40 for supplying fuel respectively. A plurality of outer nozzles 1110 are formed on the outer wall of the housing 11 at intervals along the circumferential direction, and two ends of the outer nozzles 1110 are respectively communicated with the fuel supply device 40 and the drainage channel 101, so that the fuel is sprayed into the drainage channel 101 through the plurality of outer nozzles 1110. And/or a plurality of inner nozzles 1210 which are sequentially arranged at intervals along the circumferential direction are processed on the outer wall of the rear body 121, and two ends of the inner nozzles 1210 are respectively communicated with the fuel supply device 40 and the drainage channel 101, so that the fuel is sprayed into the drainage channel 101 through the inner nozzles 1210.
the combined cycle engine of the present invention has two modes of operation: when the flying speed is lower than Ma2.5, the combined cycle engine works in an injection mode, the rocket engine 20 needs to be started at the moment, oxidant and fuel carried by the combined cycle engine respectively enter the combustion chamber 21 for isobaric combustion, and high-temperature and high-pressure combustion products are quickly discharged from the second spray pipe 22. Under the injection effect of the rocket engine 20, the incoming air enters the flow guide channel 101 after passing through the air inlet channel formed by the front body 122 of the central cone 12, and enters the detonation chamber 112 after being rapidly mixed with the fuel ejected by the outer nozzle 1110 and/or the inner nozzle 1210, the incoming air is combusted and released heat in a rotary detonation mode and is rapidly mixed with combustion products of the rocket engine 20, and the mixed combustion products are accelerated and discharged through the first nozzle 113, so that the thrust is generated. When the flight speed is greater than ma2.5, the combined cycle engine operates in the ramjet mode, with the rocket engine 20 turned off and the rotary detonation ramjet engine 10 operating alone. The incoming air enters the flow guide channel 101 after passing through the air inlet channel, and enters the detonation chamber 112 after being rapidly mixed with the fuel sprayed out through the outer nozzle 1110 and/or the inner nozzle 1210, the incoming air is combusted and released heat in the detonation chamber 112 in a rotary detonation mode, combustion products are discharged through the first spray pipe 113 in an accelerating mode, so that thrust is generated, the aircraft flies in an accelerating mode, the aircraft gradually accelerates until Ma6, and finally hypersonic cruise flight is achieved.
The combined cycle engine adopts the scheme of combining the conventional rocket engine 20 and the rotary knocking ramjet engine 10, wherein the rocket engine 20 is positioned in the central cone 12 of the air inlet channel of the rotary knocking ramjet engine 10, and compared with the conventional RBCC engine, the combined cycle engine has simpler and more compact integral structure. In the combined cycle engine, the rotary knocking ramjet engine 10 is organized to burn in a rotary knocking mode, and the conventional RBCC engine is organized to burn in an isobaric mode, so that compared with the conventional RBCC engine, the combined cycle engine has higher cycle thermal efficiency and better thrust performance, is more suitable for remote hypersonic cruise flight and has better economic performance; in addition, because the injection mode needs an injection mixing area, and the isobaric combustion heat release speed is low, the conventional RBCC needs a longer combustion chamber, so that the whole length of the engine is larger. In the invention, the detonation combustion has high heat release speed, so that the required length of the combustion chamber is short, and the whole length of the engine is shorter. The traditional RBCC engine has more working modes (including an injection mode, a sub-combustion stamping mode and a super-combustion stamping mode) and large mode conversion difficulty, and the engine has less working modes (including the injection mode and the stamping mode) and small mode conversion difficulty. In the prior art, the rotary detonation engine adopts the annular combustion chamber, and in the scheme of the application, the shell 11 is in a hollow cylinder shape with two communicated ends, and the central cone 12 only extends into the shell 11 through the rear body 121, so that the detonation chamber 112 is in a hollow cylinder shape, and compared with the annular combustion chamber, the cylindrical combustion chamber has stronger detonation combustion organization capacity, and the combustion stability of flame is good.
Optionally, as shown in fig. 4, a plurality of outer nozzles 1110 are formed on the outer wall of the housing 11, and each outer nozzle 1110 is 10mm to 20mm away from the outlet of the drainage channel 101. And/or a plurality of inner nozzles 1210 are processed on the outer wall of the rear body 121, and each inner nozzle 1210 is 10 mm-20 mm away from the outlet of the drainage channel 101. In the present invention, the gap between the outer shell 11 and the rear body 121 forms the flow guide channel 101, and the flow guide channel 101 is not only used for guiding air into the outer shell 11, but also used for isolating the influence of high back pressure in the detonation chamber 112 on the air inlet channel formed by the front body 122 of the central cone 12 in front of the flow guide channel 101, and also used for enabling the fuel sprayed by the outer nozzle 1110 and/or the inner nozzle 1210 to be quickly and fully mixed with the air and then sprayed into the detonation chamber 112. Therefore, the outer nozzle 1110 and/or the inner nozzle 1210 should be located at the upstream of the outlet of the flow guide channel 101, and the distance from the outer nozzle 1110 and/or the inner nozzle 1210 to the outlet of the flow guide channel 101 cannot be greater than 20mm, otherwise, the high back pressure in the detonation chamber 112 is easy to ignite the rotation detonation in the channel 101, namely, the phenomenon of 'backfire' occurs, so that the air inlet channel is affected; however, the distance from the outer nozzle 1110 and/or the inner nozzle 1210 to the outlet of the diversion channel 101 cannot be less than 10mm, otherwise the fuel and air cannot be sufficiently mixed before entering the detonation chamber 112, thereby affecting detonation generation in the detonation chamber 112. When the distance between each outer nozzle 1110 and/or each inner nozzle and the outlet of the flow guide channel 101 is 10 mm-20 mm, the fuel and the air can be fully mixed in the flow guide channel 101 and then injected into the detonation chamber 112, and the tempering phenomenon cannot occur.
In an embodiment of the present invention, as shown in fig. 4, a plurality of outer nozzles 1110 are formed on the outer wall of the outer casing 11, a plurality of inner nozzles 1210 are formed on the outer wall of the rear body 121, the plurality of outer nozzles 1110 and the plurality of inner nozzles 1210 are disposed in one-to-one correspondence, or the plurality of outer nozzles 1110 and the plurality of inner nozzles 1210 are disposed in a staggered manner. Because the outer wall of shell 11 is processed with a plurality of outer spouts 1110, and the outer wall of afterbody 121 is processed with a plurality of interior spouts 1210, and a plurality of outer spouts 1110 and a plurality of interior spout 1210 one-to-one set up, or a plurality of outer spouts 1110 and a plurality of interior spout 1210 staggered arrangement each other, thereby make air and fuel intensive mixing, both misce benes, be favorable to the emergence of detonation reaction in the detonation chamber 112, and when overcoming the cross-sectional height of drainage channel 101 big, when only setting up outer spouts 1110 or only setting up interior spout 1210, the problem of air can not with the fuel intensive mixing.
preferably, the first nozzle 113 and the outer wall of the detonation chamber 112 are each provided with a first wall cooling channel (not shown) connected to the fuel supply means 40. The plurality of outer nozzle holes 1110 are respectively communicated with the first wall surface cooling passage, so that the fuel after cooling the first nozzle pipe 113 and the detonation chamber 112 is injected into the flow guiding passage 101 by the plurality of outer nozzle holes 1110. The arrangement mode not only enables the structure of the engine to be simple and compact, but also fully utilizes the fuel, so that the fuel firstly serves as a coolant to cool the first spray pipe 113 and the detonation chamber 112, and then serves as the fuel to participate in the combustion of the detonation chamber.
Preferably, the second nozzle 22 and the outer wall of the combustion chamber 21 are provided with a second wall cooling channel (not shown) which is connected to the fuel supply means 40. The plurality of inner nozzles 1210 and the head of the combustion chamber 21 are respectively communicated with the second wall surface cooling channel, so that part of the fuel cooled by the second nozzle 22 and the combustion chamber 21 is injected into the flow guide channel 101 through the plurality of inner nozzles 1210, and the other part of the fuel is injected into the combustion chamber 21 through the head of the combustion chamber 21. The arrangement mode not only makes the structure of the engine simple and compact, but also fully utilizes the fuel, so that the fuel firstly serves as a coolant to cool the second spray pipe 22 and the combustion chamber 21, and then serves as the fuel to participate in rocket and detonation chamber combustion.
In actual operation, the engine is fuelled schematically as shown in figure 5. In the present invention, the propellant includes a liquid oxidizer and a liquid fuel carried by the engine itself, and the oxidizer is stored in the oxidizer-supplying device 30 and the fuel is stored in the fuel-supplying device 40. In operation, the oxidant is pumped by the oxidant supply 30 into the central cone 12 and through the supply channels in the central cone 12 and finally injected into the combustion chamber 21 from the head of the combustion chamber 21. The fuel is pumped out by the fuel supply device 40 and then supplied in two paths, wherein one path enters the first wall cooling channel to cool the first nozzle 113 and the detonation chamber 112, and absorbs heat in the first wall cooling channel to be heated, gasified or gasified and cracked, and when the outer nozzle 1110 is arranged on the shell 11, the fuel for cooling the first nozzle 113 and the detonation chamber 112 is sprayed into the drainage channel 101 by the outer nozzle 1110 to participate in combustion in the detonation chamber 112. And the other path of fuel enters the central cone 12 and flows into a second wall surface cooling channel from a supply flow channel in the central cone 12 to cool the second spray pipe 22 of the rocket engine and the wall surface of the combustion chamber 21, the heat absorption temperature rise gasification or gasification cracking is carried out in the channel, the heat absorption temperature rise fuel is sprayed into the combustion chamber 21 from the head part of the combustion chamber 21, or when the rear body 121 is provided with an inner nozzle 1210, the fuel in the second wall surface cooling channel is supplied in two paths, wherein one path of fuel enters the head part of the combustion chamber 21 to participate in combustion, and the other path of fuel is sprayed into the flow guide channel 101 through the inner nozzle 1210 to participate in the combustion in the detonation chamber 112.
preferably, as shown in fig. 4, the detonation chamber 112 is in the shape of a straight cylinder. Compared with the annular detonation chamber in the prior art, the cylindrical detonation chamber has stronger detonation combustion organization capacity, and the combustion stability of flame is good. Further, the length of the detonation chamber 112 is 400mm to 800 mm. When the length of the detonation chamber 112 is larger than 800mm, the whole length of the engine is lengthened, and when the length of the detonation chamber 112 is smaller than 400mm, the detonation combustion is insufficient, so that the thrust performance of the engine is influenced.
Alternatively, as shown in fig. 4, the rotary knocking ramjet 10 further includes a plurality of connecting ribs 13 for connecting the rear body 121 to the casing 11, the plurality of connecting ribs 13 are sequentially arranged at intervals in the circumferential direction of the flow guide passage 101, and each connecting rib 13 is connected to the inner wall of the casing 11 and the outer wall of the rear body 121. In the embodiment of the present invention, as shown in fig. 4, a plurality of connecting ribs 13 are sequentially arranged at intervals in the circumferential direction of the drainage channel 101. The upper end surface of each connecting rib 13 connected with the inner wall of the housing 11 is a curved surface matched with the inner wall of the housing 11, and the lower end surface of each connecting rib 13 connected with the outer wall of the rear body 121 is a curved surface matched with the outer wall of the rear body 121. When the upper end face of the connecting rib 13 is connected with the inner wall face of the housing 11 in a matching manner, and the lower end face of the connecting rib is connected with the outer wall face of the rear body 121 in a matching manner, the central cone 12 is stably connected with the housing 11, and the overall structural strength of the engine is strong.
Alternatively, as shown in fig. 4, the central cone 12 is arranged coaxially with the housing 11, and the rear end face of the rear body 121 is flush with the outlet of the drainage channel 101.
Optionally, as shown in fig. 4, the central cone 12 further includes a front body 122 connected to the rear body 121 along the axial direction, the front body 122 is located outside the casing 11, and the outer wall surface of the front body 122 forms an inlet for compressing air. The combustion chamber 21 is located within the precursor 122. The second nozzle 22 is connected to the outlet of the combustion chamber 21, and the second nozzle 22 extends in the axial direction of the central cone 12 to communicate with the rear end face of the rear body 121.
Alternatively, as shown in fig. 4, an oxidant flow passage (not shown) for guiding the oxidant and a fuel flow passage (not shown) for guiding the fuel are provided in the central cone 12. The inlet of the oxidant flow passage is connected to the oxidant supply device 30, the outlet of the oxidant flow passage is connected to the head of the combustion chamber 21, and the oxidant pumped out by the oxidant supply device 30 passes through the oxidant flow passage and is finally injected into the combustion chamber from the head of the combustion chamber 21. The inlet of the fuel flow path is connected to the fuel supply device 40, and the outlet of the fuel flow path is connected to the second wall surface cooling passage. Part of the fuel pumped out by the fuel supply device 40 flows into the first wall cooling channel cooling the first nozzle 113 and the detonation chamber 112, and the other part of the fuel enters the second wall cooling channel cooling the second nozzle 22 and the combustion chamber 21 through the fuel flow passage.
In the embodiment of the present invention, the first nozzle 113 is an expanding nozzle, and the second nozzle 22 is a laval nozzle, which is used to increase the flow velocity of the high-temperature and high-pressure combustion products, so that the combustion products are rapidly discharged to generate thrust.
according to another aspect of the invention, there is also provided a hypersonic aircraft comprising the combined cycle engine of the above embodiment. Experiments prove that the hypersonic aircraft has the advantages of small working mode, small mode conversion difficulty, high thermal cycle efficiency, good thrust performance and short length of a combustion chamber.
The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. a combined cycle engine comprising a rotary detonation ramjet engine (10) and a rocket engine (20) disposed therein;
The rotary detonation ramjet engine (10) comprises:
the shell (11), the said shell (11) is the hollow tube-shape that both ends communicate;
the rear body (121) of the central cone (12) extends into the shell (11) from the air inlet end of the shell (11) and is connected with the shell (11), a drainage channel (101) for introducing air into the shell (11) is formed in a gap between the shell (11) and the rear body (121), a detonation chamber (112) communicated with the drainage channel (101) and a first spray pipe (113) communicated with the detonation chamber (112) are formed in a cavity between the rear end face of the rear body (121) and the inner wall of the shell (11), and a spray opening of the first spray pipe (113) is communicated with the atmosphere;
The rocket engine (20) comprises:
A combustion chamber (21) arranged in the central cone (12), a second nozzle (22) communicating with the combustion chamber (21), the nozzle of the second nozzle (22) communicating with the detonation chamber (112), the combustion chamber (21) communicating with an oxidant supply means (30) for supplying an oxidant and a fuel supply means (40) for supplying a fuel, respectively;
A plurality of outer nozzles (1110) which are sequentially arranged at intervals along the circumferential direction of the outer wall of the shell (11) are machined on the outer wall of the shell, and two ends of each outer nozzle (1110) are respectively communicated with the fuel supply device (40) and the drainage channel (101), so that fuel is sprayed into the drainage channel (101) through the outer nozzles (1110); and/or
The outer wall of the rear body (121) is provided with a plurality of inner nozzles (1210) which are sequentially arranged at intervals along the circumferential direction, and two ends of each inner nozzle (1210) are respectively communicated with the fuel supply device (40) and the drainage channel (101) so that fuel is sprayed into the drainage channel (101) through the inner nozzles (1210).
2. The combined-cycle engine of claim 1,
Each outer nozzle (1110) is 10-20 mm away from the outlet of the drainage channel (101); and/or
Each inner nozzle (1210) is 10-20 mm away from the outlet of the drainage channel (101).
3. the combined-cycle engine of claim 2,
The plurality of outer nozzles (1110) and the plurality of inner nozzles (1210) are arranged in a one-to-one correspondence manner, or the plurality of outer nozzles (1110) and the plurality of inner nozzles (1210) are arranged in a staggered manner.
4. The combined-cycle engine of claim 3,
The outer walls of the first nozzle (113) and the detonation chamber (112) are provided with first wall surface cooling channels, and the first wall surface cooling channels are connected with the fuel supply device (40);
The outer nozzles (1110) are respectively communicated with the first wall surface cooling channel, so that the fuel which is used for cooling the first spray pipe (113) and the detonation chamber (112) is sprayed into the drainage channel (101) through the outer nozzles (1110).
5. The combined-cycle engine of claim 3,
The outer walls of the second nozzle (22) and the combustion chamber (21) are provided with second wall surface cooling channels, and the second wall surface cooling channels are connected with the fuel supply device (40);
The inner nozzles (1210) and the head of the combustion chamber (21) are respectively communicated with the second wall surface cooling channel, so that part of the fuel for cooling the second nozzle (22) and the combustion chamber (21) is injected into the flow guide channel (101) by the inner nozzles (1210), and the other part of the fuel is injected into the combustion chamber (21) by the head of the combustion chamber (21).
6. The combined-cycle engine according to any one of claims 1 to 5,
The detonation chamber (112) is in a straight cylinder shape;
the length of the detonation chamber (112) is 400-800 mm.
7. the combined-cycle engine according to any one of claims 1 to 5,
rotatory detonation ramjet (10) still including being used for making afterbody (121) with many connection rib (13) that shell (11) link to each other, many connection rib (13) are followed the circumference interval arrangement in proper order of diversion passageway (101), and each the up end of connecting rib (13) with the inner wall of shell (11) links to each other, each the lower terminal surface of connecting rib (13) with the outer wall of afterbody (121) links to each other.
8. The combined-cycle engine according to any one of claims 1 to 5,
The central cone (12) further comprises a front body (122) axially connected with the rear body (121), the front body (122) is positioned outside the shell (11), and the outer wall surface of the front body (122) forms an air inlet channel for compressing air;
the combustion chamber (21) is located within the precursor (122);
the second spray pipe (22) is connected with an outlet of the combustion chamber (21), and the second spray pipe (22) extends to be communicated with the rear end face of the rear body (121) along the axial direction of the central cone (12).
9. the combined-cycle engine of claim 5,
An oxidant flow channel for guiding oxidant and a fuel flow channel for guiding fuel are arranged in the central cone (12);
The inlet of the oxidant flow channel is connected with the oxidant supply device (30), and the outlet of the oxidant flow channel is connected with the head of the combustion chamber (21);
The inlet of the fuel flow passage is connected to the fuel supply device (40), and the outlet of the fuel flow passage is connected to the second wall surface cooling passage.
10. A hypersonic aircraft comprising a combined cycle engine according to any one of claims 1 to 9.
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| CN112627983B (en) * | 2020-12-25 | 2022-02-22 | 中国人民解放军国防科技大学 | RBCC engine inner flow channel and RBCC engine |
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| CN113123899B (en) * | 2021-04-19 | 2022-07-15 | 中国人民解放军国防科技大学 | Central injection rocket, RBCC engine and fuel injection method thereof |
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