CN112682218B - Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping - Google Patents
Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping Download PDFInfo
- Publication number
- CN112682218B CN112682218B CN202011573592.8A CN202011573592A CN112682218B CN 112682218 B CN112682218 B CN 112682218B CN 202011573592 A CN202011573592 A CN 202011573592A CN 112682218 B CN112682218 B CN 112682218B
- Authority
- CN
- China
- Prior art keywords
- rocket
- annular
- section
- channel
- central body
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Active
Links
Images
Abstract
A wide-speed-range engine based on a mixed-section confluence rocket of an annular supercharging central body comprises a central body, an air inlet channel, a concave cavity combustion chamber and a spray pipe. The embedded rocket in the conventional rocket ramjet is arranged in front to the outer side of the central body, the mixing pressurization of the rocket and air is carried out in the annular mixing channel, and then the mixture flows are merged into the central pipe flow. Because the contact area of the rocket jet and the air is increased, the mixing distance of the two airflows can be effectively shortened. By adopting the scheme, the integral thrust performance of the rocket ramjet can be effectively improved.
Description
Technical Field
The invention relates to the technical field of engines, in particular to a rocket stamping wide-speed-range engine capable of reliably working under a wide flight Mach number.
Background
The speed range of the engine is an important design index of the engine, the wide speed range means that the application range of the engine is wider, the larger the flight envelope range is, the more practical the engine is, and the wider the application prospect is. The conventional ramjet engine is only dependent on the air stamping effect, so that the engine stamping effect is insufficient and the thrust performance of the engine is low when the flying speed is low. The conventional rocket ramjet needs a longer mixing section, so the weight of the engine is larger, the thrust-weight ratio of the engine is low, and the overall performance of the engine is greatly influenced by the overlong length of the mixing section at higher Mach number.
Disclosure of Invention
Aiming at the defects in the prior art, the invention provides a wide-speed-range engine based on the combined rocket of the annular supercharging central body mixing section.
In order to achieve the technical purpose, the invention adopts the technical scheme that:
the engine comprises a central body, an air inlet passage, a concave cavity combustion chamber and a spray pipe. The central body leading edge and the inlet lip form an annular inlet. The air inlet channel is an air flow channel formed by the inner wall of the air inlet channel and the central body, and is sequentially provided with an annular isolation section channel, an annular mixing channel, an annular contraction channel and an annular diffusion channel. The middle section of the central body is provided with a rocket device, rocket nozzles of the rocket device are annularly distributed on the central body, the ejection direction of the rocket nozzles is the same as the air transmission direction in the air inlet channel, and rocket fuel gas ejected by the rocket nozzles directly enters the annular mixing channel; air captured from the annular air inlet passes through the annular isolation section channel and then is fully mixed and accelerated with rocket fuel gas ejected from the rocket nozzle in the annular mixing channel, the mixture is accelerated into supersonic mixed airflow to enter the annular contraction channel, and annular tube flow is converted into central tube flow to enter the annular diffusion channel; the tail end of the annular diffusion channel is connected with a concave cavity in the concave cavity combustion chamber, the supersonic mixed airflow is converted into subsonic velocity through the annular diffusion channel and enters the concave cavity combustion chamber, and fuel is sprayed into the concave cavity combustion chamber through the fuel nozzle and is combusted with the mixed airflow entering the concave cavity combustion chamber through the annular diffusion channel; and the rear part of the concave cavity combustion chamber is connected with a spray pipe, and the mixed gas flow is discharged through the spray pipe after being combusted. The engine can utilize the rocket to pressurize air in the annular mixing section at low flight Mach number, and then confluence is central pipe flow.
As a further improvement of the invention, the rear section of the central body is set as a central body convergence section, the cross-sectional area of the central body convergence section, namely the rear section of the central body, is continuously reduced, and an annular contraction channel is formed between the inner wall of the air inlet channel and the central body convergence section. Preferably, the convergent section of the central body is in the form of a cone.
As a further improvement, the fuel is fuel oil, the fuel nozzle is a fuel oil nozzle, and the fuel oil injection pressure of the fuel oil nozzle is adjustable. The fuel nozzles are annularly distributed on the inner wall of the annular diffusion channel close to the concave cavity combustion chamber. Furthermore, a plurality of ring fuel nozzles are arranged on the inner wall of the annular diffusion channel close to the concave cavity combustion chamber at intervals. According to different flight states, the fuel injection pressure of the fuel nozzle can be adjusted, and fuel injection rings at different positions can be selected for fuel injection.
As a further improvement of the invention, when the annular contraction channel changes the gas flow channel from an annular shape to a central pipe flow, the cross-sectional area of the inlet cross section of the annular contraction channel is A2, the cross-sectional area of the outlet cross section of the annular contraction channel is a confluence cross section, the cross-sectional area of the confluence cross section is A22, and A22 is 1-10% larger than A2, namely 1.01A2 is more than or equal to A22 and is less than or equal to 1.1A 2. The equivalent circle diameter of the inlet section area A2 of the annular contraction channel is D2, the axial length of the annular contraction channel is L12, and the length range is D2-L12-4D 2.
As a further improvement of the invention, the annular diffusion channel is an expansion channel with an equivalent expansion angle, the confluence section and the inlet section of the concave cavity combustion chamber are respectively the inlet and outlet sections of the annular diffusion channel, the cross-sectional areas of the confluence section and the inlet section of the concave cavity combustion chamber are respectively A22 and A3, and 1.05A 22-A3-2A 22. The axial length of the annular diffusion passage is L13, and L13 controls the equivalent expansion angle of the confluence section and the inlet section of the concave cavity combustion chamber, and the equivalent expansion angle of the confluence section and the inlet section of the concave cavity combustion chamber ranges from 1 degree to 8 degrees.
As a further improvement of the invention, the annular mixing channel is an equal-area annular channel or a slightly-contracted annular channel.
As a further improvement of the invention, the cross section area of the air flow channel at the lip of the air inlet channel is A1, the area of the outlet of the rocket nozzle is A0, the cross section of the outlet of the annular mixing channel is the cross section of the inlet of the annular contraction channel, the area of the cross section is A2, and the size of the cross section is A1-A2-A1 + A0. The axial length of the annular mixing channel is L11 and the range is D2 ≦ L11 ≦ 5D 2.
As a further improvement of the invention, the rocket device comprises a rocket oxidant supply source, a rocket oxidant pipeline, a rocket reductant supply source and a rocket, wherein the rocket oxidant supply source, the rocket oxidant pipeline, the rocket reductant pipeline and the rocket reductant supply source are all arranged in a central body, the rocket oxidant supply source is communicated into the rocket through a rocket oxidant pipe, the rocket reductant supply source is communicated into the rocket through a rocket reductant pipeline, and rocket fuel gas generated after the oxidant and the reductant are subjected to combustion reaction in the rocket is sprayed out through a rocket nozzle and enters an annular mixing channel.
The range of the rocket outlet Mach number is Mach 2 to Mach 3, when the total pressure of the air inlet of the engine is higher, the rocket outlet Mach number can be selected to be a smaller value, and when the total pressure of the air inlet of the engine is lower, the rocket outlet Mach number can be selected to be a larger value.
As a further improvement of the invention, the rocket in the rocket device can be a rocket with an annular rocket nozzle, or a rocket assembly formed by annularly arranging a plurality of small rockets. The rocket nozzles of the small rockets are distributed on the central body in an annular shape, and the rocket nozzle of each small rocket is responsible for a sector annular seam.
As a further improvement of the invention, the combustion chamber of the concave cavity combustion chamber is in a concave cavity layout, the flame is stabilized in a leeward area of the concave cavity, the cross-sectional area of the inlet of the concave cavity combustion chamber is A3, the cross-sectional area of the outlet of the concave cavity combustion chamber is A4, and the cross-sectional area of 1.1A3 is not less than A4 and not more than 2.5A 3.
The axial length of the concave cavity combustion chamber is L2, and when the concave cavity combustion chamber is designed, the axial length of the concave cavity combustion chamber is reasonably designed, so that the combustion effect can be improved. Preferably, the axial length of the concave cavity combustion chamber is 0.2D 2-2D 2-2.
As a further improvement of the invention, the spray pipe consists of an equal straight section and an expansion section, the inlet section is the equal straight section, the outlet section is the expansion section, the outlet area of the spray pipe is A5, and the size of the spray pipe is 1.5A 4-A5-4A 4; the axial length of the spray pipe is L3, and L3 is more than or equal to 3D2 and less than or equal to 6D 2.
The section of the airflow channel from the rocket outlet section (namely the outlet section of the rocket nozzle) to the concave cavity combustion chamber inlet section is a pressurizing section, the axial length of the pressurizing section is L1, and L1 is L11+ L12+ L13. The axial length of the pressurizing section determines the thrust performance such as the thrust-weight ratio of the rocket ramjet, and when the length of L1 is less than 4D2, the overall performance of the rocket ramjet is better.
As a further improvement of the invention, the engine can be an axisymmetrical layout engine or a two-dimensional layout engine.
The invention provides an aircraft, which is provided with any one of the confluence rocket ramjet based on the annular supercharging central body mixing section.
Compared with the prior art, the invention can obtain the following technical effects:
(1) the embedded rocket in the conventional rocket ramjet is arranged to the outer side of the central body in a front-mounted mode, the rocket is mixed with air in the annular mixing channel, and the mixing distance of two airflows can be effectively shortened due to the fact that the contact area of the rocket jet and the air is increased. By adopting the scheme, the integral thrust performance of the rocket ramjet can be effectively improved.
(2) The annular mixing channel is adopted, so that the contact area of the rocket fuel gas and air is increased, the mixing efficiency is enhanced, the length scale of the supercharging section of the rocket ramjet is greatly reduced, and the thrust-weight ratio performance of the engine is improved.
(3) When the engine works under a high flight Mach number, the rocket can be closed, and the engine can realize stamping work under the high Mach number by means of the stamping action of the engine. The annular contraction channel has small area expansion, which not only meets the requirement of high pressure ratio in a high Mach number state, but also meets the requirement of air flow mixing in a low Mach number state. Therefore, the rocket ramjet can share one inner runner to realize mode conversion at a wider Mach number.
(4) The engine has a wider low flight Mach number range than conventional ramjet engines and bimodal ramjet engines; compared with the conventional rocket ramjet engine, the structure is more compact, and the overall thrust performance of the engine is better; compared with a pure rocket engine, the rocket engine has the advantages of lower fuel consumption and high specific impulse performance.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the structures shown in the drawings without creative efforts.
The invention is further described below with reference to the accompanying drawings:
fig. 1 is a schematic structural diagram according to an embodiment of the present invention.
In the figure:
1. a central body; 2. an inlet lip; 3. an annular isolation section channel; 4. an annular mixing channel; 5. an annular mixing channel outlet cross-section; 6. an annular constricting channel; 7. a flow-merging section; 8. an annular diffuser channel; 9. a fuel nozzle; 10. the section of the inlet of the concave cavity combustion chamber; 11. a concave cavity combustion chamber; 12. the outlet section of the combustion chamber of the concave cavity combustion chamber; 13. a nozzle; 14. a rocket oxidant supply; 15. a rocket oxidizer line; 16. a rocket reductant line; 17. a rocket reductant supply; 18. a rocket nozzle; 19. a concave cavity; 20. a centerbody convergence section; 21. an equal straight section; 22. and (4) an expansion section.
Detailed description of the preferred embodiments
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that all the directional indicators (such as up, down, left, right, front, and rear … …) in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture (as shown in the drawing), and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the descriptions related to "first", "second", etc. in the present invention are only for descriptive purposes and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one such feature. In the description of the present invention, "a plurality" means at least two, e.g., two, three, etc., unless specifically limited otherwise.
In the present invention, unless otherwise expressly stated or limited, the terms "connected," "secured," and the like are to be construed broadly, and for example, "secured" may be a fixed connection, a removable connection, or an integral part; the connection can be mechanical connection, electrical connection, physical connection or wireless communication connection; they may be directly connected or indirectly connected through intervening media, or they may be connected internally or in any other suitable relationship, unless expressly stated otherwise. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
Referring to fig. 1, an embodiment of the present invention provides a hybrid merging rocket ramjet engine based on an annular supercharged central body, which comprises a central body 1, an air intake, a concave chamber combustor 11 and a nozzle 13. The central body leading edge and the inlet lip 2 form an annular inlet. The air inlet channel is an air flow channel formed by the inner wall of the air inlet channel and the central body 1, and the air inlet channel is sequentially provided with an annular isolation section channel 3, an annular mixing channel 4, an annular contraction channel 6 and an annular diffusion channel 8.
The middle section of the central body is provided with a rocket device, rocket nozzles 18 of the rocket device are annularly distributed on the central body 1, the ejection direction of the rocket nozzles 18 is the same as the air transmission direction in the air inlet channel, and rocket fuel gas ejected by the rocket nozzles 18 directly enters the annular mixing channel 4.
Referring to fig. 1, the rocket device comprises a rocket oxidant supply source 14, a rocket oxidant pipeline 15, a rocket reductant pipeline 16, a rocket reductant supply source 17 and a rocket, wherein the rocket oxidant supply source 14, the rocket oxidant pipeline 15, the rocket reductant pipeline 16 and the rocket reductant supply source 17 are all arranged inside a central body 1, the rocket oxidant supply source 14 is communicated with the rocket through the rocket oxidant pipeline 15, the rocket reductant supply source 17 is communicated with the rocket through the rocket reductant pipeline 16, and rocket fuel gas generated after combustion reaction of oxidant and reductant inside the rocket is ejected through a rocket nozzle 18 and enters an annular mixing channel 4.
The range of the rocket outlet Mach number is Mach 2 to Mach 3, when the total pressure of the air inlet of the engine is higher, the rocket outlet Mach number can be selected to be a smaller value, and when the total pressure of the air inlet of the engine is lower, the rocket outlet Mach number can be selected to be a larger value.
Air captured from the annular air inlet passes through the annular isolation section channel 3 and then is fully mixed and accelerated with rocket fuel gas ejected from the rocket nozzle 18 in the annular mixing channel 4, the mixture is accelerated into supersonic mixed gas flow to enter the annular contraction channel 6, and annular tube flow is converted into central tube flow to enter the annular diffusion channel 8; the tail end of the annular diffusion channel 8 is connected with a concave cavity 19 in the concave cavity combustion chamber 11, the supersonic mixed gas flow is converted from supersonic speed to subsonic speed after passing through the annular diffusion channel 8 and enters the concave cavity combustion chamber 11, and fuel is sprayed into the concave cavity combustion chamber 11 through the fuel nozzle 9 and is combusted with the mixed gas flow entering the concave cavity combustion chamber 11 through the annular diffusion channel 8; the concave cavity combustion chamber 11 is connected with a spray pipe 13, and the mixed gas flow is discharged through the spray pipe 13 after being combusted. The engine can utilize the rocket to pressurize air in the annular mixing section at low flight Mach number, and then confluence is central pipe flow.
Referring to fig. 1, the rear section of the center body in one embodiment of the present invention is provided with a center body converging section 20, and the cross-sectional area of the center body converging section 20, i.e., the rear section of the center body, is continuously reduced. An annular converging passage is formed between the inner wall of the inlet and the converging section 20 of the centerbody. Preferably, the central body converging section 20 may converge into a cone.
The fuel is fuel oil, the fuel nozzle 9 is a fuel oil nozzle, and the fuel oil injection pressure of the fuel oil nozzle is adjustable.
In one embodiment of the invention, the fuel nozzles are annularly distributed on the inner wall of the annular diffusion channel close to the concave cavity combustion chamber.
In one embodiment of the invention, multiple ring fuel nozzles are spaced on the inner wall of the annular diffuser passage 8 adjacent to the concave combustion chamber 11. According to different flight states, the fuel injection pressure of the fuel nozzle can be adjusted, and fuel injection rings at different positions can be selected for fuel injection.
As a further improvement of the invention, when the annular contraction channel 6 changes the gas flow channel from the annular shape to the central pipe flow, the cross-sectional area of the inlet cross section of the annular contraction channel is A2, the cross-sectional area of the outlet cross section of the annular contraction channel is the confluence cross section 7, the cross-sectional area of the confluence cross section 7 is A22, and A22 is 1-10% larger than A2, namely 1.01A2 is equal to or less than A22 and equal to or less than 1.1A 2. The equivalent circle diameter of the inlet section area A2 of the annular contraction channel is D2, the axial length of the annular contraction channel 6 is L12, and the length range is D2-L12-4D 2.
The annular diffusion channel 8 is an expansion channel with an equivalent expansion angle, the confluence section 7 and the concave cavity combustion chamber inlet section 10 are respectively the inlet and outlet sections of the annular diffusion channel 8, the cross-sectional areas of the confluence section 7 and the concave cavity combustion chamber inlet section 10 are respectively A22 and A3, and 1.05A22 is not less than A3 and not more than 2A 22. The axial length of the annular diffuser passage 8 is L13, and L13 controls the equivalent divergent angle of the converging section 7 and the concave combustion chamber inlet section 10, and the equivalent divergent angle of the converging section 7 and the concave combustion chamber inlet section 10 ranges from 1 ° to 8 °.
The annular mixing channel 4 is an equal-area annular channel or a slightly constricted annular channel. The cross section area of the airflow channel at the inlet lip 2 is A1, the area of the rocket nozzle outlet is A0, the outlet cross section 5 of the annular mixing channel is the inlet cross section of the annular contraction channel, the area of the annular contraction channel is A2, and the size of the annular contraction channel is A1-A2-A1 + A0. The axial length of the annular mixing channel 4 is L11 and ranges from D2 ≦ L11 ≦ 5D 2.
In one embodiment of the invention, the rocket in the rocket device is a rocket with a ring-shaped rocket nozzle.
In an embodiment of the invention, the rocket in the rocket device is a rocket assembly formed by annularly arranging a plurality of small rockets. The rocket nozzles of the small rockets are distributed on the central body 1 in an annular shape, and the rocket nozzle of each small rocket is responsible for a sector annular seam.
In one embodiment of the invention, the combustion chamber of the concave cavity combustion chamber is in a concave cavity layout, flame is stabilized in a leeward area of the concave cavity, the area of an inlet section 10 of the concave cavity combustion chamber is A3, the area of an outlet section 12 of the concave cavity combustion chamber is A4, and the ratio of 1.1A3 to A4 is less than or equal to 2.5A 3.
The axial length of the concave cavity combustion chamber 11 is L2, and when the concave cavity combustion chamber 11 is designed, the axial length of the concave cavity combustion chamber is reasonably designed, so that the combustion effect can be improved. In one embodiment of the present invention, the axial length of the concave chamber combustion chamber 11 is 0.2D2 ≤ L2 ≤ 2D 2.
In one embodiment of the invention, the nozzle 13 is composed of an equal straight section 21 and an expansion section 22, the inlet section is the equal straight section 21, the outlet section is the expansion section 22, the nozzle outlet area is A5, and the size is 1.5A 4-A5-4A 4; the axial length of the spray pipe 13 is L3, and L3 is more than or equal to 3D2 and less than or equal to 6D 2.
The section of the air flow passage from the rocket outlet section (i.e. the outlet section of the rocket nozzle) to the concave cavity combustion chamber inlet section 10 is a pressurizing section, the axial length of the pressurizing section is L1, and L1 is L11+ L12+ L13. The axial length of the pressurizing section determines the thrust performance such as the thrust-weight ratio of the rocket ramjet, and when the length of L1 is less than 4D2, the overall performance of the rocket ramjet is better.
In one embodiment of the invention, a hybrid merging rocket ramjet engine based on an annular supercharging central body is provided, and the engine comprises a central body, an air inlet passage, a concave cavity combustion chamber and a nozzle. When the engine is in an axisymmetrical configuration, the three-dimensional structure of the engine takes the central line of fig. 1 as a central axis, and the engine configuration can be obtained by rotating for 360 degrees. In combination with the engine trajectory, if the total pressure at the inlet of the engine is around 200kPa on average during the low-Mach-number flight (Mach number is less than 2), the Mach number at the outlet of the rocket can be Mach 2, L11 is more than or equal to D2 and less than or equal to 2D2, and L12 is more than or equal to D2 and less than or equal to 2D 2. The rocket is composed of 6 small annular rockets arranged along the circumference of the outer side wall of the central body 1, and each rocket is responsible for a 60-degree annular sector.
In one embodiment of the invention, a hybrid merging rocket ramjet engine based on an annular supercharging central body is provided, and the engine comprises a central body, an air inlet passage, a concave cavity combustion chamber and a nozzle. When the engine is in a two-dimensional configuration, the three-dimensional structure of the engine is stretched in two dimensions according to the diagram of fig. 1, and the engine configuration can be obtained. In combination with the engine trajectory, if the total pressure at the inlet of the engine is around 100kPa on average during the low-Mach-number flight (Mach number is less than 2), the Mach number at the outlet of the rocket can be Mach 3, L11 is more than or equal to 1.5D2 and is more than or equal to 3D2, and L12 is more than or equal to 1.5D2 and is less than or equal to 3D 2.
In one embodiment of the invention, an aircraft is provided with any one of the above-mentioned merging rocket ramjet engines based on an annular supercharged central body mixing section.
In summary, although the present invention has been described with reference to the preferred embodiments, it should be understood that various changes and modifications can be made by those skilled in the art without departing from the spirit and scope of the invention.
Claims (22)
1. The utility model provides a based on wide fast territory engine of annular pressure boost central part mixing section confluence rocket punching press, this engine includes central body, intake duct, concave cavity combustion chamber and spray tube, its characterized in that: the front edge of the central body and the lip of the air inlet form an annular air inlet, the air inlet is an airflow channel formed by the inner wall of the air inlet and the central body, and the air inlet is sequentially provided with an annular isolation section channel, an annular mixing channel, an annular contraction channel and an annular diffusion channel; the middle section of the central body is provided with a rocket device, rocket nozzles of the rocket device are annularly distributed on the central body, the ejection direction of the rocket nozzles is the same as the air transmission direction in the air inlet channel, and rocket fuel gas ejected by the rocket nozzles directly enters the annular mixing channel; air captured from the annular air inlet passes through the annular isolation section channel and then is fully mixed and accelerated with rocket fuel gas ejected from the rocket nozzle in the annular mixing channel, the mixture is accelerated into supersonic mixed airflow to enter the annular contraction channel, and annular tube flow is converted into central tube flow to enter the annular diffusion channel; the tail end of the annular diffusion channel is connected with a concave cavity in the concave cavity combustion chamber, the supersonic mixed airflow is converted into subsonic velocity through the annular diffusion channel and enters the concave cavity combustion chamber, and fuel is sprayed into the concave cavity combustion chamber through the fuel nozzle and is combusted with the mixed airflow entering the concave cavity combustion chamber through the annular diffusion channel; the rear part of the concave cavity combustion chamber is connected with a spray pipe, and mixed gas flow is discharged through the spray pipe after being combusted; the cross section area of the inlet section of the annular contraction channel is A2, the equivalent circle diameter of the cross section area A2 of the inlet section of the annular contraction channel is D2, the axial length of the annular mixing channel is L11, and the range of the axial length is D2-L11-5D 2.
2. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 1, wherein: the central body rear section is set as the central body convergent section, the cross-sectional area of the central body convergent section, namely the central body rear section, is continuously reduced, and an annular convergent channel is formed between the inner wall of the air inlet channel and the central body convergent section.
3. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 2, wherein: the convergent section of the central body is in the form of a cone.
4. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 2, wherein: the fuel is fuel oil, the fuel nozzle is a fuel oil nozzle, and the fuel oil injection pressure of the fuel oil nozzle is adjustable.
5. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 4, wherein: the fuel nozzles are annularly distributed on the inner wall of the annular diffusion channel close to the concave cavity combustion chamber.
6. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 5, wherein: and multiple ring fuel nozzles are arranged on the inner wall of the annular diffusion channel close to the concave cavity combustion chamber at intervals, and fuel nozzle injection rings at different positions are selected for fuel injection according to different flight states.
7. The wide-speed range engine based on annular supercharging central body mixing section confluence rocket stamping according to any one of claims 1 to 6, wherein: the section of the outlet of the annular contraction channel is a confluence section, the section area of the confluence section is A22, and A22 is 1-10% larger than A2, namely 1.01A2 is more than or equal to A22 is less than or equal to 1.1A 2.
8. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 7, wherein: the axial length of the annular contraction channel is L12, and the length range is D2 ≤ L12 ≤ 4D 2.
9. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 8, wherein: the annular diffusion channel is an expansion channel with an equivalent expansion angle, the confluence section and the inlet section of the concave cavity combustion chamber are respectively the inlet and outlet sections of the annular diffusion channel, the cross-sectional areas of the confluence section and the inlet section of the concave cavity combustion chamber are respectively A22 and A3, and 1.05A22 is not less than A3 and not more than 2A 22.
10. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 9, wherein: the axial length of the annular diffusion passage is L13, and the equivalent expansion angle of the confluence section and the inlet section of the concave cavity combustion chamber ranges from 1 degree to 8 degrees.
11. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 10, wherein: the annular mixing channel is an equal-area annular channel or a slightly-contracted annular channel.
12. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 11, wherein: the cross section area of the air flow channel at the lip of the air inlet channel is A1, the area of the outlet of the rocket nozzle is A0, the cross section of the outlet of the annular mixing channel is the inlet cross section of the annular contraction channel, the area of the annular contraction channel is A2, and the size of the annular contraction channel is A1-A2-A1 + A0.
13. The wide-speed range engine based on annular supercharging central body mixing section confluence rocket stamping according to any one of claims 1 to 6, wherein: the rocket device comprises a rocket oxidant supply source, a rocket oxidant pipeline, a rocket reductant supply source and a rocket, wherein the rocket oxidant supply source, the rocket oxidant pipeline, the rocket reductant pipeline and the rocket reductant supply source are all arranged in a central body, the rocket oxidant supply source is communicated into the rocket through a rocket oxidant pipe, the rocket reductant supply source is communicated into the rocket through a rocket reductant pipeline, and rocket fuel gas generated after the oxidant and the reductant undergo a combustion reaction in the rocket is ejected through a rocket nozzle and enters an annular mixing channel.
14. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 13, wherein: the rocket exit mach number ranges from mach 2 to mach 3.
15. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 13, wherein: the rocket in the rocket unit is a rocket with a ring-shaped rocket nozzle.
16. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 13, wherein: the rocket in the rocket device is a rocket combination formed by annularly arranging a plurality of small rockets.
17. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 12, wherein: the cross-sectional area of the inlet of the concave cavity combustion chamber is A3, the cross-sectional area of the outlet of the concave cavity combustion chamber is A4, and A4 is more than or equal to 1.1A3 and less than or equal to 2.5A 3.
18. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 17, wherein: the axial length of the concave cavity combustion chamber is L2, and the value of the axial length of the concave cavity combustion chamber is 0.2D 2-L2-2D 2.
19. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 18, wherein: the spray pipe consists of an equal straight section and an expansion section, the inlet section is the equal straight section, the outlet section is the expansion section, the outlet area of the spray pipe is A5, and the size of the spray pipe is 1.5A 4-A5-4A 4; the axial length of the spray pipe is L3, and L3 is more than or equal to 3D2 and less than or equal to 6D 2.
20. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 17, wherein: the section of the airflow channel from the rocket outlet section to the concave cavity combustion chamber inlet section is a pressurizing section, the axial length of the pressurizing section is L1, L1 is L11+ L12+ L13, and the length of L1 is less than 4D 2.
21. The wide speed range engine based on annular supercharged central body mixing section confluence rocket ram of claim 1, wherein: the engine is an axisymmetrical layout engine or a two-dimensional layout engine.
22. An aircraft comprising the wide speed range engine based on annular supercharged central body hybrid section confluence rocket ram as claimed in claim 1.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011573592.8A CN112682218B (en) | 2020-12-24 | 2020-12-24 | Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011573592.8A CN112682218B (en) | 2020-12-24 | 2020-12-24 | Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN112682218A CN112682218A (en) | 2021-04-20 |
| CN112682218B true CN112682218B (en) | 2022-03-29 |
Family
ID=75452146
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011573592.8A Active CN112682218B (en) | 2020-12-24 | 2020-12-24 | Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112682218B (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114109650B (en) * | 2021-10-27 | 2023-02-28 | 厦门大学 | Integral liquid rocket punching combined power device |
Family Cites Families (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2656382B1 (en) * | 1989-12-21 | 1994-07-08 | Europ Propulsion | HIGH ADAPTABILITY COMBINED PROPULSION ENGINE FOR AIRCRAFT OR SPACE AIRCRAFT. |
| US8256203B1 (en) * | 2007-01-26 | 2012-09-04 | The University Of Alabama In Huntsville | Rocket based combined cycle propulsion unit having external rocket thrusters |
| KR101126861B1 (en) * | 2009-12-23 | 2012-03-23 | 한국항공우주연구원 | Combined cycle engine for hypersonic air-breathing and it's engine mode |
| CN102374071A (en) * | 2011-09-15 | 2012-03-14 | 西北工业大学 | Center rocket type bimodal ramjet with circular structure |
| CN104295406B (en) * | 2014-05-26 | 2016-08-24 | 清华大学 | A kind of rocket punching press combined engine with ring-like ejection structure |
| CN111664022B (en) * | 2020-06-16 | 2021-06-18 | 中国人民解放军国防科技大学 | Combustion chamber of rotary detonation ramjet engine with fuel injection |
-
2020
- 2020-12-24 CN CN202011573592.8A patent/CN112682218B/en active Active
Also Published As
| Publication number | Publication date |
|---|---|
| CN112682218A (en) | 2021-04-20 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN112682219B (en) | Wide-speed-range engine based on tail confluence rocket of annular supercharging central body | |
| CN111664022B (en) | Combustion chamber of rotary detonation ramjet engine with fuel injection | |
| CN109854379B (en) | Turbofan engine | |
| CN104654362B (en) | Large scale scramjet engine and three-dimensional petal cross section combustor | |
| CN111577480A (en) | Low detectable integration spray tube suitable for self-adaptation engine | |
| CN109139296B (en) | Rocket-based combined cycle engine | |
| CN113374595B (en) | Design method of elliptical throat offset type pneumatic vectoring nozzle | |
| CN109339975B (en) | Stealth exhaust pipe with adjustable conical cavity | |
| US3012400A (en) | Nozzle | |
| CN110131074B (en) | Bipropellant air turbine rocket propulsion system | |
| CN112727635B (en) | Double-culvert engine | |
| CN214403792U (en) | Series-parallel three-power combined engine | |
| CN112682218B (en) | Wide-speed-range engine based on annular supercharging central body mixed section confluence rocket stamping | |
| CN107013334A (en) | A kind of double combustion chamber's Scramjet Inlet and air intake control method | |
| CN109339875A (en) | A kind of mixing diffuser of band bypass bleed | |
| CN112780615A (en) | Supersonic air ejector | |
| CN109236759A (en) | A kind of Supersonic Ejector of multiple-unit honeycomb composite structure | |
| CN112682217B (en) | Rocket stamping wide-speed-range engine based on annular supercharging central body rear edge expansion | |
| US20160090174A1 (en) | Reaction drive blade tip with turning vanes | |
| CN106321282A (en) | Convergent-divergent nozzle with stealth function | |
| US10377475B2 (en) | Nozzles for a reaction drive blade tip with turning vanes | |
| CN103987949A (en) | Convergent-divergent turbo-machine nozzle | |
| CN212177294U (en) | Three-duct convergent-divergent type tail nozzle structure | |
| CN214577972U (en) | Supersonic air ejector | |
| CN212177293U (en) | Double-contraction-expansion type tail nozzle structure with three ducts for injection |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |