CN110131074B - Bipropellant air turbine rocket propulsion system - Google Patents
Bipropellant air turbine rocket propulsion system Download PDFInfo
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- CN110131074B CN110131074B CN201910441184.8A CN201910441184A CN110131074B CN 110131074 B CN110131074 B CN 110131074B CN 201910441184 A CN201910441184 A CN 201910441184A CN 110131074 B CN110131074 B CN 110131074B
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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
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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
- F02K9/74—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof combined with another jet-propulsion plant
- F02K9/78—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof combined with another jet-propulsion plant with an air-breathing jet-propulsion plant
Abstract
The invention relates to an aircraft power system, in particular to a bipropellant air turbine rocket propulsion system, which solves the technical problem that the existing aircraft power system cannot meet the requirement of a hypersonic aircraft on the power system. The technical solution of the invention is as follows: a bipropellant air turbine rocket propulsion system comprises an engine shell, a secondary combustion chamber and a tail nozzle which are sequentially connected along a gas injection direction; the inner cavities of the engine shell and the secondary combustion chamber are provided with a fairing cone, an axial flow fan, a connecting shaft, a primary combustion chamber, a gas turbine and a mixing device; the outlet end of the rectifying cone is connected with the inlet of the axial flow fan; a primary combustion chamber is arranged between the axial flow fan and the gas turbine; the inlet of the primary combustion chamber is connected with a propellant supply pipeline, and the outlet of the primary combustion chamber is connected with the inlet of the gas turbine; the outlet of the gas turbine is communicated with an inner flow passage in the mixing device; an outer flow passage formed between the mixing device and the secondary combustion chamber is communicated with an outlet of the axial flow fan.
Description
Technical Field
The invention relates to an aircraft power system, in particular to a bipropellant air turbine rocket propulsion system.
Background
The aircraft is the core of aerospace technology and is a highly comprehensive product of modern science and technology. The hypersonic aircraft is a winged or wingless aircraft with flight speed more than 5 times of sound speed, such as airplanes, missiles and shells, has the characteristic of high penetration success rate, has great military value and potential economic value, and is an important research and development direction in the field of aerospace.
The hypersonic aircraft not only needs to fly at hypersonic speed (over 5 times of sound speed), but also needs to take off and land horizontally, and also needs to be reusable, so that a power system of the hypersonic aircraft needs to have the capability of working at wide Mach number. Mach number refers to the ratio of the speed of the aircraft in air to the speed of sound in undisturbed air far ahead of the altitude.
The existing aircraft power systems are mainly divided into piston engines, turbojet engines, ramjet engines, rocket engines and the like. Rocket engines cannot be repeated, and the working Mach numbers of piston engines, turbojet engines and ramjet engines are narrow; the upper limit of the working Mach number of the piston engine and the turbojet engine is generally not more than Mach 2, so that the requirement of the novel hypersonic aircraft on hypersonic flight cannot be met; the ramjet has a lower limit of working Mach number, the working range cannot be lower than Mach 3, and the requirement of horizontal take-off and landing of the novel hypersonic aircraft cannot be met.
Disclosure of Invention
The invention provides a double-component air turbine rocket propulsion system, aiming at solving the technical problem that the existing aircraft power system cannot meet the requirement of a hypersonic aircraft on the power system.
The technical solution of the invention is as follows:
a bipropellant air turbine rocket propulsion system is characterized in that:
the device comprises an engine shell, a secondary combustion chamber and a tail nozzle which are sequentially connected along the gas injection direction;
the inner cavities of the engine shell and the secondary combustion chamber are sequentially provided with a rectifying cone, an axial flow fan, a connecting shaft, a primary combustion chamber, a gas turbine and a mixing device along the gas injection direction;
the outlet end of the rectifying cone is connected to the inlet of the axial flow fan; one end of the connecting shaft is fixedly sleeved in an inner hole of the axial flow fan, and the other end of the connecting shaft is fixedly connected with the gas turbine; a primary combustion chamber is arranged between the axial flow fan and the gas turbine; the inlet of the primary combustion chamber is connected with a propellant supply pipeline, and the outlet of the primary combustion chamber is connected with the inlet of the gas turbine; the outlet of the gas turbine is fixedly connected with the mixing device and is communicated with an inner flow passage in the mixing device; an outer flow passage formed between the mixing device and the secondary combustion chamber is communicated with an outlet of the axial flow fan, and an outlet of the mixing device is positioned in the secondary combustion chamber.
Further, in order to reduce the gas pressure loss, reduce the heat radiation of the engine and improve the performance of the engine, the primary combustion chamber comprises an annular injector and an annular body part which is hermetically and fixedly connected to one end of the annular injector;
the annular injector comprises an outer bottom, an inner bottom and an inner bottom which are sequentially and fixedly connected in a sealing manner along the gas injection direction; a first cavity is formed between the outsole and the midsole; a second cavity is formed between the middle sole and the inner bottom;
the annular body part is of a double-layer structure and comprises an annular inner body part and an annular outer body part which is coaxial with the annular inner body part; one ends of the annular inner body part and the annular outer body part, which are far away from the annular injector, are both flared ports; a third cavity is formed among the annular inner body part, the annular outer body part and the inner bottom;
a plurality of first nozzles communicated with the first cavity and the third cavity are arranged on the middle bottom; a plurality of second nozzles and third nozzles which are communicated with the second cavity and the third cavity are arranged on the inner bottom;
a plurality of partition plates are arranged in the third cavity; a plurality of the baffle is followed the circumference equipartition of body portion in the annular, just the baffle is fixed on body portion and interior bottom in the annular.
Further, in order to improve the penetration capacity of the fuel-rich gas in the air, shorten the air flow mixing distance, and enhance the uniformity of air-fuel-rich gas mixing and the flame stabilizing capacity of the mixing device, the mixing device comprises a chamfered lobe mixer and a cover plate;
the chamfered lobe mixer is of a hollow structure, the front end of the chamfered lobe mixer is of a cylindrical structure, and a plurality of protrusions with hollow interiors are arranged at the rear end of the chamfered lobe mixer along the circumferential direction; one side of the bulge close to the tail nozzle is provided with an opening, and the opening is in a chrysanthemum petal-shaped structure;
the shape of the cover plate is matched with that of the opening and is arranged in the opening;
an inner flow channel is formed between the inclined lobe mixer and the cover plate and is communicated with the secondary combustion chamber through a plurality of injection holes formed in the cover plate;
and an outer flow passage is formed between the outer surface of the plurality of bulges and the inner wall of the secondary combustion chamber.
Further, the cross section of the protrusion is triangular.
Furthermore, heat insulation lining sleeves are arranged on the inner walls of the secondary combustion chamber and the tail nozzle.
Further, the axial fan is a five-stage axial fan, wherein the first-stage fan and the third-stage fan are fixedly connected with the connecting shaft, and the second-stage fan, the fourth-stage fan and the fifth-stage fan are fixedly connected with the engine shell.
Further, the outlet of the primary combustion chamber is welded with the inlet of the gas turbine.
Furthermore, the material of the primary combustion chamber is GH 230.
Furthermore, the material of the blending device is GH 230.
Compared with the prior art, the invention has the beneficial effects that:
1. according to the invention, a rocket engine and a turbojet engine are combined, a primary combustion chamber is arranged between an axial flow fan and a gas turbine, an oxidant and fuel are combusted in the primary combustion chamber to generate rich combustion gas, the rich combustion gas expands to do work to drive the gas turbine to rotate, the gas turbine drives the axial flow fan to rotate, parameters such as flow and pressure of the axial flow fan and the gas turbine are decoupled, and the working Mach number of the engine is widened; the hypersonic aircraft has the capability of working within the Mach number range of 0-4, and meets the requirements of repeated use, horizontal take-off and landing and cruise flight of the hypersonic aircraft.
2. The invention has compact structure, the primary combustion chamber is positioned in the engine, the layout that the traditional primary combustion chamber is positioned outside the engine is changed, the gas pressure loss is effectively reduced, the heat radiation of the engine is reduced, and the performance of the engine is improved.
3. The flow channel of the mixing device is divided into an inner flow channel and an outer flow channel, the mixing device is provided with injection holes communicating the inner flow channel and the outer flow channel, the rich fuel gas flowing through the inner flow channel and the air flowing through the outer flow channel are efficiently mixed through the mixing device and then enter a secondary combustion chamber for secondary combustion, secondary differentiated flow division can be realized, the penetration capacity of the rich fuel gas in the air is improved, the air flow mixing distance is shortened, the mixing uniformity of the air and the rich fuel gas and the flame stability of the mixing device are enhanced, and the stable and efficient combustion of the air and the rich fuel gas in the combustion chamber of the combined power engine under the condition of low pressure loss (the total pressure recovery coefficient of the rich fuel gas in the mixer is lower than 0.80, and the total pressure recovery coefficient of an air path caused by the mixer is about 0.98) can be realized under the condition of Ma 0-3.5 incoming flows.
4. According to the invention, the cover plate is arranged at the outlet of the inner channel of the mixing device to collect the rich fuel gas, the secondary differential shunting is realized by using the injection holes, the residual gas coefficient of the whole secondary combustion chamber can be balanced by adjusting the size, the position and the number of the injection holes, so that the combustion heat release position can be controlled, a low-speed stable flame region can be formed near the axis of the combustion chamber behind the mixing device, the stable and efficient combustion of the air and the rich fuel gas is realized, the local high-temperature region on the outer wall of the combustion chamber can be avoided, and the thermal protection difficulty of the combustion chamber is reduced.
5. According to the invention, the cover plate is arranged at the outlet of the inner flow passage of the mixing device, so that the integral rigidity of the structure is increased, and the mixing device can meet the working requirement by selecting a relatively small wall thickness; the mixing and flame stabilizing device adopts an integrated design, has simple and compact structure and high reliability, does not need to additionally arrange a flame stabilizer device in a combustion chamber, has small air flow loss, and is beneficial to shortening the length of an engine and lightening the weight.
Drawings
FIG. 1 is a schematic block diagram of an embodiment of the present invention;
FIG. 2 is a schematic structural view of a primary combustion chamber in the embodiment;
FIG. 3 is an enlarged view of a portion of FIG. 2;
FIG. 4 is a schematic view of the blending device in this embodiment;
FIG. 5 is a right side view of the blending apparatus of this embodiment;
FIG. 6 is a schematic diagram of the construction of the chamfered lobe mixer in this embodiment;
FIG. 7 is a schematic structural view of the cover plate in this embodiment;
the reference signs are:
1-a fairing cone, 2-an axial fan, 3-a shell, 4-a connecting shaft, 5-a primary combustion chamber, 501-an outer sole, 502-a middle sole, 503-an inner sole, 504-a first cavity, 505-a second cavity, 506-an annular inner body, 507-an annular outer body, 508-a third cavity, 509-a first nozzle, 510-a second nozzle, 511-a third nozzle, 512-a partition plate, 6-a gas turbine, 7-a mixing device, 701-a chamfered lobe mixer, 702-a cover plate, 703-an injection hole, 704-a bulge, 8-a secondary combustion chamber and 9-a tail nozzle.
Detailed Description
The invention is further described with reference to the following figures and specific embodiments.
Referring to fig. 1, a two-component air turbine rocket propulsion system includes an engine case 3, a secondary combustion chamber 8 and a tail nozzle 9 which are connected in sequence along a gas injection direction.
The inner cavities of the engine shell 3 and the secondary combustion chamber 8 are sequentially provided with a fairing cone 1, an axial flow fan 2, a connecting shaft 4, a primary combustion chamber 5, a gas turbine 6 and a mixing device 7 along the gas injection direction. In order to reduce heat loss, heat insulation lining sleeves are arranged on the inner walls of the secondary combustion chamber 8 and the tail nozzle 9.
The outlet end of the rectifying cone 1 is connected to the inlet of the axial flow fan 2. One end of the connecting shaft 4 is sleeved in an inner hole of the axial flow fan 2 and fixedly connected with the axial flow fan 2, and the other end of the connecting shaft 4 is fixedly connected with the gas turbine 6. The axial fan 2 in this embodiment is a five-stage axial fan, in which the first-stage fan and the third-stage fan are fixedly connected to the connecting shaft 4 by a key, and the second-stage fan, the fourth-stage fan and the fifth-stage fan are fixedly connected to the engine case 3 by bolts. A primary combustion chamber 5 is provided between the axial flow fan 2 and the gas turbine 6.
Referring to fig. 2 and 3, the primary combustion chamber 5 comprises an annular injector and an annular body sealingly attached at one end of the annular injector. The annular injector comprises an outer bottom 501, an inner bottom 502 and an inner bottom 503 which are hermetically and fixedly connected in sequence along the gas injection direction; a first cavity 504 is formed between the outsole 501 and the midsole 502; the midsole 502 and the insole 503 define a second cavity 505 therebetween.
The annular body part is of a double-layer structure and comprises an annular inner body part 506 and an annular outer body part 507 which is coaxially arranged with the annular inner body part 506. The ends of the annular inner body part 506 and the annular outer body part 507 far away from the annular injector are both flared; a third cavity 508 is formed between the annular inner body portion 506, the annular outer body portion 507 and the insole 503.
The midsole 502 is provided with a plurality of first nozzles 509 which communicate the first cavity 504 with the third cavity 508; the insole 503 is provided with a plurality of second nozzles 510 and third nozzles 511 communicating the second chamber 505 with the third chamber 508.
A plurality of partition plates 512 are arranged in the third cavity 508; the partition plates 512 are uniformly distributed along the circumferential direction of the annular inner body portion 506, and the partition plates 512 are fixed on the annular inner body portion 506 and the inner bottom 503. The third cavity 508 may also house other components according to the overall needs of the engine.
Referring to fig. 1, the primary combustion chamber 5 has an inlet connected to a propellant supply line and an outlet welded to the inlet of the gas turbine 6. The outlet of the gas turbine 6 is welded to the inlet of the blending device 7.
Referring to fig. 5-7, the blending apparatus 7 includes a chamfered lobe mixer 701 and a cover plate 702;
the chamfered lobe mixer 701 has a hollow structure, the front end of the chamfered lobe mixer is a cylindrical structure, and the rear end of the chamfered lobe mixer is provided with a plurality of protrusions 704 with hollow interiors along the circumferential direction; one side of the bulge 704 close to the tail nozzle is provided with an opening, and the opening is in a chrysanthemum petal-shaped structure; the cover plate 702 has an outer shape matching the outer shape of the opening and is fitted in the opening.
An inner flow channel is formed between the oblique lobe mixer 701 and the cover plate 702 and is communicated with the secondary combustion chamber 8 through a plurality of injection holes 703 formed in the cover plate 702; the outer surfaces of the plurality of protrusions 704 and the inner wall of the secondary combustion chamber 8 form an outer flow passage therebetween. The cross-section of the protrusion 704 is triangular.
The chamfer angle of the lobes 704 of the chamfered lobe mixer 701 may range from 20-45 degrees and the number of lobes 704 may be between 12 and 24.
The number of the projections 704 is 12 in this embodiment, the chamfer angle is 30 degrees, and the diameter of the cross section of the trailing edge of the chamfered lobe mixer 701 is 90 mm. The shape of the injection holes 703 is not limited to a circle, and the size, position, direction and number thereof may be selected according to the incoming flow parameters. The parameter design principle is as follows: the method has the advantages that the uniform mixing of air and rich-combustion gas behind the lobe of the mixing device 7 is ensured, the residual gas coefficient of a low-speed backflow zone near the axis behind the mixing device 7 is ensured to be within a reasonable range (usually between 0.8 and 1.2), a low-speed flame stabilizing zone is further formed behind the mixing device 7, and the efficient and stable combustion of fuel in the secondary combustion chamber 8 is ensured. In addition, the total area of the injection holes 703 should be selected to ensure that the airflow pressure loss meets the overall performance design requirements of the engine. In the embodiment, the injection holes 703 are circular, have different diameters of 0.5-8 mm, are parallel to the axial direction, and have a total area of 8% of the total area of the fuel-rich gas inlet of the flow channel in the mixing device 7.
The rich fuel gas in the primary combustion chamber 5 enters the secondary combustion chamber 8 through the inner runner, the air compressed by the axial flow fan 2 enters the secondary combustion chamber 8 through the outer runner, and simultaneously the rich fuel gas and the air are mixed and combusted in the secondary combustion chamber 8.
In this embodiment, the fuel-rich gas is collected by an internal flow channel formed between the chamfered lobe mixer 701 and the cover plate 702, and is divided into two flows through the chamfered lobe mixer 701 and the injection holes 703 on the cover plate 702. The parameters such as the number of lobes, the chamfer angle of the tail edge of the lobe, the size and the relative position of injection holes, the direction and the number of the lobes of the chamfered lobe mixer 701, the cross section diameter of the tail end of an outlet of the mixer and the like are optimized according to the dimension of a combustion chamber and the parameters of inlet inflow, so that the enhanced mixing and the stable and efficient combustion of the air compressed by the outer axial flow fan 2 and the rich fuel gas are realized under the condition that the air flow pressure loss is low (the total pressure recovery coefficient of the rich fuel gas in the mixer is not less than 0.80).
The material of the primary combustion chamber 5 and the material of the blending device 7 are GH 230.
The range of the design point of the axial flow fan 2 is 2-5, and the efficiency of the design point is not lower than 0.88; the range of the design point of the gas turbine 6 is 8-30, and the efficiency of the design point is not lower than 0.8; the efficiency of the primary combustion chamber is not lower than 0.9; the combustion efficiency of the secondary combustion chamber 8 is not lower than 0.95, and the ignition delay is not more than 10 ms; the air flow resistance loss is not more than 10 percent, and the gas flow resistance loss is not more than 20 percent; the flow coefficient of the tail nozzle 9 is not lower than 0.98, and the area ratio of the throat part to the nozzle outlet is 0.6; the specific impulse can reach 900 s.
The working principle of the system is as follows:
the oxidizer and the fuel respectively enter the first cavity 504 and the second cavity 505 after passing through the propellant supply pipeline, the oxidizer in the first cavity 504 enters the third cavity 508 through the first nozzle 509, and the fuel in the second cavity 505 enters the third cavity 508 through the second nozzle 510 and the third nozzle 511. The oxidizer and the fuel are mixed and combusted in the third chamber 508, wherein part of the fuel sprayed from the second nozzle 510 and the third nozzle 511 forms a liquid film to cool the annular body, and a partition 512 is provided in the third chamber 508 to suppress combustion instability and increase the rigidity of the annular body.
The fuel enters the primary combustion chamber 5, rich combustion gas is generated after the natural or ignition combustion of the primary combustion chamber 5, the rich combustion gas expands through the gas turbine 6 to do work, then the rich combustion gas enters the secondary combustion chamber 8 through an inner flow channel of the mixing device 7, meanwhile, the gas turbine 6 drives the axial flow fan 2 to rotate, incoming air enters the secondary combustion chamber 8 through an outer flow channel of the mixing device after being compressed through the axial flow fan 2, the rich combustion gas and the air are mixed and combusted in the secondary combustion chamber 8, and the fuel gas is finally sprayed out through the tail spray pipe 9 to generate thrust.
Since the cover plate 702 is installed at the outlet of the inner flow channel of the chamfered lobe mixer 701 of the blending device in this embodiment, the rich fuel gas of the inner flow channel is collected, and is secondarily distributed into the secondary combustion chamber 8 by using the chamfered lobe mixer 701 and the injection holes 703 on the cover plate 702, so as to improve the injection pressure of the rich fuel gas, and further improve the penetration capability of the rich fuel gas in the gaseous environment. The size, the relative position, the direction and the number of the injection holes are adjusted according to the inflow parameters, differential shunting is realized, the mass exchange rate of air and rich fuel gas can be improved under the condition of low pressure loss, the mixed gas pressure, the temperature and the speed at the outlet of the inclined-cutting lobe mixer 701 are uniform, the distribution of a combustion area is wider, and the combustion is more violent. Meanwhile, the injection holes 703 are arranged on the chamfered lobe mixer 701, so that on one hand, part of the fuel-rich gas and air can be premixed before the outlet of the mixing device 7; on the other hand, the mixing device is beneficial to improving the turbulence of air, strengthening the air flow mixing effect and shortening the mixing distance behind the mixing device 7. Further, the arrangement of the cover plate 702 can increase the range of the low-speed recirculation zone near the axis of the secondary combustion chamber 8 behind the blending device 7. By adjusting the size, position distribution and number of the injection holes, the residual gas coefficient of the low-speed area can be in a reasonable range, and a low-speed stable flame area is further formed near the axis of the secondary combustion chamber 8 behind the mixing device 7, so that stable combustion of the mixed gas in the secondary combustion chamber 8 is ensured.
Because the system is provided with the primary combustion chamber 5, parameters such as flow, pressure and the like of the axial flow fan 2 and the gas turbine 6 are decoupled, and the working Mach number of the engine is widened; the hypersonic aircraft using the system has the capability of working within the Mach number range of 0-4, can perform cruise flight at hypersonic speed, can take off and land horizontally, and can be reused.
The above description is only an embodiment of the present invention, and is not intended to limit the scope of the present invention, and all equivalent structural changes made by using the contents of the present specification and the drawings, or applied directly or indirectly to other related technical fields, are included in the scope of the present invention.
Claims (8)
1. A two-component air turbine rocket propulsion system, characterized in that:
comprises an engine shell (3), a secondary combustion chamber (8) and a tail nozzle (9) which are connected in sequence along the gas injection direction;
the inner cavities of the engine shell (3) and the secondary combustion chamber (8) are sequentially provided with a rectifying cone (1), an axial flow fan (2), a connecting shaft (4), a primary combustion chamber (5), a gas turbine (6) and a mixing device (7) along the gas injection direction;
the outlet end of the rectifying cone (1) is connected to the inlet of the axial flow fan (2); one end of the connecting shaft (4) is fixedly sleeved in an inner hole of the axial flow fan (2), and the other end of the connecting shaft (4) is fixedly connected with the gas turbine (6); a primary combustion chamber (5) is arranged between the axial flow fan (2) and the gas turbine (6); the inlet of the primary combustion chamber (5) is connected with a propellant supply pipeline, and the outlet of the primary combustion chamber is connected with the inlet of the gas turbine (6); an outlet of the gas turbine (6) is fixedly connected with the blending device (7) and is communicated with an inner flow passage in the blending device (7); an outer flow channel formed between the blending device (7) and the secondary combustion chamber (8) is communicated with an outlet of the axial flow fan (2), and an outlet of the blending device (7) is positioned in the secondary combustion chamber (8);
the primary combustion chamber (5) comprises an annular injector and an annular body part which is hermetically and fixedly connected with one end of the annular injector;
the annular injector comprises an outer bottom (501), an inner bottom (502) and an inner bottom (503) which are sequentially and hermetically and fixedly connected along the gas injection direction; a first cavity (504) is formed between the outsole (501) and the midsole (502); a second cavity (505) is formed between the midsole (502) and the insole (503);
the annular body part is of a double-layer structure and comprises an annular inner body part (506) and an annular outer body part (507) which is coaxial with the annular inner body part (506); one ends of the annular inner body part (506) and the annular outer body part (507) far away from the annular injector are both flared ports; a third cavity (508) is formed among the annular inner body part (506), the annular outer body part (507) and the inner bottom (503);
a plurality of first nozzles (509) which are communicated with the first cavity (504) and the third cavity (508) are arranged on the middle sole (502); a plurality of second nozzles (510) and third nozzles (511) which are communicated with the second cavity (505) and the third cavity (508) are arranged on the inner bottom (503);
a plurality of partition plates (512) are arranged in the third cavity (508); the plurality of the partition plates (512) are uniformly distributed along the circumferential direction of the annular inner body part (506), and the partition plates (512) are fixed on the annular inner body part (506) and the inner bottom (503).
2. A two-component air turbine rocket propulsion system according to claim 1, wherein:
the blending device (7) comprises a chamfered lobe mixer (701) and a cover plate (702);
the oblique lobe mixer (701) is of a hollow structure, the front end of the oblique lobe mixer is of a cylindrical structure, and a plurality of protrusions (704) with hollow interiors are arranged at the rear end of the oblique lobe mixer along the circumferential direction; one side of the bulge (704) close to the tail nozzle is provided with an opening, and the opening is in a chrysanthemum petal-shaped structure;
the shape of the cover plate (702) is matched with that of the opening and is arranged in the opening;
an inner flow channel is formed between the oblique lobe mixer (701) and the cover plate (702), and the inner flow channel is communicated with a secondary combustion chamber (8) through a plurality of injection holes (703) formed in the cover plate (702);
an outer flow passage is formed between the outer surface of the plurality of protrusions (704) and the inner wall of the secondary combustion chamber (8).
3. A two-component air turbine rocket propulsion system according to claim 2, wherein: the cross section of the bulge (704) is triangular.
4. A two-component air turbine rocket propulsion system according to claim 3, wherein: and heat-insulating lining sleeves are arranged on the inner walls of the secondary combustion chamber (8) and the tail spray pipe (9).
5. A two-component air turbine rocket propulsion system according to claim 4, wherein: the axial flow fan (2) is a five-stage axial flow fan, wherein a first-stage fan and a third-stage fan are fixedly connected with the connecting shaft (4), and a second-stage fan, a fourth-stage fan and a fifth-stage fan are fixedly connected with the engine shell (3).
6. A two-component air turbine rocket propulsion system according to claim 5, wherein: and the outlet of the primary combustion chamber (5) is welded with the inlet of the gas turbine (6).
7. A two-component air turbine rocket propulsion system according to claim 6, wherein: the material of the primary combustion chamber (5) is GH 230.
8. A two-component air turbine rocket propulsion system according to claim 7, wherein: the material of the blending device (7) is GH 230.
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| CN110939529B (en) * | 2019-11-25 | 2020-11-06 | 西安航天动力研究所 | Integrated gas supply device |
| CN111852691B (en) * | 2020-08-13 | 2024-02-06 | 北京星际荣耀空间科技股份有限公司 | Integrated double-component injector, manufacturing method thereof and aerospace equipment |
| RU2757437C1 (en) * | 2020-12-22 | 2021-10-15 | Владимир Дмитриевич Куликов | Turbojet engine |
| CN113108654B (en) * | 2021-04-27 | 2021-11-23 | 西北工业大学 | Air water flushing combined cross-medium anti-ship anti-submarine missile |
| CN115419917A (en) * | 2022-07-29 | 2022-12-02 | 西安航天动力研究所 | Integrated heterogeneous multiphase flow mixing flame stabilizing device and combined power engine combustion chamber |
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