CN112431692B - Cooperation air-breathing liquid rocket engine propellant supply system - Google Patents
Cooperation air-breathing liquid rocket engine propellant supply system Download PDFInfo
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- CN112431692B CN112431692B CN202011285056.8A CN202011285056A CN112431692B CN 112431692 B CN112431692 B CN 112431692B CN 202011285056 A CN202011285056 A CN 202011285056A CN 112431692 B CN112431692 B CN 112431692B
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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
- F02K9/44—Feeding propellants
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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
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- 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
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Abstract
The invention discloses a synergistic air-breathing liquid rocket engine propellant supply system which comprises a liquid oxygen storage tank, a liquid hydrogen storage tank, a liquid oxygen pipeline, a liquid hydrogen pipeline, a high-temperature incoming flow air pipeline, a helium pipeline and a rocket thrust chamber. A rocket belly is formed between the liquid hydrogen storage tank and the rocket thrust chamber; the liquid hydrogen pipeline is positioned at the abdomen of the rocket and is arranged in a linear way; the high-temperature incoming flow air pipeline and the helium pipeline are wound on the outer side of the liquid hydrogen pipeline in a parallel mode, and the helium pipeline is positioned between the two pipelines; the side wall of the air inlet is provided with a telescopic arm with telescopic length; when the telescopic arm is elongated, the air inlet is pushed out from the abdomen of the rocket, so that an air suction working mode is realized; when the length of the telescopic arm is shortened, the air inlet is stored in the belly of the rocket, and a pure rocket working mode is realized. The invention can utilize oxygen in the atmosphere, optimizes the layout of the internal pipelines of the traditional liquid rocket engine, and ensures that the traditional liquid rocket engine has high specific impulse and large thrust-weight ratio, thereby reducing the launching cost.
Description
Technical Field
The invention relates to the field of aerospace power, in particular to a synergistic air-breathing liquid rocket engine propellant supply system.
Background
In various space power devices, liquid rocket engines are widely applied due to the characteristics of large specific impulse, easy control and the like. The oxyhydrogen rocket engine is a liquid rocket engine which takes liquid oxygen as an oxidant and liquid hydrogen as a combustion agent, the liquid hydrogen and the liquid oxygen are the only combination with specific impulse exceeding 400s in a liquid propellant, and the oxyhydrogen rocket engine has the advantage of high specific impulse, can greatly reduce the takeoff weight of the rocket and is widely used for carrier rockets. However, the oxyhydrogen rocket engine in China starts late, has lower level in the aspects of thrust and specific impulse compared with foreign countries, and how to improve the performance of the oxyhydrogen rocket engine is the research focus in the aerospace power field in China.
Conventional liquid rocket engines can be divided into two types, namely a pump type and a squeeze type, according to the propellant supply mode, however, the two types have some structural problems: taking the pumping pressure type as an example, the use of a large number of elements such as valves, pumps and the like needs to be driven by a power supply, so that the power supply subsystem of the rocket is burdened, the limited power supply capacity is not enough to drive the propellant to supply, and the pumping pressure type liquid rocket engine is complicated in structure and easy to break down due to the fact that the propellant is supplied by a small engine such as a fuel gas generator and a fuel starting oil tank. The extrusion type liquid rocket engine is simple in structure, needs to carry high-pressure gas cylinders and other components, is huge in structural mass, and is generally suitable for a propulsion system with low thrust and low total impact requirements or an orbit attitude control engine for a satellite. These propellant supply system problems also exist for oxyhydrogen engines.
Therefore, how to design a more optimized propellant delivery system of the oxyhydrogen rocket engine becomes a very potential research direction.
Disclosure of Invention
The technical problem to be solved by the invention is to provide a cooperative air-breathing liquid rocket engine propellant supply system aiming at the defects of the prior art, the cooperative air-breathing liquid rocket engine propellant supply system can utilize oxygen in the atmosphere, and the layout of the internal pipelines of the traditional liquid rocket engine is optimized, so that the traditional liquid rocket engine has high specific impulse and large thrust-weight ratio, and the launching cost is reduced.
In order to solve the technical problems, the invention adopts the technical scheme that:
a synergistic air-breathing liquid rocket engine propellant supply system comprises a liquid oxygen storage tank, a liquid hydrogen storage tank, a liquid oxygen pipeline, a liquid hydrogen pipeline, a high-temperature incoming flow air pipeline, a helium pipeline and a rocket thrust chamber.
The liquid oxygen storage tank and the liquid hydrogen storage tank are arranged outside the head of the rocket thrust chamber, and a rocket belly is formed between the liquid hydrogen storage tank and the rocket thrust chamber.
The liquid inlet end of the liquid oxygen pipeline passes through the liquid hydrogen storage tank and then is communicated with the liquid oxygen storage tank; the liquid outlet end of the liquid oxygen pipeline is communicated with the rocket thrust chamber.
The liquid hydrogen pipeline is positioned at the belly of the rocket and is arranged in a linear mode and used for transmitting and injecting liquid hydrogen in the liquid hydrogen storage tank into the rocket thrust chamber.
The high-temperature incoming flow air pipeline and the helium pipeline are wound on the outer side of the liquid hydrogen pipeline in a parallel mode, and the helium pipeline is positioned between the high-temperature incoming flow air pipeline and the liquid hydrogen pipeline; the high-temperature incoming flow air pipeline is a flexible pipe, an air inlet is formed in the air inlet end of the high-temperature incoming flow air pipeline, the air outlet end of the high-temperature incoming flow air pipeline is connected with an air compressor, and the air outlet of the air compressor is connected with the rocket thrust chamber.
The side wall of the air inlet is provided with a telescopic arm, the other end of the telescopic arm is fixed, and the length of the telescopic arm can be extended and contracted; when the telescopic arm is elongated, the air inlet is pushed out from the abdomen of the rocket, and can absorb air in the atmosphere; when the length of the telescopic arm is shortened, the air inlet is stored in the belly of the rocket, and air in the atmosphere is prevented from entering the high-temperature incoming flow air pipeline.
The liquid oxygen storage tank, the liquid hydrogen storage tank and the rocket thrust chamber are located on the same axis, and the liquid hydrogen storage tank is located between the liquid oxygen storage tank and the rocket thrust chamber.
The liquid outlet end of the liquid oxygen pipeline is spirally wound on the side wall of the cylinder of the rocket thrust chamber, and liquid oxygen is injected into the rocket thrust chamber from the tail end of the side wall of the cylinder of the rocket thrust chamber.
A precooler is arranged in the air inlet.
The air inlet is a conical bell mouth, and the included angle between the conical generatrix of the conical bell mouth and the central line is 20 degrees.
When the flying speed of the rocket is 0-5 Mach numbers, the length of the telescopic arm is extended, and the air inlet is pushed out from the abdomen of the rocket to absorb air in the atmosphere; and then, the windward posture of the air inlet is controlled by adjusting the posture of the telescopic arm, so that the overflow and resistance increase of supersonic airflow are avoided.
The air compressor is an air turbine compressor with the compression ratio larger than 100:1 and supplies compressed air for the rocket thrust chamber.
The helium pipeline is a helium circulating pipeline, and a helium air pump is arranged in the helium circulating pipeline.
The device also comprises a liquid oxygen branch pipeline, an evaporator, an oxygen reflux pipeline and an oxygen waste discharge pipeline; one end of the liquid oxygen branch pipeline is connected to the middle part of the liquid oxygen pipeline, and the other end of the liquid oxygen branch pipeline is connected to a liquid inlet of the evaporator; the steam outlet of the evaporator is connected with an oxygen reflux pipeline, and the other end of the oxygen reflux pipeline is communicated with the liquid oxygen storage tank; the oxygen exhaust pipeline is connected with the exhaust outlet of the evaporator.
The tail end of the rocket thrust chamber is coaxially provided with a Laval nozzle.
The invention has the following beneficial effects:
1. compared with the traditional liquid rocket engine, the invention has higher specific impulse, larger thrust-weight ratio and simpler internal structure under the condition of the same flight Mach number. The whole propellant transport system has small power consumption and high energy utilization rate. The air suction working mode can greatly save the demand of the oxidant carried by the air suction working mode, thereby greatly reducing the emission cost.
2. Compared with the existing air-breathing engine fuel supply system, the invention does not need to subversive change to the inside of the existing liquid rocket engine, fully utilizes the original system structure, and fully utilizes the temperature difference between different liquids and gases to pre-cool and preheat.
3. The invention does not need to increase too many valves and can be well applied to the aerospace propulsion device.
4. The liquid hydrogen pipeline is linear, namely a straight pipe, the helium pipeline and the high-temperature incoming flow air pipeline are wound around the liquid hydrogen pipeline and are parallel to each other, and the helium pipeline is positioned between the high-temperature incoming flow air pipeline and the liquid hydrogen pipeline. The invention utilizes the characteristic that the liquid hydrogen has extremely low temperature to cool the high-temperature air precooled by the precooler again, and can also preheat the liquid hydrogen at the same time. However, the temperature difference between the liquid hydrogen and the high-temperature air is too large, and the helium gas is adopted for neutralization and buffering, so that the hydrogen embrittlement effect can be avoided.
5. The liquid outlet end of the liquid oxygen pipeline is spirally wound on the side wall of the barrel body of the rocket thrust chamber, and the liquid oxygen can be preheated while the rocket thrust chamber is cooled.
6. The air inlet and the telescopic arm are arranged, and the air suction mode and the carrier rocket mode are switched through the extension and contraction of the telescopic arm. In addition, the windward posture of the air inlet can be controlled by adjusting the posture of the telescopic arm, and the overflow and resistance increase of supersonic airflow are avoided.
Drawings
FIG. 1 shows a schematic diagram of a synergistic aspirated liquid rocket engine propellant supply system according to the present invention.
FIG. 2 shows a schematic winding of a liquid hydrogen pipeline, a helium pipeline and a high temperature incoming air pipeline.
FIG. 3 shows a schematic winding layout of liquid oxygen pipeline on rocket thrust chamber.
Among them are:
10. a liquid oxygen storage tank;
11. a liquid oxygen pipeline; 12. liquid oxygen branch pipeline; 13. an evaporator; 14. an oxygen return line; 15. an oxygen exhaust pipeline;
20. a liquid hydrogen storage tank; 21. a liquid hydrogen pipeline;
30. a high temperature incoming air duct;
31. an air inlet; 32. a telescopic arm; 33. an air compressor;
40. a helium gas conduit;
50. a rocket thrust chamber;
60. a laval nozzle.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and specific preferred embodiments.
In the description of the present invention, it is to be understood that the terms "left side", "right side", "upper part", "lower part", etc., indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of describing the present invention and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and that "first", "second", etc., do not represent an important degree of the component parts, and thus are not to be construed as limiting the present invention. The specific dimensions used in the present example are only for illustrating the technical solution and do not limit the scope of protection of the present invention.
The invention only aims at the bipropellant liquid hydrogen and liquid oxygen as the propellant, and is mainly based on the following two points:
(1) compared with hydrocarbon fuels such as kerosene and the like, the liquid hydrogen has the advantages of higher combustion flame speed, larger range and high heat value, and has higher flight speed and more stable combustion working condition when the system works in an air suction state;
(2) various gases and liquids in the internal pipeline can be treated in a cooling way by virtue of the low temperature of the liquid hydrogen, so that the space is fully utilized, and the power consumption is reduced.
The high temperatures referred to in the present invention are generally referred to as high-temperature incoming air since the air temperature is normal earth atmospheric temperature for liquid hydrogen, but is higher than liquid hydrogen.
As shown in fig. 1, a co-breathing liquid rocket engine propellant supply system comprises a liquid oxygen storage tank 10, a liquid hydrogen storage tank 20, a liquid oxygen pipeline 11, a liquid hydrogen pipeline 21, a high-temperature incoming flow air pipeline 30, a helium pipeline 40, a liquid oxygen branch pipeline 12, an evaporator 13, an oxygen return pipeline 14, an oxygen exhaust pipeline 15, a rocket thrust chamber 50 and a laval nozzle 60.
The liquid oxygen tank 10 and the liquid hydrogen tank 20 are disposed outside the head of the rocket thrust chamber 50, preferably on the same axis, and the liquid hydrogen tank is disposed between the liquid oxygen tank and the rocket thrust chamber, and the liquid hydrogen tank and the rocket thrust chamber form a rocket belly therebetween. The invention can realize static stability in the rocket body without producing bending moment on the integral structure of the rocket.
The laval nozzle is coaxially arranged at the tail end of the rocket thrust chamber, and the rocket can fly at supersonic speed under the air suction state.
The liquid inlet end of the liquid oxygen pipeline passes through the liquid hydrogen storage tank and then is communicated with the liquid oxygen storage tank; as shown in fig. 3, the liquid outlet end of the liquid oxygen pipeline is preferably spirally wound on the barrel sidewall of the rocket thrust chamber, and injects the liquid oxygen into the rocket thrust chamber from the tail end of the barrel sidewall of the rocket thrust chamber. Therefore, the liquid oxygen can be preheated while the rocket thrust chamber is cooled.
One end of the liquid oxygen branch pipeline is connected to the middle part of the liquid oxygen pipeline, and the other end of the liquid oxygen branch pipeline is connected to a liquid inlet of the evaporator; the steam outlet of the evaporator is connected with an oxygen reflux pipeline, and the other end of the oxygen reflux pipeline is communicated with the liquid oxygen storage tank; the oxygen exhaust pipeline is connected with the exhaust outlet of the evaporator. The liquid oxygen is vaporized and pressurized in the evaporator and enters the liquid oxygen storage tank from the oxygen return pipe 14 in a gas form to promote the whole circulation, and the oxygen exhaust pipe is used for exhausting waste gas.
The liquid hydrogen pipeline is positioned at the belly of the rocket and is arranged in a linear mode and used for transmitting and injecting liquid hydrogen in the liquid hydrogen storage tank into the rocket thrust chamber.
As shown in fig. 2, the high-temperature incoming flow air pipeline and the helium pipeline are wound on the outer side of the liquid hydrogen pipeline in a parallel manner (i.e. in a mutually parallel manner), and the helium pipeline is positioned between the high-temperature incoming flow air pipeline and the liquid hydrogen pipeline.
The helium pipeline is preferably a helium circulating pipeline, a helium gas pump is arranged in the helium circulating pipeline, a helium storage tank is not required to be added, the helium is introduced into the pipeline in advance and is circulated by one helium gas pump all the time, and certain power can be provided for the helium gas pipeline when the helium gas pipeline passes through a gas compressor in the circulating process.
Furthermore, the helium circulating pipeline is preferably connected with an air compressor, helium flow is driven by a helium pump, kinetic energy of the helium flow is utilized to drive a turbine in the air turbine compressor, and then certain power is provided for the air turbine compressor.
The high-temperature incoming flow air pipeline is a flexible pipe, an air inlet 31 is formed in the air inlet end of the high-temperature incoming flow air pipeline, the air outlet end of the high-temperature incoming flow air pipeline is connected with an air compressor 33, and the air outlet of the air compressor is connected with the rocket thrust chamber.
Preferably, a precooler is arranged in the air inlet, the air inlet is preferably a conical bell mouth, and the included angle between the conical generatrix of the conical bell mouth and the central line is preferably 20 degrees.
The telescopic arm 32 is attached to the side wall of the air inlet, and the other end of the telescopic arm is fixed in position, and is preferably attached to a liquid hydrogen pipe.
The air compressor is preferably an air turbine compressor with a compression ratio of more than 100:1, and the supply of compressed air to the rocket thrust chamber (or combustion chamber) can improve the combustion efficiency of fuel.
After the high-temperature incoming air is pre-cooled by the pre-cooler (the temperature is usually-130 ℃ to-150 ℃ after being cooled), the high-temperature incoming air is cooled by the low temperature of the liquid hydrogen. In order to avoid overlarge difference between the cold and the heat, the helium pipeline and the high-temperature incoming flow air pipeline are wound in parallel, and a neutralization buffer effect is realized between the helium pipeline and the high-temperature incoming flow air pipeline. The carrier rocket can be switched between an air suction working mode and a pure rocket working mode.
One, air suction working mode
When the carrier rocket is launched and lifted off, the telescopic arm is extended in length, the air inlet is pushed out from the rocket abdomen, and air in the atmosphere can be absorbed and converted into an air suction mode.
The sucked high-temperature incoming flow air is cooled by the precooler and then flows into an air compressor which is a high-compression-ratio turbine compressor with the compression ratio of more than 100:1, and the compressed incoming flow air enters the combustion chamber without consuming liquid oxygen. As shown in figure 2, the high-temperature incoming flow air pipeline is wound on the liquid hydrogen pipeline, so that the liquid hydrogen can be cooled again by utilizing the characteristic that the temperature of the liquid hydrogen is extremely low, the reasonable outlet temperature of the compressor can be obtained, and the power is reduced. Simultaneously, the liquid hydrogen can also be preheated.
However, because the temperature difference between the liquid hydrogen and the high-temperature incoming flow air is too large, a helium pipeline is added into the two pipelines to avoid the hydrogen embrittlement effect, and because helium has the characteristic of high specific heat ratio, the winding mode is the same as that of the high-temperature incoming flow air pipeline.
The liquid oxygen pipeline is divided into two paths, wherein one path enters the evaporator, waste gas is discharged after passing through the evaporator, and steam enters the liquid oxygen storage tank again to pressurize the liquid oxygen storage tank. As shown in fig. 3, the other path of liquid oxygen is wound around the combustion chamber and then enters the combustion chamber, which can cool the combustion chamber or preheat the liquid oxygen.
In addition, in the state, the windward posture of the air inlet is controlled by adjusting the posture of the telescopic arm, so that the overflow and resistance increase of supersonic airflow are avoided.
Second, rocket working mode
When the flying height rises, the oxygen content in the atmosphere reduces (oxygen content reduces and judges according to flying height usually, according to the flight panel board, reach mach number 5 when the rocket flies, it is better selection to convert pure rocket mode of operation to about the flying height 26km, actually can adopt the formula of breathing in mode flight under higher altitude), flexible arm length shortens, air inlet accomodates in the rocket abdomen, closes the intake duct, the air in the separation atmosphere gets into high temperature incoming flow air duct, convert the carrier rocket mode into.
The design mode of the invention can greatly reduce the consumption of the oxidant. For example, a launch vehicle may require about 100 to 300 seconds to launch from launch to jettison fairing, while the main engine nozzle throat may consume about 1.5-3 tons of propellant per second of combustion, with about 1 ton of oxidizer per second. If the air suction mode of the present invention is adopted at the initial stage of emission, hundreds of tons of oxidizer are saved. The engine tank can reduce the volume and oxidant carrying capacity: (1) the engine can be reduced in structural weight, and can carry more loads and improve the thrust-weight ratio. (2) The emission cost is reduced due to the reduction of the consumption of the oxidant, and the emission of nitrogen oxides can be reduced by adopting a RQL (Rich-Quench-Lean) method in a suction mode.
Compared with the traditional liquid rocket engine, the liquid rocket engine has higher specific impulse, larger thrust-weight ratio and simpler internal structure under the condition of the same flight Mach number. Specifically, referring to a new generation of liquid hydrogen/liquid oxygen rocket engine in China, the ground specific impulse is 310s, and under the air suction working mode, when the Mach number is 2 theoretically, the maximum specific impulse can reach 3200s, and the thrust-weight ratio can reach 14. The invention fully utilizes the temperature difference between different liquid propellants and gas media to carry out heat exchange, thereby reducing the requirements on other heat exchange devices carried by the rocket; in addition, the adopted novel pipeline winding type layout more efficiently utilizes the internal space of the rocket body, and the saved space can be used for carrying more fuel or effective load, so that the rocket has stronger cruising ability and higher space utilization rate; in addition, the weight of the arrow body is reduced by reducing the use of internal structural parts and connecting parts, so that the purpose of increasing the thrust-weight ratio is achieved.
Although the preferred embodiments of the present invention have been described in detail, the present invention is not limited to the details of the embodiments, and various equivalent modifications can be made within the technical spirit of the present invention, and the scope of the present invention is also within the scope of the present invention.
Claims (10)
1. A synergistic aspirated liquid rocket engine propellant supply system, characterized by: comprises a liquid oxygen storage tank, a liquid hydrogen storage tank, a liquid oxygen pipeline, a liquid hydrogen pipeline, a high-temperature incoming flow air pipeline, a helium pipeline and a rocket thrust chamber;
the liquid oxygen storage tank and the liquid hydrogen storage tank are arranged on the outer side of the head of the rocket thrust chamber, and a rocket belly is formed between the liquid hydrogen storage tank and the rocket thrust chamber;
the liquid inlet end of the liquid oxygen pipeline passes through the liquid hydrogen storage tank and then is communicated with the liquid oxygen storage tank; the liquid outlet end of the liquid oxygen pipeline is communicated with the rocket thrust chamber;
the liquid hydrogen pipeline is positioned at the belly of the rocket and is arranged in a linear manner, and is used for transmitting and injecting liquid hydrogen in the liquid hydrogen storage tank into the rocket thrust chamber;
the high-temperature incoming flow air pipeline and the helium pipeline are wound on the outer side of the liquid hydrogen pipeline in a parallel mode, and the helium pipeline is positioned between the high-temperature incoming flow air pipeline and the liquid hydrogen pipeline; the high-temperature incoming flow air pipeline is a flexible pipe, an air inlet is formed in the air inlet end of the high-temperature incoming flow air pipeline, the air outlet end of the high-temperature incoming flow air pipeline is connected with an air compressor, and the air outlet of the air compressor is connected with the rocket thrust chamber;
the side wall of the air inlet is provided with a telescopic arm, the other end of the telescopic arm is fixed, and the length of the telescopic arm can be extended and contracted; when the telescopic arm is elongated, the air inlet is pushed out from the abdomen of the rocket, and can absorb air in the atmosphere; when the length of the telescopic arm is shortened, the air inlet is stored in the belly of the rocket, and air in the atmosphere is prevented from entering the high-temperature incoming flow air pipeline.
2. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the liquid oxygen storage tank, the liquid hydrogen storage tank and the rocket thrust chamber are located on the same axis, and the liquid hydrogen storage tank is located between the liquid oxygen storage tank and the rocket thrust chamber.
3. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the liquid outlet end of the liquid oxygen pipeline is spirally wound on the side wall of the cylinder of the rocket thrust chamber, and liquid oxygen is injected into the rocket thrust chamber from the tail end of the side wall of the cylinder of the rocket thrust chamber.
4. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: a precooler is arranged in the air inlet.
5. The co-breathing liquid rocket engine propellant supply system of claim 4, wherein: the air inlet is a conical bell mouth, and the included angle between the conical generatrix of the conical bell mouth and the central line is 20 degrees.
6. The co-breathing liquid rocket engine propellant supply system of claim 5, wherein: when the flying speed of the rocket is 0-5 Mach numbers, the length of the telescopic arm is extended, and the air inlet is pushed out from the abdomen of the rocket to absorb air in the atmosphere; and then, the windward posture of the air inlet is controlled by adjusting the posture of the telescopic arm, so that the overflow and resistance increase of supersonic airflow are avoided.
7. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the air compressor is an air turbine compressor with the compression ratio larger than 100:1 and supplies compressed air for the rocket thrust chamber.
8. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the helium pipeline is a helium circulating pipeline, and a helium air pump is arranged in the helium circulating pipeline.
9. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the device also comprises a liquid oxygen branch pipeline, an evaporator, an oxygen reflux pipeline and an oxygen waste discharge pipeline; one end of the liquid oxygen branch pipeline is connected to the middle part of the liquid oxygen pipeline, and the other end of the liquid oxygen branch pipeline is connected to a liquid inlet of the evaporator; the steam outlet of the evaporator is connected with an oxygen reflux pipeline, and the other end of the oxygen reflux pipeline is communicated with the liquid oxygen storage tank; the oxygen exhaust pipeline is connected with the exhaust outlet of the evaporator.
10. The co-breathing liquid rocket engine propellant supply system of claim 1, wherein: the tail end of the rocket thrust chamber is coaxially provided with a Laval nozzle.
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| CN114412662B (en) * | 2022-01-21 | 2023-05-02 | 五邑大学 | Liquid level stabilizing device of oxidant box and oxyhydrogen rocket |
| CN114348243B (en) * | 2022-03-18 | 2022-05-17 | 四川凯德源科技有限公司 | Hydrogenation type multistage combustion and explosion propelling device |
Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4771601A (en) * | 1986-05-30 | 1988-09-20 | Erno Raumfahrttechnik Gmbh | Rocket drive with air intake |
| JPH01309899A (en) * | 1988-06-06 | 1989-12-14 | Mitsubishi Heavy Ind Ltd | Engine |
| US5101622A (en) * | 1983-12-23 | 1992-04-07 | Rolls-Royce Plc | Aerospace propulsion |
| US6052987A (en) * | 1997-11-14 | 2000-04-25 | Trw Inc. | Non-propellant fluid cooled space craft rocket engine |
| CN103711610A (en) * | 2013-12-18 | 2014-04-09 | 中国航天科技集团公司第六研究院第十一研究所 | RBCC fuel gas generator integrated supplying and adjusting system based on liquid oxygen |
| CN107152353A (en) * | 2017-06-15 | 2017-09-12 | 葛明龙 | New rocket and airship series hydrogen peroxide thrust chamber |
| CN109026444A (en) * | 2017-06-09 | 2018-12-18 | 范学军 | Composite engine |
| CN109630316A (en) * | 2018-12-13 | 2019-04-16 | 西安航天动力研究所 | A kind of dynamical system for carrier rocket Upper Stage and orbit transfer vehicle |
| CN109630315A (en) * | 2019-02-25 | 2019-04-16 | 中国人民解放军国防科技大学 | Solid rocket scramjet engine, arc-shaped gas generator and central injection device |
| CN110985238A (en) * | 2019-12-31 | 2020-04-10 | 西北工业大学 | Variable working condition rocket engine capable of realizing height compensation |
| CN111207010A (en) * | 2020-01-19 | 2020-05-29 | 上海交通大学 | Ground test device and test method for directly pressurizing cold helium in liquid oxygen temperature zone |
| CN111794878A (en) * | 2020-08-06 | 2020-10-20 | 北京环境特性研究所 | Rocket engine cooling and stealth design device |
Family Cites Families (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20120060464A1 (en) * | 2007-07-24 | 2012-03-15 | James Robert Grote | Systems, methods and apparatus for propulsion |
| FR3000995B1 (en) * | 2013-01-11 | 2015-07-24 | Snecma | ERGOL SUPPLY CIRCUIT AND COOLING METHOD |
| US20140260186A1 (en) * | 2013-03-15 | 2014-09-18 | Patrick R.E. Bahn | Rocket engine systems with an independently regulated cooling system |
| EA201991483A1 (en) * | 2013-03-15 | 2019-11-29 | STARTING SYSTEM AND METHOD OF STARTING USEFUL LOAD |
-
2020
- 2020-11-17 CN CN202011285056.8A patent/CN112431692B/en active Active
Patent Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5101622A (en) * | 1983-12-23 | 1992-04-07 | Rolls-Royce Plc | Aerospace propulsion |
| US4771601A (en) * | 1986-05-30 | 1988-09-20 | Erno Raumfahrttechnik Gmbh | Rocket drive with air intake |
| JPH01309899A (en) * | 1988-06-06 | 1989-12-14 | Mitsubishi Heavy Ind Ltd | Engine |
| US6052987A (en) * | 1997-11-14 | 2000-04-25 | Trw Inc. | Non-propellant fluid cooled space craft rocket engine |
| CN103711610A (en) * | 2013-12-18 | 2014-04-09 | 中国航天科技集团公司第六研究院第十一研究所 | RBCC fuel gas generator integrated supplying and adjusting system based on liquid oxygen |
| CN109026444A (en) * | 2017-06-09 | 2018-12-18 | 范学军 | Composite engine |
| CN107152353A (en) * | 2017-06-15 | 2017-09-12 | 葛明龙 | New rocket and airship series hydrogen peroxide thrust chamber |
| CN109630316A (en) * | 2018-12-13 | 2019-04-16 | 西安航天动力研究所 | A kind of dynamical system for carrier rocket Upper Stage and orbit transfer vehicle |
| CN109630315A (en) * | 2019-02-25 | 2019-04-16 | 中国人民解放军国防科技大学 | Solid rocket scramjet engine, arc-shaped gas generator and central injection device |
| CN110985238A (en) * | 2019-12-31 | 2020-04-10 | 西北工业大学 | Variable working condition rocket engine capable of realizing height compensation |
| CN111207010A (en) * | 2020-01-19 | 2020-05-29 | 上海交通大学 | Ground test device and test method for directly pressurizing cold helium in liquid oxygen temperature zone |
| CN111794878A (en) * | 2020-08-06 | 2020-10-20 | 北京环境特性研究所 | Rocket engine cooling and stealth design device |
Non-Patent Citations (2)
| Title |
|---|
| 推进剂入口压力响应时间对发动机起动过程的影响;王迪等;《上海航天》;20171025;83-87 * |
| 舱外推进剂管路冗余电加热技术;刘小旭等;《导弹与航天运载技术》;20181010;27-30 * |
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