CN113944568B - HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine - Google Patents
HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine Download PDFInfo
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- CN113944568B CN113944568B CN202111190960.5A CN202111190960A CN113944568B CN 113944568 B CN113944568 B CN 113944568B CN 202111190960 A CN202111190960 A CN 202111190960A CN 113944568 B CN113944568 B CN 113944568B
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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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid 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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/26—Burning control
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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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—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
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
- F02K9/34—Casings; Combustion chambers; Liners thereof
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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
- F02K9/95—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by starting or ignition means or arrangements
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E20/00—Combustion technologies with mitigation potential
- Y02E20/16—Combined cycle power plant [CCPP], or combined cycle gas turbine [CCGT]
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Testing Of Engines (AREA)
Abstract
The invention relates to a unit propellant-based support plate injection rocket-based combined cycle power engine which comprises an air inlet channel, a central support plate unit rocket, a powder fuel supply system, a hydroxylamine nitrate propellant supply system, a high-pressure air source, a combustion chamber and a spray pipe. The hydroxylamine nitrate monocomponent propellant which has good normal temperature storage performance and can be quickly catalyzed and started at normal temperature is used as the propellant of the support plate rocket in the engine, so that the structure of the support plate rocket system is simplified, and the reliability of the engine is improved; aluminum powder is adopted to replace the traditional hydrocarbon fuel, so that the density ratio of the engine is improved, the fuel dissociation phenomenon in the scramjet mode is overcome, and the performance of the engine is maintained. The engine realizes mode conversion in the flight process by adjusting the flow of the hydroxylamine nitrate propellant.
Description
Technical Field
The invention belongs to the technical field of aerospace power, and relates to a powder fuel support plate rocket-based combined cycle engine based on HAN unit propellant.
Background
The rocket-based combined cycle power system is a propulsion system combining a rocket engine and a ramjet engine, the working process of the rocket-based combined cycle power system comprises four modes of injection mode, sub-combustion mode, super-combustion mode and rocket mode, and the respective advantages of the rocket engine and the ramjet engine can be fully exerted by adjusting the working modes of the engines, so that an aircraft is always in an optimal thermodynamic cycle state under different flight parameters, and the rocket-based combined cycle power system has high specific impulse and high thrust-weight ratio.
The rocket-based combined cycle power system mainly realizes the mutual conversion among different modes by controlling the oxygen-fuel ratio of the injection rocket in the working process, and the propulsion system is mainly based on aviation kerosene and a liquid oxygen system (Zhang Qian and the like, rocket propulsion, 2014,40 (5): 1-7). Because the boiling point temperature of oxygen is very low, the oxygen can be evaporated due to heating in the working process of the aircraft, and long-time storage is difficult. When the cryogenic liquid oxygen evaporates to a certain extent, the pressure in the storage tank is increased, and the safety of the storage tank is greatly threatened (Cheng Jinjie, low-temperature engineering, 2012, (06): 46-50), so that the reliability of the aircraft is reduced. When the flying speed of the aircraft is greater than 7Ma, the aircraft is influenced by high total incoming flow temperature, so that hydrocarbon fuel is irreversibly dissociated, the fuel utilization rate is reduced (Chen Jun and the like, propulsion technology, 2013,34 (10): 1345-1352); in addition, the higher total incoming flow temperature can also raise the temperature of the combustion chamber of the engine, so that the temperature is higher than the adiabatic combustion temperature of hydrocarbon fuel, and the ram air flow cannot be continuously expanded to work, thereby reducing the performance of the engine.
Disclosure of Invention
Aiming at the defects of the prior art, the invention provides a novel power combination mode with excellent performance, simple structure and high reliability for a rocket-based combined cycle power system.
In order to solve the problem of engine performance reduction caused by dissociation of fuel in a scramjet mode of a rocket-based combined cycle power system, the invention provides a powder fuel support plate injection rocket-based combined cycle engine based on a hydroxylamine nitrate (HAN) unit propellant, and the HAN unit propellant and aluminum powder are respectively used as fuel and oxidant in the power system, so that efficient application of the powder fuel in the rocket-based combined cycle power system is realized.
The idea of the invention is as follows: the aluminum powder and hydroxylamine nitrate monocomponent propellant are used as fuel and oxidant of the rocket-based combined cycle engine, so that the normal-temperature stable storage of the propellant can be realized. Compared with hydrocarbon fuel, aluminum powder is used as fuel, most of combustion products of the hydrocarbon fuel take condensed phase aluminum oxide as a main material, the structure is stable, dissociation effect can not occur under the high temperature condition, stable combustion and heat release of the fuel can be ensured to be maintained under the full mode, the volume heat value of the aluminum powder is far higher than that of the traditional hydrocarbon fuel, the density specific impulse of an engine can be remarkably improved, and in addition, the adiabatic combustion temperature of the aluminum powder is higher than that of the hydrocarbon fuel under the same equivalent ratio, so that the influence caused by high incoming flow total temperature can be overcome. The hydroxylamine nitrate monocomponent propellant can realize catalytic ignition in the catalytic bed, and the propellant can realize stable decomposition and heat release of the propellant at normal temperature, so that an engine does not need to carry an additional ignition system. The decomposition gas temperature of the hydroxylamine nitrate propellant can reach more than about 1500K, and the hydroxylamine nitrate propellant contains a large amount of water vapor, nitrogen dioxide and other oxidizing atmospheres required by aluminum powder combustion, so that a high-temperature oxidant is provided for aluminum powder combustion, and the rapid ignition and efficient combustion of the aluminum powder in high-speed airflow are promoted. Under the injection mode, the flow rates of the engine fuel and the oxidant are determined according to the chemical reaction equivalent ratio of the aluminum powder to the oxidizing atmosphere in the complete decomposition product of the HAN-based propellant; in the sub-combustion and super-combustion stamping modes, aluminum powder fuel is mainly used, hydroxylamine nitrate propellant and aluminum powder are supplied in a small amount to conduct pre-combustion, and the combustion efficiency of the fuel in a combustion chamber is improved; in the pure rocket mode, the aluminum powder fuel supply system is closed, and the hydroxylamine nitrate single-component rocket engine is utilized to power the aircraft. The hydroxylamine nitrate propellant can be used as an oxidant and an aluminum powder ignition heat source, and can be used as a gesture control power propellant of an aircraft, and rocket mode, oxidant, fuel ignition heat source and gesture control power of the aircraft can be finished by using a single propellant, so that the negative quality of an engine can be greatly reduced, and the reliability is improved.
The invention is realized by the following technical scheme:
the powder fuel supporting plate injection rocket-based combined cycle engine based on HAN unit propellant is characterized by comprising an air inlet channel 1, a central supporting plate unit rocket 2, a powder fuel supply system 3, a hydroxylamine nitrate propellant supply system 4, a combustion chamber 5, a spray pipe 6 and a high-pressure air source 7;
the central support plate single-unit rocket 2 comprises a propellant inlet 8, a capillary tube 9, a front bed 10, a rear bed 11 and a convergent nozzle 12 which are sequentially arranged from the front end to the rear; the powder fuel supply system 3 comprises a driving cavity 14, a piston 15, a powder storage tank 16, a fluidization cavity 17 and a depressurization element 18 for controlling the high-pressure air source, which are sequentially arranged at the inlet of the high-pressure air source, and a powder switch valve 19 is arranged at the outlet of the fluidization cavity;
an air inlet channel 1 is positioned at the front end of the engine and leads to a central support plate unit rocket 2, one path of high-pressure air source 7 is connected with the inlet of a hydroxylamine nitrate propellant supply system 4, the outlet of the hydroxylamine nitrate propellant supply system 4 is connected with a propellant inlet 8 of the central support plate unit rocket 2, and a convergent nozzle 12 of the central support plate unit rocket is communicated with a combustion chamber 5; the other path of the high-pressure air source 7 is connected with a driving cavity 14 of the powder fuel supply system 3 through a solenoid valve 13, and the powder switch valve 19 is connected with the combustion chamber 5; a lance 6 is located downstream of the combustion chamber 5.
Advantageous effects
The engine structure of the invention is suitable for hydroxylamine nitrate as a propellant and aluminum powder as a fuel. The engine can be started by normal-temperature catalytic ignition without carrying an ignition component, so that the reliability of the engine is improved, and the engine can be used as an attitude and orbit control power system of an aircraft, so that the negative quality of the aircraft is reduced; meanwhile, dissociation of the traditional hydrocarbon fuel in the scramjet mode can be effectively inhibited, and density specific impulse of the engine is improved.
Drawings
FIG. 1 is a schematic diagram of a pulverized fuel panel-ejector rocket-based combined cycle engine based on a hydroxylamine nitrate (HAN) unit propellant.
Wherein, 1: inlet channel, 2: center extension board unit rocket, 3: pulverized fuel feed system, 4: hydroxylamine nitrate propellant supply system, 5: combustion chamber, 6: jet pipe, 7: and a high-pressure air source.
Fig. 2 is a schematic structural view of a central support plate single-component rocket.
Wherein, 8: propellant inlet, 9: capillary tube, 10: front bed, 11: back bed, 12: and (5) converging the spray pipe.
Fig. 3 is a schematic view of the powder feeding system.
Wherein, 14: drive chamber, 15: piston, 16) powder reservoir, 17: fluidization chamber, 18: pressure drop element, 19: powder switching valve.
Figure 4 (a) is the engine performance in the overfire mode of the ai pulverized fuel system,
FIG. 4 (b) is the engine performance in the overfire mode of the RP-1 kerosene fuel system.
Detailed Description
The invention is further illustrated by the following detailed description and examples, and by the accompanying drawings.
The HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine is shown in figures 1, 2 and 3.
Ram air enters a combustion chamber 5 from an air inlet channel 1 through channels on two sides of a central support plate single-component rocket 2; the engine oxidant is extruded and conveyed to a propellant inlet 8 of a central support plate single-component rocket 2 by a hydroxylamine nitrate propellant supply system 4 through a high-pressure air source 7, the propellant enters a catalytic bed through a capillary 9, contacts with a propellant decomposition catalyst in a front bed 10 and a rear bed 11 and undergoes decomposition reaction to generate high-temperature vapor, carbon dioxide, nitric oxide and other oxidizing atmospheres, and is discharged into a combustion chamber through an engine converging spray pipe 12; the fuel aluminum powder in the engine is fluidized and conveyed to the combustion chamber 5 by the powder supply system 3 through the high-pressure air source 7, the powder is conveyed by adopting an air pressure driving piston type powder supply scheme, firstly, the high-pressure air source flows into the driving cavity 14 through the electromagnetic valve 13 to drive the piston 15 to move forwards, the aluminum powder in the powder fuel storage tank 16 is extruded to flow 17 into the fluidization cavity, meanwhile, the fluidization air enters the fluidization cavity 17 through the pressure drop element 18, the aluminum powder in the rolling fluidization cavity 17 enters the combustion chamber 5 through the powder switch valve 19, and the combustion products of the powder and the central support plate single-component rocket 2 are mixed and pre-combusted near the outlet of the engine converging nozzle 12. The accurate control of the oxygen-fuel ratio of the engine can be realized by adjusting the flow of the HAN-based single-component propellant; the on-off of the aluminum powder fuel supply can be realized by controlling the powder supply switching valve 19 and the switching of the driving gas and the fluidizing gas, and the conversion of the engine under different modes can be realized by the adjustment.
Examples
In the starting process of the aircraft, the engine is in an injection state and is supplied according to the equivalent ratio of aluminum powder to hydroxylamine nitrate propellant complete decomposition products, the HAN-based single-component rocket engine is started by catalysis, and the stable operation of the HAN-based single-component rocket engine can be realized by directly injecting the propellant into the fore bed and the aft bed. With the increase of the speed of the aircraft, when the flying speed reaches 2Ma, the engine enters a sub-combustion ram mode, the flow of the propellant of the HAN-based single-component rocket engine begins to gradually decrease, and when the back pressure in the single-component rocket engine is lower than the ambient pressure, the single-component rocket engine is shut down, and aluminum powder and ram air are combusted in a combustion chamber. When the flying speed is greater than 5Ma, the engine enters a scramjet mode, the aircraft enters a nearby space, the HAN-based single-component rocket engine restarts to work, and the flow of the hydroxylamine nitrate propellant is 10% of the flow of the aluminum powder. In the punching mode, the ratio of the aluminum powder flow to the air flow is always constant at 1:6. And after the aircraft enters the orbit, closing the aluminum powder supply system, and using the HAN-based single-component rocket engine to power the aircraft.
To verify the performance advantages of the present invention, thermal calculations were performed on the engine performance at 8Ma and 30km altitude to obtain the density ratio and the combustor adiabatic temperature variation with the combustion equivalence ratio for the aluminum powder-fueled engine and the conventional kerosene (RP-1) -fueled engine, respectively, as shown in fig. 4 (a) and 4 (b). According to the graph, the density ratio of the rocket-based combined cycle engine can be effectively improved, the adiabatic combustion temperature in the combustion chamber of the engine can be improved, and the fuel utilization rate in the overfire mode is improved. Compared with a rocket-based combined cycle power system of a traditional hydrocarbon fuel and liquid oxygen system, the power system provided by the invention can reduce carrying of a low-temperature liquid oxygen storage tank and a rocket-jet ignition system in actual use, reduce the negative quality of an engine and improve the reliability.
Claims (1)
1. The powder fuel support plate injection rocket-based combined cycle engine based on HAN unit propellant is characterized by comprising an air inlet channel (1), a central support plate unit rocket (2), a powder fuel supply system (3), a hydroxylamine nitrate propellant supply system (4), a combustion chamber (5), a spray pipe (6) and a high-pressure air source (7);
the central support plate single-unit rocket (2) comprises a propellant inlet (8), a capillary tube (9), a front bed (10), a rear bed (11) and a convergent nozzle (12) which are sequentially arranged from the front end to the rear; the powder fuel supply system (3) comprises a driving cavity (14), a piston (15), a powder storage tank (16), a fluidization cavity (17) and a depressurization element (18) for controlling the high-pressure air source, which are sequentially arranged at the inlet of the high-pressure air source, and a powder switch valve (19) is arranged at the outlet of the fluidization cavity;
an air inlet channel (1) is arranged at the front end of the engine and leads to a central support plate single-component rocket (2), one path of high-pressure air source (7) is connected with the inlet of a hydroxylamine nitrate propellant supply system (4), the outlet of the hydroxylamine nitrate propellant supply system (4) is connected with a propellant inlet (8) of the central support plate single-component rocket (2), and a convergent nozzle (12) of the central support plate single-component rocket is communicated with a combustion chamber (5); the other path of the high-pressure air source (7) is connected with a driving cavity (14) of the powder fuel supply system (3) through an electromagnetic valve (13), and the powder switch valve (19) is connected with the combustion chamber (5); a lance (6) is located downstream of the combustion chamber (5).
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| CN202111190960.5A CN113944568B (en) | 2021-10-13 | 2021-10-13 | HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine |
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| CN202111190960.5A CN113944568B (en) | 2021-10-13 | 2021-10-13 | HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine |
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| CN113944568B true CN113944568B (en) | 2023-09-05 |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| IL292537B2 (en) * | 2022-04-26 | 2024-02-01 | Powder Fuels Ltd | Rocket Motor Using Powder Monopropellant |
| CN114837849B (en) * | 2022-05-10 | 2023-03-28 | 南京航空航天大学 | Powder-liquid stamping combined engine and control method |
Citations (3)
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| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN109322763A (en) * | 2018-09-19 | 2019-02-12 | 中国人民解放军国防科技大学 | Solid rocket powder scramjet engine |
| CN111594346A (en) * | 2020-05-01 | 2020-08-28 | 西北工业大学 | Mesoscale rocket-based combined cycle engine |
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| US7966805B2 (en) * | 2007-05-15 | 2011-06-28 | Raytheon Company | Hydroxyl amine based staged combustion hybrid rocket motor |
| US8024918B2 (en) * | 2008-04-29 | 2011-09-27 | Honeywell International Inc. | Rocket motor having a catalytic hydroxylammonium (HAN) decomposer and method for combusting the decomposed HAN-based propellant |
| US11124464B2 (en) * | 2016-09-09 | 2021-09-21 | Terves, Llc | High density hybrid rocket motor |
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Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107503862A (en) * | 2017-10-10 | 2017-12-22 | 北京航空航天大学 | A kind of hybrid rocket combination circulation propulsion system and its control method |
| CN109322763A (en) * | 2018-09-19 | 2019-02-12 | 中国人民解放军国防科技大学 | Solid rocket powder scramjet engine |
| CN111594346A (en) * | 2020-05-01 | 2020-08-28 | 西北工业大学 | Mesoscale rocket-based combined cycle engine |
Non-Patent Citations (1)
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| RBCC的可实现性方案-DRBCC分析;张倩;王兵;张耘隆;张会强;;火箭推进(05);第1-7页 * |
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