CN111594346A - Mesoscale rocket-based combined cycle engine - Google Patents
Mesoscale rocket-based combined cycle engine Download PDFInfo
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- CN111594346A CN111594346A CN202010367767.3A CN202010367767A CN111594346A CN 111594346 A CN111594346 A CN 111594346A CN 202010367767 A CN202010367767 A CN 202010367767A CN 111594346 A CN111594346 A CN 111594346A
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 141
- 239000000446 fuel Substances 0.000 claims abstract description 64
- 239000007924 injection Substances 0.000 claims abstract description 59
- 238000002347 injection Methods 0.000 claims abstract description 59
- 238000002955 isolation Methods 0.000 claims abstract description 31
- 238000002156 mixing Methods 0.000 claims abstract description 7
- 239000007788 liquid Substances 0.000 claims description 7
- 235000015842 Hesperis Nutrition 0.000 claims description 5
- 235000012633 Iberis amara Nutrition 0.000 claims description 5
- 238000003466 welding Methods 0.000 claims description 3
- 239000000567 combustion gas Substances 0.000 claims description 2
- 230000000694 effects Effects 0.000 abstract description 4
- 230000035939 shock Effects 0.000 abstract description 4
- 230000008602 contraction Effects 0.000 description 4
- 239000007789 gas Substances 0.000 description 4
- 230000009471 action Effects 0.000 description 3
- 230000015572 biosynthetic process Effects 0.000 description 3
- 238000004401 flow injection analysis Methods 0.000 description 3
- 239000000243 solution Substances 0.000 description 3
- 230000000087 stabilizing effect Effects 0.000 description 3
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 2
- 238000000889 atomisation Methods 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 239000002737 fuel gas Substances 0.000 description 2
- 230000008520 organization Effects 0.000 description 2
- 239000001301 oxygen Substances 0.000 description 2
- 229910052760 oxygen Inorganic materials 0.000 description 2
- 230000006641 stabilisation Effects 0.000 description 2
- 238000011105 stabilization Methods 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 230000001133 acceleration Effects 0.000 description 1
- 238000001704 evaporation Methods 0.000 description 1
- 230000008020 evaporation Effects 0.000 description 1
- 239000003350 kerosene Substances 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 239000007800 oxidant agent Substances 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
- 230000035515 penetration Effects 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
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Classifications
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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
- F02K7/00—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
- F02K7/10—Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines
- F02K7/18—Composite ram-jet/rocket engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/14—Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/042—Air intakes for gas-turbine plants or jet-propulsion plants having variable geometry
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/22—Fuel supply systems
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/26—Control of fuel supply
- F02C9/28—Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
- F02K1/78—Other construction of jet pipes
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Geometry (AREA)
- Testing Of Engines (AREA)
Abstract
The invention provides a medium-scale rocket-based combined cycle engine which comprises an air inlet, a support plate rocket, an isolation section, a first-stage combustion chamber, a second-stage combustion chamber, a tail nozzle, a first rocket unit, a fuel injection hole and a second rocket unit, wherein the air inlet, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are sequentially connected, and air flow flows into the isolation section, the first-stage combustion chamber and the second-stage combustion chamber from the air inlet and then is discharged outwards from the tail nozzle; the support rocket, the plasma generating body around the support rocket and the support fuel injection points are used for improving the mixing combustion efficiency of incoming flow, and have the functions of injection and providing part of fuel; the isolation section is stabilized the intake duct and is kept apart with first order combustion chamber and lay for hold the precombustion shock wave cluster that forms under the combustion chamber pressure effect, prevent that the air current from taking place great air current fluctuation and then influencing combustion efficiency at the intake section, improve the matching stability of intake duct and rocket unit.
Description
Technical Field
The invention relates to the field of air-breathing combined propulsion systems, in particular to a medium-scale rocket-based combined cycle engine.
Background
A Rocket-Based Combined Cycle (RBCC) engine organically integrates a Rocket engine with a high thrust-weight ratio and an air-breathing ramjet engine with a high specific impulse into the same runner, can be compatible with injection, sub-combustion, super-combustion and pure Rocket modes, and realizes high-performance work in a wide speed range and a large airspace. How to ensure that the same engine can realize good work of each mode in such a wide Mach number range and smooth transition among different modes is the key for determining whether the RBCC engine can succeed or not. The configuration and the mode of operation of the RBCC engine play a decisive role. Especially, how to ensure the reasonable matching of rocket layout and ramjet runner layout, combustion organization strategies of a combustion chamber and the like is the key technology in research.
At present, the commonly used RBCC engine structural scheme is 'variable geometry air inlet channel + central/side rocket or liquid fuel injection + fixed geometry second stage combustion chamber + variable geometry tail nozzle', and is multi-purpose to small scale (the air inlet capture area is smaller than that of the air inlet capture area)0.1m2) The principle level scheme or prototype of (1) to carry out the relevant research. For the air inlet capture area is between 0.2m2To 0.4m2The medium-scale RBCC engine has a combustion chamber with limited penetration depth of fuel injection and relatively large difficulty of combustion organization. In order to solve the problems and improve the combustion efficiency of the engine, the invention adopts a secondary fuel injection mode combining wall injection and central support plate injection, and a plasma generator is assisted around the central support plate fuel injection hole to enhance the atomization and mixing combustion of the fuel. Meanwhile, a second-stage rocket is arranged in the engine, and the two stages of rockets arranged at different positions work in a matching way, so that the two stages of rockets are respectively and pertinently used for high-efficiency injection and suction of air and effective formation of thermal congestion, and the aim of improving the working performance of the RBCC engine is fulfilled.
Disclosure of Invention
In view of the above, the technical problems to be solved by the present invention are: the medium-scale and medium-scale rocket-based combined cycle engine can improve the working performance of the engine in a wide flight range by arranging wall jetting and central support plate jetting combination and arranging a secondary rocket.
To achieve the above object, the present invention provides a mesoscale rocket-based combined cycle engine, comprising: intake duct, extension board rocket, isolation section, first order combustion chamber, second level combustion chamber, exhaust nozzle, first rocket unit, fuel injection hole and second rocket unit, wherein:
the air inlet channel, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are sequentially connected, air flows in from the air inlet channel, and is discharged outwards from the tail nozzle after passing through the isolation section, the first-stage combustion chamber and the second-stage combustion chamber;
the first rocket unit is arranged at the joint of the isolation section and the first-stage combustion chamber, the fuel injection hole is arranged at the joint of the first-stage combustion chamber and the second-stage combustion chamber, and the second rocket unit is arranged at the joint of the second-stage combustion chamber and the tail nozzle;
the first-stage combustion chamber is cylindrical, the outer diameter of the vertical section of the first-stage combustion chamber in the axial direction of the first-stage combustion chamber is the same, and the first-stage combustion chamber is used for completing rapid mixing between rocket jet flow and jet air under an injection mode and efficient combustion under a super-combustion mode;
the tail nozzle is in a flaring shape along the flowing direction of the airflow, and the flowing speed of the flowing airflow is improved by the tail nozzle;
the support plate rocket is arranged on a central support plate in the isolation section, a plasma generation body and support plate fuel injection points are further arranged on the central support plate, the plasma generation body is arranged around the periphery of the support plate rocket, and the support plate fuel injection points are arranged around the periphery of the plasma generation body.
Furthermore, the air inlet channel, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are connected through flanges or welding.
Further, the air inlet is a variable structure air inlet.
Furthermore, the direction of the rocket nozzle of the support plate is parallel to the direction of air flow in the air inlet.
Further, the first rocket unit is configured to introduce airflow from the intake port into the first stage combustion chamber or the second stage combustion chamber.
Further said fuel injection orifices are for providing combustion gases to said first stage combustion chamber and said second stage combustion chamber.
Further, the first rocket unit and the second rocket unit are liquid fuel rockets.
Further, the inclination of the top wall surface of the second-stage combustion chamber in the air flow injection direction ranges from (1:14) to (1: 10).
Further, the capture area of the air inlet channel ranges from 0.2m to 0.4m2。
Compared with the prior art, the invention provides a medium-scale rocket-based combined cycle engine which comprises an air inlet, a support plate rocket, an isolation section, a first-stage combustion chamber, a second-stage combustion chamber, a tail nozzle, a first rocket unit, a fuel injection hole and a second rocket unit, wherein the air inlet, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are sequentially connected, the support plate rocket is arranged on the support plate in the air inlet, and air flow flows into the isolation section, the first-stage combustion chamber and the second-stage combustion chamber from the air inlet and then is discharged outwards from the tail nozzle; the air inlet is used for decelerating and boosting the airflow flowing in from the air inlet; the support rocket, the plasma generator on the support and the support fuel injection point are used for improving the mixing combustion efficiency of the incoming flow and providing a part of fuel; the isolation section is used for stably isolating and distributing the air inlet channel and the first-stage combustion chamber, accommodating pre-combustion shock wave strings formed under the action of the pressure of the combustion chamber, preventing air flow from generating larger air flow fluctuation in the air inlet section to influence the combustion efficiency, and improving the matching stability of the air inlet channel and the rocket unit; the first-stage combustion chamber is cylindrical, and the outer diameters of vertical sections of the first-stage combustion chamber along the axial direction of the first-stage combustion chamber are the same or gradually increase at a smaller angle; the inclination of the top wall surface of the second-stage combustion chamber in the air flow injection direction ranges from (1:14) to (1: 10); the tail nozzle is in a flaring shape along the flowing direction of the airflow, and the flowing speed of the flowing airflow is improved by the tail nozzle.
Drawings
FIG. 1 is a schematic vertical cross-sectional view of a mesoscale rocket-based combined cycle engine in an ejector mode according to an embodiment of the invention;
FIG. 2 is a schematic vertical cross-sectional view of a mesoscale rocket-based combined cycle engine in a sub-burn/hyper-burn mode in an embodiment of the present invention;
FIG. 3 is a schematic structural diagram of a mesoscale rocket-based combined cycle engine in an embodiment of the invention.
FIG. 4 is a schematic structural diagram of an air inlet, a support plate rocket and an isolation section of a medium-scale rocket-based combined cycle engine in an embodiment of the invention.
FIG. 5 is a cross-sectional view of a strut rocket exit of a mesoscale rocket-based combined cycle engine in an embodiment of the invention.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and specific embodiments.
Referring to fig. 1-5, the invention provides a medium-scale rocket-based combined cycle engine, which comprises an air inlet 11, an isolation section 12, a first-stage combustion chamber 13, a second-stage combustion chamber 14, a tail nozzle 15, a first rocket unit 21, a fuel injection hole 22, a second rocket unit 23, a rocket plate 31, a plasma generator 32 and a rocket plate fuel injection point 33.
The air inlet 11, the isolation section 12, the first-stage combustion chamber 13, the second-stage combustion chamber 14 and the tail pipe 15 are connected in sequence, and specifically, the air inlet 11, the isolation section 12, the first-stage combustion chamber 13, the second-stage combustion chamber 14 and the tail pipe 15 are connected through flanges or welding. The air flow flows from the air inlet 11, passes through the isolation section 12, the first stage combustion chamber 13, and the second stage combustion chamber 14, and is discharged from the tail pipe 15, and the outflow direction of the air flow is shown in the direction of a in fig. 1.
The air inlet 11 is a variable geometry air inlet, and has the functions of improving the static temperature and static pressure of the incoming high-speed incoming flow and reducing the incoming flow speed at the same time, and the working state of the air inlet in an injection mode is shown in fig. 1; the operation in the sub-burn/super-burn mode is shown in fig. 2. A variable geometry air inlet is adopted, the injection mode adopts a two-channel configuration, the contraction ratio is 2-4, a small contraction ratio and a large throat area are formed, and sufficient air injection amount and good starting capability are ensured; the sub-burn and super-burn modes employ a single-channel configuration with a contraction ratio greater than 6, providing sufficient compression capacity for the incoming flow through a large contraction ratio.
The capture area of the air inlet 11 is 0.2-0.4 m2The second-stage combustion chamber 14 adopts secondary fuel wall injection, and fuel sprayed from the fuel injection holes is atomized and evaporated and then is combusted with oxygen in air sucked by an air inlet to form thrust.
The nozzle direction of the rocket plate 31 is parallel to the air flow direction in the air inlet 11 (refer to the direction A in figure 1).
A central support plate 34 is arranged in the square section of the inlet isolation section 12 and is used for assisting in compressing incoming air and providing a mounting position for the support plate rocket 31 and the like.
The strut rocket 31, the plasma generator 32 and the strut fuel injection point 33 are used for improving the mixing combustion efficiency of the incoming flow and providing a part of fuel.
The structure of the integrated central body consisting of the strut rocket 31, the plasma generating body 32 and the strut fuel injection points 33 is arranged from the inside to the outside in sequence with reference to fig. 5. The support 34 is also provided with a plasma generation body 32 and a support fuel injection point 33, the plasma generation body 32 is arranged around the periphery of the support rocket 31, and the support fuel injection point 33 is arranged around the periphery of the plasma generation body 32, so that the working stability and the output uniformity of the plasma generation body 32 and the support fuel injection point 33 are ensured.
In the injection mode, the support plate rocket 31 is used for sucking injection air, and the sub-combustion mode and the super-combustion mode can be used as a small-flow ignition flame stabilizing torch. The circumferential strut fuel injection points 33 on the central strut 34 can inject corresponding flow rates according to the actual requirements of different modes.
The isolation section 12 is used for stably isolating and distributing the air inlet 11 and the first-stage combustion chamber 13, accommodating pre-combustion shock wave strings formed under the action of the pressure of the combustion chamber, preventing air flow from generating large air flow fluctuation at the air inlet section to influence the combustion efficiency, and enabling air entering the first-stage combustion chamber 13 and the second-stage combustion chamber 14 to be more suitable for being combusted with rich-combustion plumes and fuel provided by the first rocket unit 21 and the fuel injection holes 22.
The first-stage combustor 13 is cylindrical, and the outer diameter of the vertical section of the first-stage combustor 13 in the axial direction thereof is the same in size or gradually increases with a small inclination (1:115) - (1: 57).
Further, the inclination of the top wall surface of the second-stage combustion chamber 14 in the air flow injection direction (see direction a in fig. 1) ranges from (1:14) to (1:10), wherein the top wall surface of the second-stage combustion chamber 14 means the top wall surface of a section in the vertical direction (see direction B in fig. 1).
The jet nozzle 15 is flared in the direction of outflow of the gas flow (indicated in the direction of a in fig. 1), and the velocity of the outflow gas flow is increased by the jet nozzle 15.
The rocket motors 31 of the first and second rocket units 21 and 23 are liquid fuel rockets, and the magnitude of the working flow rate is controlled by adjusting the flow rate of the fuel supply.
The fuel injection holes 22 and the strut fuel injection points 33 provide liquid fuel at a flow rate that can be controllably adjusted based on engine conditions.
The first rocket unit 21 is arranged at the joint of the isolation section 12 and the first-stage combustion chamber 13, the direction of a spray pipe of the first rocket unit 21 is parallel to incoming flow of an air inlet passage, the first rocket unit operates in an equivalence ratio state under an injection mode, the effect of sucking and introducing incoming flow air is mainly played, and sufficient oxygen is provided for combustion tissues of the combustion chamber; the fuel can be closed in the sub-combustion mode and the super-combustion mode, and can also work in a low-flow rich combustion state, so that the effect of ignition and stable combustion is achieved; the aircraft can work in a large flow state in a sub-combustion mode and a super-combustion mode, and the effects of improving the thrust of an engine and improving the acceleration capability and maneuverability of the aircraft are achieved.
The fuel injection holes 22 are arranged on the circumferential wall surfaces of the first-stage combustion chamber 13 and the second-stage combustion chamber 14, the fuel injection holes 22 are mainly used for providing fuel required by combustion for the second-stage combustion chamber 14 under different modes, the fuel can be kerosene or other liquid fuel, and the sprayed liquid fuel is subjected to evaporation atomization and mixing combustion in an enhancement mode such as microwave discharge.
A second rocket unit 23 is disposed at the junction of the second stage combustion chamber 14 and the jet nozzle 15. The second rocket unit 23 operates at a stoichiometric ratio, i.e. the oxidant and the fuel in the rocket react completely, the fuel gas does not burn with the air any more, but the aerodynamic profile of the flow channel is changed by high-pressure jet flow, so that the function of an aerodynamic nozzle is formed, the direction of the nozzle deviates into the flow channel, the angle is small, and the angle is selected to be 30 degrees in the example.
Specifically, the first rocket unit 21, the fuel injection hole 22 and the second rocket unit 23 work in a mutually matched mode in different states under different working modes, two groups of different rocket units in different positions are adopted, and the functions of air injection, ignition and flame stabilization and formation of a pneumatic throat are sequentially realized through the built-in rocket combinations in different states to realize matching with the working requirements of each mode of the combined engine.
Specifically, in the ejection mode, the support plate rocket 31 and the first rocket unit 21 work in a large-flow equivalence ratio state, and mainly play a role in ejecting air; the fuel injection holes 22 provide fuel required for combustion; the second rocket unit 23 works in a large-flow equivalence ratio state and plays a role in generating a pneumatic throat;
in a sub-combustion mode, the support plate rocket 31 and the first rocket unit 21 work in a low-flow rich combustion or closed state to play a role of igniting and stabilizing a flame torch; the fuel injection holes 22 are used for injecting secondary fuel required by combustion according to the required flow; the second rocket unit 23 works in a large flow state and plays a role in generating a pneumatic throat. The pneumatic throat is formed by the plume of the second rocket unit 23, the high-temperature high-pressure fuel gas in the second-stage combustion chamber is accelerated to the sound velocity through the compression of the plume of the second rocket unit 23 in the subsonic velocity state, and then is continuously accelerated to the supersonic velocity through the expansion section of the tail nozzle;
in the over-burning state, the support plate rocket 31 and the first rocket unit 21 work in a low-flow rich burning or closing state to play a role of igniting and stabilizing a flame torch; the fuel injection points 33 on the central support plate 34 and the fuel injection holes 22 on the wall surface inject secondary fuel required by combustion according to the required flow; in the super-combustion mode, the gas in the second-stage combustion chamber is combusted under supersonic flow, so that the gas is directly sprayed out through the expansion nozzle without being compressed by a pneumatic throat formed by the second rocket unit 23 when entering the tail nozzle, and the second rocket unit 23 works in a small flow or closed state.
Under different working states of injection, sub-combustion and super-combustion, the fuel injection holes 22 inject secondary fuel according to different requirements, and the secondary fuel is combusted in the secondary combustion chamber 14 to form thrust; in the injection and sub-combustion modes, the second rocket unit 23 operates in a stoichiometric state to form a pneumatic throat and simultaneously generate thrust.
In conclusion, the invention provides a medium-scale rocket-based combined cycle engine, which comprises an air inlet, a support plate rocket, an isolation section, a first-stage combustion chamber, a second-stage combustion chamber, a tail nozzle, a first rocket unit, a fuel injection hole and a second rocket unit, wherein the air inlet, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are sequentially connected, and air flow flows into the isolation section, the first-stage combustion chamber and the second-stage combustion chamber from the air inlet and then is discharged outwards from the tail nozzle; the air inlet is used for decelerating and boosting the airflow flowing in from the air inlet; the support plate rocket is arranged on the support plate in the air inlet channel; the isolation section is used for stably isolating and distributing the air inlet channel and the first-stage combustion chamber, accommodating pre-combustion shock wave strings formed under the action of the pressure of the combustion chamber, preventing air flow from generating larger air flow fluctuation in the air inlet section to influence the combustion efficiency, and improving the matching stability of the air inlet channel and the rocket unit; the first-stage combustion chamber is cylindrical, and the outer diameters of vertical sections of the first-stage combustion chamber along the axial direction of the first-stage combustion chamber are the same or gradually increase at a smaller angle; the tail nozzle is in a flaring shape along the flowing direction of the airflow, and the flowing speed of the flowing airflow is improved by the tail nozzle. The first rocket unit 21, the fuel injection holes 22 and the second rocket unit 23 are arranged to work in different states under different working modes, a plurality of groups of different rocket units and fuel injection holes at different positions are adopted, and the functions of air injection, ignition and flame stabilization and formation of a pneumatic throat are sequentially realized by the built-in rocket combination in different states to match with the working requirements of each mode of the combined engine.
Finally, it should be noted that: the above examples are only intended to illustrate the technical solution of the present invention, but not to limit it; although the present invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equivalently replaced; and such modifications or substitutions do not depart from the spirit and scope of the corresponding technical solutions of the embodiments of the present invention.
Claims (9)
1. A mesoscale rocket based combined cycle engine comprising: intake duct, extension board rocket, isolation section, first order combustion chamber, second level combustion chamber, exhaust nozzle, first rocket unit, fuel injection hole and second rocket unit, wherein:
the air inlet channel, the isolation section, the first-stage combustion chamber, the second-stage combustion chamber and the tail nozzle are sequentially connected, air flows in from the air inlet channel, and is discharged outwards from the tail nozzle after passing through the isolation section, the first-stage combustion chamber and the second-stage combustion chamber;
the first rocket unit is arranged at the joint of the isolation section and the first-stage combustion chamber, the fuel injection hole is arranged at the joint of the first-stage combustion chamber and the second-stage combustion chamber, and the second rocket unit is arranged at the joint of the second-stage combustion chamber and the tail nozzle;
the first-stage combustion chamber is cylindrical, the outer diameter of the vertical section of the first-stage combustion chamber in the axial direction of the first-stage combustion chamber is the same, and the first-stage combustion chamber is used for completing rapid mixing between rocket jet flow and jet air under an injection mode and efficient combustion under a super-combustion mode;
the tail nozzle is in a flaring shape along the flowing direction of the airflow, and the flowing speed of the flowing airflow is improved by the tail nozzle;
the support plate rocket is arranged on a central support plate in the isolation section, a plasma generation body and support plate fuel injection points are further arranged on the central support plate, the plasma generation body is arranged around the periphery of the support plate rocket, and the support plate fuel injection points are arranged around the periphery of the plasma generation body.
2. The mesoscale rocket based combined cycle engine defined in claim 1, wherein said inlet duct, said isolated section, said first stage combustion chamber, said second stage combustion chamber and said jet nozzle are connected by flange connections or by welding.
3. The mesoscale rocket based combined cycle engine defined in claim 1, wherein said air scoop is a variable structure air scoop.
4. The mesoscale rocket based combined cycle engine defined in claim 1, wherein said rocket nozzle direction is parallel to the air flow direction in the air intake.
5. The mesoscale rocket based combined cycle engine as recited in claim 1, wherein said first rocket unit is adapted to introduce airflow from said intake port into said first stage combustion chamber or said second stage combustion chamber.
6. The mesoscale rocket based combined cycle engine as recited in claim 1, wherein said fuel injection orifices are for providing combustion gases to said first stage combustion chamber and said second stage combustion chamber.
7. The mesoscale rocket based combined cycle engine as recited in claim 1, wherein said first rocket unit and said second rocket unit are liquid fuel rockets.
8. The mesoscale rocket based combined cycle engine defined in claim 1, wherein the inclination of the top wall surface of said second stage combustion chamber in the direction of jet flow is in the range of (1:14) - (1: 10).
9. The mesoscale rocket based combined cycle engine as recited in claim 1, wherein said air intake has a capture area in the range of 0.2-0.4 m2。
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010367767.3A CN111594346A (en) | 2020-05-01 | 2020-05-01 | Mesoscale rocket-based combined cycle engine |
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| CN112816218A (en) * | 2021-01-07 | 2021-05-18 | 中国人民解放军国防科技大学 | Rich fuel gas generation control system and control method |
| CN113944568A (en) * | 2021-10-13 | 2022-01-18 | 华东理工大学 | Powder fuel support plate ejection rocket-based combined cycle engine based on HAN monopropellant |
| CN114215653A (en) * | 2021-12-13 | 2022-03-22 | 哈尔滨工业大学 | Combustion mode conversion method for ramjet engine |
| CN115419521A (en) * | 2022-08-23 | 2022-12-02 | 西北工业大学 | Double-component support plate rocket device for rocket-based combined cycle engine |
| CN115434826A (en) * | 2022-08-23 | 2022-12-06 | 西北工业大学 | Embedded support plate rocket device for rocket-based combined cycle engine |
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| CN113944568B (en) * | 2021-10-13 | 2023-09-05 | 华东理工大学 | HAN unit propellant-based powder fuel support plate rocket-based combined cycle engine |
| CN114215653A (en) * | 2021-12-13 | 2022-03-22 | 哈尔滨工业大学 | Combustion mode conversion method for ramjet engine |
| CN115419521A (en) * | 2022-08-23 | 2022-12-02 | 西北工业大学 | Double-component support plate rocket device for rocket-based combined cycle engine |
| CN115434826A (en) * | 2022-08-23 | 2022-12-06 | 西北工业大学 | Embedded support plate rocket device for rocket-based combined cycle engine |
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