CN111594344A - Small-scale two-stage rocket combined ramjet engine - Google Patents
Small-scale two-stage rocket combined ramjet engine Download PDFInfo
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- CN111594344A CN111594344A CN202010367625.7A CN202010367625A CN111594344A CN 111594344 A CN111594344 A CN 111594344A CN 202010367625 A CN202010367625 A CN 202010367625A CN 111594344 A CN111594344 A CN 111594344A
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- 238000002485 combustion reaction Methods 0.000 claims abstract description 146
- 238000002347 injection Methods 0.000 claims abstract description 43
- 239000007924 injection Substances 0.000 claims abstract description 43
- 239000000446 fuel Substances 0.000 claims abstract description 41
- 238000002955 isolation Methods 0.000 claims abstract description 27
- 239000007788 liquid Substances 0.000 claims description 6
- 235000015842 Hesperis Nutrition 0.000 claims description 4
- 235000012633 Iberis amara Nutrition 0.000 claims description 4
- 238000002156 mixing Methods 0.000 claims description 4
- 238000003466 welding Methods 0.000 claims description 3
- 239000000567 combustion gas Substances 0.000 claims description 2
- 150000001875 compounds Chemical class 0.000 claims 2
- 230000000694 effects Effects 0.000 abstract description 4
- 230000035939 shock Effects 0.000 abstract description 4
- 230000008859 change Effects 0.000 abstract description 2
- 108091053398 TRIM/RBCC family Proteins 0.000 abstract 1
- 102000011408 Tripartite Motif Proteins Human genes 0.000 abstract 1
- 239000007789 gas Substances 0.000 description 5
- 230000008602 contraction Effects 0.000 description 4
- 230000009471 action 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
- 230000001133 acceleration Effects 0.000 description 2
- QVGXLLKOCUKJST-UHFFFAOYSA-N atomic oxygen Chemical compound [O] QVGXLLKOCUKJST-UHFFFAOYSA-N 0.000 description 2
- 230000015572 biosynthetic process Effects 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 239000002737 fuel gas Substances 0.000 description 2
- 239000001301 oxygen Substances 0.000 description 2
- 229910052760 oxygen Inorganic materials 0.000 description 2
- 230000003068 static effect Effects 0.000 description 2
- 238000000889 atomisation Methods 0.000 description 1
- 230000002708 enhancing effect 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
- 230000008520 organization Effects 0.000 description 1
- 239000007800 oxidant agent Substances 0.000 description 1
- 230000001590 oxidative effect Effects 0.000 description 1
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- 230000006641 stabilisation Effects 0.000 description 1
- 238000011105 stabilization Methods 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
- 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 small-scale two-stage rocket combined ramjet engine which comprises an air inlet, 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 channel provides required capture air for the RBCC engine through the change of the molded surface under different flight Mach numbers and working modes; 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 small-scale two-stage rocket combined ramjet 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 the rocket layout and the ramjet runner layout is the key technology in research.
At present, the commonly used RBCC engine structural scheme is 'variable geometry air inlet channel + middle/side rocket or liquid fuel injection + fixed geometry second stage combustion chamber + variable geometry tail nozzle', for the RBCC engine with the air inlet channel capture area less than 0.1m2, if the built-in rocket is arranged at the front end of the engine, when the built-in rocket works in a large-flow high-chamber pressure state to contribute to the main thrust of the engine, the starting capability of the air inlet channel in an injection mode can be seriously influenced, and the thermal congestion generation difficulty in the combustion chamber is higher under a lower flight Mach number; if the built-in rocket is arranged at the tail part of the engine, the built-in rocket can play a role of boosting, but the air injection suction capacity provided under the injection mode is greatly reduced. In order to improve the working efficiency of the built-in rocket, a second-stage rocket can be arranged in the RBCC engine, and the two stages of rockets arranged at different positions work in a matching way and are respectively and pertinently used for effectively forming air injection suction and thermal congestion, so that the injection modal performance, the starting capability and the acceleration capability of the whole engine are comprehensively improved.
Disclosure of Invention
In view of the above, the technical problems to be solved by the present invention are: how to provide a small-scale rocket-based combined cycle engine which can improve the comprehensive working performance of the engine by arranging a secondary rocket.
In order to achieve the above object, the present invention provides a small-scale two-stage rocket combined ramjet engine, comprising: intake duct, isolation section, first stage combustion chamber, second stage combustion chamber, exhaust nozzle, first rocket unit, fuel injection hole 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 holes are annularly distributed at the joint of the first-stage combustion chamber and the second-stage combustion chamber after the outlet of the first rocket unit rocket nozzle, and the second rocket unit is arranged at the joint of the first-stage combustion chamber and the second-stage combustion chamber and 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 quick mixing between rocket jet flow and jet air in an injection mode and efficient combustion in 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.
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.
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, the second rocket unit is used for providing gas for the second stage combustion chamber.
Further, the first rocket unit and the second rocket unit are liquid fuel rockets.
Further, the length dimension of the second-stage combustion chamber in the air flow injection direction ranges from 500-.
Further, the capture area of the air inlet is less than 0.1m2。
Compared with the prior art, the invention provides a rocket-based combined cycle engine, which comprises an air inlet, 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 channel provides required capture air for the RBCC engine through the change of the molded surface under different flight Mach numbers and working modes; the isolation section is used for stably isolating and distributing the air inlet channel and the first-stage combustion chamber and accommodating pre-combustion shock wave strings formed under the action of the pressure of the combustion chamber, so that the air flow is prevented from generating large air flow fluctuation in the air inlet section to further influence the combustion efficiency, and the matching stability of the air inlet channel and the rocket unit is improved; 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 or gradually increases with a smaller angle, and the first-stage combustion chamber is mainly used for completing quick mixing between rocket jet flow and jet air in an injection mode and efficient combustion in 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 two rocket units and a series of fuel injection holes are matched with each other in different states under different working modes, a plurality of groups of different rocket units and fuel injection holes in different positions are adopted, and the built-in rocket combinations in different states sequentially play roles of injection and pneumatic throat formation and fuel supply through the fuel injection holes to realize matching with working requirements of various modes of the combined engine.
Drawings
FIG. 1 is a schematic vertical cross-sectional view of a rocket-based combined cycle engine in an ejector mode according to an embodiment of the present invention;
FIG. 2 is a schematic vertical cross-sectional view of a rocket-based combined cycle engine in a sub-combustion/super-combustion mode in accordance with an embodiment of the present invention;
FIG. 3 is a schematic view of a rocket-based combined cycle engine according to an embodiment of the present invention.
Detailed Description
The present invention will be described in further detail with reference to the accompanying drawings and specific embodiments.
Referring to fig. 1-3, the present invention provides a rocket-based combined cycle engine, which includes an intake port 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, and a second rocket unit 23.
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, the small contraction ratio and the large throat area are ensured, and the larger air injection amount and the better 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 less than 0.1m2The first-stage combustion chamber 13 and the second-stage combustion chamber 14 adopt secondary fuel injection, and the 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 isolation section 12 is used for stably isolating and arranging 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 the 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 provided by the first rocket unit 21 and the second rocket unit 23.
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 or gradually increases in the linear direction at an 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 length dimension of the second-stage combustion chamber in the air flow injection direction ranges from 500 mm to 800 mm.
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 first rocket unit 21 and the second rocket unit 23 are liquid fuel rockets, and the magnitude of the operating flow rate thereof is controlled by adjusting the flow rate of the fuel supply.
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 valve can be closed in a sub-combustion mode, and can also work in a low-flow rich combustion state to play a role in igniting and stabilizing combustion; the device works in a rich combustion state in a super-combustion mode, and provides rich combustion gas with required flow rate for combustion organization; 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.
The second rocket unit 23 is disposed behind the nozzle exit of the first rocket unit 21 and at the junction of the second stage combustion chamber 14 and the jet nozzle 15. The fuel injection holes 22 are annularly arranged along the axial direction of the first-stage combustion chamber 13 and the second-stage combustion chamber 14, so that the uniformity of fuel injection can be ensured. The second rocket unit 23 operates at a stoichiometric ratio, that is, the oxidant and the fuel in the rocket completely react, the fuel gas does not continuously combust with the air any more, but the aerodynamic profile of the flow channel is changed by high-pressure jet flow, so that the pneumatic jet pipe is formed, the boosting effect is achieved, the direction of the jet pipe deviates into the flow channel, the angle is small, and the angle is 15 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, pneumatic throat and thrust lifting are sequentially realized through the built-in rocket combinations in different states to be matched with the working requirements of each mode of the combined engine.
Specifically, in the ejection mode, the first rocket unit 21 works in a large-flow equivalence ratio state, and mainly plays 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 and boosting;
in the sub-combustion mode, the first rocket unit 21 works 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 operates in a large flow state, and plays a role in generating a pneumatic throat and enhancing thrust. 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 state of super-combustion, the first rocket unit 21 works in a rich combustion state with large flow or small flow to play a role of igniting and stabilizing a flame torch; the fuel injection holes 22 inject secondary fuel required for 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, but the second rocket unit 23 can work at a certain flow rate to play a role in thrust enhancement.
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 rocket-based combined cycle engine, which comprises an air inlet, 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 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 injection, ignition and flame stabilization, auxiliary formation of a pneumatic throat and thrust enhancement are sequentially realized through 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 (8)
1. A small-scale two-stage rocket combined ramjet engine is characterized by comprising: intake duct, isolation section, first stage combustion chamber, second stage combustion chamber, exhaust nozzle, first rocket unit, fuel injection hole, 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 holes are annularly distributed at the joint of the first-stage combustion chamber and the second-stage combustion chamber after the outlet of the first rocket unit rocket nozzle, and the second rocket unit is arranged at the joint of the first-stage combustion chamber and the second-stage combustion chamber and 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 quick mixing between rocket jet flow and jet air in an injection mode and efficient combustion in 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.
2. A small scale two-stage rocket combination ramjet engine as recited 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 connection or welding.
3. A small scale two-stage rocket combination ramjet engine as recited in claim 1, wherein said air scoop is a variable structure air scoop.
4. A small scale two-stage rocket combination ramjet engine as recited in claim 1, wherein said first rocket unit is adapted to introduce air flow from said air intake into said first stage combustion chamber or said second stage combustion chamber.
5. A small scale two-stage rocket combination ramjet as recited in claim 1, wherein said second rocket unit is adapted to provide combustion gases to said second stage combustion chamber.
6. A small scale two-stage rocket compound ramjet engine as recited in claim 1, wherein said first rocket unit and said second rocket unit are liquid fuel rockets.
7. The small-scale two-stage rocket combination ramjet engine according to claim 1, wherein the length dimension of said second-stage combustion chamber in the air jet direction is in the range of 500-800mm, and the inclination of the top wall surface of said second-stage combustion chamber in the air jet direction is in the range of (1:14) - (1: 10).
8. A small scale two-stage rocket compound ramjet engine as recited in claim 1, wherein said air scoop has a capture area of less than 0.1m2。
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Cited By (7)
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| CN112345529A (en) * | 2021-01-11 | 2021-02-09 | 中国人民解放军国防科技大学 | Image processing-based experimental test system and method for rectangular thermal state temperature difference mixing layer |
| CN112627983A (en) * | 2020-12-25 | 2021-04-09 | 中国人民解放军国防科技大学 | RBCC engine inner flow channel and RBCC engine |
| CN113090416A (en) * | 2021-04-27 | 2021-07-09 | 西北工业大学 | Simulation experiment device for rocket stamping combined air inlet channel |
| CN113188747A (en) * | 2021-04-28 | 2021-07-30 | 中山大学 | Variable mach number supersonic wind tunnel |
| CN114109650A (en) * | 2021-10-27 | 2022-03-01 | 厦门大学 | Integral liquid rocket punching combined power device |
| CN114941582A (en) * | 2022-03-18 | 2022-08-26 | 华中科技大学 | RBCC ejection rocket adopting multi-thrust-chamber engine and control method thereof |
| CN115434823A (en) * | 2022-08-31 | 2022-12-06 | 西安航天动力研究所 | Rocket stamping combined engine with parallel compressor runners |
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Application publication date: 20200828 |