CN108757179A - Combined cycle engine and hypersonic aircraft - Google Patents
Combined cycle engine and hypersonic aircraft Download PDFInfo
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- CN108757179A CN108757179A CN201810529474.3A CN201810529474A CN108757179A CN 108757179 A CN108757179 A CN 108757179A CN 201810529474 A CN201810529474 A CN 201810529474A CN 108757179 A CN108757179 A CN 108757179A
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- shell
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- combustion chamber
- fuel
- drainage channel
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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
- 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
- F02C3/32—Inducing air flow by fluid jet, e.g. ejector action
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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
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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
- F02C7/12—Cooling of plants
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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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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Fuel-Injection Apparatus (AREA)
Abstract
The invention discloses a combined cycle engine and a hypersonic aircraft, which comprise a rotary knocking ramjet engine and a rocket engine arranged in the rotary knocking ramjet engine. A rotary detonation ramjet engine comprising: the shell is in a hollow cylinder shape with two communicated ends. The rear body of the central cone extends into the shell from the air inlet end of the shell and is connected with the shell, a drainage channel for introducing air into the shell is formed in a gap between the shell and the rear body, a detonation chamber communicated with the drainage channel is formed by a cavity between the rear end face of the rear body and the inner wall of the shell, and a first spray pipe communicated with the detonation chamber. The rocket engine includes: the combustion chamber is arranged in the central cone, the second spray pipe is communicated with the combustion chamber, and a nozzle of the second spray pipe is communicated with the detonation chamber. The outer wall of the shell is provided with an outer nozzle, and two ends of the outer nozzle are respectively communicated with the fuel supply device and the drainage channel. And/or a plurality of inner nozzles are processed on the outer wall of the rear body, and two ends of each inner nozzle are respectively communicated with the fuel supply device and the drainage channel.
Description
Technical field
The present invention relates to aero-engine fields, particularly, are related to a kind of combined cycle engine.In addition, the present invention is also
It is related to a kind of hypersonic aircraft including said combination cycle engine.
Background technology
Hypersonic aircraft flying speed is more than 5 times of velocities of sound, flying height usually at 18-35 kilometers, be have it is long-range,
Quickly, novel " trump card " weapon of precision strike capability is the focus of the current competition of aerospace field in the world, core
It is hypersonic Push Technology.Scramjet engine is a kind of hypersonic propulsion device being widely studied at present,
It is burnt with isobaric pattern tissue.Continuous rotation pinking punching engine is a kind of novel punching promotion program proposed in the recent period, with
Detonation mode tissue burns, and has the higher efficiency of cycle compared with isobaric combustion, it is expected to provide performance for hypersonic aircraft
Higher power plant.Document【Wang Chao,Liu Weidong,Liu Shijie,Jiang Luxin,Lin
Zhiyong.Experimental Verification of Air-breathing Continuous Rotating
Detonation Fueled by Hydrogen,International Journal of Hydrogen Energy,2015,
40:9530-9538.】With【Shijie Liu,Weidong Liu,Yi Wang,Zhiyong Lin,Free Jet Test of
Continuous Rotating Detonation Ramjet Engine,AIAA 2017-2282.】Respectively by it is direct-connected with
Free jet test demonstrates the feasibility of continuous rotation pinking punching engine.
Although the key technology of scramjet engine has been broken through at present, continuous rotation pinking punching engine can
Row has been verified, but these engines all can only operate in punching press mode.To realize that zero velocity starts, hypersonic patrols
Boat flight, it is necessary to punching engine and other power are combined, combined engine technical research is carried out.Two kinds of common groups
Conjunction form is rocket base combined engine (RBCC) and turbine base combined engine (TBCC), and two kinds of assembled schemes respectively have quality.
RBCC is organically blended to rocket engine technology and Ramjet Technique, and punching engine high specific impulse is had both
With the advantage of the high thrust-weight ratio of rocket engine, have All Speed Range, full spatial domain ability to work.The original intention of RBCC is as Single Stage To Orbit
Dynamical system, but can also be used for zero velocity startup, hypersonic cruising flight dynamical system.
Traditional RBCC engines are mainly made of Rocket ejector and punching engine runner, can as needed be pacified rocket
Suitable position in runner.For hypersonic cruise RBCC engines, three kinds of operation modes can be divided into, as shown in Figure 1,
As Ma=0-3, it is operated in Ejector Mode.Under injection and speed punching press double action, air enters flow channel, in injection
Mixed zone is blended with the high-temperature fuel gas that Rocket ejector generates, and fuel is sprayed into mixed gas and carries out secondary afterburning, Rocket ejector
Thrust can be generated with the afterburning in downstream.As shown in Fig. 2, as Ma=3-5, RBCC is operated in sub- combustion mode, closes at this time
The high stagnation pressure subsonic airflow of Rocket ejector, air intake duct capture is restored in combustion chamber upstream pressure, is sprayed in combustion chamber downstream
It notes fuel and carries out subsonic combustion, final combustion product is discharged by jet pipe.As shown in figure 3, as Ma5 or more, engine
It is transferred to ultra-combustion ramjet mode, still closes Rocket ejector at this time, the supersonic flow of air intake duct capture enters combustion chamber, in combustion chamber
Upstream position spray fuel is to realize that supersonic combustion, final combustion product are discharged by jet pipe.
Traditional RBCC engine operation modes are more, and mode conversion difficulty is big；RBCC engines are organized in an isobaric manner
Burning, the efficiency of cycle is low, thrust poor performance；In addition, since Ejector Mode needs injection mixed zone, isobaric combustion heat release rate
Slowly, therefore longer combustion chamber is needed, so that engine entire length is larger.
Invention content
The present invention provides a kind of combined cycle engine and hypersonic aircrafts, to solve traditional RBCC engines
Operation mode is more, mode conversion difficulty is big, the efficiency of cycle is low, the technical problem of thrust poor performance and chamber length length.
The technical solution adopted by the present invention is as follows：
A kind of combined cycle engine includes the rocket engine of rotation pinking punching engine and setting in the inner；Rotation
Turning pinking punching engine includes：Shell, shell are in the hollow tubular of both ends connection；Conical inner body, the rear body of conical inner body by
The inlet end of shell stretches into shell and is connected with shell, and the gap between shell and rear body forms and introduced in shell for air
Drainage channel, cavity between rear body rear end face and outer casing inner wall forms the detonation chamber being connected to drainage channel, and with it is quick-fried
Shake the first jet pipe of room connection, the spout and atmosphere of the first jet pipe；Rocket engine includes：It is set in conical inner body
Combustion chamber, with the second jet pipe of combustion chamber, the spout of the second jet pipe be connected to detonation chamber, combustion chamber and is used to provide oxidation
The oxidation agent feeding device of agent and the fuel supply system for providing fuel are respectively communicated with；It is machined with along it on the outer wall of shell
Circumferential multiple outer spouts spaced successively, outer spout both ends are connected to fuel supply system and drainage channel respectively, so that
Fuel is sprayed by multiple outer spouts in drainage channel；And/or it is machined on the outer wall of rear body circumferential spaced successively along it
Multiple interior spouts, interior spout both ends are connected to fuel supply system and drainage channel respectively, so that fuel is sprayed by multiple interior spouts
Enter in drainage channel.
Further, outlet 10mm~20mm of each outer spout away from drainage channel；And/or each interior spout is away from drainage channel
Export 10mm~20mm.
Further, multiple outer spouts and multiple interior spouts are arranged in a one-to-one correspondence or multiple outer spouts and multiple interior sprays
Mouth misses one another arrangement.
Further, the outer wall of the first jet pipe and detonation chamber is equipped with the first wall cooling channel, and the first wall cooling is logical
Road is connected with fuel supply system；Multiple outer spouts are connected to the first wall cooling channel respectively, for cooling first jet pipe and
Fuel after detonation chamber is sprayed by multiple outer spouts in drainage channel.
Further, the second jet pipe and the outer wall of combustion chamber are equipped with the second wall cooling channel, and the second wall cooling is logical
Road is connected with fuel supply system；Multiple interior spouts and the head spout of combustion chamber are connected to the second wall cooling channel respectively,
So that for the fuel behind cooling second jet pipe and combustion chamber, a part is sprayed by multiple interior spouts in drainage channel, another part
It is sprayed into combustion chamber by the head of combustion chamber.
Further, detonation chamber is in straight-tube shape；The length of detonation chamber is 400mm~800mm.
Further, it further includes that more for making rear body and shell be connected connect rib to rotate pinking punching engine,
Circumferential direction successively arranged for interval, and the inner wall phase of the upper surface and shell of each connection rib of the more connection ribs along drainage channel
Even, each lower face for connecting rib is connected with the outer wall of rear body.
Further, conical inner body further includes the precursor being connected in an axial direction with rear body, and precursor is located at outside shell, and precursor
Outside wall surface constitutes the air intake duct for being compressed to air；Combustion chamber is located in precursor；The outlet of second jet pipe and combustion chamber
It is connected, and the second jet pipe is connected to along axially extending to for conical inner body with the rear end face of rear body.
Further, it is equipped in conical inner body and the oxidant flow channel of water conservancy diversion is carried out, for being led to fuel to oxidant
The fuel flow channel of stream；The flow inlet of oxidant flow channel is connected with oxidation agent feeding device, the outflow port of oxidant flow channel and burning
The head of room is connected；The flow inlet of fuel flow channel is connected with fuel supply system, and the outflow port of fuel flow channel and the second wall surface are cold
But channel is connected.
According to another aspect of the present invention, a kind of hypersonic aircraft is additionally provided, includes the group such as any of the above-described
Close cycle engine.
The invention has the advantages that：
The combined cycle engine of the present invention uses conventional rocket engine and rotation pinking punching engine is combined
Scheme, wherein rocket engine be located at rotation pinking ramjet engine air inlet conical inner body in, with conventional RBCC engines
It compares, combined cycle engine overall structure of the invention is simpler, compact.In the combined cycle engine of the present invention, rotation
Turn pinking punching engine to rotate the burning of detonation mode tissue, and conventional RBCC engines are burnt with isobaric pattern tissue,
Compared with conventional RBCC engines, cycle of engine thermal efficiency higher of the invention, thrust performance are more excellent, remote to be more suitable for
The hypersonic cruising flight of journey, economic performance are more excellent；In addition, for conventional RBCC, since Ejector Mode needs injection to mix
Area, and isobaric combustion heat release rate is slow, therefore longer combustion chamber is needed, so that engine entire length is larger.And this hair
In bright, due to detonating combustion, heat release rate is fast, short to the length of required combustion chamber, and then the entirety of the engine of the present invention is long
It spends shorter.Traditional RBCC engine operation modes are more (including Ejector Mode, sub- combustion punching press mode and ultra-combustion ramjet mode),
Mode conversion difficulty is big, and the application engine operation mode is less (including Ejector Mode and punching press mode), and turns between mode
It is small to change difficulty.In the prior art, rotation detonation engine is all made of the combustion chamber of circular ring shape, and in application scheme, due to outer
Shell is in the hollow tubular of both ends connection, and the only rear body of conical inner body stretches into shell, so detonation chamber is in hollow tubular, compared to circle
The combustion chamber of the combustion chamber of annular, tubular has stronger detonating combustion organizational capacity, the flameholding performance of flame good；
The hypersonic aircraft operation mode of the present invention is small, mode conversion difficulty is small, the efficiency of cycle is high, thrust performance
Good and chamber length is short.
Other than objects, features and advantages described above, the present invention also has other objects, features and advantages.
Below with reference to figure, the present invention is described in further detail.
Description of the drawings
The attached drawing constituted part of this application is used to provide further understanding of the present invention, schematic reality of the invention
Example and its explanation are applied for explaining the present invention, is not constituted improper limitations of the present invention.In the accompanying drawings：
Fig. 1 is existing RBCC engines Ejector Mode operating diagram；
Fig. 2 is existing RBCC engines Asia combustion punching press mode operating diagram；
Fig. 3 is existing RBCC engines ultra-combustion ramjet Modality work schematic diagram；
Fig. 4 is the schematic cross-sectional view of the combined cycle engine of the preferred embodiment of the present invention；
Fig. 5 is combined cycle engine Propellant Supply schematic diagram in Fig. 4.
Marginal data
10, pinking punching engine is rotated；101, drainage channel；11, shell；1110, outer spout；112, detonation chamber；
113, the first jet pipe；12, conical inner body；121, rear body；1210, interior spout；122, precursor；13, rib is connected；20, rocket is sent out
Motivation；21, combustion chamber；22, the second jet pipe；30, agent feeding device is aoxidized；40, fuel supply system.
Specific implementation mode
The embodiment of the present invention is described in detail below in conjunction with attached drawing, but the present invention can be defined by the claims
Implement with the multitude of different ways of covering.
With reference to Fig. 4, the preferred embodiment of the present invention provides a kind of combined cycle engine, including rotation pinking punching press hair
The rocket engine 20 of motivation 10 and setting in the inner.Rotating pinking punching engine 10 includes：Shell 11, shell 11 are in both ends
The hollow tubular of connection.Further include conical inner body 12, the rear body 121 of conical inner body 12 stretches into shell 11 by the inlet end of shell 11
It is interior and be connected with shell 11, and the gap between shell 11 and rear body 121 forms the drainage channel introduced for air in shell 11
101, the cavity between 121 rear end face of rear body and 11 inner wall of shell forms the detonation chamber 112 being connected to drainage channel 101, and
The first jet pipe 113 being connected to detonation chamber 112, the spout and atmosphere of the first jet pipe 113.Rocket engine 20 includes：If
The combustion chamber 21 being placed in conical inner body 12, the second jet pipe 22 being connected to combustion chamber 21, the spout of the second jet pipe 22 and pinking
Room 112 is connected to, and combustion chamber 21 is supplied with the oxidation agent feeding device 30 for providing oxidant and the fuel for providing fuel
Device 40 is respectively communicated with.It is machined on the outer wall of shell 11 along its circumferential multiple outer spout 1110 spaced successively, outer spray
1110 both ends of mouth are connected to fuel supply system 40 and drainage channel 101 respectively, so that fuel is sprayed by multiple outer spouts 1110
In drainage channel 101.And/or be machined with along its circumferential multiple interior spout 1210 spaced successively on the outer wall of rear body 121,
1210 both ends of interior spout are connected to fuel supply system 40 and drainage channel 101 respectively, so that fuel is by multiple interior spouts 1210
It sprays into drainage channel 101.
There are two types of operation modes for the combined cycle engine tool of the present invention：When flying speed is less than Ma2.5, which follows
Ring engine is operated in Ejector Mode, needs to open rocket engine 20 at this time, oxidant entrained by combined cycle engine and
Fuel respectively enters combustion chamber 21 and carries out isobaric combustion, and high temperature and pressure combustion product is quickly discharged from the second jet pipe 22.In rocket
Under the ejector action of engine 20, incoming air enters drainage channel after the air intake duct that 12 precursor 122 of conical inner body is constituted
In 101, and enter detonation chamber 112 with after the fuel of outer spout 1110 and/or the ejection of interior spout 1210 quickly mixes, with rotation
Detonation mode combustion heat release is simultaneously quickly blended with 20 combustion product of rocket engine, and the combustion product after blending is through the first jet pipe
113 Accelerating Removals, to generate thrust.When flying speed is more than Ma2.5, combined cycle engine is operated in punching press mode,
Rocket engine 20 is closed at this time, and rotation pinking punching engine 10 works independently.Incoming air enters after air intake duct and draws
In circulation road 101, and enter detonation chamber 112 with after the fuel of outer spout 1110 and/or the ejection of interior spout 1210 quickly mixes,
To rotate detonation mode combustion heat release in detonation chamber 112, combustion product is pushed away through 113 Accelerating Removal of the first jet pipe to generate
Power, aircraft accelerates flight under punching press mode, ramps up until Ma6, finally realizes hypersonic cruising flight.
The combined cycle engine of the present invention uses 10 phase of conventional rocket engine 20 and rotation pinking punching engine
The scheme of combination, wherein rocket engine 20 is located in 10 air inlet passage center cone 12 of rotation pinking punching engine, with routine
RBCC engines are compared, and combined cycle engine overall structure of the invention is simpler, compact.The combination cycle hair of the present invention
In motivation, rotation pinking punching engine 10 is to rotate the burning of detonation mode tissue, and conventional RBCC engines are to wait pressing molds
Formula tissue burns, and compared with conventional RBCC engines, cycle of engine thermal efficiency higher of the invention, thrust performance are more excellent,
To be more suitable for long-range hypersonic cruising flight, economic performance is more excellent；In addition, since Ejector Mode needs injection mixed zone,
And isobaric combustion heat release rate is slow, therefore routine RBCC needs longer combustion chamber, so that engine entire length is larger.And
In the present invention, due to detonating combustion, heat release rate is fast, short to the length of required combustion chamber, and then the engine of the present invention is whole
Body length is shorter.Traditional RBCC engine operation modes more (including Ejector Mode, sub- combustion punching press mode and ultra-combustion ramjet mould
State), mode conversion difficulty is big, and the application engine operation mode is less (including Ejector Mode and punching press mode), and mode
Between conversion difficulty it is small.In the prior art, rotation detonation engine is all made of the combustion chamber of circular ring shape, and in application scheme, by
In shell 11 in the hollow tubular of both ends connection, and only body 121 stretches into shell 11 conical inner body 12 afterwards, so detonation chamber 112
In hollow tubular, the combustion chamber of circular ring shape is compared, the combustion chamber of tubular has stronger detonating combustion organizational capacity, the combustion of flame
It is good to burn stability.
Optionally, as shown in figure 4, being machined with multiple outer spouts 1110 on the outer wall of shell 11, each outer spout 1110 is away from drawing
Outlet 10mm~20mm of circulation road 101.And/or it is machined with multiple interior spouts 1210, each interior spout on the outer wall of rear body 121
1210 away from the outlet of drainage channel 101 10mm~20mm.In the present invention, the gap between shell 11 and rear body 121 forms drainage
Channel 101, drainage channel 101 are applied not only to enter in shell 11 for air conduction, the high back-pressure being additionally operable in isolation detonation chamber 112
Influence to the air intake duct formed by the precursor 122 of conical inner body 12 before drainage channel 101, and also make by outer spout
1110 and/or interior spout 1210 fuel and the air that spray quickly, be sufficiently mixed after spray into again in detonation chamber 112.So outer spray
Mouth 1110 and/or interior spout 1210 should be located at the upstream that drainage channel 101 exports, and outer spout 1110 and/or interior spout 1210
20mm cannot be more than away from the distance that drainage channel 101 exports, otherwise the high back-pressure in detonation chamber 112 is easy to draw in channel 101
Combustion rotation pinking, that is, occur " tempering " phenomenon, to be had an impact to air intake duct；But outer spout 1110 and/or interior spout 1210
10mm cannot be less than away from the distance that drainage channel 101 exports, otherwise fuel and air cannot be fully mixed before entering detonation chamber 112
It closes, and then influences the pinking in detonation chamber 112 and occur.When each outer spout 1110 and/or each interior spout are away from drainage channel 101
Outlet is 10mm~20mm, and fuel and air spray into after capable of being sufficiently mixed in drainage channel 101 in detonation chamber 112 again, and
It not will produce tempering phenomenon.
In the specific embodiment of the invention, as shown in figure 4, multiple outer spouts 1110 are machined on the outer wall of shell 11, and after
Multiple interior spouts 1210 are machined on the outer wall of body 121, multiple outer spouts 1110 and multiple interior spouts 1210 are arranged in a one-to-one correspondence,
Or multiple outer spouts 1110 and multiple interior spouts 1210 miss one another arrangement.It is multiple outer due to being machined on the outer wall of shell 11
Spout 1110, and it is machined with multiple interior spouts 1210, and multiple outer spouts 1110 and multiple interior spouts on the outer wall of rear body 121
1210 are arranged in a one-to-one correspondence or multiple outer spouts 1110 and multiple interior spouts 1210 miss one another arrangement, to make air with
Fuel is sufficiently mixed, and the two is uniformly mixed, and is conducive to the generation that pinking is reacted in detonation chamber 112, and overcome drainage channel 101
When depth of section is big, when outer spout 1110 or the only interior spout 1210 of setting are only set, air cannot with fuel is well-mixed asks
Topic.
Preferably, the outer wall of the first jet pipe 113 and detonation chamber 112 is equipped with the first wall cooling channel (not shown), the
One wall cooling channel is connected with fuel supply system 40.Multiple outer spouts 1110 are connected to the first wall cooling channel respectively,
So that the fuel after cooling first jet pipe 113 and detonation chamber 112 is sprayed by multiple outer spouts 1110 in drainage channel 101.Pass through
First wall cooling channel is set in the outer wall of the first jet pipe 113 and detonation chamber 112, so as to introduce the first jet pipe of fuel pair
113 and detonation chamber 112 cooled down, and multiple outer spouts 1110 are connected to the first wall cooling channel respectively, to cooling
The fuel of first jet pipe 113 and detonation chamber 112 can be used as the fuels sources that fuel is provided for outer spout 1110, this kind of set-up mode,
Not only make the structure of engine simple, compact, and make full use of fuel, it is made to be used as coolant to cool down the first jet pipe 113 first
With detonation chamber 112, then it is re-used as fuel and participates in detonation chamber burning.
Preferably, the second jet pipe 22 and the outer wall of combustion chamber 21 are equipped with the second wall cooling channel (not shown), and second
Wall cooling channel is connected with fuel supply system 40.The head of multiple interior spouts 1210 and combustion chamber 21 respectively with the second wall surface
Cooling duct is connected to, so that the fuel part behind cooling second jet pipe 22 and combustion chamber 21 is drawn by the penetrating of multiple interior spouts 1210
In circulation road 101, another part is sprayed by the head of combustion chamber 21 in combustion chamber 21.By in the second jet pipe 22 and combustion chamber 21
Outer wall be arranged the second wall cooling channel, cooled down so as to introduce the second jet pipe of fuel pair 22 and combustion chamber 21, and more
The head of a interior spout 1210 and combustion chamber 21 is connected to the second wall cooling channel respectively, to 22 He of cooling second jet pipe
The fuel of combustion chamber 21 can be used as the fuels sources that fuel is provided for the head of interior spout 1210 and combustion chamber 21, this kind of setting side
Formula not only makes the structure of engine simple, compact, and makes full use of fuel, it is made to be used as coolant to cool down the second jet pipe first
22 and combustion chamber 21, it is then re-used as fuel and participates in rocket and detonation chamber burning.
In actual work, the fuel supply schematic diagram of engine is as shown in Figure 5.In the present invention, propellant includes engine
Self-contained liquid oxidant and liquid fuel, and oxidant is stored in oxidation agent feeding device 30, fuel
It is stored in fuel supply system 40.When work, oxidant enters after being pumped out by oxidation agent feeding device 30 in conical inner body 12,
And it is finally sprayed into combustion chamber 21 by the head of combustion chamber 21 by the supply runner in conical inner body 12.Fuel is supplied by fuel
Device 40 divides two-way to be supplied after pumping out, wherein entering in the first wall cooling channel cooling first jet pipe 113 and quick-fried all the way
Room 112, and heat absorption heating gasification or gasification cracking in the first wall cooling channel are shaken, when shell 11 is equipped with outer spout
When 1110, the fuel of cooling first jet pipe 113 and detonation chamber 112 is sprayed by outer spout 1110 participates in pinking in drainage channel 101
Burning in room 112.Another way fuel enters in conical inner body 12, and flows into the second wall by the supply runner in conical inner body 12
In the cooling duct of face, the wall surface of the second jet pipe 22 and combustion chamber 21 to rocket engine cools down, and absorbs heat in the channel
Heating gasification or gasification cracking, heat absorption heating after fuel by combustion chamber 21 head spray into combustion chamber 21 in, or when after
When body 121 is equipped with interior spout 1210, the fuel in the second wall cooling channel divides two-way to be supplied again, wherein entering all the way
The head of combustion chamber 21 participates in burning, and another way is sprayed into drainage channel 101 via interior spout 1210 and participated in detonation chamber 112
Burning.
Preferably, as shown in figure 4, detonation chamber 112 is in straight-tube shape.The detonation chamber of circular ring shape in compared with prior art, circle
The detonation chamber of tubular has stronger detonating combustion organizational capacity, the flameholding performance of flame good.Further, detonation chamber
112 length is 400mm~800mm.When the length of detonation chamber 112 is more than 800mm, the entire length of engine will be lengthened, it is quick-fried
When shaking the length of room 112 less than 400mm, detonating combustion is insufficient, and then influences the thrust performance of engine.
Optionally, as shown in figure 4, rotation pinking punching engine 10 further includes for making rear body 121 be connected with shell 11
More connection ribs 13, more connection ribs 13 and respectively connect rib 13 along the circumferential direction arranged for interval successively of drainage channel 101
It is connected respectively with the outer wall of the inner wall of shell 11 and rear body 121.In the specific embodiment of the invention, as shown in figure 4, more connection ribs
Circumferential direction successively arranged for interval of the item 13 along drainage channel 101.The upper surface being connected with the inner wall of shell 11 on each connection rib 13
In with the matched curved surface of the inner wall of shell 11, it is each connect the lower face that is connected with the outer wall of rear body 121 on rib 13 in and rear body
The 121 matched curved surface of outer wall.When the upper surface of connection rib 13 matches connection with the internal face of shell 11, and connection rib
When lower face matches connection with the outside wall surface of rear body 121, conical inner body 12 and 11 stable connection of shell, the overall structure of engine
Intensity is strong.
Optionally, as shown in figure 4, conical inner body 12 is coaxially disposed with shell 11, and the rear end face of rear body 121 and drainage are logical
The outlet in road 101 flushes.
Optionally, as shown in figure 4, conical inner body 12 further includes the precursor 122 being connected in an axial direction with rear body 121, precursor 122
Outside shell 11, and the outside wall surface of precursor 122 constitutes the air intake duct for being compressed to air.Combustion chamber 21 is located at precursor
In 122.Second jet pipe 22 is connected with the outlet of combustion chamber 21, and the second jet pipe 22 axially extending to after along conical inner body 12
The rear end face of body 121 is connected to.
Optionally, the oxidant flow channel of oxidant progress water conservancy diversion (is schemed not as shown in figure 4, being equipped in conical inner body 12
Show), for fuel carry out water conservancy diversion fuel flow channel (not shown).The flow inlet of oxidant flow channel and oxidation agent feeding device 30
It is connected, the outflow port of oxidant flow channel is connected with the head of combustion chamber 21, and the oxidant pumped out by oxidation agent feeding device 30 leads to
Peroxide agent runner is finally sprayed by the head of combustion chamber 21 in combustion chamber.The flow inlet of fuel flow channel and fuel supply system 40
It is connected, the outflow port of fuel flow channel is connected with the second wall cooling channel.The fuel meat stream pumped out by fuel supply system 40
Enter in the first wall cooling channel cooling first jet pipe 113 and detonation chamber 112, another part fuel enters the by fuel flow channel
Cooling second jet pipe 22 and combustion chamber 21 in two wall cooling channels.
In the specific embodiment of the invention, the first jet pipe 113 is diverging nozzle, and the second jet pipe 22 is Laval nozzle, is used
In the flow velocity for increasing high temperature and pressure combustion product, so that it is quickly discharged and generate thrust.
According to another aspect of the present invention, a kind of hypersonic aircraft is additionally provided, the combination of above-described embodiment is included
Cycle engine.By experimental verification, hypersonic aircraft operation mode of the invention is small, mode conversion difficulty is small, thermal cycle
It is efficient, thrust performance is good and chamber length is short.
The foregoing is only a preferred embodiment of the present invention, is not intended to restrict the invention, for the skill of this field
For art personnel, the invention may be variously modified and varied.All within the spirits and principles of the present invention, any made by repair
Change, equivalent replacement, improvement etc., should all be included in the protection scope of the present invention.
Claims (10)
1. a kind of combined cycle engine, which is characterized in that in the inner including rotation pinking punching engine (10) and setting
Rocket engine (20)；
The rotation pinking punching engine (10) includes：
Shell (11), the shell (11) are in the hollow tubular of both ends connection；
The rear body (121) of conical inner body (12), the conical inner body (12) is stretched by the inlet end of the shell (11) outside described
It is connected in shell (11) and with the shell (11), and the gap between the shell (11) and the rear body (121) is formed for sky
Gas introduces the drainage channel (101) in the shell (11), it is described after between body (121) rear end face and the shell (11) inner wall
Cavity form the detonation chamber (112) that is connected to the drainage channel (101), and the be connected to the detonation chamber (112)
One jet pipe (113), the spout and atmosphere of first jet pipe (113)；
The rocket engine (20) includes：
The combustion chamber (21) being set in the conical inner body (12), the second jet pipe (22) being connected to the combustion chamber (21),
The spout of second jet pipe (22) is connected to the detonation chamber (112), the combustion chamber (21) with for providing oxidant
Oxidation agent feeding device (30) and the fuel supply system (40) for providing fuel are respectively communicated with；
It is machined on the outer wall of the shell (11) along its circumferential multiple outer spout (1110) spaced successively, the outer spray
Mouth (1110) both ends are connected to the fuel supply system (40) and the drainage channel (101) respectively, so that fuel is by multiple
The outer spout (1110) sprays into the drainage channel (101)；And/or
It is machined on the outer wall of body (121) after described along its circumferential multiple interior spout (1210) spaced successively, it is described interior
Spout (1210) both ends are connected to the fuel supply system (40) and the drainage channel (101) respectively, so that fuel is by more
A interior spout (1210) sprays into the drainage channel (101).
2. combined cycle engine according to claim 1, which is characterized in that
Each outer outlet 10mm~20mm of the spout (1110) away from the drainage channel (101)；And/or
Each interior outlet 10mm~20mm of the spout (1210) away from the drainage channel (101).
3. combined cycle engine according to claim 2, which is characterized in that
Multiple outer spouts (1110) and multiple interior spouts (1210) are arranged in a one-to-one correspondence or multiple outer spouts
(1110) it misses one another arrangement with multiple interior spouts (1210).
4. combined cycle engine according to claim 3, which is characterized in that
The outer wall of first jet pipe (113) and the detonation chamber (112) is equipped with the first wall cooling channel, first wall
Face cooling duct is connected with the fuel supply system (40)；
Multiple outer spouts (1110) are connected to first wall cooling channel respectively, for cooling first jet pipe
(113) it is sprayed into the drainage channel (101) by multiple outer spouts (1110) with the fuel after the detonation chamber (112).
5. combined cycle engine according to claim 3, which is characterized in that
The outer wall of second jet pipe (22) and the combustion chamber (21) is equipped with the second wall cooling channel, second wall surface
Cooling duct is connected with the fuel supply system (40)；
The head of multiple interior spouts (1210) and the combustion chamber (21) is connected to second wall cooling channel respectively,
So that for the fuel behind cooling second jet pipe (22) and the combustion chamber (21), a part is by multiple interior spouts
(1210) it sprays into the drainage channel (101), another part sprays into the combustion chamber by the head of the combustion chamber (21)
(21) in.
6. combined cycle engine according to any one of claim 1 to 5, which is characterized in that
The detonation chamber (112) is in straight-tube shape；
The length of the detonation chamber (112) is 400mm~800mm.
7. combined cycle engine according to any one of claim 1 to 5, which is characterized in that
The rotation pinking punching engine (10) further include for make it is described after body (121) be connected with the shell (11) it is more
Root connects rib (13), the more connection ribs (13) along the drainage channel (101) circumferential direction arranged for interval successively, and respectively
It is described connection rib (13) upper surface be connected with the inner wall of the shell (11), it is each it is described connect rib (13) lower face and
The outer wall of body (121) is connected after described.
8. combined cycle engine according to any one of claim 1 to 5, which is characterized in that
The conical inner body (12) further includes the precursor (122) being connected in an axial direction with the rear body (121), the precursor (122)
Outside positioned at the shell (11), and the outside wall surface of the precursor (122) constitutes the air intake duct for being compressed to air；
The combustion chamber (21) is located in the precursor (122)；
Second jet pipe (22) is connected with the outlet of the combustion chamber (21), and second jet pipe (22) is along the center cone
Axially extending to for body (12) is connected to the rear end face of the rear body (121).
9. combined cycle engine according to any one of claim 1 to 5, which is characterized in that
It is equipped in the conical inner body (12) and the oxidant flow channel of water conservancy diversion is carried out, for the combustion to fuel progress water conservancy diversion to oxidant
Stream road；
The flow inlet of the oxidant flow channel is connected with the oxidation agent feeding device (30), the outflow port of the oxidant flow channel
It is connected with the head of the combustion chamber (21)；
The flow inlet of the fuel flow channel is connected with the fuel supply system (40), the outflow port of the fuel flow channel with it is described
Second wall cooling channel is connected.
10. a kind of hypersonic aircraft, which is characterized in that include the combination as described in any one of the claims 1 to 9
Cycle engine.
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| CN109779784A (en) * | 2018-12-14 | 2019-05-21 | 西安航天动力研究所 | A kind of RBCC engine inner flow passage of the preposition center layout of rocket |
| CN109887616A (en) * | 2019-02-18 | 2019-06-14 | 中国人民解放军国防科技大学 | Coolant-free first wall part of fusion reactor based on heat conduction of carbon nano tube |
| CN109931185A (en) * | 2019-03-20 | 2019-06-25 | 西北工业大学 | A kind of monoblock type pinking punching engine |
| CN109931628A (en) * | 2019-03-27 | 2019-06-25 | 北京理工大学 | It is a kind of based on the ring cavity eddy flow of the combustion chamber RDE to spray structure |
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| CN109887616A (en) * | 2019-02-18 | 2019-06-14 | 中国人民解放军国防科技大学 | Coolant-free first wall part of fusion reactor based on heat conduction of carbon nano tube |
| CN109887616B (en) * | 2019-02-18 | 2024-02-20 | 中国人民解放军国防科技大学 | Coolant-free fusion reactor first wall part based on carbon nano tube heat conduction |
| CN109931185A (en) * | 2019-03-20 | 2019-06-25 | 西北工业大学 | A kind of monoblock type pinking punching engine |
| CN109931185B (en) * | 2019-03-20 | 2021-02-19 | 西北工业大学 | Integral detonation ramjet engine |
| CN109931628A (en) * | 2019-03-27 | 2019-06-25 | 北京理工大学 | It is a kind of based on the ring cavity eddy flow of the combustion chamber RDE to spray structure |
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| 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 |
| CN112627983B (en) * | 2020-12-25 | 2022-02-22 | 中国人民解放军国防科技大学 | RBCC engine inner flow channel and RBCC engine |
| CN112879178A (en) * | 2021-01-22 | 2021-06-01 | 南京理工大学 | Solid rocket ramjet based on detonation combustion |
| CN112879178B (en) * | 2021-01-22 | 2022-11-04 | 南京理工大学 | Solid rocket ramjet based on detonation combustion |
| CN113123899A (en) * | 2021-04-19 | 2021-07-16 | 中国人民解放军国防科技大学 | Central injection rocket, RBCC engine and fuel injection method thereof |
| CN114321980A (en) * | 2022-01-07 | 2022-04-12 | 中国人民解放军国防科技大学 | Jet mixing control method and device for self-sustaining and stable detonation in supersonic airflow |
| CN114777162A (en) * | 2022-06-15 | 2022-07-22 | 清航空天(北京)科技有限公司 | Continuous rotation knocking ramjet engine with radial oil supply and air supply |
| CN115183276A (en) * | 2022-07-25 | 2022-10-14 | 清航空天(北京)科技有限公司 | Fuel supply assembly, engine combustion chamber structure and engine |
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