CN109812352A - Rocket ejector and its thermal protection structure and thermal protection method - Google Patents
Rocket ejector and its thermal protection structure and thermal protection method Download PDFInfo
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- CN109812352A CN109812352A CN201910166190.7A CN201910166190A CN109812352A CN 109812352 A CN109812352 A CN 109812352A CN 201910166190 A CN201910166190 A CN 201910166190A CN 109812352 A CN109812352 A CN 109812352A
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Abstract
The invention discloses a kind of using discharge and sweating Compound cooling to the thermal protection structure and method of TRRE Rocket ejector, is classified as two parts of leading edge nose cone and thrust chamber and is cooled down respectively.For leading edge nose cone part, thermal protection is carried out using laminate Sweat coolling.Using the cooling Compound cooling mode with Sweat coolling of discharge, discharge is cooling to use spiral shape cooling channel structure for thrust chamber part.Discharge is cooling simultaneously to cool down burning outdoor face and inner surface, and nozzle portion cooling effect is bad, carries out porous Sweat coolling again to nozzle portion.Thermal protection method and structure provided by the invention, good cooling results, required coolant be few, high reliablity, and can reach reusable purpose.
Description
Technical field
The present invention relates to the thermal protection structure and method of Rocket ejector in a kind of TRRE engine, specially using discharge and
Sweating Compound cooling carries out thermal protection to Rocket ejector in TRRE engine.
Background technique
For a long time, countries in the world are increasing always utilization and exploitation to space resources, but are more that activity exists
On atmosphere and Earth's orbit within 20km, near space (about 20-100km) between the two but pay close attention to compared with
It is few.With the development of scientific and technological level and aviation integral Concept of Operations, various countries begin to focus on the near space between empty day.
For near space vehicle as the important tool quickly travelled to and fro between the world, one of core technology is exactly dynamical system, and
And to it, more stringent requirements are proposed, needs to have high-performance, low cost, wide fast domain, reusable, lightweight, high motor-driven
The features such as ability.Turbine ancillary rocket enhances punching press combined cycle engine (Turbo-aided Rocket-augmented
Ramjet Combined Cycle Engine, TRRE) it is a kind of combined engine that can satisfy requirement.
TRRE engine be it is a kind of by turbine, rocket and punching engine by structure height integrated, thermodynamic cycle and work
The highly integrated air suction type combined cycle engine making the organic assembling of process and being formed.Rocket ejector is in effect wherein
It is to provide thrust when aircraft carries out speed change and thrust deficiency or lighted a fire using Rocket Exhausted Jet.Rocket ejector is put
It sets in high-speed channel i.e. punching press channel, in the channels, Rocket ejector leading edge is faced with washing away for hypersonic air-flow.And
When high engine speeds channels operation, scramjet combustor fuel combustion, Rocket ejector will have been at high-temperature fuel gas
In, fuel gas temperature range is up to 2000~3000K.Hot-fluid and temperature when Rocket ejector oneself work, in rocket thrust room
It is also that current material is unaffordable.At such high temperatures, in order to enable Rocket ejector not to be burned out, keep shape
And structural integrity, it is necessary to thermal protection be carried out to it by the way of active thermal protection.
At present for the thermal protection method of Rocket ejector are as follows: leading edge portion can refer to aircraft leading edge, mainly use ablation
The thermal protection method of thermal protection, thermal barrier coating and Sweat coolling.Ablating heat shield is only capable of first use, should not be reusable
Engine in use, and ablation layer consume after can change engine models face.Thermal barrier coating and Sweat coolling application are wide
It is general, and Sweat coolling is considered as most promising novel cooling technology, have strong cooling capacity, good cooling results,
It is the effective technology for solving the high hot-fluid of Rocket ejector leading edge high temperature using the less feature of coolant.But it is traditional from sweating
With porous media Sweat coolling in hot-spot, causes local flow resistance to increase cooling medium in turn and do not pass through from this, cause
Expansion and deterioration at hot-spot.And since Rocket ejector is in stamping combustion chamber, the high-temperature fuel gas pole in combustion chamber
The blocking for easily causing cooling duct, influences cooling effect.For rocket thrust chamber part, can using the method for re-generatively cooled
It enough meets the requirements, but working time of the Rocket ejector in the TRRE entire duty cycle is not long, the cooling after absorbing heat
Fuel, which enters, will will cause waste in rocket chamber.And for nozzle throat because there are high hot-fluid, cooling capacity is insufficient.
Summary of the invention
The purpose of the present invention is for TRRE Rocket ejector complicated thermal environment provide it is a kind of using discharge and sweating it is compound
The cooling thermal protection structure to TRRE Rocket ejector.Rocket ejector leading edge nose cone uses laminate Sweat coolling, solves because in combustion gas
Particulate matter blocks the hot-spot problem that cooling duct generates；Thrust chamber makes coolant be fully used using discharge cooling,
Nozzle portion is further added by porous Sweat coolling and solves the problems, such as that throat is cooling insufficient.
To solve the above problems, the technical solution adopted by the present invention are as follows: the present invention provides a kind of multiple using discharge and sweating
Close the cooling thermal protection structure to TRRE Rocket ejector, the TRRE Rocket ejector includes leading edge nose cone and thrust chamber, described to push away
Power room includes combustion chamber and jet pipe, and the jet pipe is divided into jet pipe internal layer and jet pipe outer layer；The thermal protection structure includes that sweating is cold
But structure and discharge cooling structure, the Sweat coolling structure includes laminate Sweat coolling structure and porous Sweat coolling structure；
The leading edge nose cone uses laminate Sweat coolling structure, the combustion chamber and jet pipe outer layer using discharge cooling structure, the spray
Inner tube layer uses porous Sweat coolling structure.
Further, the laminate Sweat coolling structure are as follows: shape is welded by diffusion welding (DW) using the laminate of multiple and different sizes
At the leading edge nose cone, the leading edge nose cone being welded is centrally formed coolant channel；The leading edge nose cone is logical by coolant
Road sweating is to form laminate sweating boundary film in the outer surface of the leading edge nose cone；
The laminate is centrally formed hollow cavity, when the laminate of multiple and different sizes is welded into leading edge nose cone by diffusion welding (DW)
Afterwards, the center cavity of different size of laminate forms the coolant channel.
Further, the laminate is ring structure, 0.1~1mm of thickness；The laminate successively processes circumferential direction from interior outward
It distributes area, control runner and spreads runner, the circumferential distribution area, control runner and distribution runner are equal along the circumferencial direction of laminate
Even distribution.
Further, the circumferential distribution area is annular channel, and control runner is elongate structure, and distribution runner is rectangle knot
Structure.
Further, jet pipe outer layer is set as the discharge cooling structure with Rocket ejector combustion chamber one, and jet pipe internal layer is by gold
Belong to base or ceramic matric composite is made.
Further, the discharge cooling structure are as follows: the jet pipe outer layer is provided with discharge-channel；The combustion chamber and institute
The heat for stating jet pipe outer layer is discharged by the discharge-channel, and the coolant flow pass is set on the combustion chamber.
Further, fuel flow channel and ejector filler, the end of ejector filler are provided between the leading edge nose cone and the thrust chamber
Portion is equipped with oxidant inlet, is equipped with fuel nozzle and oxidant nozzle, the oxidant nozzle and oxidant in the ejector filler
Entrance be connected, the fuel flow channel include feed inlet, along feed inlet to the split channel of Rocket ejector internal stretch, along feed inlet
Enter channel to the coolant that jet pipe outer layer extends to the fuel nozzle of combustion chamber direction extension and along feed inlet, it is described cold
But agent enters channel and is connected with the discharge-channel.
Further, the porous Sweat coolling structure are as follows: storage room, the storage are provided with above the throat of the jet pipe
It deposits room and is connected with storage room channel, storage room channel is connected with discharge-channel, and the storage room gives jet pipe to store coolant
The porous Sweat coolling of internal layer provides coolant, and inner surface of the jet pipe internal layer by sweating in the jet pipe internal layer forms spray
Tube edge limitans.
Further, the discharge-channel uses spiral shape cooling channel structure.
The present invention also provides a kind of using discharge with sweating Compound cooling to the thermal protection method of Rocket ejector, and use is above-mentioned
Any thermal protection structure；The leading edge nose cone uses laminate Sweat coolling, the combustion chamber and jet pipe outer layer using row
It lets cool but, the jet pipe internal layer uses porous Sweat coolling.
The present invention also provides a kind of Rocket ejector, the Rocket ejector uses thermal protection method described above, described to draw
Rocket is penetrated to be installed in TRRE stamping combustion chamber.
Beneficial effects of the present invention:
1, the present invention provide it is a kind of using discharge and sweating Compound cooling to the thermal protection structure of TRRE Rocket ejector.Injection
Rocket leading edge nose cone uses laminate Sweat coolling, solves to ask because of the hot-spot that particulate matter blocking cooling duct generates in combustion gas
Topic；Thrust chamber makes coolant be fully used using discharge cooling, and nozzle portion is further added by porous Sweat coolling and solves throat
Cooling insufficient problem.Thermal protection structure provided by the invention and method, good cooling results, required coolant be few, reliability
Height, and reusable purpose can be reached.
Rocket ejector nose cone part uses the technology of laminate Sweat coolling, and cooling capacity is strong, good cooling results, required cold
But agent is few, high reliablity, and can reach reusable purpose；The combustion chamber technology cooling using discharge, reduces coolant
Waste, improve the performance of engine；Jet pipe fully ensures that nozzle portion using the cooling Compound cooling with Sweat coolling of discharge
Cooling effect.
2, TRRE Rocket ejector can be divided into two parts of leading edge nose cone and thrust chamber and be cooled down.For leading edge nose cone portion
Point, thermal protection is carried out using laminate Sweat coolling.Entire leading edge nose cone part is welded by laminate by diffusion welding (DW), and laminate is whole
Body is ring-type, and the nose cone central part being welded just forms coolant channel.Coolant is inhaled when through the runner of laminate
Heat is received, liquid film is formed when reaching leading edge nose cone surface or air film (laminate sweating boundary film) blocks heat transfer, reach thermal protection
Purpose.Using the cooling Compound cooling mode with Sweat coolling of discharge, discharge is cooling to be led to using spiral shape discharge for thrust chamber part
Road carries out discharge cooling.Discharge cooling of the invention can simultaneously cool down burning outdoor face and inner surface, spout portion
Divide cooling effect bad, porous Sweat coolling is carried out again to nozzle portion.Nozzle portion is by Metal Substrate or ceramic matric composite
It is made up of incomplete densification process (common process), coolant can reach jet pipe inner surface by hole.
3, nose cone of the present invention is welded by laminate by diffusion welding (DW), the method that laminate passes through chemical etching or photoetching
Circumferential distribution area, control runner are processed on the surface thereof and spread runner.Coolant first enters laminate by coolant channel
Circumferential distribution area, flows successively through control runner and spreads runner, carries out heat convection in this section, absorbs the combustion gas of high temperature mainstream
Heat.When flowing out laminate arrival leading edge nose cone surface, one layer of boundary film (laminate sweating boundary film) is formed, reaches barrier heat transfer
Effect.Traditional porous Sweat coolling structure be easy to cause duct to block, causes hot-spot in stamping combustion chamber, part
Flow resistance at overheat increases, and coolant is flowed through without at hot-spot from duct is communicated, the expansion that causes hot-spot to go out and
Deteriorate.Laminate of the present invention is by can accurately control to circumferential distribution area, control runner and the rational design for spreading runner
The coolant flow in each region.And control runner be slim-lined construction, the flow resistance of entire laminate nearly all in the region, and
Runner is spread as heat affected area, even if there is hot-spot, control runner, the coolant flow of inflow will not be influenced
Will not be impacted, the problem of expanding at the hot-spot that porous Sweat coolling structure occurs and deteriorating thus can be properly settled,
It also can reach the purpose of reuse.
4, the compound cooling structure that thrust chamber of the present invention uses, entire combustion chamber use discharge cooling structure.Row
It lets cool but using spiral shape cooling channel structure, can be improved coolant flow speed, reinforce cooling effect.Discharge is cooling and raw and cold again
But principle is similar, is one kind of heat convection.When traditionally application discharge is cooling, the waste of coolant will cause, but in TRRE
In stamping combustion chamber, coolant is emitted from jet pipe tail portion, can also be burnt in stamping combustion chamber, to improve engine
Performance.And Rocket ejector has the function of increasing thrust and injection igniting when designing, coolant can fire when flowing out in punching press
Burn full combustion in room.Compared with re-generatively cooled, discharges cooling coolant and enter stamping combustion chamber, the coolant of re-generatively cooled
Into Rocket ejector combustion chamber, the working time of Rocket ejector is much smaller than punching engine, but the thermal protection of Rocket ejector with
The working time of punching engine is consistent, so using cooling more preferable, the coolant utilization rate than using re-generatively cooled effect of discharge
It is higher.
5, nozzle portion of the present invention is using the cooling structure for discharging cooling and Sweat coolling Compound cooling.Since it is desired that simultaneously
It is cooling simultaneously to jet pipe outer surface and inner surface, and nozzle throat is because heat flow density is big, cooling capacity is insufficient, therefore individually adopts
It is bad with discharge cooling effect.Entire jet pipe is divided into two parts of internal layer and outer layer, and outer layer is the row with rocket chamber one
Cooling structure is put, internal layer is the jet pipe made of incomplete densification process Metal Substrate or ceramic matric composite.It is spraying
Above pipe throat, the storage room of a coolant is designed, provides coolant to Sweat coolling.Coolant passes through by Metal Substrate or pottery
The active porosity channel that porcelain based composites are formed by incomplete densification process (existing technique), using its high porosity,
The characteristics of high-specific surface area, absorbs the heat spread out of inside jet pipe.Coolant is flowed out from hole, and then is formed in jet pipe inner surface
Boundary film, barrier heat transfer.The porous Sweat coolling structure of jet pipe internal layer can increase cooling capacity, solve the cooling energy of nozzle throat
Hypodynamic problem improves jet pipe cooling effect and reliability.
Detailed description of the invention
Fig. 1 is a kind of schematic perspective view of Rocket ejector provided in an embodiment of the present invention；
Fig. 2 is a kind of half sectional view of Rocket ejector provided in an embodiment of the present invention, shows injection of the embodiment of the present invention
The thermal protection structure that rocket uses；
Fig. 3 is partial enlargement diagram in Fig. 2；
Fig. 4 is the working environment schematic diagram of Rocket ejector in TRRE stamping combustion chamber in the embodiment of the present invention；
Fig. 5 is middle plate of embodiment of the present invention schematic perspective view；
Fig. 6 is leading edge of embodiment of the present invention nose cone part laminate Sweat coolling coolant flow schematic diagram；
Fig. 7 is jet pipe of embodiment of the present invention Sweat coolling coolant flow schematic diagram；
Fig. 8 is that oxidant flows into ejector filler diagrammatic cross-section in the embodiment of the present invention.
Appended drawing reference: 1 leading edge nose cone；10 laminates；100 circumferential distribution areas；101 control runners；102 stroll runners；In 103
The chambers of the heart；11 coolant channels；12 laminate sweating boundary films；2 thrust chambers；20 combustion chambers；21 jet pipes；210 jet pipe internal layers；2100 sprays
Tube edge limitans；211 jet pipe outer layers；2110 discharge-channels；22 coolant flow pass；23 oxidant inlets；24 oxidant nozzles；
3 coolant channels；4 storage rooms；5 feed inlets；6 split channels；7 fuel nozzles；8 coolants enter channel；9 storage room channels；
13 ejector fillers.
Specific embodiment
In order to make those skilled in the art more fully understand this technology, we are in conjunction with attached drawing and specific implementation example to this
Invention is described in further detail.
The embodiment of the present invention provides a kind of thermal protection structure using discharge with sweating Compound cooling to TRRE Rocket ejector,
The TRRE Rocket ejector includes leading edge nose cone 1 and thrust chamber 2, and the thrust chamber 2 includes combustion chamber 20 and jet pipe 21, the spray
Pipe 21 divides for jet pipe internal layer 210 and jet pipe outer layer 211；The thermal protection structure includes Sweat coolling structure and the cooling knot of discharge
Structure, the Sweat coolling structure include laminate Sweat coolling structure and porous Sweat coolling structure；The leading edge nose cone 1 uses layer
Plate Sweat coolling structure, the combustion chamber 20 and jet pipe outer layer 211 are used using discharge cooling structure, the jet pipe internal layer 210
Porous Sweat coolling structure.
Further, the laminate Sweat coolling structure are as follows: welded using the laminate 10 of multiple and different sizes by diffusion welding (DW)
The leading edge nose cone 1 is formed, the leading edge nose cone 1 being welded is centrally formed coolant channel 11；The leading edge nose cone 1 passes through cold
But 11 sweating of agent channel is to form laminate sweating boundary film 12 in the outer surface of the leading edge nose cone 1；
The laminate 10 is centrally formed hollow cavity 103, when the laminate 10 of multiple and different sizes is welded by diffusion welding (DW)
After leading edge nose cone 1, the center cavity 103 of different size of laminate 1 forms the coolant channel 11.
Referring to Fig. 5, the laminate 10 is ring structure, 0.1~1mm of thickness；The laminate 10 is successively processed outward from interior
Circumferential distribution area 100, control runner 101 and distribution runner 102 out, the circumferential distribution area 100, control runner 101 and distribution
Runner 102 is uniformly distributed along the circumferencial direction of laminate 10.
Traditional porous Sweat coolling structure be easy to cause duct to block, causes hot-spot, office in stamping combustion chamber
Flow resistance at portion's overheat increases, and coolant is flowed through without at hot-spot from duct is communicated, the expansion for causing hot-spot to go out
And deterioration.Laminate of the present invention is by can accurately control to circumferential distribution area, control runner and the rational design for spreading runner
Make the coolant flow in each region.And controlling runner is slim-lined construction, the flow resistance of entire laminate nearly all in the region,
And runner is spread as heat affected area, even if there is hot-spot, control runner, the coolant flow of inflow will not be influenced
Will not be impacted, it thus can properly settle and expand at the hot-spot that porous Sweat coolling structure occurs and what is deteriorated asks
Topic, also can reach the purpose of reuse.
Further, the circumferential distribution area 100 is annular channel, and control runner 101 is elongate structure, spreads runner
102 be rectangular configuration.
Further, jet pipe outer layer 211 is set as the discharge cooling structure with 20 one of Rocket ejector combustion chamber, jet pipe internal layer
210 are made of Metal Substrate or ceramic matric composite.
Further, the discharge cooling structure are as follows: the jet pipe outer layer 211 is provided with discharge-channel 2110；The combustion
The heat for burning room 20 and the jet pipe outer layer 211 is discharged by the discharge-channel 2110, and the coolant flow pass 22 is set
In on the combustion chamber 2.
Further, fuel flow channel and ejector filler 13, ejector filler are provided between the leading edge nose cone 1 and the thrust chamber 2
13 end is equipped with oxidant inlet 23, and fuel nozzle 7 and oxidant nozzle 24, the oxidant spray are equipped in the ejector filler
Mouth 24 is connected with oxidant inlet 23, the fuel flow channel include feed inlet 5, along feed inlet 5 to Rocket ejector internal stretch
Split channel 6, the fuel nozzle 7 extended to 20 direction of combustion chamber along feed inlet 5 and prolong along feed inlet 5 to jet pipe outer layer 211
The coolant stretched enters channel 8, and the coolant enters channel 8 and is connected with the discharge-channel 2110.
Further, the porous Sweat coolling structure are as follows: storage room 4, institute are provided with above the throat of the jet pipe 21
It states storage room 4 and is connected with storage room channel 9, storage room channel 9 is connected with discharge-channel 2110, and the storage room 4 is to store
Coolant provides coolant to the porous Sweat coolling of jet pipe internal layer, and the jet pipe internal layer 210 is by sweating in the jet pipe
The inner surface of layer 210 forms jet pipe boundary film 2100.
Further, the discharge-channel 2110 uses spiral shape cooling channel structure.
The another aspect of the embodiment of the present invention provides a kind of anti-to the heat of Rocket ejector using discharge and sweating Compound cooling
Maintaining method, using above-mentioned any thermal protection structure；The leading edge nose cone 1 uses laminate Sweat coolling, the combustion chamber
20 and jet pipe outer layer 211 using discharge cooling, the jet pipe internal layer 210 uses porous Sweat coolling.
Another embodiment of the present invention provides a kind of Rocket ejector, the Rocket ejector uses thermal protection side described above
Method, the Rocket ejector are installed in TRRE stamping combustion chamber.
The working principle of the embodiment of the present invention is as follows: as shown in Fig. 2, the embodiment of the present invention uses TRRE Rocket ejector,
TRRE Rocket ejector works in stamping combustion chamber, by the effect of high-temperature fuel gas.Fuel enters Rocket ejector by feed inlet 5, point
It works for three roads:
First via fuel: a part of fuel enters coolant as coolant and enters channel 8, using discharge-channel
2110, into after the discharge-channel 2110 of spiral, a part of coolant is expelled to TRRE punching by coolant flow pass 22
Compression ignition burns interior and carries out burning recycling.It is cold in discharge-channel 2110 since discharge-channel 2110 is connected with storage room channel 9
But agent some by storage room channel 9 enter storage room 4, to jet pipe Sweat coolling provide coolant.Coolant passes through spray
Heat is absorbed when the hole of inner tube layer, jet pipe inner surface is eventually arrived at and forms jet pipe boundary film 2100, barrier heat transfer.
Second and third road fuel: another part fuel passes through split channel 6, and a part of fuel enters cold in leading edge nose cone 1
But (the second tunnel) in agent channel 3, the leading edge nose cone 1 that fuel is formed by coolant channel 3 into laminate 10 as coolant, stream
Circumferential distribution area 100, the control runner 101, distribution channels 102 for crossing laminate 10, utilize laminate large specific surface area in this area
The characteristics of carry out heat convection, be finally reached leading edge nose cone 1 surface formed laminate sweating boundary film 12 barrier heat transfer.There are also one
Part of fuel pass through split channel 6, into ejector filler 13 by fuel nozzle 7 spray (third road), with from oxidant inlet 23 into
Enter, the oxidant mixing that oxidant nozzle 24 sprays, and then burns within the combustion chamber 20.
Into ejector filler 13, the fuel flow rate of coolant channel 3 and discharge-channel 2110, controlled by the size in channel
System.The flow area area in channel is smaller, and channel is longer, and flow resistance is bigger when fuel passes through, by flow with regard to smaller.It can be pre-
Estimate the fuel mass flow rates into ejector filler 13, coolant channel 3 and discharge-channel 2110, and then each channel of careful design is big
Small and length controls flow.By accurately controlling uninterrupted everywhere, it both can fully ensure that cooling effect, can not also cause
The waste of fuel, because the fuel of engine institute band is limited.
And fuel flow rate needed for laminate Sweat coolling is smaller, to smaller for burning and discharging cooling influence.
It should be noted that the fuel that coolant uses TRRE engine to carry in the embodiment of the present invention.
The embodiment of the present invention is not to limit the present invention, all within the spirits and principles of the present invention, made any to repair
Change, equivalent replacement, improvement etc., should all be included in the protection scope of the present invention.
Claims (10)
1. a kind of utilize the thermal protection structure discharged with sweating Compound cooling to TRRE Rocket ejector, which is characterized in that described
TRRE Rocket ejector includes leading edge nose cone (1) and thrust chamber (2), and the thrust chamber (2) includes combustion chamber (20) and jet pipe (21),
The jet pipe (21) is divided into jet pipe internal layer (210) and jet pipe outer layer (211)；The thermal protection structure include Sweat coolling structure and
Cooling structure is discharged, the Sweat coolling structure includes laminate Sweat coolling structure and porous Sweat coolling structure；The leading edge
Nose cone (1) uses laminate Sweat coolling structure, the combustion chamber (20) and jet pipe outer layer (211) using discharge cooling structure, institute
It states jet pipe internal layer (210) and uses porous Sweat coolling structure.
2. thermal protection structure according to claim 1, which is characterized in that the laminate Sweat coolling structure are as follows: using more
A different size of laminate (10) welds to form the leading edge nose cone (1) by diffusion welding (DW), in the leading edge nose cone (1) being welded
It is formed centrally coolant channel (11)；The leading edge nose cone (1) is by coolant channel (11) sweating in the leading edge nose cone
(1) outer surface forms laminate sweating boundary film (12)；
The laminate (10) is centrally formed hollow cavity (103), when the laminate (10) of multiple and different sizes is welded by diffusion welding (DW)
After leading edge nose cone (1), the center cavity (103) of different size of laminate (1) forms the coolant channel (11).
3. thermal protection structure according to claim 2, which is characterized in that the laminate (10) is ring structure, thickness 0.1
~1mm；The laminate (10) successively processes circumferential distribution area (100), control runner (101) and distribution runner from interior outward
(102), the circumferential distribution area (100), control runner (101) and distribution runner (102) are equal along the circumferencial direction of laminate (10)
Even distribution.
4. thermal protection structure according to claim 3, which is characterized in that the circumferential distribution area (100) is annular channel,
Controlling runner (101) is elongate structure, and spreading runner (102) is rectangular configuration.
5. thermal protection structure according to claim 1, which is characterized in that jet pipe outer layer (211) is set as and Rocket ejector
The discharge cooling structure of combustion chamber (20) one, jet pipe internal layer (210) are made of Metal Substrate or ceramic matric composite.
6. thermal protection structure according to claim 1, which is characterized in that the discharge cooling structure are as follows: in the jet pipe
Outer layer (211) is provided with discharge-channel (2110)；The heat of the combustion chamber (20) and the jet pipe outer layer (211) passes sequentially through
The discharge-channel (2110), coolant flow pass (22) discharge；The coolant flow pass (22) is set to the burning
On room (2).
7. thermal protection structure according to claim 1, which is characterized in that the leading edge nose cone (1) and the thrust chamber (2)
Between be provided with fuel flow channel and ejector filler (13), the end of ejector filler (13) is equipped with oxidant inlet (23), the ejector filler
(13) fuel nozzle (7) and oxidant nozzle (24) are equipped in, the oxidant nozzle (24) is connected with oxidant inlet (23),
The fuel flow channel include feed inlet (5), along feed inlet (5) to the split channel (6) of Rocket ejector internal stretch, along feed inlet
(5) fuel nozzle (7) extended to combustion chamber (20) direction and the cooling extended along feed inlet (5) to jet pipe outer layer (211)
Agent enters channel (8), and the coolant enters channel (8) and is connected with the discharge-channel (2110).
8. thermal protection structure according to claim 1, which is characterized in that the porous Sweat coolling structure are as follows: described
It is provided with storage room (4) above the throat of jet pipe (21), the storage room (4) is connected with storage room channel (9), storage room channel
(9) it is connected with discharge-channel (2110), the storage room (4) is to store coolant, to the porous Sweat coolling of jet pipe internal layer
Coolant is provided, inner surface of the jet pipe internal layer (210) by sweating in the jet pipe internal layer (210) forms jet pipe boundary film
(2100)。
9. a kind of utilize the thermal protection method discharged with sweating Compound cooling to Rocket ejector, which is characterized in that using such as right
It is required that any thermal protection structure of 1-8；The leading edge nose cone (1) use laminate Sweat coolling, the combustion chamber (20) and
Jet pipe outer layer (211) uses porous Sweat coolling using discharge cooling, the jet pipe internal layer (210).
10. a kind of Rocket ejector, which is characterized in that the Rocket ejector uses thermal protection method as claimed in claim 9, described
Rocket ejector is installed in TRRE stamping combustion chamber.
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Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110566999A (en) * | 2019-09-20 | 2019-12-13 | 清华大学 | Combustion chamber thermal protection wall surface structure utilizing fuel oil self-suction sweating cooling |
| CN110594037A (en) * | 2019-10-09 | 2019-12-20 | 湖南云顶智能科技有限公司 | Integrated injection rocket engine assembly and engine thereof |
| CN110594036A (en) * | 2019-08-23 | 2019-12-20 | 西北工业大学 | Active cooling double-nozzle support plate ejection rocket of rocket-based combined cycle engine |
| CN111874273A (en) * | 2020-07-01 | 2020-11-03 | 北京坤飞航天科技有限公司 | Propellant filling system and propellant filling method |
| CN114483372A (en) * | 2022-03-18 | 2022-05-13 | 华中科技大学 | Wide-adjustment-ratio ejection rocket for RBCC engine and control method thereof |
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Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2283965C1 (en) * | 2004-02-12 | 2006-09-20 | Юнайтид Текнолоджиз Копэрейшн | Cooling system of engine combustion chamber |
| CN103192978A (en) * | 2013-04-02 | 2013-07-10 | 中国人民解放军国防科学技术大学 | Laminate type sweating and reverse-jetting combined cooling nose cone |
| CN106894918A (en) * | 2017-03-13 | 2017-06-27 | 西北工业大学 | A kind of mode pre-cooling cycle system suitable for assembly power cycle engine |
| CN107719630A (en) * | 2017-09-25 | 2018-02-23 | 中国人民解放军国防科技大学 | Laminate type gas sweating nose cone |
| CN107956601A (en) * | 2017-06-15 | 2018-04-24 | 葛明龙 | One group of major diameter rocket thrust chamber with discharge re-generatively cooled partition plate |
| CN108438205A (en) * | 2018-03-13 | 2018-08-24 | 中国科学技术大学 | Adaptive local activates the hypersonic leading edge thermal protection method of Sweat coolling |
| CN108843460A (en) * | 2018-06-28 | 2018-11-20 | 厦门大学 | Heat to electricity conversion and pushing method is pre-chilled in turbo ramjet engine |
| CN209990561U (en) * | 2019-03-06 | 2020-01-24 | 中南大学 | Thermal protection structure of TRRE ejection rocket and ejection rocket thereof |
-
2019
- 2019-03-06 CN CN201910166190.7A patent/CN109812352A/en active Pending
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2283965C1 (en) * | 2004-02-12 | 2006-09-20 | Юнайтид Текнолоджиз Копэрейшн | Cooling system of engine combustion chamber |
| CN103192978A (en) * | 2013-04-02 | 2013-07-10 | 中国人民解放军国防科学技术大学 | Laminate type sweating and reverse-jetting combined cooling nose cone |
| CN106894918A (en) * | 2017-03-13 | 2017-06-27 | 西北工业大学 | A kind of mode pre-cooling cycle system suitable for assembly power cycle engine |
| CN107956601A (en) * | 2017-06-15 | 2018-04-24 | 葛明龙 | One group of major diameter rocket thrust chamber with discharge re-generatively cooled partition plate |
| CN107719630A (en) * | 2017-09-25 | 2018-02-23 | 中国人民解放军国防科技大学 | Laminate type gas sweating nose cone |
| CN108438205A (en) * | 2018-03-13 | 2018-08-24 | 中国科学技术大学 | Adaptive local activates the hypersonic leading edge thermal protection method of Sweat coolling |
| CN108843460A (en) * | 2018-06-28 | 2018-11-20 | 厦门大学 | Heat to electricity conversion and pushing method is pre-chilled in turbo ramjet engine |
| CN209990561U (en) * | 2019-03-06 | 2020-01-24 | 中南大学 | Thermal protection structure of TRRE ejection rocket and ejection rocket thereof |
Non-Patent Citations (1)
| Title |
|---|
| ""高超声速组合循环发动机综合热管理技术需求分析"" * |
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110594036A (en) * | 2019-08-23 | 2019-12-20 | 西北工业大学 | Active cooling double-nozzle support plate ejection rocket of rocket-based combined cycle engine |
| CN110594036B (en) * | 2019-08-23 | 2020-06-02 | 西北工业大学 | Active cooling double-nozzle support plate ejection rocket of rocket-based combined cycle engine |
| CN110566999A (en) * | 2019-09-20 | 2019-12-13 | 清华大学 | Combustion chamber thermal protection wall surface structure utilizing fuel oil self-suction sweating cooling |
| CN110566999B (en) * | 2019-09-20 | 2020-07-28 | 清华大学 | Combustion chamber thermal protection wall surface structure utilizing fuel oil self-suction sweating cooling |
| CN110594037A (en) * | 2019-10-09 | 2019-12-20 | 湖南云顶智能科技有限公司 | Integrated injection rocket engine assembly and engine thereof |
| CN111874273A (en) * | 2020-07-01 | 2020-11-03 | 北京坤飞航天科技有限公司 | Propellant filling system and propellant filling method |
| CN114483372A (en) * | 2022-03-18 | 2022-05-13 | 华中科技大学 | Wide-adjustment-ratio ejection rocket for RBCC engine and control method thereof |
| CN114483372B (en) * | 2022-03-18 | 2023-07-04 | 华中科技大学 | Rocket with wide regulating ratio for RBCC engine and control method thereof |
| CN114810424A (en) * | 2022-04-29 | 2022-07-29 | 西北工业大学 | Engine active cooling cavity structure based on spray cooling |
| CN114810424B (en) * | 2022-04-29 | 2024-02-02 | 西北工业大学 | Engine active cooling concave cavity structure based on spray cooling |
| CN117514523A (en) * | 2023-11-20 | 2024-02-06 | 北京智创联合科技股份有限公司 | Rocket engine combustion chamber with inner wall made of sweating material and manufacturing method thereof |
| CN117514523B (en) * | 2023-11-20 | 2024-08-09 | 北京智创联合科技股份有限公司 | Rocket engine combustion chamber with inner wall made of sweating material and manufacturing method thereof |
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