CN109736974A - A kind of temperature barrier and liquid launch vehicle - Google Patents
A kind of temperature barrier and liquid launch vehicle Download PDFInfo
- Publication number
- CN109736974A CN109736974A CN201910155308.6A CN201910155308A CN109736974A CN 109736974 A CN109736974 A CN 109736974A CN 201910155308 A CN201910155308 A CN 201910155308A CN 109736974 A CN109736974 A CN 109736974A
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- Prior art keywords
- cooling
- temperature
- launch vehicle
- liquid
- propellant
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- 239000007788 liquid Substances 0.000 title claims abstract description 65
- 230000004888 barrier function Effects 0.000 title claims abstract description 28
- 238000001816 cooling Methods 0.000 claims abstract description 76
- 239000003380 propellant Substances 0.000 claims abstract description 38
- 239000000758 substrate Substances 0.000 claims description 12
- 210000005239 tubule Anatomy 0.000 claims description 5
- 238000001514 detection method Methods 0.000 claims description 3
- 230000001681 protective effect Effects 0.000 abstract description 4
- 239000000463 material Substances 0.000 description 9
- 101100204059 Caenorhabditis elegans trap-2 gene Proteins 0.000 description 5
- 230000006870 function Effects 0.000 description 5
- 238000010438 heat treatment Methods 0.000 description 5
- 238000010586 diagram Methods 0.000 description 4
- 239000007789 gas Substances 0.000 description 4
- 238000000034 method Methods 0.000 description 4
- 238000002679 ablation Methods 0.000 description 3
- 229910001069 Ti alloy Inorganic materials 0.000 description 1
- 239000011204 carbon fibre-reinforced silicon carbide Substances 0.000 description 1
- 239000000919 ceramic Substances 0.000 description 1
- 239000000567 combustion gas Substances 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000001808 coupling effect Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 230000009977 dual effect Effects 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000004744 fabric Substances 0.000 description 1
- 239000011152 fibreglass Substances 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 238000001513 hot isostatic pressing Methods 0.000 description 1
- 239000007769 metal material Substances 0.000 description 1
- 230000008450 motivation Effects 0.000 description 1
- 238000000465 moulding Methods 0.000 description 1
- 230000000704 physical effect Effects 0.000 description 1
- 238000002360 preparation method Methods 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000001737 promoting effect Effects 0.000 description 1
- 239000003870 refractory metal Substances 0.000 description 1
- 238000005476 soldering Methods 0.000 description 1
- 239000010935 stainless steel Substances 0.000 description 1
- 229910001220 stainless steel Inorganic materials 0.000 description 1
- 239000002699 waste material Substances 0.000 description 1
Abstract
The present invention provides a kind of temperature barrier and liquid launch vehicles, temperature barrier includes the first cooling structure for being set to liquid launch vehicle bottom, first cooling structure one end is connected to propellant transfer system, the other end is connected to the thrust chamber of liquid launch vehicle engine, carries out heat convection by the propellant and the liquid launch vehicle that flow through first cooling structure.It is designed in this way, using the propellant low-temperature characteristics of liquid launch vehicle engine, active cooling is carried out to liquid launch vehicle bottom, the significant increase thermal protective performance of carrier rocket bottom, the thermal environment range being applicable in is more than traditional passive type thermal protection structure.The cooling propellant completed enters engine system and burns, and realizes the reasonable utilization of propellant.
Description
Technical field
The present invention relates to aerospace equipment technical fields, and in particular to a kind of temperature barrier and liquid embarkation fire
Arrow.
Background technique
The Aerodynamic Heating and radiant heating that liquid launch vehicle bottom is generated due to liquid launch vehicle jet cutting car flow, liquid
Liquid launch vehicle engine pipelines, electrical equipment etc. are arranged at body carrier rocket bottom, need to take thermal protection measure.It is existing primary
Property generally uses the composite materials such as glass reinforced plastic, flexible high temperature adiabatic cloth that bottom thermal protection knot is made using liquid launch vehicle bottom
Structure.The dual shadow that reusable liquid launch vehicle is heated during returning and reentering by atmosphere Aerodynamic Heating and jet flow
It rings, thermal environment is apparently higher than traditional disposable liquid launch vehicle, and Base Heat protection is faced with huge technological challenge.
Traditional passive type thermally protective materials and structure utilize the high temperature resistant of material ablation principle or material itself, low-heat
Conductance characteristic realizes thermal protection.Ablation class material will appear under the conditions of the high hot-fluid for the process that reenters, high dynamic pressure serious ablation,
Problem is degraded, to meet reentry stage thermal protective performance, needs to improve thermal protection structure thickness, brings greatly promoting for weight, and
It needs to replace thermal protection shield after returning every time, cannot achieve reusable.The non-ablative heat-resisting material of high-performance, such as
Novel C/SiC, C/C, superhigh temperature ceramics etc., processing difficulty is big, and cost is very high, and it is small that there are damage tolerances, and brittleness is larger
The problems such as, it is easily damaged under the conditions of reuse.
Summary of the invention
Therefore, the technical problem to be solved in the present invention is that overcoming thermal protection structure thermal protection efficiency in the prior art
It is low, the insufficient defect of reusable performance difference, to provide a kind of temperature barrier and liquid launch vehicle.
A kind of temperature barrier provided by the invention, the first cooling structure including being set to liquid launch vehicle bottom,
First cooling structure one end is connected to propellant transfer system, and the other end is connected to the thrust of liquid launch vehicle engine
Room carries out heat convection by the propellant and the liquid launch vehicle that flow through first cooling structure.
Optionally, first cooling structure includes substrate, the first cooling line for being set to the substrate and is located at
The first cooling line arrival end and the entrance liquid trap of outlet end and outlet liquid trap.
Optionally, first cooling line is provided with more, and more first cooling lines are intensively arranged.
Optionally, first cooling line is extended around the central axes of the liquid launch vehicle bottom.
Optionally, first cooling line is to be set to inside the substrate.
Optionally, first cooling line is thin with the integrally formed circular cross-section tubule of the substrate, rectangular section
Pipe or U-shaped cross-sectional passage.
It optionally, further include the second cooling structure for being set to the thrust chamber, the outlet end of first cooling body
Be connected by the second connecting line with the arrival end of second cooling structure, the outlet end of second cooling structure with it is described
Engine connection.
It optionally, further include the 4th connection being set between the propellant transfer system and second connecting line
Pipeline, the outlet end for being connected to first cooling structure connect with the 5th connecting line of the engine and corresponding described second
Take over road, the 4th connecting line and respectively arranged second control valve of the 5th connecting line, the 4th control valve and the
Five control valves.
Optionally, further include be set to first cooling structure exit temperature-detecting device and with the temperature
The controller of detection device electrical connection, the temperature information that the controller detect according to the temperature-detecting device control described the
Two control valves, the 4th control valve and the 5th control valve.
Present invention simultaneously provides a kind of liquid launch vehicles, including temperature barrier described in any of the above embodiments.
Technical solution of the present invention has the advantages that
A kind of temperature barrier provided by the invention, it is right using the propellant low-temperature characteristics of liquid launch vehicle engine
Liquid launch vehicle bottom carries out active cooling, the significant increase thermal protective performance of carrier rocket bottom, the hot ring that is applicable in
Border range is more than traditional passive type thermal protection structure.The cooling propellant completed enters engine system and burns, and realizes
The reasonable utilization of propellant.
Detailed description of the invention
It, below will be to specific in order to illustrate more clearly of the specific embodiment of the invention or technical solution in the prior art
Embodiment or attached drawing needed to be used in the description of the prior art be briefly described, it should be apparent that, it is described below
Attached drawing is some embodiments of the present invention, for those of ordinary skill in the art, before not making the creative labor
It puts, is also possible to obtain other drawings based on these drawings.
Fig. 1 is the structural schematic diagram of the first cooling structure provided in the first embodiment of the invention；
Fig. 2 is a kind of working principle diagram A of the temperature barrier provided in the first embodiment of the invention；
Fig. 3 is a kind of working principle diagram B of the temperature barrier provided in the first embodiment of the invention；
Fig. 4 is a kind of working principle diagram C of the temperature barrier provided in the first embodiment of the invention.
Description of symbols:
1- entrance liquid trap, 2- outlet liquid trap, the first cooling line of 3-, 4- substrate, the first connecting line of 5-,
The first cooling structure of 6-, 7- thrust chamber, the second cooling structure of 8-, the second connecting line of 9-, 10- third connecting line,
The 5th connecting line of 11-, the 4th connecting line of 12-.
Specific embodiment
Technical solution of the present invention is clearly and completely described below in conjunction with attached drawing, it is clear that described implementation
Example is a part of the embodiment of the present invention, instead of all the embodiments.Based on the embodiments of the present invention, ordinary skill
Personnel's every other embodiment obtained without making creative work, shall fall within the protection scope of the present invention.
Fig. 1 to Fig. 4 shows a kind of temperature barrier embodiment provided by the invention.
The temperature barrier is applied to liquid launch vehicle.Liquid launch vehicle includes propellant transfer system and engine
System, engine system include thrust chamber 7, precombustion chamber and gas generator.Temperature barrier includes being set to liquid embarkation fire
First cooling structure 6 of arrow bottom and the second cooling structure 8 for being set to thrust chamber 7.
Referring to FIG. 1, the first cooling structure 6 uses flat re-generatively cooled structure, including substrate 4, the first cooling line 3
And two liquid traps.Substrate 4 is the plate structure for being set to liquid launch vehicle bottom, and the first cooling line 3 is set to base
Inside plate 4, integral panel is formed with substrate 4.First cooling line 3 can take various forms, including but not limited to ripple
Plate and bundled tube structure, such as circular cross-section tubule, rectangular section tubule, U-shaped cross-sectional passage.Liquid trap includes being located at the
One cooling line, 3 arrival end and the entrance liquid trap 1 of outlet end and outlet liquid trap 2.
In the present embodiment, the first cooling line 3 is the annulus line extended around liquid launch vehicle bottom central axes, and
More are provided with, more annulus lines intensively arrange that the annulus line intensively arranged can play the function of ribs, and realization is held
Carry function.
Entrance liquid trap 1 is connect by the first connecting line 5 with the propellant transfer system of liquid launch vehicle, is realized and is divided
Flow purpose.Propellant transfer system conveys propellant to the first cooling line 3 by the first connecting line 5 and entrance liquid trap 1.
Propellant is cryogenic media, takes away the gas of liquid launch vehicle bottom by way of heat convection in the first cooling line 3
Dynamic heating and radiant heating.It can be seen from the above description that the present invention realizes liquid launch vehicle bottom using the above structure
Portion's thermal protection is integrated with aerodynamic force bearing function.
It exports liquid trap 2 and realizes interflow function, and be connected by the second connecting line 9 with the second cooling structure 8.Second is cold
But pipeline is set to thrust chamber 7, is cooled down by the heat convection realization of propellant and thrust chamber 7 to thrust chamber 7, second is cold
But structure 8 is connected by third connecting line 10 with thrust chamber 7.
The temperature barrier further includes the 4th connecting line 12 for being connected to propellant transfer system and the second connecting line 9,
And the 5th connecting line 11 of connection outlet liquid trap 2 and thrust chamber 7.Corresponding second connecting line 9, the 4th connecting line
12 and the 5th connecting line 11 be respectively arranged with the second control valve, the 4th control valve and the 5th control valve.The temperature barrier is also
Including the controller for being set to the temperature-detecting device in 6 exit of the first cooling structure and being electrically connected with temperature-detecting device,
Controller controls the second control valve, the 4th control valve and the 5th control valve according to the temperature information that temperature-detecting device detects.Tool
Body control is as follows:
Propellant enters the first cooling line 3 by entrance liquid trap 1 by propellant transport system pipeline, to liquid embarkation
After rocket bottom structure carries out convection current cooling, from outlet, liquid trap 2 flows out.According to the cooling from first of temperature-detecting device detection
The propellant temperature that structure 6 flows out, there are three types of schemes for flow direction:
1 when the propellant temperature of outflow is lower, and controller controls the second control valve and opens, the 4th control valve and the 5th control
Valve processed is closed.Referring to FIG. 2, propellant enters the second cooling knot by the second connecting line 9 after the outflow of the first cooling structure 6
Structure 8；The propellant flowed out from the second cooling structure 8 enters 7/ pre-burning of thrust chamber of engine system by third connecting line 10
Room/gas generator participates in burning；I.e. liquid launch vehicle bottom first relatively low to thermal environment cools down, then to thermal environment
More harsh motor power room 7 is cooled down.
2 when the propellant temperature of outflow is higher, is unable to satisfy the cooling requirement of thrust chamber 7, the 5th control of controller control
Valve is opened, and the second control valve and the 4th control valve are closed, referring to FIG. 3, the propellant of the first cooling structure 6 outflow passes through the 5th
7/ precombustion chamber of thrust chamber/gas generator that connecting line 11 is directly entered engine system participates in burning；
3 when the propellant temperature of outflow is higher, also can control the second control valve and the conducting of the 4th control valve, the 5th control
Valve processed is closed, referring to FIG. 4, the lower propellant of temperature from propellant transport system is made to pass through the 4th connecting line 12
It is blended with the higher propellant of temperature flowed out from the first cooling structure 6, subsequently into the second cooling structure 8.From second
The propellant that cooling structure 8 flows out enters 7/ precombustion chamber of thrust chamber/combustion gas hair of engine system by third connecting line 10
Raw device participates in burning；
First cooling structure 6 uses the refractory metal materials such as titanium alloy, stainless steel, specific according to liquid launch vehicle
The suitable structural material of the selections such as aerodynamic force bearing requirements, thermal environment condition and propellant physical property.First cooling line 3 and substrate
4, which can choose identical material, can also use compatible different materials.According to different pipeline forms using molding, soldering, expansion
It dissipates the techniques such as weldering, hot isostatic pressing weldering and realizes structure preparation.
The present invention is designed in this way, and the efficient active thermal protection of liquid launch vehicle bottom, energy are realized by active cooling
The bottom structure temperature of liquid launch vehicle reentry stage is enough effectively reduced.And metal material is used, damage strong with bearing capacity
Hurt the high feature of tolerance, repeatable making under power thermal coupling effect for liquid launch vehicle bottom thermal protection structure may be implemented
With.First cooling line 3 itself has ribs effect, can be realized the integrated function of thermal protection carrying.It is additionally useful for heat
The propellant of exchange burns after completing cooling into engine system, and propellant is avoided to waste.
As another embodiment, the first cooling line 3 may include single direction, may also comprise multiple directions, more
The arrangement for flowing through liquid trap in a direction is realized.
As another embodiment, the flat re-generatively cooled panel an of entirety can be used in the first cooling structure 6,
Multiple panel splicing forms can be used, each panel have independent pipeline, respectively with propellant transport system and independent hair
Motivation pipeline is connected.
As another embodiment, which only includes the first cooling structure 6, and outlet liquid trap 2 passes through pipe
Line structure is directly conducted to thrust chamber 7.
The present invention can be used in reusable liquid launch vehicle and disposable liquid launch vehicle, and
The earth/planetary atmosphere reentry vehicle of braking technique is regurgitated using liquid launch vehicle engine.
Invention also provides a kind of liquid launch vehicle embodiments comprising any of the above-described temperature barrier.
Obviously, the above embodiments are merely examples for clarifying the description, and does not limit the embodiments.It is right
For those of ordinary skill in the art, can also make on the basis of the above description it is other it is various forms of variation or
It changes.There is no necessity and possibility to exhaust all the enbodiments.And it is extended from this it is obvious variation or
It changes still within the protection scope of the invention.
Claims (10)
1. a kind of temperature barrier, it is characterised in that: the first cooling structure (6) including being set to liquid launch vehicle bottom,
Described first cooling structure (6) one end is connected to propellant transfer system, and the other end is connected to liquid launch vehicle engine
Thrust chamber (7) carries out heat convection by the propellant and the liquid launch vehicle that flow through first cooling structure (6).
2. temperature barrier according to claim 1, it is characterised in that: first cooling structure (6) includes substrate
(4), it is set to the first cooling line (3) of the substrate (4) and is located at the first cooling line (3) arrival end and outlet
The entrance liquid trap (1) at end and outlet liquid trap (2).
3. temperature barrier according to claim 2, it is characterised in that: first cooling line (3) is provided with more,
More first cooling lines (3) are intensively arranged.
4. temperature barrier according to claim 1, it is characterised in that: first cooling line (3) is around the liquid
The central axes of body carrier rocket bottom are extended.
5. temperature barrier according to claim 4, it is characterised in that: first cooling line (3) is is set to
It is internal to state substrate (4).
6. temperature barrier according to claim 5, it is characterised in that: first cooling line (3) be and the base
The integrally formed circular cross-section tubule of plate (4), rectangular section tubule or U-shaped cross-sectional passage.
7. temperature barrier according to claim 1 to 6, it is characterised in that: further include being set to described push away
The second cooling structure (8) of power room (7), the outlet end of first cooling body pass through the second connecting line (9) and described the
The arrival end of two cooling structures (8) is connected, and the outlet end of second cooling structure (8) is connected to the engine.
8. temperature barrier according to claim 7, it is characterised in that: further include being set to the propellant transfer system
With the 4th connecting line (12) between second connecting line (9), be connected to the outlet end of first cooling structure (6) with
5th connecting line (11) of the engine and corresponding second connecting line (9), the 4th connecting line (12)
With respectively arranged second control valve of the 5th connecting line (11), the 4th control valve and the 5th control valve.
9. temperature barrier according to claim 8, it is characterised in that: further include being set to first cooling structure
(6) temperature-detecting device in exit and the controller being electrically connected with the temperature-detecting device, the controller is according to institute
The temperature information for stating temperature-detecting device detection controls second control valve, the 4th control valve and the 5th control
Valve.
10. a kind of liquid launch vehicle, it is characterised in that: including temperature barrier of any of claims 1-9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910155308.6A CN109736974B (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910155308.6A CN109736974B (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN109736974A true CN109736974A (en) | 2019-05-10 |
| CN109736974B CN109736974B (en) | 2024-03-29 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201910155308.6A Active CN109736974B (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
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| Country | Link |
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| CN (1) | CN109736974B (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| CN112012852A (en) * | 2020-09-02 | 2020-12-01 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2013122163A (en) * | 2013-05-14 | 2014-11-20 | Открытое акционерное общество "Конструкторское бюро химавтоматики" | BOTTOM PROTECTION OF THE TAILING BODY OF THE CARRIER ROCKET |
| CN105468846A (en) * | 2015-11-24 | 2016-04-06 | 北京宇航系统工程研究所 | Method for determining bottom heat flow of rocket by radiation form factor |
| RU2610624C1 (en) * | 2016-01-20 | 2017-02-14 | Владислав Юрьевич Климов | Liquid-propellant rocket engine chamber |
| CN106640424A (en) * | 2016-10-26 | 2017-05-10 | 湖北航天技术研究院总体设计所 | Combustion chamber of liquid rocket engine |
| CN210106023U (en) * | 2019-02-28 | 2020-02-21 | 北京星际荣耀空间科技有限公司 | Thermal protection device and liquid carrier rocket |
-
2019
- 2019-02-28 CN CN201910155308.6A patent/CN109736974B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2013122163A (en) * | 2013-05-14 | 2014-11-20 | Открытое акционерное общество "Конструкторское бюро химавтоматики" | BOTTOM PROTECTION OF THE TAILING BODY OF THE CARRIER ROCKET |
| CN105468846A (en) * | 2015-11-24 | 2016-04-06 | 北京宇航系统工程研究所 | Method for determining bottom heat flow of rocket by radiation form factor |
| RU2610624C1 (en) * | 2016-01-20 | 2017-02-14 | Владислав Юрьевич Климов | Liquid-propellant rocket engine chamber |
| CN106640424A (en) * | 2016-10-26 | 2017-05-10 | 湖北航天技术研究院总体设计所 | Combustion chamber of liquid rocket engine |
| CN210106023U (en) * | 2019-02-28 | 2020-02-21 | 北京星际荣耀空间科技有限公司 | Thermal protection device and liquid carrier rocket |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| CN112012852A (en) * | 2020-09-02 | 2020-12-01 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
| CN112012852B (en) * | 2020-09-02 | 2021-06-18 | 西安航天动力测控技术研究所 | Reverse-injection protection and collection device and method for solid rocket engine |
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| Publication number | Publication date |
|---|---|
| CN109736974B (en) | 2024-03-29 |
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