CN109736974B - Thermal protection device and liquid carrier rocket - Google Patents
Thermal protection device and liquid carrier rocket Download PDFInfo
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- CN109736974B CN109736974B CN201910155308.6A CN201910155308A CN109736974B CN 109736974 B CN109736974 B CN 109736974B CN 201910155308 A CN201910155308 A CN 201910155308A CN 109736974 B CN109736974 B CN 109736974B
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- 239000007788 liquid Substances 0.000 title claims abstract description 63
- 238000001816 cooling Methods 0.000 claims abstract description 84
- 239000003380 propellant Substances 0.000 claims abstract description 38
- 238000001514 detection method Methods 0.000 claims description 10
- 239000000758 substrate Substances 0.000 claims description 6
- 210000005239 tubule Anatomy 0.000 claims description 6
- 238000002485 combustion reaction Methods 0.000 abstract description 6
- 239000000463 material Substances 0.000 description 8
- 238000010438 heat treatment Methods 0.000 description 6
- 101100204059 Caenorhabditis elegans trap-2 gene Proteins 0.000 description 5
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- 235000012633 Iberis amara Nutrition 0.000 description 2
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- 230000001172 regenerating effect Effects 0.000 description 2
- 230000003014 reinforcing effect Effects 0.000 description 2
- 238000003466 welding Methods 0.000 description 2
- 229910001069 Ti alloy Inorganic materials 0.000 description 1
- 230000009471 action Effects 0.000 description 1
- 238000005219 brazing Methods 0.000 description 1
- 239000011203 carbon fibre reinforced carbon Substances 0.000 description 1
- 239000011204 carbon fibre-reinforced silicon carbide Substances 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000008878 coupling Effects 0.000 description 1
- 238000010168 coupling process Methods 0.000 description 1
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- 238000009792 diffusion process Methods 0.000 description 1
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- 238000005516 engineering process Methods 0.000 description 1
- 239000004744 fabric Substances 0.000 description 1
- 239000011152 fibreglass Substances 0.000 description 1
- 238000001513 hot isostatic pressing Methods 0.000 description 1
- 230000010354 integration Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000000704 physical effect Effects 0.000 description 1
- 238000002360 preparation method Methods 0.000 description 1
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- 239000010935 stainless steel Substances 0.000 description 1
- 229910001220 stainless steel Inorganic materials 0.000 description 1
- 239000011215 ultra-high-temperature ceramic Substances 0.000 description 1
- 239000002699 waste material Substances 0.000 description 1
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- Cooling, Air Intake And Gas Exhaust, And Fuel Tank Arrangements In Propulsion Units (AREA)
- Plasma Technology (AREA)
Abstract
The invention provides a heat protection device and a liquid carrier rocket, wherein the heat protection device comprises a first cooling structure arranged at the bottom of the liquid carrier rocket, one end of the first cooling structure is communicated with a propellant conveying system, the other end of the first cooling structure is communicated to a thrust chamber of an engine of the liquid carrier rocket, and the first cooling structure performs convection heat exchange with the liquid carrier rocket through the propellant flowing through the first cooling structure. By means of the design, the propellant low-temperature characteristic of the liquid carrier rocket engine is utilized to actively cool the bottom of the liquid carrier rocket, so that the thermal protection performance of the bottom of the carrier rocket is greatly improved, and the applicable thermal environment range exceeds that of a traditional passive thermal protection structure. And the cooled propellant enters an engine system for combustion, so that reasonable utilization of the propellant is realized.
Description
Technical Field
The invention relates to the technical field of aerospace equipment, in particular to a heat protection device and a liquid carrier rocket.
Background
The bottom of the liquid carrier rocket is provided with liquid carrier rocket engine pipelines, electric equipment and the like due to pneumatic heating and radiation heating generated by jet flow of the liquid carrier rocket engine, and thermal protection measures need to be adopted. The bottom of the existing disposable liquid carrier rocket is generally made of glass fiber reinforced plastic, high-flexibility Wen Jue heat cloth and other composite materials. The reusable liquid carrier rocket is subjected to double influences of atmospheric pneumatic heating and jet flow heating in the returning and re-entering process, the thermal environment is obviously higher than that of the traditional disposable liquid carrier rocket, and the bottom thermal protection faces a huge technical challenge.
Traditional passive heat protection materials and structures utilize the material ablation principle or the high temperature resistance and low heat conductivity characteristics of the material to realize heat protection. The ablation material has serious ablation and ablation problems under the conditions of high heat flow and high dynamic pressure in the reentry process, the thickness of the thermal protection structure needs to be increased to meet the thermal protection performance of the reentry section, the weight is greatly increased, and the thermal protection layer needs to be replaced after each return, so that the thermal protection layer cannot be reused. The high-performance non-ablative high-temperature resistant materials, such as novel C/SiC, C/C, ultrahigh-temperature ceramics and the like, have the problems of high processing and production difficulty, high cost, small damage tolerance, high brittleness and the like, and are easy to damage under the condition of repeated use.
Disclosure of Invention
Therefore, the technical problem to be solved by the invention is to overcome the defects of low heat protection efficiency and poor reusability of the heat protection structure in the prior art, thereby providing the heat protection device and the liquid carrier rocket.
The invention provides a heat protection device, which comprises a first cooling structure arranged at the bottom of a liquid carrier rocket, wherein one end of the first cooling structure is communicated with a propellant conveying system, the other end of the first cooling structure is communicated with a thrust chamber of a liquid carrier rocket engine, and the first cooling structure and the liquid carrier rocket are subjected to heat convection through a propellant flowing through the first cooling structure.
Optionally, the first cooling structure includes a substrate, a first cooling pipeline disposed on the substrate, and an inlet liquid collector and an outlet liquid collector disposed at an inlet end and an outlet end of the first cooling pipeline.
Optionally, the first cooling pipelines are provided with a plurality of first cooling pipelines, and the plurality of first cooling pipelines are densely arranged.
Optionally, the first cooling pipeline extends around the central axis of the bottom of the liquid carrier rocket.
Optionally, the first cooling pipeline is disposed inside the substrate.
Optionally, the first cooling pipeline is a circular section tubule, a rectangular section tubule or a U-shaped section channel integrally formed with the substrate.
Optionally, the engine further comprises a second cooling structure arranged in the thrust chamber, wherein the outlet end of the first cooling structure is connected with the inlet end of the second cooling structure through a second connecting pipeline, and the outlet end of the second cooling structure is communicated with the engine.
Optionally, the system further comprises a fourth connecting pipeline arranged between the propellant conveying system and the second connecting pipeline, a fifth connecting pipeline for communicating the outlet end of the first cooling structure with the engine, and a second control valve, a fourth control valve and a fifth control valve which are respectively arranged corresponding to the second connecting pipeline, the fourth connecting pipeline and the fifth connecting pipeline.
Optionally, the device further comprises a temperature detection device arranged at the outlet of the first cooling structure and a controller electrically connected with the temperature detection device, wherein the controller controls the second control valve, the fourth control valve and the fifth control valve according to temperature information detected by the temperature detection device.
The invention also provides a liquid carrier rocket, which comprises the heat protection device.
The technical scheme of the invention has the following advantages:
according to the thermal protection device provided by the invention, the propellant low-temperature characteristic of the liquid carrier rocket engine is utilized to actively cool the bottom of the liquid carrier rocket, so that the thermal protection performance of the bottom of the carrier rocket is greatly improved, and the applicable thermal environment range exceeds that of a traditional passive thermal protection structure. And the cooled propellant enters an engine system for combustion, so that reasonable utilization of the propellant is realized.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings that are needed in the description of the embodiments or the prior art will be briefly described, and it is obvious that the drawings in the description below are some embodiments of the present invention, and other drawings can be obtained according to the drawings without inventive effort for a person skilled in the art.
Fig. 1 is a schematic structural view of a first cooling structure provided in a first embodiment of the present invention;
fig. 2 is a schematic diagram a illustrating the operation of a thermal protection device according to a first embodiment of the present invention;
fig. 3 is a schematic diagram B of a thermal protection device according to a first embodiment of the present invention;
fig. 4 is a schematic diagram C of a thermal protection device according to a first embodiment of the present invention.
Reference numerals illustrate:
1-inlet liquid collector, 2-outlet liquid collector, 3-first cooling pipeline, 4-base plate, 5-first connecting pipeline, 6-first cooling structure, 7-thrust chamber, 8-second cooling structure, 9-second connecting pipeline, 10-third connecting pipeline, 11-fifth connecting pipeline, 12-fourth connecting pipeline.
Detailed Description
The following description of the embodiments of the present invention will be made apparent and fully in view of the accompanying drawings, in which some, but not all embodiments of the invention are shown. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Fig. 1 to 4 show an embodiment of a thermal protection device provided by the present invention.
The heat protection device is applied to a liquid carrier rocket. The liquid carrier rocket comprises a propellant delivery system and an engine system comprising a thrust chamber 7, a prechamber and a gas generator. The thermal protection device comprises a first cooling structure 6 arranged at the bottom of the liquid carrier rocket and a second cooling structure 8 arranged at the thrust chamber 7.
Referring to fig. 1, the first cooling structure 6 is a flat-plate type regenerative cooling structure, and includes a substrate 4, a first cooling pipe 3, and two liquid collectors. The base plate 4 is a plate-shaped structure arranged at the bottom of the liquid carrier rocket, and the first cooling pipeline 3 is arranged inside the base plate 4 and forms an integrated panel with the base plate 4. The first cooling circuit 3 may take a variety of forms including, but not limited to, corrugated plate and tube bundle structures such as circular cross-section tubules, rectangular cross-section tubules, U-section channels, etc. The liquid trap comprises an inlet liquid trap 1 and an outlet liquid trap 2 at the inlet end and the outlet end of the first cooling line 3, respectively.
In this embodiment, the first cooling pipeline 3 is an annular pipeline extending around the central axis of the bottom of the liquid carrier rocket, and is provided with a plurality of annular pipelines, the plurality of annular pipelines are densely arranged, and the densely arranged annular pipelines can play a role of reinforcing ribs, so that a bearing function is realized.
The inlet liquid collector 1 is connected with a propellant conveying system of the liquid carrier rocket through a first connecting pipeline 5, so that the purpose of diversion is realized. The propellant delivery system delivers propellant to the first cooling circuit 3 via the first connection circuit 5 and the inlet accumulator 1. The propellant is a low-temperature medium, and the pneumatic heating and the radiation heating of the bottom of the liquid carrier rocket are carried away in the first cooling pipeline 3 in a convection heat exchange mode. From the above description, the invention adopts the structure to realize the integration of the bottom heat protection and aerodynamic bearing functions of the liquid carrier rocket.
The outlet liquid trap 2 performs a converging function and is connected to the second cooling structure 8 via a second connecting line 9. The second cooling pipeline is arranged in the thrust chamber 7, the thrust chamber 7 is cooled through convection heat exchange between the propellant and the thrust chamber 7, and the second cooling structure 8 is connected with the thrust chamber 7 through a third connecting pipeline 10.
The thermal protection device further comprises a fourth connecting line 12 communicating the propellant delivery system with the second connecting line 9 and a fifth connecting line 11 communicating the outlet liquid trap 2 with the thrust chamber 7. The second, fourth and fifth control valves are provided corresponding to the second, fourth and fifth connecting lines 9, 12 and 11, respectively. The heat protection device further comprises a temperature detection device arranged at the outlet of the first cooling structure 6 and a controller electrically connected with the temperature detection device, wherein the controller controls the second control valve, the fourth control valve and the fifth control valve according to temperature information detected by the temperature detection device. The specific control is as follows:
propellant enters the first cooling pipeline 3 through the inlet liquid trap 1 by the propellant transportation system pipeline, and flows out of the outlet liquid trap 2 after convective cooling of the bottom structure of the liquid carrier rocket. Based on the temperature of the propellant flowing out of the first cooling structure 6 detected by the temperature detection means, there are three schemes for its flow direction:
and 1, when the temperature of the outflowing propellant is lower, the controller controls the second control valve to be opened, and the fourth control valve and the fifth control valve to be closed. Referring to fig. 2, the propellant flows out of the first cooling structure 6 and then enters the second cooling structure 8 through the second connecting pipe 9; propellant flowing out of the second cooling structure 8 enters the thrust chamber 7/prechamber/gas generator of the engine system via the third connecting line 10 for combustion; namely, the bottom of the liquid carrier rocket with relatively low thermal environment is cooled firstly, and then the engine thrust chamber 7 with more severe thermal environment is cooled.
2 when the temperature of the flowing out propellant is higher, the cooling requirement of the thrust chamber 7 cannot be met, the controller controls the fifth control valve to be opened, the second control valve and the fourth control valve to be closed, and referring to fig. 3, the propellant flowing out of the first cooling structure 6 directly enters the thrust chamber 7/the precombustor/the gas generator of the engine system through the fifth connecting pipeline 11 to participate in combustion;
3 when the temperature of the exiting propellant is higher, the second control valve and the fourth control valve can also be controlled to be turned on, the fifth control valve is turned off, please refer to fig. 4, so that the propellant with a lower temperature from the propellant transportation system is mixed with the propellant with a higher temperature exiting the first cooling structure 6 through the fourth connecting line 12 and then enters the second cooling structure 8. Propellant flowing out of the second cooling structure 8 enters the thrust chamber 7/prechamber/gas generator of the engine system via the third connecting line 10 for combustion;
the first cooling structure 6 is made of high-temperature resistant metal materials such as titanium alloy and stainless steel, and the appropriate structural materials are selected according to specific aerodynamic force bearing requirements, thermal environment conditions, physical properties of the propellant and the like of the liquid carrier rocket. The first cooling line 3 and the base plate 4 may be made of the same material or of compatible different materials. And adopting processes such as mould pressing, brazing, diffusion welding, hot isostatic pressing welding and the like according to different pipeline forms to realize structure preparation.
The invention is designed in such a way, and the high-efficiency active heat protection of the bottom of the liquid carrier rocket is realized through active cooling, so that the temperature of the bottom structure of the reentry section of the liquid carrier rocket can be effectively reduced. And the metal material is adopted, so that the device has the characteristics of strong bearing capacity and high damage tolerance, and can realize the reusability of the bottom thermal protection structure of the liquid carrier rocket under the action of force thermal coupling. The first cooling pipeline 3 has a reinforcing rib effect, and can realize the integrated function of heat protection bearing. In addition, the propellant for heat exchange enters the engine system for combustion after being cooled, so that the propellant waste is avoided.
As another embodiment, the first cooling line 3 may comprise a single direction, and may also comprise multiple directions of flow through an arrangement of liquid traps.
As another embodiment, the first cooling structure 6 may be an integral flat regenerative cooling panel, or may be a plurality of panels spliced, each panel having separate pipes connected to the propellant transportation system and separate engine pipes, respectively.
As another embodiment, the thermal protection device comprises only the first cooling structure 6, the outlet liquid trap 2 being directly connected to the thrust chamber 7 by a piping structure.
The invention can be applied to reusable liquid carrier rockets, disposable liquid carrier rockets and earth/planet atmosphere reentry vehicles adopting liquid carrier rocket engine back-spraying deceleration technology.
The invention also provides a liquid carrier rocket embodiment, which comprises any one of the heat protection devices.
It is apparent that the above examples are given by way of illustration only and are not limiting of the embodiments. Other variations or modifications of the above teachings will be apparent to those of ordinary skill in the art. It is not necessary here nor is it exhaustive of all embodiments. While still being apparent from variations or modifications that may be made by those skilled in the art are within the scope of the invention.
Claims (6)
1. A thermal protection device, characterized by: the device comprises a first cooling structure (6) arranged at the bottom of a liquid carrier rocket, wherein one end of the first cooling structure (6) is communicated with a propellant conveying system, the other end of the first cooling structure is communicated with a thrust chamber (7) of a liquid carrier rocket engine, and the first cooling structure (6) and the liquid carrier rocket perform heat convection through the propellant flowing through the first cooling structure (6);
the engine is characterized by further comprising a second cooling structure (8) arranged on the thrust chamber (7), wherein the outlet end of the first cooling structure is connected with the inlet end of the second cooling structure (8) through a second connecting pipeline (9), and the outlet end of the second cooling structure (8) is communicated with the engine;
the first cooling structure (6) comprises a base plate (4), a first cooling pipeline (3) arranged on the base plate (4), and an inlet liquid collector (1) and an outlet liquid collector (2) which are positioned at the inlet end and the outlet end of the first cooling pipeline (3);
the first cooling pipeline (3) is arranged around the central axis of the bottom of the liquid carrier rocket in an extending way;
the first cooling pipelines (3) are provided with a plurality of cooling pipelines, and the plurality of first cooling pipelines (3) are densely arranged.
2. The thermal protection device of claim 1, wherein: the first cooling pipeline (3) is arranged inside the base plate (4).
3. The thermal protection device of claim 2, wherein: the first cooling pipeline (3) is a circular section tubule, a rectangular section tubule or a U-shaped section channel which is integrally formed with the substrate (4).
4. A thermal protection device according to any one of claims 1-3, wherein: the device also comprises a fourth connecting pipeline (12) arranged between the propellant conveying system and the second connecting pipeline (9), a fifth connecting pipeline (11) which is communicated with the outlet end of the first cooling structure (6) and corresponds to the engine, and a second control valve, a fourth control valve and a fifth control valve which are respectively arranged on the second connecting pipeline (9), the fourth connecting pipeline (12) and the fifth connecting pipeline (11).
5. The thermal protection device of claim 4, wherein: the cooling system further comprises a temperature detection device arranged at the outlet of the first cooling structure (6) and a controller electrically connected with the temperature detection device, wherein the controller controls the second control valve, the fourth control valve and the fifth control valve according to temperature information detected by the temperature detection device.
6. A liquid launch vehicle, characterized by: comprising a thermal protection device according to any one of claims 1-5.
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 |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910155308.6A CN109736974B (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
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| Publication Number | Publication Date |
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| CN109736974A CN109736974A (en) | 2019-05-10 |
| CN109736974B true CN109736974B (en) | 2024-03-29 |
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| CN201910155308.6A Active CN109736974B (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
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Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
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| 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 |
| 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 |
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 |
-
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 |
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