CN210106023U - Thermal protection device and liquid carrier rocket - Google Patents
Thermal protection device and liquid carrier rocket Download PDFInfo
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- CN210106023U CN210106023U CN201920261387.4U CN201920261387U CN210106023U CN 210106023 U CN210106023 U CN 210106023U CN 201920261387 U CN201920261387 U CN 201920261387U CN 210106023 U CN210106023 U CN 210106023U
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- carrier rocket
- cooling structure
- liquid carrier
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Abstract
The utility model provides a hot protector and liquid carrier rocket, hot protector is including setting up in the first cooling structure of liquid carrier rocket bottom, first cooling structure one end and propellant conveying system intercommunication, the other end intercommunication extremely the thrust room of liquid carrier rocket engine, through flowing through the propellant of first cooling structure with liquid carrier rocket carries out the convection heat transfer. By 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, 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. The cooled propellant enters an engine system for combustion, and the reasonable utilization of the propellant is realized.
Description
Technical Field
The utility model relates to an aerospace equipment technical field, concretely relates to hot protector and liquid carrier rocket.
Background
The bottom of the liquid carrier rocket is provided with a liquid carrier rocket engine pipeline, electrical equipment and the like due to pneumatic heating and radiation heating generated by jet flow of a liquid carrier rocket engine, and thermal protection measures are required. The bottom of the existing disposable liquid carrier rocket is generally made of composite materials such as glass fiber reinforced plastics, flexible high-temperature heat-insulating cloth and the like to form a bottom heat protection structure. The reusable liquid carrier rocket is subjected to the double effects of atmospheric pneumatic heating and jet flow heating in the process of returning and re-entering, the thermal environment is obviously higher than that of the traditional disposable liquid carrier rocket, and the bottom thermal protection faces huge technical challenges.
The traditional passive thermal protection material and structure realize thermal protection by utilizing the material ablation principle or the characteristics of high temperature resistance and low thermal conductivity of the material. The ablation material can have serious ablation and denudation problems under the conditions of high heat flow and high dynamic pressure in the reentry process, the thickness of a thermal protection structure needs to be improved to meet the thermal protection performance of the reentry section, the weight is greatly improved, the thermal protection layer needs to be replaced after the reentry section returns at each time, and the repeated use cannot be realized. High-performance non-ablative high-temperature resistant materials, such as novel C/SiC, C/C, ultra-high 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.
SUMMERY OF THE UTILITY MODEL
Therefore, the to-be-solved technical problem of the utility model lies in overcoming the hot protective structure heat protection inefficiency among the prior art, but not enough defect that reusability is poor to a hot protector and liquid carrier rocket are provided.
The utility model provides a pair of hot protector, including setting up in the first cooling structure of liquid carrier rocket bottom, first cooling structure one end and propellant conveying system intercommunication, the other end communicates to the thrust room of liquid carrier rocket engine, through flowing through the propellant of first cooling structure with liquid carrier rocket carries out the convection heat transfer.
Optionally, the first cooling structure includes a base plate, a first cooling pipeline disposed on the base plate, and an inlet liquid trap and an outlet liquid trap located 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 a 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, an outlet end of the first cooling structure is connected with an inlet end of the second cooling structure through a second connecting pipeline, and an outlet end of the second cooling structure is communicated with the engine.
Optionally, the engine 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 cooling system further comprises a temperature detection device arranged at an 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 utility model discloses provide a liquid carrier rocket simultaneously, including above-mentioned arbitrary hot protector.
The utility model discloses technical scheme has following advantage:
the utility model provides a pair of hot protector utilizes the propellant low temperature characteristic of liquid carrier rocket engine, carries out the initiative cooling to liquid carrier rocket bottom, has greatly promoted the hot barrier propterty of carrier rocket bottom, and the hot environmental scope that is suitable for surpasss traditional passive form thermal protection structure. The cooled propellant enters an engine system for combustion, and the 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 used in the embodiments or the technical solutions in the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to these drawings without creative efforts.
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 working schematic diagram a of a thermal protection device according to a first embodiment of the present invention;
fig. 3 is a working schematic diagram B of a thermal protection device according to a first embodiment of the present invention;
fig. 4 is a working schematic diagram C of a thermal protection device according to a first embodiment of the present invention.
Description of reference numerals:
1-an inlet liquid collector, 2-an outlet liquid collector, 3-a first cooling pipeline, 4-a base plate, 5-a first connecting pipeline, 6-a first cooling structure, 7-a thrust chamber, 8-a second cooling structure, 9-a second connecting pipeline, 10-a third connecting pipeline, 11-a fifth connecting pipeline and 12-a fourth connecting pipeline.
Detailed Description
The technical solution of the present invention will be described clearly and completely with reference to the accompanying drawings, and obviously, the described embodiments are some, but not all embodiments of the present invention. Based on the embodiments in the present invention, all other embodiments obtained by a person skilled in the art without creative work belong to the protection scope of the present invention.
Fig. 1 to 4 show an embodiment of a thermal protection device provided by the present invention.
The thermal protection device is applied to a liquid carrier rocket. The liquid launch vehicle comprises a propellant delivery system and an engine system comprising a thrust chamber 7, a pre-combustion chamber and a gas generator. The thermal protection device comprises a first cooling structure 6 arranged at the bottom of the liquid launch vehicle and a second cooling structure 8 arranged at the thrust chamber 7.
Referring to fig. 1, the first cooling structure 6 is a flat regenerative cooling structure, and includes a substrate 4, a first cooling pipeline 3, and two liquid collectors. The base plate 4 is a plate-like structure provided at the bottom of the liquid carrier rocket, and the first cooling line 3 is provided inside the base plate 4 and forms an integrated panel with the base plate 4. The first cooling line 3 may take a variety of forms including, but not limited to, corrugated plate and tube bundle type structures such as a circular cross-section tubule, a rectangular cross-section tubule, a U-shaped cross-section channel, and the like. The liquid trap comprises an inlet liquid trap 1 and an outlet liquid trap 2 located 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 which are densely arranged, and the densely arranged annular pipelines can play a role of reinforcing ribs to realize a bearing function.
The inlet liquid collector 1 is connected with a propellant conveying system of the liquid carrier rocket through a first connecting pipeline 5, and the purpose of shunting is achieved. The propellant feed system feeds propellant to the first cooling line 3 via the first connecting line 5 and the inlet liquid trap 1. The propellant is a low-temperature medium, and takes away pneumatic heating and radiation heating at the bottom of the liquid carrier rocket in the form of convection heat exchange in the first cooling pipeline 3. As can be seen from the above description, the utility model adopts the above structure to realize the integration of liquid carrier rocket bottom thermal protection and aerodynamic force bearing function.
The outlet liquid collector 2 realizes a confluence function and is connected with the second cooling structure 8 through a second connecting pipeline 9. The second cooling pipeline is arranged in the thrust chamber 7, the thrust chamber 7 is cooled through convective heat transfer 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. A second control valve, a fourth control valve and a fifth control valve are respectively provided corresponding to the second connecting pipeline 9, the fourth connecting pipeline 12 and the fifth connecting pipeline 11. The thermal 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, and 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 control is as follows:
the propellant enters the first cooling pipeline 3 through the inlet liquid collector 1 by a propellant transport system pipeline, and flows out of the outlet liquid collector 2 after the bottom structure of the liquid carrier rocket is cooled in a convection manner. According to the temperature of the propellant flowing out of the first cooling structure 6 detected by the temperature detection device, the flow direction has three schemes:
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 are 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 pipeline 9; the propellant flowing out of the second cooling structure 8 enters a thrust chamber 7/a pre-combustion chamber/a gas generator of the engine system through a third connecting pipeline 10 to participate in combustion; namely, the bottom of the liquid carrier rocket with relatively low thermal environment is cooled, and then the engine thrust chamber 7 with more severe thermal environment is cooled.
2 when the temperature of the outflowing propellant is higher, the cooling requirement of the thrust chamber 7 cannot be met, the controller controls the fifth control valve to be opened, and the second control valve and the fourth control valve are closed, please refer to fig. 3, the propellant outflowing from the first cooling structure 6 directly enters the thrust chamber 7/precombustion chamber/gas generator of the engine system through the fifth connecting pipeline 11 to participate in combustion;
3 when the temperature of the outflowing propellant is higher, the second control valve and the fourth control valve can also be controlled to be switched on, and the fifth control valve is switched off, please refer to fig. 4, so that the propellant with lower temperature in the propellant transportation system is mixed with the propellant with higher temperature outflowing from the first cooling structure 6 through the fourth connecting pipeline 12 and then enters the second cooling structure 8. The propellant flowing out of the second cooling structure 8 enters a thrust chamber 7/a pre-combustion chamber/a gas generator of the engine system through a third connecting pipeline 10 to participate in combustion;
the first cooling structure 6 is made of high-temperature resistant metal materials such as titanium alloy and stainless steel, and suitable structural materials are selected according to the specific aerodynamic bearing requirements, the thermal environment conditions, the physical properties of the propellant and the like of the liquid carrier rocket. The first cooling channels 3 and the base plate 4 may be made of the same material or may be made of different compatible materials. And the structure is prepared by adopting the processes of die pressing, brazing, diffusion welding, hot isostatic pressing welding and the like according to different pipeline forms.
The utility model discloses so design, through the high-efficient initiative thermal protection of active cooling in order to realize liquid carrier rocket bottom, can effectively reduce the bottom structure temperature of liquid carrier rocket reentry section. And the metal material is adopted, so that the liquid carrier rocket bottom thermal protection structure has the characteristics of strong bearing capacity and high damage tolerance, and can be repeatedly used under the action of force-heat coupling. The first cooling pipeline 3 has a reinforcing rib effect, and the integrated function of heat protection bearing can be realized. In addition, the propellant for heat exchange enters an engine system for combustion after being cooled, so that the waste of the propellant is avoided.
As another embodiment, the first cooling circuit 3 may comprise a single direction, or may comprise multiple directions, with the flow in multiple directions being achieved by the arrangement of the liquid trap.
As another embodiment, the first cooling structure 6 may be an integral flat regenerative cooling panel, or may be a plurality of panels, each having separate pipes connected to the propellant transport system and separate engine pipes.
As another embodiment, the thermal protection device comprises only the first cooling structure 6, the outlet liquid collector 2 leading directly to the thrust chamber 7 through a pipe structure.
The utility model discloses can be used in repeatedly usable liquid carrier rocket and disposable liquid carrier rocket to and adopt the earth/planet atmosphere reentry aircraft of liquid carrier rocket engine backspray deceleration technique.
The utility model discloses provide a liquid carrier rocket embodiment simultaneously, it includes the hot protector of above-mentioned arbitrary item.
It should be understood that the above examples are only for clarity of illustration and are not intended to limit the embodiments. Other variations and modifications will be apparent to persons skilled in the art in light of the above description. And are neither required nor exhaustive of all embodiments. And obvious variations or modifications can be made without departing from the scope of the invention.
Claims (10)
1. A thermal protection device characterized by: the liquid carrier rocket engine cooling structure comprises a first cooling structure (6) arranged at the bottom of a liquid carrier rocket, 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 to a thrust chamber (7) of the liquid carrier rocket engine, and convective heat exchange is carried out between the propellant flowing through the first cooling structure (6) and the liquid carrier rocket.
2. The thermal shield apparatus of claim 1, wherein: 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).
3. The thermal shield apparatus of claim 2, wherein: the first cooling pipelines (3) are arranged in a plurality of rows, and the first cooling pipelines (3) are arranged densely.
4. The thermal shield apparatus of claim 2, wherein: the first cooling pipeline (3) is arranged around the central axis of the bottom of the liquid carrier rocket in an extending mode.
5. The thermal shield apparatus of claim 4, wherein: the first cooling pipeline (3) is arranged inside the substrate (4).
6. The thermal shield apparatus of claim 5, wherein: the first cooling pipeline (3) is a circular section thin pipe, a rectangular section thin pipe or a U-shaped section channel which is integrally formed with the substrate (4).
7. The thermal protection device according to any one of claims 1-6, wherein: the engine further comprises a second cooling structure (8) arranged in the thrust chamber (7), the outlet end of the first cooling structure (6) 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.
8. The thermal shield apparatus of claim 7, wherein: the cooling system is characterized by further comprising a fourth connecting pipeline (12) arranged between the propellant conveying system and the second connecting pipeline (9), and a second control valve, a fourth control valve and a fifth control valve, wherein the outlet end of the first cooling structure (6) is communicated with a fifth connecting pipeline (11) of the engine and corresponds to the second connecting pipeline (9), the fourth connecting pipeline (12) and the fifth connecting pipeline (11) are respectively arranged.
9. The thermal shield apparatus of claim 8, wherein: the cooling system further comprises a temperature detection device arranged at an 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.
10. A liquid launch vehicle characterized by: comprising the thermal protection device according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201920261387.4U CN210106023U (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 |
|---|---|---|---|
| CN201920261387.4U CN210106023U (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
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| Publication Number | Publication Date |
|---|---|
| CN210106023U true CN210106023U (en) | 2020-02-21 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201920261387.4U Active CN210106023U (en) | 2019-02-28 | 2019-02-28 | Thermal protection device and liquid carrier rocket |
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| CN (1) | CN210106023U (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109736974A (en) * | 2019-02-28 | 2019-05-10 | 北京星际荣耀空间科技有限公司 | A kind of temperature barrier and liquid launch vehicle |
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2019
- 2019-02-28 CN CN201920261387.4U patent/CN210106023U/en active Active
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109736974A (en) * | 2019-02-28 | 2019-05-10 | 北京星际荣耀空间科技有限公司 | A kind of temperature barrier and liquid launch vehicle |
| CN109736974B (en) * | 2019-02-28 | 2024-03-29 | 北京星际荣耀空间科技股份有限公司 | Thermal protection device and liquid carrier rocket |
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Address after: 100045 1-14-214, 2nd floor, 136 Xiwai street, Xicheng District, Beijing Patentee after: Beijing Star glory Space Technology Co.,Ltd. Address before: 329, floor 3, building 1, No. 9, Desheng South Street, Daxing Economic and Technological Development Zone, Beijing 100176 Patentee before: BEIJING XINGJIRONGYAO SPACE TECHNOLOGY Co.,Ltd. |