CN109131948B - Spacecraft tail flame protection and heat insulation device and spacecraft - Google Patents
Spacecraft tail flame protection and heat insulation device and spacecraft Download PDFInfo
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- CN109131948B CN109131948B CN201811034568.XA CN201811034568A CN109131948B CN 109131948 B CN109131948 B CN 109131948B CN 201811034568 A CN201811034568 A CN 201811034568A CN 109131948 B CN109131948 B CN 109131948B
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/52—Protection, safety or emergency devices; Survival aids
- B64G1/58—Thermal protection, e.g. heat shields
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- Aviation & Aerospace Engineering (AREA)
- Thermal Insulation (AREA)
- Cooling Or The Like Of Electrical Apparatus (AREA)
Abstract
The invention relates to a spacecraft tail flame protection and heat insulation device and a spacecraft, wherein the spacecraft tail flame protection and heat insulation device is mainly applied to protection and heat insulation of a main engine tail flame of a spacecraft with an outer bearing cylinder type configuration. The spacecraft tail flame protection heat insulation device disclosed by the invention has the beneficial effects of high temperature resistance, good heat insulation, light weight and the like.
Description
Technical Field
The invention relates to a space thermal control assembly, in particular to a spacecraft tail flame protection and heat insulation device and a spacecraft.
Background
Due to the need for orbital transfer or orbital maneuvers, spacecraft such as manned spacecraft, cargo spacecraft, deep space exploration spacecraft and the like are typically equipped with high thrust main engines for orbital control. When the main engine with high thrust is in operation, a spray pipe of the main engine can directly generate high-temperature tail flame, and the structure of a combustion chamber part of the main engine can radiate high-temperature heat flow. If the high-temperature tail flame and the high-temperature heat flow are not protected, adverse effects can be caused to instruments and equipment inside the spacecraft, the instruments and equipment are easily subjected to high-temperature radiation, the temperature is too high, and the reliability and the service life of the instruments and equipment are reduced.
In order to ensure that instruments and equipment on a spacecraft can reliably operate in a certain temperature environment, the spacecraft needs to be subjected to thermal control design in the design of the spacecraft. The spacecraft is generally provided with a plurality of cabins, a closed space needs to be arranged for the thermal control environment of instruments and equipment in the cabin, so that the instruments and equipment in the cabin are isolated from the severe space environment outside the cabin, and the temperature environment in the cabin meets the range requirement of normal operation of the instruments and equipment through an active and passive thermal control means.
The spacecraft such as manned spacecraft, cargo ship, deep space exploration spacecraft and the like generally adopts a propulsion cabin with an outer force bearing cylinder structure, the tail of the propulsion cabin is in an open form, and in order to prevent instruments and equipment in the propulsion cabin of the spacecraft from being influenced by high-temperature tail flame and high-temperature heat flow of a main engine, a heat insulation device is required to be arranged to enable the propulsion cabin to form a closed heat insulation space, so that the instruments and equipment installed in the propulsion cabin have a good temperature environment.
Disclosure of Invention
The invention aims to provide a spacecraft tail flame protection and heat insulation device and a spacecraft, and aims to solve the technical problem that instrument equipment in an outer force bearing cylinder type propulsion cabin is influenced by high-temperature tail flame and high-temperature heat flow of a main engine in the prior art.
In order to solve the problems, the invention provides a spacecraft tail flame protection and heat insulation device which comprises a frame structure, a heat insulation plate assembly and a heat insulation screen assembly, wherein the frame structure is used as a supporting structure of the whole device, the heat insulation plate assembly is arranged between the frame structure and the heat insulation screen assembly, and the heat insulation screen assembly is covered and installed on the outer sides of the frame structure and the heat insulation plate assembly to form a closed barrier layer.
Preferably, the frame structure comprises a support inner ring, a support middle ring, a support outer ring, a plurality of first reinforcement bars, a plurality of second reinforcement bars, and a plurality of third reinforcement bars, wherein,
the support inner ring is connected with the support middle ring through the first reinforcing rod to form an inverted frustum frame structure I;
the supporting middle ring is connected with the supporting outer ring through the second reinforcing rod to form a first circular truncated cone frame structure;
the support inner ring is connected with the support outer ring through the third reinforcing rod to form a circular truncated cone frame structure II.
The first round platform frame structure, the second round platform frame structure and the first inverted round platform frame structure are coaxial, and the second round platform frame structure and the first inverted round platform frame structure are located inside the first round platform frame structure.
Preferably, the first reinforcing rod, the second reinforcing rod and the third reinforcing rod are T-shaped stringer profiles.
Preferably, the heat shield assembly forms a complete wrap around the outside of the frame structure along an inside surface of the inverted frustum frame structure and an outside surface of the frustum frame structure.
Preferably, the support inner ring is composed of a plurality of first arc strips, the support middle ring is composed of a plurality of second arc strips, and the support outer ring is composed of a plurality of third arc strips.
Preferably, the first arc strip, the second arc strip and the third arc strip are factory-shaped stringer profiles.
Preferably, the heat insulation plate assembly is formed by splicing a plurality of aluminum alloy thin-wall plates, and a main engine spray pipe mounting hole is formed in the splicing position.
Preferably, the heat shield assembly is formed by stacking a plurality of thin-wall metal plates.
Preferably, the frame structure further comprises a plurality of first joints and a plurality of second joints for connecting with the spacecraft, the first joints are uniformly distributed on the support inner ring, and the second joints are uniformly distributed on the support outer ring.
Preferably, the support outer ring further comprises a mounting seat assembly, and each second joint is connected with the propulsion cabin structure ring frame of the spacecraft through the mounting seat assembly so as to realize the connection between the support outer ring and the propulsion cabin structure ring frame of the spacecraft.
Preferably, the mounting seat assembly comprises a plurality of mounting plate heat insulation cushion blocks, each second joint is in heat insulation connection with the mounting seat assembly through one mounting plate heat insulation cushion block, and the mounting seat assembly is connected with a propulsion cabin structure ring frame of the spacecraft.
Preferably, the aircraft further comprises auxiliary support rod assemblies, and each first joint is connected with a propulsion module on a propulsion cabin of the spacecraft through the auxiliary support rod assemblies so as to realize the connection of the support inner ring and the propulsion module on the propulsion cabin of the spacecraft.
Preferably, the auxiliary supporting rod assembly comprises a plurality of connecting rods, a plurality of third joints and a plurality of connecting rod supports, the number of the connecting rods is the same as that of the connecting rod supports, the number of the third joints is 2 times that of the connecting rods, each connecting rod is fixedly connected with the two ends of each connecting rod, each third joint at one end of each connecting rod is connected with a propulsion module on a propulsion cabin of the spacecraft through one connecting rod support, and the third joint at the other end is connected with one first joint.
Preferably, the auxiliary support rod assembly further comprises a plurality of connecting rod support heat insulation cushion blocks, and each connecting rod support is in heat insulation connection with a propulsion module on a propulsion cabin of the spacecraft through one connecting rod support heat insulation cushion block.
Compared with the prior art, the invention has the following technical effects:
1. the heat shield assembly and the spacecraft propulsion cabin structure form a closed heat insulation space, and high-temperature tail flames and high-temperature heat flows of the main engine can be simultaneously protected.
2. The welding structure design of factory-shaped and T-shaped stringer profiles is adopted, and the structural rigidity and the light weight design requirement are guaranteed.
3. The spacecraft tail flame protection heat insulation device has the beneficial effects of high temperature resistance, good heat insulation, light weight and the like.
Of course, it is not necessary for any product in which the invention is practiced to achieve all of the above-described advantages at the same time.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings needed to be used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts. In the drawings:
fig. 1 is an exploded view of the preferred embodiment of the present invention.
Fig. 2 is a schematic structural composition diagram of a frame structure according to a preferred embodiment of the present invention.
Fig. 3 is a schematic view showing the structural composition of the heat shield assembly according to the preferred embodiment of the present invention.
Fig. 4 is a schematic structural view of an auxiliary support rod assembly according to a preferred embodiment of the present invention.
Figure 5 is a schematic view of the installation of the preferred embodiment of the invention on a propulsion pod of an aircraft.
Fig. 6 is an enlarged view at I in fig. 5.
Fig. 7 is an enlarged view at II in fig. 5.
Detailed Description
The spacecraft tail flame protection and thermal insulation device provided by the invention will be described in detail with reference to the accompanying drawings, the embodiment is implemented on the premise of the technical scheme of the invention, and a detailed implementation mode and a specific operation process are given, but the protection scope of the invention is not limited to the following embodiment, and a person skilled in the art can modify and decorate the device within the scope of not changing the spirit and content of the invention.
Referring to fig. 1, a spacecraft tail flame protection and heat insulation device adopts a frame rod system structure, and includes a frame structure 1, a heat insulation plate assembly 2 and a heat insulation screen assembly 3. The frame structure 1 serves as a supporting structure of the whole device, is formed by welding in a connecting mode, and has good overall rigidity and bearing capacity. The heat insulation plate assembly 2 is arranged between the frame structure 1 and the heat insulation plate assembly 3, the heat insulation plate assembly 3 is covered and installed on the outer sides of the frame structure 1 and the heat insulation plate assembly 2 to form a closed barrier layer, and the closed barrier layer and the spacecraft propulsion cabin structure 6 form a closed cabin body together to block the main engine tail flame and high-temperature heat flow.
Referring to fig. 2, the frame structure 1 includes a support inner ring 101, a support middle ring 102, a support outer ring 103, a plurality of first stiffeners 104, a plurality of second stiffeners 105, and a plurality of third stiffeners 106, wherein,
the supporting inner ring 101 is connected with the supporting middle ring 102 through a first reinforcing rod 104 to form a first inverted frustum frame structure, wherein the first reinforcing rods 104 are uniformly distributed along the circumferential direction of the first inverted frustum frame structure;
the supporting middle ring 102 is connected with the supporting outer ring 103 through a second reinforcing rod 105 to form a first circular truncated cone frame structure, wherein the second reinforcing rods 105 are uniformly distributed along the circumferential direction of the first circular truncated cone frame structure;
the support inner ring 101 is connected with the support outer ring 103 through a third reinforcing rod 106 to form a second circular truncated cone frame structure, wherein the third reinforcing rods 106 are uniformly distributed along the circumferential direction of the second circular truncated cone frame structure.
The first round platform frame structure, the second round platform frame structure and the first inverted round platform frame structure are coaxial, and the second round platform frame structure and the first inverted round platform frame structure are located inside the first round platform frame structure.
As an example, the present device uses 16 first reinforcing bars 104, 16 second reinforcing bars 105, and 8 third reinforcing bars 106.
As an example, the first stiffener 104, the second stiffener 105 and the third stiffener 106 use T-shaped stringer profiles.
Referring to fig. 2 in conjunction with fig. 1, the heat shield assembly 3 is completely wrapped around the outer side of the frame structure 1 along an inner surface of the inverted frustum frame structure and an outer surface of the frustum frame structure.
As an embodiment, the inner supporting ring 101 is formed by welding a plurality of first arc bars; the supporting middle ring 102 is formed by welding a plurality of second arc bars; the support outer ring 103 is formed by welding a plurality of third arc bars.
Preferably, the present embodiment uses 4 first arc bars, 2 second arc bars, and 16 third arc bars.
Preferably, the first arc strip, the second arc strip and the third arc strip are factory-shaped stringer profiles.
Referring to fig. 3, the heat insulation plate assembly 2 is formed by installing and connecting 2 outer plates 201 and 1 middle plate 202 through supporting plate nuts and screws, the outer plates 201 and the middle plate 202 are aluminum alloy thin-wall plates, rigidity is guaranteed by punching reinforcing grooves, 2 main engine nozzle installation holes are respectively arranged at the splicing positions of the 2 outer plates 201 and the middle plate 202, certain safety distance between the edges of the installation holes and a combustion chamber of a spacecraft engine is guaranteed when the installation holes are arranged, and meanwhile, installation of the engine is facilitated.
As an example, the heat shield assembly 3 is formed by stacking a plurality of thin-walled metal plates.
Referring to fig. 2, the frame structure 1 further includes a plurality of first joints 107 and a plurality of second joints 108 for connecting with the spacecraft, the first joints 107 are uniformly distributed and welded on the inner support ring 101, and the second joints 108 are uniformly distributed and welded on the outer support ring 103.
Preferably, the present device uses 4 first connectors 107 and 16 second connectors 108.
As an embodiment, the apparatus further includes a mount assembly 4, and each second joint 108 on the support outer ring 103 of the frame structure 1 is connected to the propulsion cabin structure ring frame 9 of the spacecraft through the mount assembly 4, respectively, so as to connect the support outer ring 103 to the propulsion cabin structure ring frame 9 of the spacecraft.
Referring to fig. 6 in conjunction with fig. 5, the mounting base assembly 4 includes a plurality of mounting plate heat insulation blocks 402, each second joint 108 is fixedly mounted on a lower surface of a mounting plate 401 through a bolt, a mounting plate heat insulation block 402 is disposed between the second joint 108 and the mounting plate 401, and the mounting plate 401 is fixedly mounted on an upper surface of the propulsion cabin structure ring frame 9 of the spacecraft through a bolt, wherein the number of the mounting plates 401 is the same as that of the second joints 108, and the mounting plates 401 are connected to the second joints 108 in a one-to-one correspondence manner.
As an example, the device further comprises an auxiliary support rod assembly 5, and each first joint 107 on the support inner ring 101 of the frame structure 1 is connected with a propulsion module 7 on the propulsion nacelle structure 6 of the spacecraft through the auxiliary support rod assembly 5, respectively, so as to realize the connection between the support inner ring 101 and the propulsion module 7 on the propulsion nacelle structure 6 of the spacecraft.
Referring to fig. 4, the auxiliary support rod assembly 5 includes a connecting rod 501, a third joint 502, a connecting rod support 503, a connecting rod support insulating block 504, and a pin 505. Connecting rod 501 is carbon-fibre composite pipe, and third joint 502 is the aluminum alloy and connects, and connecting rod 501 both ends are glued third joint 502 to reinforce through round pin 505. Referring to fig. 7 in conjunction with fig. 5, each first joint 107 is connected to one end of a link 501 through a hinge, the other end of the link 501 is connected to a link support 503 through a hinge, and then the link support 503 is fixedly mounted on the lower surface of the propulsion module 7 on the propulsion cabin structure 6 of the spacecraft through a screw, wherein the number of the links 501 is the same as that of the first joints 107, and the links 501 and the first joints 107 are connected in a one-to-one correspondence manner. In order to facilitate the adaptive adjustment of the installation position during installation, the third joint 502 is hinged to the first joint 107 and the connecting rod support 503.
Preferably, the auxiliary strut assembly 5 further comprises a link support insulating block 504, and the link support 503 is connected with the propulsion module 7 on the propulsion nacelle structure 6 of the spacecraft in an insulating manner through the link support insulating block 504.
The disclosure above is only one specific embodiment of the present application, but the present application is not limited thereto, and any variations that can be made by those skilled in the art are intended to fall within the scope of the present application.
Claims (10)
1. The spacecraft tail flame protection and heat insulation device is characterized by comprising a frame structure, a heat insulation plate assembly and a heat insulation screen assembly, wherein the frame structure is used as a supporting structure of the whole device, the heat insulation plate assembly is arranged between the frame structure and the heat insulation screen assembly, and the heat insulation screen assembly is covered and installed on the outer sides of the frame structure and the heat insulation plate assembly to form a closed barrier layer; the frame structure comprises a supporting inner ring, a supporting middle ring, a supporting outer ring, a plurality of first reinforcing rods, a plurality of second reinforcing rods and a plurality of third reinforcing rods, wherein,
the support inner ring is connected with the support middle ring through the first reinforcing rod to form an inverted frustum frame structure I;
the supporting middle ring is connected with the supporting outer ring through the second reinforcing rod to form a first circular truncated cone frame structure;
the support inner ring is connected with the support outer ring through the third reinforcing rod to form a circular truncated cone frame structure II;
the first round platform frame structure, the second round platform frame structure and the first inverted round platform frame structure are coaxial, and the second round platform frame structure and the first inverted round platform frame structure are positioned inside the first round platform frame structure;
the supporting inner ring consists of a plurality of first arc strips, the supporting middle ring consists of a plurality of second arc strips, and the supporting outer ring consists of a plurality of third arc strips;
the frame structure further comprises a plurality of first joints and a plurality of second joints, wherein the first joints and the second joints are used for being connected with the spacecraft, the first joints are uniformly distributed on the support inner ring, and the second joints are uniformly distributed on the support outer ring;
the first joint is connected with a propulsion module on a propulsion cabin of the spacecraft through the auxiliary support rod assembly so as to realize the connection of the support inner ring and the propulsion module on the propulsion cabin of the spacecraft;
the support outer ring is connected with the propulsion cabin structure ring frame of the spacecraft through the mounting seat assemblies.
2. A spacecraft tail flame protection and thermal insulation apparatus as claimed in claim 1, wherein said first stiffener, second stiffener and third stiffener are T-shaped stringer profiles.
3. A spacecraft tail flame shield insulation as set forth in claim 1, wherein said heat shield assembly forms a complete wrap around said frame structure outboard along an inner surface of said inverted frustum frame structure and an outer surface of said frustum frame structure.
4. A spacecraft tail flame protection and thermal insulation device as claimed in claim 1, wherein the first arc strip, the second arc strip and the third arc strip are made of factory-shaped stringer profiles.
5. A spacecraft tail flame protection and thermal insulation device according to claim 1, wherein the thermal insulation plate assembly is formed by splicing a plurality of aluminum alloy thin-wall plates, and a main engine jet pipe mounting hole is formed at the spliced part.
6. A spacecraft tail flame shield insulation as claimed in claim 1, wherein said heat shield assembly is formed by stacking a plurality of thin-walled metal sheets.
7. A spacecraft tail flame protection and thermal insulation device according to claim 1, wherein the mounting assembly comprises a plurality of mounting plate thermal insulation blocks, each second joint is in thermal insulation connection with the mounting assembly through a mounting plate thermal insulation block, and the mounting assembly is connected with a propulsion cabin structure ring frame of the spacecraft.
8. A spacecraft tail flame protection and thermal insulation device according to claim 1, wherein the auxiliary support rod assembly comprises a plurality of connecting rods, a plurality of third joints and a plurality of connecting rod supports, the number of the connecting rods is equal to that of the connecting rod supports, the number of the third joints is 2 times of that of the connecting rods, the third joints are fixedly connected to two ends of each connecting rod, the third joint at one end of each connecting rod is connected with a propulsion module on a propulsion cabin of a spacecraft through one connecting rod support, and the third joint at the other end of each connecting rod is connected with one first joint.
9. A spacecraft tail flame protection and thermal insulation device according to claim 8, wherein the auxiliary support rod assembly further comprises a plurality of connecting rod support thermal insulation cushion blocks, and each connecting rod support is in thermal insulation connection with a propulsion module on a propulsion cabin of the spacecraft through one connecting rod support thermal insulation cushion block.
10. A spacecraft comprising a spacecraft tail flame protection and insulation arrangement as claimed in any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811034568.XA CN109131948B (en) | 2018-09-04 | 2018-09-04 | Spacecraft tail flame protection and heat insulation device and spacecraft |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811034568.XA CN109131948B (en) | 2018-09-04 | 2018-09-04 | Spacecraft tail flame protection and heat insulation device and spacecraft |
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| Publication Number | Publication Date |
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| CN109131948A CN109131948A (en) | 2019-01-04 |
| CN109131948B true CN109131948B (en) | 2021-08-10 |
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| CN201811034568.XA Active CN109131948B (en) | 2018-09-04 | 2018-09-04 | Spacecraft tail flame protection and heat insulation device and spacecraft |
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Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN110306687A (en) * | 2019-07-04 | 2019-10-08 | 刘军 | A kind of concreting net wall structure |
| CN113253688B (en) * | 2021-06-11 | 2021-10-01 | 四川航天长征装备制造有限公司 | Servo mechanism flame exhaust pipe digital assembly manufacturing method |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104859869A (en) * | 2015-02-12 | 2015-08-26 | 上海卫星装备研究所 | Method for mounting high-temperature heat shield of spacecraft |
| CN105292522A (en) * | 2015-11-09 | 2016-02-03 | 上海卫星装备研究所 | Spacecraft high-temperature thermal insulation screen installation device and method |
| CN205345360U (en) * | 2015-08-26 | 2016-06-29 | 上海宇航系统工程研究所 | Light -duty multi -functional heat accuse multilayer support of moon exploration spacecraft |
| WO2018011753A1 (en) * | 2016-07-13 | 2018-01-18 | Ali S.C.A.R.L. | Decelerator for landing bodies in aerospace field |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US9475591B2 (en) * | 2013-11-19 | 2016-10-25 | Arthur Mckee Dula | Space launch apparatus |
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Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104859869A (en) * | 2015-02-12 | 2015-08-26 | 上海卫星装备研究所 | Method for mounting high-temperature heat shield of spacecraft |
| CN205345360U (en) * | 2015-08-26 | 2016-06-29 | 上海宇航系统工程研究所 | Light -duty multi -functional heat accuse multilayer support of moon exploration spacecraft |
| CN105292522A (en) * | 2015-11-09 | 2016-02-03 | 上海卫星装备研究所 | Spacecraft high-temperature thermal insulation screen installation device and method |
| WO2018011753A1 (en) * | 2016-07-13 | 2018-01-18 | Ali S.C.A.R.L. | Decelerator for landing bodies in aerospace field |
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| CN109131948A (en) | 2019-01-04 |
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