CN116163855A - Reusable space engine thermal control device - Google Patents
Reusable space engine thermal control device Download PDFInfo
- Publication number
- CN116163855A CN116163855A CN202211633294.2A CN202211633294A CN116163855A CN 116163855 A CN116163855 A CN 116163855A CN 202211633294 A CN202211633294 A CN 202211633294A CN 116163855 A CN116163855 A CN 116163855A
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- Prior art keywords
- semicircular structure
- thermal control
- control device
- main body
- transition
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/96—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof characterised by specially adapted arrangements for testing or measuring
Abstract
The invention provides a reusable space engine thermal control device, which comprises a main body part and a transition part arranged at the rear side of the main body part, wherein the main body part comprises a first semicircular structure and a second semicircular structure, the first semicircular structure and the second semicircular structure are symmetrically arranged on the outer surface of the rear end of an engine, and main body shells are arranged on the outer sides of the first semicircular structure and the second semicircular structure; the first semicircular structure is provided with a first heating loop, and the second semicircular structure is provided with a second heating loop and a temperature measuring loop; the front end of the transition part is connected with the main body part, the rear end of the transition part extends towards the rear end of the engine, the transition part comprises a transition shell and an outgoing line arranged in the transition shell, and the outgoing line is led out from the rear end of the transition shell. The invention relates to a heating and temperature measurement integrated design, which solves the technical problem that the heating and temperature measurement cannot be carried out simultaneously by only considering the on-orbit working performance or adopting an indirect temperature measurement mode of a thermal control and temperature measurement device applied to the prior space engine.
Description
Technical Field
The invention relates to the technical field of thermal control devices, in particular to a reusable space engine thermal control device, and especially relates to a reusable space engine heating and temperature measuring integrated thermal control device.
Background
The reusable return spacecraft is an important development direction of the return spacecraft at home and abroad, and the main purposes are to improve the emission efficiency and reduce the emission cost.
For the space engines at home and abroad capable of being recycled, the thermal control device needs to break through key technologies such as full-task complex thermal environment control, so that the on-orbit multi-time temperature control work requirement of the space engine is met, and the on-orbit work-recycling-on-orbit work repeatability use requirement of the space engine is ensured for a plurality of times. When the space engine works on the track, the starting temperature of the engine body part before each ignition is higher than a required value by a heating and temperature measuring device, and meanwhile, the temperature of the engine body part is higher and can reach 700-1100 ℃ due to heat generated in the combustion process of the propellant when the engine works. The performances of the heater, the temperature sensor and the heat conduction silicone grease on the surface of the engine at high temperature are obviously reduced, and the repeated use cannot be ensured after the engine falls on the earth. The heat control and temperature measurement device applied to the prior space engine generally only considers the on-orbit working performance, adopts an indirect temperature measurement mode, or cannot carry out heating and temperature measurement integrated design, and has a defect in the aspect of repeated use research.
Patent document with publication number CN203067074U discloses a general engine shutter type thermal control cooling regulator, shutter blades are installed between a fan cover and an engine cylinder, the shutter blades are connected with a pull rod, and an expansion wax thermal controller is fixed on the engine cylinder; the heat generated by the engine cylinder body can be rapidly and timely acted on the expansion wax thermal controller, the expansion wax is arranged in the expansion wax thermal controller, and the expansion wax expands after being heated and contracts after being cooled, so that the control on the opening and closing of the shutter blades is realized by acting on the shutter blades through the pull rod connected with the expansion wax, the occurrence of the phenomenon that the engine cannot be excessively cooled or overheated is ensured, and the abrasion and the efficiency of the cylinder of the engine are improved. However, the thermal control device cannot adapt to a high-temperature environment and is not suitable for a space engine.
Disclosure of Invention
In view of the shortcomings in the prior art, the invention aims to provide a reusable space engine thermal control device.
The reusable space engine thermal control device comprises a main body part and a transition part arranged at the rear side of the main body part, wherein the main body part comprises a first semicircular structure and a second semicircular structure, the first semicircular structure and the second semicircular structure are symmetrically arranged on the outer surface of the rear end of an engine, and main body shells are arranged on the outer sides of the first semicircular structure and the second semicircular structure;
the first semicircular structure is provided with a first heating loop, and the second semicircular structure is provided with a second heating loop and a temperature measuring loop;
the front end of the transition part is connected with the main body part, the rear end of the transition part extends towards the rear end of the engine, the transition part comprises a transition shell and an outgoing line arranged inside the transition shell, and the outgoing line is led out from the rear end of the transition shell.
Preferably, the main body shell is made of a high-temperature alloy material, and an insulating structure is arranged on the outer side of the main body shell.
Preferably, the two ends of the first semicircle structure are provided with first lugs with holes, the two ends of the second semicircle structure are provided with second lugs with holes, the first lugs are arranged opposite to the second lugs, and the first lugs are connected with the second lugs through connecting pieces.
Preferably, the first semicircular structure and the second semicircular structure are both fixed on the outer surface of the engine through bolts.
Preferably, a metal foil is filled between the first semicircular structure and the second semicircular structure and the outer surface of the engine.
Preferably, the first heating circuit comprises a first main part heating circuit and a backup heating circuit, and the second heating circuit comprises a second main part heating circuit;
the power of the first main part heating loop is the same as that of the second main part heating loop.
Preferably, the device further comprises a controller, wherein the controller is connected with the temperature measuring loop and can receive the temperature fed back by the temperature measuring loop, so that the first heating loop and the second heating loop are controlled.
Preferably, the thicknesses of the first semicircle structure and the second semicircle structure are 10-15 mm.
Preferably, the metal foil comprises a nickel foil.
Preferably, the transition shell is of a variable-section cylinder structure and made of stainless steel materials, and insulating materials are arranged on the outer sides of the outgoing lines.
Compared with the prior art, the invention has the following beneficial effects:
1. the invention has simple structure and convenient operation, adopts the technical means of heating and temperature measurement integrated design, and solves the technical problem that the heat control and temperature measurement device applied to the prior space engine usually only considers the on-orbit working performance or adopts an indirect temperature measurement mode and cannot simultaneously carry out heating and temperature measurement.
2. The invention adopts the technical means that the high-temperature alloy material shell is arranged on the outer side of the main body part, and solves the technical problems that the performances of the existing heater, the temperature measuring sensor and the heat conduction silicone grease on the surface of the engine are obviously reduced in a high-temperature environment, and the repeated use cannot be ensured after the engine falls on the earth.
Drawings
Other features, objects and advantages of the present invention will become more apparent upon reading of the detailed description of non-limiting embodiments, given with reference to the accompanying drawings in which:
FIG. 1 is a schematic diagram of a front view of the present invention;
FIG. 2 is a schematic top view of the present invention;
fig. 3 is a schematic view of the structure of the present invention when mounted on an engine.
The figure shows:
First half-round structure 3 first main heating loop 9
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the present invention, but are not intended to limit the invention in any way. It should be noted that variations and modifications could be made by those skilled in the art without departing from the inventive concept. These are all within the scope of the present invention.
The invention discloses a reusable space engine thermal control device, which is particularly suitable for a space-earth repeatedly reciprocating recoverable attitude rail control engine body thermal control design, and can solve the problem that the thermal control device of an attitude and rail control engine of an aerospace craft can repeatedly work along with the engine repeatedly reciprocating for a plurality of times, and simultaneously ensure the on-orbit performance of the engine.
According to the reusable space engine thermal control device provided by the invention, as shown in fig. 1-3, the reusable space engine thermal control device comprises a main body part 1 and a transition part 2 arranged at the rear side of the main body part 1, wherein the main body part 1 comprises a first semicircular structure 3 and a second semicircular structure 4 which are of semicircular sheet type structures with symmetrical left-right geometric dimensions, the first semicircular structure 3 and the second semicircular structure 4 are symmetrically arranged on a mounting groove at the rear end of an engine body part, main body shells are arranged at the outer sides of the first semicircular structure 3 and the second semicircular structure 4, and in order to adapt to a high-temperature environment, the main body shells are made of high-temperature alloy materials and are ensured to be reusable after falling down the earth, and insulating structures 11 are arranged at the outer sides of the main body shells. Preferably, the temperature to which the main body housing means is subjected is selected from stainless steel or a superalloy.
As shown in fig. 1, the two ends of the first semicircular structure 3 are provided with first lugs 6 with holes, the two ends of the second semicircular structure 4 are provided with second lugs 7 with holes, the first lugs 6 and the second lugs 7 are opposite, the first lugs 6 and the second lugs 7 are connected with each other through connecting pieces, preferably, bolts and nuts pass through the adjacent first lugs 6 and second lugs 7, and torque is applied to the bolts, so that the distance between the lugs at the adjacent positions of the first semicircular structure 3 and the second semicircular structure 4 is reduced as much as possible, and the device is tightly attached to the engine or other applicable structure surfaces. Preferably, the thickness of the first ear piece 6 and the second ear piece 7 is 3-5 mm, and the length is 8-12 mm; the width is 5-6 mm. When the first ear piece 6 and the second ear piece 7 are installed, the gap between the two is 1-3 mm. The thicknesses of the first semicircular structure 3 and the second semicircular structure 4 are 10-15 mm.
As shown in fig. 1, the first semicircular structure 3 and the second semicircular structure 4 are both fixed on the outer surface of the engine through bolts. Preferably, the first semicircular structure 3 and the second semicircular structure 4 are respectively installed on the outer surface of the engine by using titanium bolts of M3 or M4 as fasteners. And metal foils are filled between the first semicircular structure 3 and the second semicircular structure 4 and the outer surface of the engine, so that heat conduction is enhanced. Preferably, the metal foil comprises a nickel foil.
As shown in fig. 1, the first semicircular structure 3 is provided with a first heating circuit, and the second semicircular structure 4 is provided with a second heating circuit and a temperature measuring circuit 12; the temperature measuring element in the temperature measuring circuit 12 is capable of withstanding the high temperature environment of the engine. The first heating circuit comprises a first main part heating circuit 9 and a backup heating circuit 8, and the second heating circuit comprises a second main part heating circuit 10; the first main part heating circuit 9 and the second main part heating circuit 10 have the same power. Preferably, the temperature measuring circuit 12 is further connected to a controller, and the controller can receive the temperature fed back by the temperature measuring circuit 12, so as to control the first heating circuit and the second heating circuit.
As shown in fig. 2, the front end of the transition part 2 is connected to the main body part 1, the rear end of the transition part 2 extends toward the rear end of the engine, the transition part 2 includes a transition housing and an outgoing line 5 provided inside the transition housing, and the outgoing line 5 is led out from the rear end of the transition housing. The transition shell is of a variable-section cylinder structure and made of stainless steel materials, and insulating materials are arranged on the outer sides of the outgoing lines 5. As shown in fig. 3, a flange at the rear end of the engine is provided with openings for the passage of the transition section 2, and a transition housing is provided through the flange.
In summary, the thermal control device provided by the invention can break through key technologies such as full-task complex thermal environment control, so that not only is the requirement of the space engine for completing on-orbit multiple works met, but also the repeated use requirement of the space engine for on-orbit work, recovery and on-orbit work for multiple times is ensured. The thermal control device is suitable for the thermal control design of the body part of the space-earth multi-round trip recoverable attitude and orbit control engine, can solve the problem that the thermal control device of the attitude and orbit control engine of an aerospace craft can repeatedly work along with the engine for a plurality of round trips, and simultaneously ensures the on-orbit performance of the engine.
In the description of the present application, it should be understood that the terms "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like indicate orientations or positional relationships based on the orientations or positional relationships illustrated in the drawings, merely to facilitate description of the present application and simplify the description, and do not indicate or imply that the devices or elements being referred to must have a specific orientation, be configured and operated in a specific orientation, and are not to be construed as limiting the present application.
The foregoing describes specific embodiments of the present invention. It is to be understood that the invention is not limited to the particular embodiments described above, and that various changes or modifications may be made by those skilled in the art within the scope of the appended claims without affecting the spirit of the invention. The embodiments of the present application and features in the embodiments may be combined with each other arbitrarily without conflict.
Claims (10)
1. The reusable space engine thermal control device is characterized by comprising a main body part (1) and a transition part (2) arranged at the rear side of the main body part (1), wherein the main body part (1) comprises a first semicircular structure (3) and a second semicircular structure (4), the first semicircular structure (3) and the second semicircular structure (4) are symmetrically arranged on the outer surface of the rear end of an engine, and main body shells are arranged on the outer sides of the first semicircular structure (3) and the second semicircular structure (4);
a first heating loop is arranged on the first semicircular structure (3), and a second heating loop and a temperature measuring loop (12) are arranged on the second semicircular structure (4);
the front end of the transition part (2) is connected with the main body part (1), the rear end of the transition part (2) extends towards the rear end of the engine, the transition part (2) comprises a transition shell and an outgoing line (5) arranged inside the transition shell, and the outgoing line (5) is led out from the rear end of the transition shell.
2. The reusable space engine thermal control device according to claim 1, wherein the main body housing is made of superalloy material, and an insulating structure (11) is provided on the outside of the main body housing.
3. The reusable space engine thermal control device according to claim 1, wherein both ends of the first semicircular structure (3) are provided with first lugs (6) with holes, both ends of the second semicircular structure (4) are provided with second lugs (7) with holes, the first lugs (6) are arranged opposite to the second lugs (7), and the first lugs (6) are connected with the second lugs (7) through connecting pieces.
4. The reusable space engine thermal control device according to claim 1, wherein the first semicircular structure (3) and the second semicircular structure (4) are both fixed on the outer surface of the engine by bolts.
5. The reusable space engine thermal control device according to claim 1, characterized in that the first (3) and second (4) semicircular structures are each filled with a metal foil with the engine outer surface.
6. The reusable space engine thermal control device according to claim 1, characterized in that the first heating circuit comprises a first primary heating circuit (9) and a backup heating circuit (8), the second heating circuit comprises a second primary heating circuit (10);
the power of the first main part heating loop (9) is the same as that of the second main part heating loop (10).
7. The reusable space engine thermal control device of claim 1, further comprising a controller connected to the thermometry circuit (12) and capable of receiving temperature feedback from the thermometry circuit (12) to control the first and second heating circuits.
8. The reusable space engine thermal control device according to claim 1, wherein the thickness of the first semicircular structure (3) and the second semicircular structure (4) are both 10-15 mm.
9. The reusable space engine thermal control device of claim 5, wherein the metal foil comprises nickel foil.
10. The reusable space engine thermal control device according to claim 1, wherein the transition housing is of variable cross-section cylindrical structure and is made of stainless steel material, and the outer side of the outgoing line (5) is provided with an insulating material.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211633294.2A CN116163855A (en) | 2022-12-19 | 2022-12-19 | Reusable space engine thermal control device |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211633294.2A CN116163855A (en) | 2022-12-19 | 2022-12-19 | Reusable space engine thermal control device |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN116163855A true CN116163855A (en) | 2023-05-26 |
Family
ID=86417388
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202211633294.2A Pending CN116163855A (en) | 2022-12-19 | 2022-12-19 | Reusable space engine thermal control device |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN116163855A (en) |
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2022
- 2022-12-19 CN CN202211633294.2A patent/CN116163855A/en active Pending
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