CN112462714A - Processing method of spacecraft thrust chamber double-layer structure - Google Patents

Processing method of spacecraft thrust chamber double-layer structure Download PDF

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Publication number
CN112462714A
CN112462714A CN202011381643.7A CN202011381643A CN112462714A CN 112462714 A CN112462714 A CN 112462714A CN 202011381643 A CN202011381643 A CN 202011381643A CN 112462714 A CN112462714 A CN 112462714A
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China
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wall
thrust chamber
spacecraft
wall structure
double
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CN202011381643.7A
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CN112462714B (en
Inventor
杨瑞康
宣智超
常克宇
袁宇
黄乐
周涛
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Landspace Technology Co Ltd
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Shaanxi Landspace Co ltd
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Priority to GB2307694.6A priority patent/GB2616149B/en
Priority to PCT/CN2021/134260 priority patent/WO2022116955A1/en
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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
    • F02K9/97—Rocket nozzles
    • F02K9/972—Fluid cooling arrangements for nozzles
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B19/00—Programme-control systems
    • G05B19/02—Programme-control systems electric
    • G05B19/418—Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control [DNC], flexible manufacturing systems [FMS], integrated manufacturing systems [IMS] or computer integrated manufacturing [CIM]
    • G05B19/41875—Total factory control, i.e. centrally controlling a plurality of machines, e.g. direct or distributed numerical control [DNC], flexible manufacturing systems [FMS], integrated manufacturing systems [IMS] or computer integrated manufacturing [CIM] characterised by quality surveillance of production
    • 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
    • B64G1/40—Arrangements or adaptations of propulsion systems
    • 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
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/62—Combustion or thrust chambers
    • F02K9/64—Combustion or thrust chambers having cooling arrangements
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05B—CONTROL OR REGULATING SYSTEMS IN GENERAL; FUNCTIONAL ELEMENTS OF SUCH SYSTEMS; MONITORING OR TESTING ARRANGEMENTS FOR SUCH SYSTEMS OR ELEMENTS
    • G05B2219/00—Program-control systems
    • G05B2219/30—Nc systems
    • G05B2219/32—Operator till task planning
    • G05B2219/32368—Quality control
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02P—CLIMATE CHANGE MITIGATION TECHNOLOGIES IN THE PRODUCTION OR PROCESSING OF GOODS
    • Y02P90/00—Enabling technologies with a potential contribution to greenhouse gas [GHG] emissions mitigation
    • Y02P90/02—Total factory control, e.g. smart factories, flexible manufacturing systems [FMS] or integrated manufacturing systems [IMS]

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion & Propulsion (AREA)
  • Chemical & Material Sciences (AREA)
  • Mechanical Engineering (AREA)
  • Manufacturing & Machinery (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Quality & Reliability (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Remote Sensing (AREA)
  • Pressure Vessels And Lids Thereof (AREA)

Abstract

The invention discloses a processing method of a double-layer structure of a spacecraft thrust chamber, which comprises the steps of providing a first wall structure, a second wall structure and convex ribs; the first wall structure and the second wall structure are close to each other and are closed at the end parts, so that a combined structure which is composed of the first wall structure, the second wall structure and the convex ribs and is provided with a closed space inside is formed; and (3) making the sealed space be negative pressure, then making the combined structure undergo the first pressure treatment, then making the combined structure undergo the second pressure treatment under the condition of that the sealed space is communicated with exterior, in which the maximum pressure of the second pressure treatment is greater than that of the first pressure treatment so as to make the combined structure be welded into one body so as to obtain double-layer structure. The method has the advantages of simple process, short manufacturing period and low cost.

Description

Processing method of spacecraft thrust chamber double-layer structure
Technical Field
The invention relates to the technical field of spacecraft engines, in particular to a method for processing a double-layer structure of a spacecraft thrust chamber.
Background
The technology of spacecraft engines is rapidly upgraded with the development of the aerospace industry. As a main component of the engine, the thrust chamber is a key component for performing energy conversion of the propellant and generating thrust. The thrust chamber body is a component which is used for mixing and burning fuel in the spacecraft engine to generate high-temperature and high-pressure fuel gas, and then the fuel gas is accelerated and discharged through the throat part to obtain reverse thrust. The body of the thrust chamber is of a Laval profile structure, and the thrust chamber can be cooled by adopting a regenerative cooling technology. The thrust chamber is composed of a milling groove inner wall and a milling groove outer wall, and a plurality of cooling channels are arranged between the milling groove inner wall and the milling groove outer wall. Normally, the two have no leak defects inside under pressures up to 60 MPa.
At present, the inner wall and the outer wall of a milling groove are connected with the following two methods, one method adopts the processes of transient liquid phase diffusion brazing and nickel electroforming, but the process has the defects of complex and expensive process and long period. The other is that the thrust chamber adopts copper steel xenogenesis alloy hot isostatic pressing diffusion welding at the preparation process, however, in the thrust chamber manufacturing process, often can take place to be located the fin on the inner wall and can't bear the high pressure and be buckled by the pressure, and then make the phenomenon that the passageway collapses, and pressure undersize can't accomplish the fin again and reliably be connected with the outer wall.
In order to solve the problems, the invention provides a welding processing method of a thrust chamber body component and a spacecraft thrust chamber, which have the advantages of simple process, short manufacturing period, cost saving, mass production and capacity improvement.
Disclosure of Invention
The invention aims to provide a processing method of a double-layer structure of a spacecraft thrust chamber, which has the advantages of simple process, short manufacturing period, cost saving, mass production, capacity improvement and the like.
In order to achieve the purpose, the invention provides the following technical scheme: a method of fabricating a bilayer structure for a thrust cell of a spacecraft, wherein the bilayer structure comprises an outer wall and an inner wall forming at least a part of a body portion of the thrust cell of the spacecraft, and a spacer portion secured between the outer wall and the inner wall and arranged to provide a passage for coolant flow between the outer wall and the inner wall, the method comprising the steps of:
(1) providing a first wall structure, a second wall structure and a rib, wherein the first wall structure and the second wall structure are close to each other in an inside-outside nesting mode to form at least one part of a spacecraft thrust chamber body, and the rib is positioned between the first wall structure and the second wall structure;
(2) closing the first wall structure and the second wall structure at the ends, thereby forming a combined structure consisting of the first wall structure, the second wall structure and the ribs, the combined structure having a closed space inside;
(3) and (2) making the closed space be negative pressure, then carrying out first pressurization treatment on the combined structure from the outside, and then carrying out second pressurization treatment under the condition that the closed space is communicated with the outside, wherein the maximum pressure of the second pressurization treatment is greater than that of the first pressurization treatment, so that the combined structure is welded into a whole, and the double-layer structure is obtained.
According to the processing method of the double-layer structure of the spacecraft thrust chamber, preferably, the pressure A during the first pressurization treatment is more than or equal to 1MPa and less than or equal to 20MPa, the pressurization time B is more than or equal to 0.2h and less than or equal to 10h, and the temperature C is more than or equal to 300 ℃ and less than or equal to 1300 ℃.
According to the processing method of the double-layer structure of the spacecraft thrust chamber, preferably, the pressure D during the second pressurization treatment is more than or equal to 2MPa and less than or equal to 120MPa, the pressurization time E is more than or equal to 0.1h and less than or equal to 10h, and the temperature F is more than or equal to 300 ℃ and less than or equal to 1400 ℃.
According to the processing method of the double-layer structure of the spacecraft thrust chamber, the pressure of the closed space is preferably controlled through an air guide assembly, the air guide assembly comprises an air guide pipe and an annular groove which is communicated with the tail end of the air guide pipe and is arranged on one side of the first wall structure, and the processing method of the double-layer structure of the spacecraft thrust chamber preferably further comprises a vacuumizing device which is connected with the air guide pipe.
The method for processing the double-layer structure of the spacecraft thrust chamber preferably further comprises the step of cutting off the air guide component.
According to the processing method of the spacecraft thrust chamber bilayer structure, preferably, the bilayer structure comprises a first bilayer structure and a second bilayer structure which form a thrust chamber body part, and the first bilayer structure and the second bilayer structure are respectively and independently a combustion chamber or an expansion section, and the processing method comprises a first processing step of obtaining the first bilayer structure and a second processing step of obtaining the second bilayer structure, wherein the first processing step and the second processing step are carried out simultaneously or respectively and sequentially.
According to the processing method of the spacecraft thrust chamber double-layer structure, in the step (1), preferably, one side of the convex rib is fixed to the second wall structure, and the other side of the convex rib is tightly attached to the surface of the first wall structure, so that the convex rib is formed into the spacing part after the step (3).
The method for processing a spacecraft thrust chamber bilayer structure according to the present invention preferably further comprises the steps of providing an end cap and removing the end cap, the end cap being provided so as to be connected to the ends of the first and second wall structures so as to close the first and second wall structures at the ends, the step of removing the end cap comprising cutting away the end cap in a radial direction of the bilayer structure.
According to the processing method of the spacecraft thrust chamber double-layer structure, one end of the first wall structure or one end of the second wall structure is provided with a sealing part, and end face sealing of the first wall structure and the second wall structure is achieved through the sealing part.
According to the processing method of the spacecraft thrust chamber double-layer structure, the combined structure is preferably subjected to pressurization treatment through a high-pressure container.
Compared with the prior art, the invention has the beneficial effects that: the combined structure consisting of the outer wall, the inner wall and the end cover after vacuumizing is placed in a high-pressure container for first pressurization treatment, so that the convex ribs are connected with the outer wall. Through the air duct makes the passageway keeps unblocked with the outside gas, the outer wall with when the inner wall is put into high-pressure vessel again and is carried out secondary pressurization, high-pressure gas passes through the air duct and gets into in the passageway, is used for supporting the fin on the one hand, prevents that the fin from causing because of pressure is too big and cave in, and on the other hand, through being greater than the maximum pressure of primary pressurization with the maximum pressure of secondary pressurization for bubble in the clearance is extruded when fin and outer wall connection, makes fin and outer wall connection inseparabler, firm. And cutting off the end cover and the part of the outer wall matched with the air duct to obtain a thrust chamber body structure consisting of the outer wall and the inner wall. The whole method has the advantages of simple process, short manufacturing period, cost saving, mass production and capacity improvement.
Drawings
FIG. 1 is a schematic view of a rocket engine thrust chamber component of the present invention;
FIG. 2 is a schematic pre-weld view of a combustor assembly of the present invention;
FIG. 3 is a perspective view of the combustion chamber of the present invention;
FIG. 4 is a schematic view of the expanding segment of the present invention;
FIG. 5 is a perspective view of the combustion chamber of the present invention connected to an expansion section;
FIG. 6 is a schematic view of the structure of the present invention in which the combustion chamber is connected to the expansion section;
FIG. 7 is a schematic structural view of the outer wall ring and the connecting pipe member being thinned after the combustion chamber and the expansion section are connected according to the present invention;
FIG. 8 is a schematic cross-sectional view of the outer wall, inner wall, air duct and rib connection of the present invention;
FIG. 9 is a schematic view of the present invention tangential in the radial direction of the combustion chamber;
FIG. 10 is a perspective view of the outer wall and annular groove of the present invention;
FIG. 11 is a schematic view showing the connection of the connecting pipe member with the combustion chamber and the expansion section according to the present invention.
FIG. 12 is a flowchart of the inventive procedure.
Description of reference numerals:
1 outer wall 2 inner wall
3 air duct 4 convex rib
5 end cover 6 combustion chamber
7 expanding section 8 ring groove
9 outer wall ring 10 connecting pipe fitting
Detailed Description
For the purpose of promoting a clear understanding of the objects, aspects and advantages of the embodiments of the invention, reference will now be made to the drawings and detailed description, wherein there are shown in the drawings and described in detail, various modifications of the embodiments described herein, and other embodiments of the invention will be apparent to those skilled in the art.
The exemplary embodiments of the present invention and the description thereof are provided to explain the present invention and not to limit the present invention. Additionally, the same or similar numbered elements/components used in the drawings and the embodiments are used to represent the same or similar parts.
As used herein, the terms "first," "second," …, etc., do not denote any order or sequence, nor are they used to limit the present invention, but rather are used to distinguish one element from another or from another element or operation described in the same technical language.
With respect to directional terminology used herein, for example: up, down, left, right, front or rear, etc., are simply directions with reference to the drawings. Accordingly, the directional terminology used is intended to be illustrative and is not intended to be limiting of the present teachings.
As used herein, the terms "comprising," "including," "having," "containing," and the like are open-ended terms that mean including, but not limited to.
As used herein, "and/or" includes any and all combinations of the described items.
As used herein, the terms "substantially", "about" and the like are used to modify any slight variation in quantity or error that does not alter the nature of the variation. Generally, the range of slight variations or errors modified by such terms may be 20% in some embodiments, 10% in some embodiments, 5% in some embodiments, or other values. It should be understood by those skilled in the art that the aforementioned values can be adjusted according to actual needs, and are not limited thereto.
Certain words used to describe the present application are discussed below or elsewhere in this specification to provide additional guidance to those skilled in the art in describing the present application.
The embodiment of the invention provides a method for welding a thrust chamber body component, as shown in fig. 1, fig. 2, fig. 3, fig. 4, fig. 5, fig. 6, fig. 8, fig. 9 and fig. 12, an outer wall 1, an inner wall 2 and an air duct 3 are provided, wherein a convex rib 4 is arranged on the outer side of the inner wall 2, the other side of the convex rib 4 is used for being in close contact connection with the inner surface of the outer wall 1, and after the outer wall 1 is connected with the inner wall 2, the convex rib 4 defines a plurality of channels for flowing coolant between the outer wall 1 and the inner wall 2, and the method comprises the following specific steps:
s1: end covers are arranged at two ends of the outer wall 1 and the inner wall 2 to form a combined structure, so that a closed space is formed between the outer wall 1 and the inner wall 2;
s2: vacuumizing the closed space through the air duct 3;
s3: putting the combined structure consisting of the outer wall 1, the inner wall 2 and the end cover 5 which are vacuumized into a high-pressure container for primary pressurization treatment;
s4: taking out the combined structure from the high-pressure container, and keeping the channel and the external air unblocked through the air duct 3;
s5: putting the outer wall 1 and the inner wall 2 into a high-pressure container again for secondary pressurization treatment, wherein the maximum pressure of the secondary pressurization treatment is greater than that of the primary pressurization treatment;
s6: and taking out the combined structure subjected to the secondary pressurization treatment, and cutting off the end cover 5 and the part of the outer wall 1 matched with the air guide pipe 3 to obtain a thrust chamber body structure consisting of the outer wall 1 and the inner wall 2.
Specifically, the combined structure of the outer wall 1, the inner wall 2 and the end 5 cap after vacuum pumping is put into a high-pressure container for first pressurization treatment, so that the convex ribs 4 are connected with the outer wall 1. Make passageway and outside gas keep unblocked through air duct 3, when putting into high pressure vessel again and carrying out pressurization treatment for the second time in outer wall 1 and inner wall 2, high-pressure gas passes through air duct 3 and gets into in the passageway, on the one hand, be used for supporting fin 4, prevent that fin 4 from causing because of pressure is too big and collapsing, the coolant of being convenient for evenly circulates in the passageway, on the other hand, be greater than the maximum pressure of pressurization treatment for the second time, make fin 4 and outer wall 1 be connected when having the bubble in the clearance and extruded, make fin 4 and outer wall 1 be connected inseparabler, it is firm. And cutting off the end cover 5 and the part of the outer wall 1 matched with the air duct 5 to obtain a thrust chamber body structure consisting of the outer wall 1 and the inner wall 2. The whole method has the advantages of simple process, short manufacturing period, cost saving, mass production and capacity improvement.
It needs to say that, during the first pressurization treatment, in order to make the outer wall 1 and the convex rib 4 be connected closely, it is fixed firm, to the pressure in the high-pressure container, the simulation experiment is carried out for many times with the temperature, when the pressure in the high-pressure container is A, and satisfies that 1MPa is not less than A and not more than 20MPa, the pressurization time is B, and satisfies that 0.2h is not less than B and not more than 10h, the temperature in the high-pressure container is C, and satisfies that 300 ℃ is not less than C and not more than 1300 ℃, it can make the outer wall 1 and the convex rib 4 be connected closely, make the constituent atoms of the outer wall 1 and the convex rib 4 can be diffused fast, it is convenient for both to fix together, be favorable to improving.
In order to prevent the convex rib 4 from collapsing due to overlarge pressure, the air duct 3 enables the channel to be kept smooth with outside air, and in the secondary pressurization process, air in the channel plays a role in supporting and fixing the convex rib 4, so that the convex rib 4 is prevented from collapsing due to the overlarge pressure. In addition, through the maximum pressure that the pressurization was handled for the second time is greater than the maximum pressure that the pressurization was handled for the first time, can be with outer wall 1 and fin 4 have the bubble in the clearance when the pressurization is connected for the first time and be extruded for fin 4 is connected inseparabler with outer wall 1, and is more firm, reaches thrust chamber service standard.
It is to be mentioned that, when the secondary pressurization treatment is performed, for example, the pressure in the high-pressure container is set to be D, and 2MPa or more and D or less than 120MPa are satisfied, the pressurization time is set to be E, and 0.1h or less and E or less than 10h are satisfied, the temperature in the high-pressure container is set to be F, and 300 ℃ or less and F or less than 1400 ℃, and the secondary pressurization parameters are set, so that the combination quality of the inner wall convex rib and the outer wall can be improved, and the quality reliability of the engine can be improved. In addition, in order to reduce air bubbles in the gap between the outer wall 1 and the rib 4 when they are connected, for example, the secondary pressurization treatment may be performed a plurality of times. In addition, the pressure, the pressurizing time and the temperature in the multiple pressurizing processes can be adjusted, so that the outer wall 1 and the convex ribs 4 are tightly connected and firmly fixed.
In this embodiment, as shown in fig. 1, 8 and 10, in order to ensure that the air in the channel is rapidly pumped out, for example, when the sealed space is evacuated through the air duct 3 in the annular groove 8 on the inner side of the outer wall 1, one end of the air duct 3 is connected and communicated with the annular groove 8 on the inner side of the outer wall 1, so that the channel is communicated with the air duct 3, which is beneficial to pumping out the air in the channel. The vacuumizing treatment process is characterized in that the vacuumizing treatment process is realized by connecting the vacuumizing equipment with the air guide tube to vacuumize the closed space, and air in the channel is discharged from the air guide tube 3 through the annular groove 8.
In the present embodiment, as shown in fig. 1, 2, 3, and 4, the thrust chamber body structure includes a combustion chamber 6 and an expansion section 7. The welding method for the thrust chamber body assembly will now be described.
Taking the combustion chamber 6 as an example, the preparation process of the combustion chamber specifically comprises:
sealing the outer wall 1 and the inner wall 2 of the combustion chamber 6 by using the end covers 5, and vacuumizing to obtain a first combined structure; carrying out first pressurization treatment on the first combined structure; and (3) communicating the channel of the first combined structure with the outside through the gas guide pipe 3, and carrying out secondary pressurization treatment on the first combined structure subjected to primary pressurization treatment to obtain the combustion chamber 6. For example, the outer wall inboard of first integrated configuration can set up annular groove to through annular groove respectively with the back of interior outer wall channel and air duct UNICOM, carry out evacuation processing through evacuating device.
Illustrated by way of example in fig. 1, 4 and 5 is an expansion segment 7, the method of manufacture of which comprises: firstly, sealing an outer wall 1 and an inner wall 2 of an expansion section 7 by using an end cover 5, and then vacuumizing to obtain a second combined structure; carrying out first pressurization treatment on the second combined structure; and (3) communicating the channel of the second combined structure with the outside through the gas guide tube 3, and carrying out secondary pressurization treatment on the second combined structure subjected to the primary pressurization treatment to obtain an expansion section 7. For example, the outer wall of the second combined structure can be provided with an annular groove on the inner side so as to be respectively communicated with the inner wall and the outer wall through the annular groove, and then vacuumized through a vacuumizing device.
As shown in fig. 2 and fig. 3, in order to facilitate one end of the inner wall 2 of the combustion chamber (the position of the throat portion of the thrust chamber on the inner wall) to penetrate through the outer wall, for example, the outer diameter of the outer wall 1 may be designed to be larger than the maximum diameter of the throat portion, in order to ensure that the surface of the inner wall 2 (the position of the throat portion of the thrust chamber on the inner wall) is normally used, before welding end covers at two ends of the combustion chamber 6, the end of the combustion chamber 6 is close to the expansion section 7, so that one end of the inner wall 2 (the position of the throat portion of the thrust chamber on the inner wall) penetrates through the outer wall 1 and is partially exposed near the expansion section, and the circumferential surface of the inner wall. The outer wall ring 9 plays a role of an outer wall, and for convenience of installation, the outer wall ring 9 can be formed by butt joint of two semicircular ring structures which are symmetrical to each other. When the combustion chamber is used, after one end (the position of the throat part of the thrust chamber on the inner wall) of the inner wall 2 of the combustion chamber penetrates through the outer wall, firstly, the outer wall ring 9 is sleeved on the outer side of the convex rib 4, the inner side of the outer wall ring 9 is tightly connected with the convex rib 4, and one end of the outer wall ring 9 is connected with one end of the outer wall 1 in an abutting mode.
In addition, as shown in fig. 1, fig. 2, fig. 3, fig. 4, fig. 5, fig. 9, fig. 10 and fig. 11, in order to facilitate the connection of the combustion chamber 6 with the expansion section 7, for example, after the combustion chamber 6 is subjected to the secondary pressurization treatment, the end cover 5, the gas guide tube 3 and the annular groove 8 are cut out in the radial direction of the combustion chamber 6, and the outer wall ring 9 of the combustion chamber 6 near the end of the expansion section 7 is cut out in the radial direction, so that the combustion chamber to be welded is obtained.
Similarly, after the expansion section 7 is subjected to secondary pressurization treatment, the end cover, the gas guide pipe and the annular groove are cut along the radial direction of the expansion section 7, and the outer wall 1 of the expansion section 7 close to the combustion chamber 6 is cut along the radial direction. By cutting off the outer wall ring 9 of the part of the combustion chamber 6 and by cutting off the outer wall of the expansion section close to the end of the combustion chamber 6 in a matching manner, it is ensured that the combustion chamber and the expansion section are matched in size with each other.
After the partial cutting of the outer wall ring 9 and the partial cutting of the outer wall 1 of the expansion section, the processing method further comprises the following steps: the inner walls of the combustion chamber 6 and the expansion section 7 near each other are welded, and the outer wall ring 9 having the cut close to each other is welded to the outer wall 1 through the connecting pipe 10, thereby obtaining the thrust chamber body assembly. Whole design benefit for when not influencing original effect, make combustion chamber 6 and expansion section 7 be connected inseparabler. In order to reduce the weight of the thrust chamber assembly and to ensure an aesthetic appearance, for example, the surfaces of the connecting pipe member 10 and the outer wall ring 9 may be subjected to a thinning process.
It should be noted that, in order to ensure that the outer wall 1, the inner wall 2 and the ribs 4 are clean and tidy, the influence of impurities on the welding strength is reduced, for example, the surfaces need to be cleaned before the outer wall 1, the inner wall 2 and the ribs 4 are used.
The thrust chamber body structure of the present embodiment is mainly described with reference to the combustion chamber 6 and the expansion section 7, and in practical application, the thrust chamber body structure may further include a third portion, a fourth portion, and the like, and the molding process of each portion is the same as that of the combustion chamber 6 or the expansion section 7.
The above embodiments may be combined with each other with corresponding technical effects.
The invention also provides a spacecraft thrust chamber which is prepared by adopting any one of the above welding processing methods for the body component of the thrust chamber.
The foregoing is merely an illustrative embodiment of the present invention, and any equivalent changes and modifications made by those skilled in the art without departing from the spirit and principle of the present invention should fall within the protection scope of the present invention.

Claims (10)

1. A method of manufacturing a double-layer structure for a thrust cell of a spacecraft, wherein the double-layer structure comprises an outer wall and an inner wall forming at least a part of a body part of the thrust cell of the spacecraft, and a spacer part fixed between the outer wall and the inner wall and arranged to provide a passage for a coolant to flow between the outer wall and the inner wall; the processing method comprises the following steps:
(1) providing a first wall structure, a second wall structure and a rib, wherein the first wall structure and the second wall structure are close to each other in an inner-outer nesting mode to form at least one part of a spacecraft thrust chamber body, and the rib is located between the first wall structure and the second wall structure;
(2) closing the first wall structure and the second wall structure at the ends, thereby forming a combined structure consisting of the first wall structure, the second wall structure and the ribs, the combined structure having a closed space inside;
(3) and (2) making the closed space under negative pressure, then carrying out first pressurization treatment on the combined structure from the outside, and then carrying out second pressurization treatment on the combined structure from the outside under the condition that the closed space is communicated with the outside, wherein the maximum pressure of the second pressurization treatment is greater than that of the first pressurization treatment, so that the combined structure is integrated to obtain the double-layer structure.
2. The processing method of the double-layer structure of the thrust chamber of the spacecraft as claimed in claim 1, wherein the pressure A during the first pressurization treatment is 1MPa or more and 20MPa or less, the pressurization time B is 0.2h or more and 10h or less, and the temperature C is 300 ℃ or more and 1300 ℃ or less.
3. The method for processing the double-layer structure of the thrust chamber of the spacecraft as claimed in claim 1, wherein the pressure D during the second pressurization treatment is 2MPa or more and 120MPa or less, the pressurization time E is 0.1h or more and 10h or less, and the temperature F is 300 ℃ or more and 1400 ℃ or less.
4. The method for processing the double-layer structure of the spacecraft thrust chamber according to claim 1, wherein the pressure of the closed space is controlled to be negative pressure through an air guide assembly, the air guide assembly comprises an air guide pipe and an annular groove which is communicated with the tail end of the air guide pipe and is arranged on one side of the first wall structure, and preferably the method further comprises a vacuum pumping device which is used for being connected with the air guide pipe.
5. The method of fabricating a spacecraft thrust chamber bilayer structure of claim 4 further comprising the step of cutting out an air guide.
6. The method of fabricating a spacecraft thrust chamber bilayer structure of claim 1 comprising a first bilayer structure and a second bilayer structure forming a thrust chamber body portion, and wherein the first bilayer structure and the second bilayer structure are each independently a combustor or an expansion section, the method comprising a first fabrication step to obtain the first bilayer structure and a second fabrication step to obtain the second bilayer structure, wherein the first fabrication step and the second fabrication step are performed simultaneously or sequentially, respectively.
7. A method for processing a spacecraft thrust chamber double-layer structure according to claim 1, wherein in the step (1), one side of the convex rib is fixed to the second wall structure in advance, and the other side of the convex rib is tightly attached to the surface of the first wall structure, so that the convex rib is formed into the spacer after the step (3).
8. A method of fabricating a spacecraft thrust chamber bilayer structure according to claim 1, further comprising the steps of providing an end cap and removing the end cap, the end cap being provided in connection with the ends of the first and second wall structures so as to close the first and second wall structures at the ends, the step of removing the end cap comprising cutting away the end cap in a radial direction of the bilayer structure.
9. A method of manufacturing a spacecraft thrust chamber bilayer structure according to claim 1, wherein one end of the first wall structure or one end of the second wall structure has a seal by which end face sealing of the first wall structure and the second wall structure is achieved.
10. A method of fabricating a spacecraft thrust chamber bilayer structure as claimed in claim 1 wherein said composite structure is externally pressurised by means of a high pressure vessel.
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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113500439A (en) * 2021-08-20 2021-10-15 宁夏东方钽业股份有限公司 Clamping device and method for machining welding surface of thin-wall niobium alloy thrust chamber
CN113894453A (en) * 2021-09-15 2022-01-07 蓝箭航天空间科技股份有限公司 Method for processing combustion chamber structure of dissimilar metal and combustion chamber
CN113982786A (en) * 2021-10-29 2022-01-28 中国人民解放军战略支援部队航天工程大学 Red copper-stainless steel rocket engine thrust chamber avoiding dissimilar metal welding
WO2022116955A1 (en) * 2020-12-01 2022-06-09 蓝箭航天空间科技股份有限公司 Method for processing double-layer structure of thrust chamber, and spacecraft thrust chamber
CN116038171A (en) * 2023-03-30 2023-05-02 北京星河动力装备科技有限公司 Regenerative cooling thrust chamber and manufacturing method thereof

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102950351A (en) * 2012-11-02 2013-03-06 首都航天机械公司 Diffusion brazing manufacture method for hydrogen peroxide thrust chamber body
CN104924036A (en) * 2015-06-16 2015-09-23 湖北三江航天江北机械工程有限公司 Manufacturing method of liquid-propellant rocket engine combustion chamber
JP5823069B1 (en) * 2015-01-23 2015-11-25 三菱重工業株式会社 Method for manufacturing rocket engine combustor, rocket engine combustor, and rocket engine
CN106078098A (en) * 2016-06-29 2016-11-09 航天材料及工艺研究所 A kind of preparation method of motor power room
CN111531321A (en) * 2020-07-09 2020-08-14 蓝箭航天空间科技股份有限公司 Method for repairing spacecraft thrust chamber jacket and spacecraft thrust chamber

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6205661B1 (en) * 1999-04-15 2001-03-27 Peter John Ring Method of making a rocket thrust chamber
ATE367519T1 (en) * 2001-12-18 2007-08-15 Volvo Aero Corp COMPONENT FOR EXPOSED TO HIGH THERMAL LOAD DURING OPERATION AND METHOD FOR PRODUCING SUCH A COMPONENT
US8584988B1 (en) * 2010-09-22 2013-11-19 The Boeing Company Mechanically fastened large pressurized spacecraft structure
CN206206011U (en) * 2016-10-31 2017-05-31 北京航天动力研究所 A kind of liquid rocket engine thrust chamber sandwich for strengthening bushing with throat
CN106438109A (en) * 2016-10-31 2017-02-22 北京航天动力研究所 Liquid-propellant rocket engine thrust chamber interlayer structure with throat reinforcing bushing
CN109759789B (en) * 2019-01-24 2020-07-03 蓝箭航天空间科技股份有限公司 Method for manufacturing pressure-resistant interlayer of thrust chamber of liquid rocket engine
CN209811495U (en) * 2019-03-04 2019-12-20 蓝箭航天空间科技股份有限公司 Composite tool for welding thrust chamber
CN209724523U (en) * 2019-03-13 2019-12-03 蓝箭航天技术有限公司 A kind of thrust chamber's structure
CN209850261U (en) * 2019-03-15 2019-12-27 蓝箭航天技术有限公司 Composite tool for thrust chamber preparation process
CN110202320B (en) * 2019-03-15 2024-06-04 蓝箭航天技术有限公司 Composite tool for thrust chamber preparation process and thrust chamber preparation process
CN112462714B (en) * 2020-12-01 2021-08-24 陕西蓝箭航天技术有限公司 Processing method of spacecraft thrust chamber double-layer structure
CN112177799B (en) * 2020-12-01 2021-03-19 蓝箭航天空间科技股份有限公司 Welding processing method of spacecraft thrust chamber body component and spacecraft thrust chamber

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102950351A (en) * 2012-11-02 2013-03-06 首都航天机械公司 Diffusion brazing manufacture method for hydrogen peroxide thrust chamber body
JP5823069B1 (en) * 2015-01-23 2015-11-25 三菱重工業株式会社 Method for manufacturing rocket engine combustor, rocket engine combustor, and rocket engine
CN104924036A (en) * 2015-06-16 2015-09-23 湖北三江航天江北机械工程有限公司 Manufacturing method of liquid-propellant rocket engine combustion chamber
CN106078098A (en) * 2016-06-29 2016-11-09 航天材料及工艺研究所 A kind of preparation method of motor power room
CN111531321A (en) * 2020-07-09 2020-08-14 蓝箭航天空间科技股份有限公司 Method for repairing spacecraft thrust chamber jacket and spacecraft thrust chamber

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2022116955A1 (en) * 2020-12-01 2022-06-09 蓝箭航天空间科技股份有限公司 Method for processing double-layer structure of thrust chamber, and spacecraft thrust chamber
GB2616149A (en) * 2020-12-01 2023-08-30 Landspace Science & Tech Co Ltd Method for processing double-layer structure of thrust chamber, and spacecraft thrust chamber
GB2616149B (en) * 2020-12-01 2024-07-17 Landspace Science & Tech Co Ltd Method for processing double-layer structure of thrust chamber, and spacecraft thrust chamber
CN113500439A (en) * 2021-08-20 2021-10-15 宁夏东方钽业股份有限公司 Clamping device and method for machining welding surface of thin-wall niobium alloy thrust chamber
CN113894453A (en) * 2021-09-15 2022-01-07 蓝箭航天空间科技股份有限公司 Method for processing combustion chamber structure of dissimilar metal and combustion chamber
CN113894453B (en) * 2021-09-15 2022-06-17 蓝箭航天空间科技股份有限公司 Method for processing combustion chamber structure of dissimilar metal and combustion chamber
CN113982786A (en) * 2021-10-29 2022-01-28 中国人民解放军战略支援部队航天工程大学 Red copper-stainless steel rocket engine thrust chamber avoiding dissimilar metal welding
CN113982786B (en) * 2021-10-29 2023-06-06 中国人民解放军战略支援部队航天工程大学 Rocket engine thrust chamber of red copper-stainless steel for avoiding dissimilar metal welding
CN116038171A (en) * 2023-03-30 2023-05-02 北京星河动力装备科技有限公司 Regenerative cooling thrust chamber and manufacturing method thereof

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