A rocket nozzle and method of manufacturing the same

Abstract

The application discloses a rocket nozzle and a manufacturing method thereof, comprising a thrust chamber core, a split pressure-bearing shell, an upper sealing cover plate and a lower fastening ring. The application adopts a stainless steel forged hot extrusion integrated thrust chamber core and at least two split shell cold assembly structures, which can adapt to the industrial demand of commercial aerospace rapid iteration, high-frequency test and batch production, with low equipment investment and short processing cycle. The thrust chamber core is a thin-walled Laval tube structure with multiple cooling ribs, which not only reduces the structural dead weight, but also shortens the heat exchange interval between the gas side and the cooling flow channel, greatly improves the heat exchange efficiency, and the cooling ribs can improve the structural rigidity and deformation resistance of the thrust chamber core, so that the rocket nozzle can withstand higher chamber pressure and higher heat flow. The upper sealing cover plate and the lower fastening ring are respectively limited and locked at the top and bottom of the rocket nozzle, which further enhances the structural rigidity of the rocket nozzle, so that the rocket nozzle can further improve the chamber temperature and the chamber pressure, and the working efficiency is higher.

Description

Technical Field

[0001] This invention relates to rockets. Background Technology

[0002] With the rapid development of the commercial aerospace industry, liquid rocket engines are gradually upgrading towards higher thrust, higher combustion chamber pressure, shorter development cycle, lower cost, and multiple reuse. As the core hot-end component that withstands high-temperature gas and high heat flux load, the manufacturing process of the regenerative cooling structure of the combustion chamber directly determines the mass production efficiency, operational reliability, and reuse life of the engine. The industry has put forward stringent requirements on the processing cycle, manufacturing cost, component density, and batch consistency of the combustion chamber forming process.

[0003] The existing mainstream forming schemes for regenerative cooling combustion chambers are divided into two categories: integral milling and vacuum brazing process and metal 3D printing integrated forming process. Both of these schemes have inherent technical defects that make them difficult to adapt to the needs of large-scale commercial aerospace development, rapid iteration, and high-frequency testing.

[0004] Metal 3D printing, a one-piece molding process, can directly form an integral combustion chamber with built-in irregularly shaped regenerative cooling channels, eliminating complex milling and multi-layer assembly processes. It can overcome the constraints of traditional machining structure design. However, the industrial application of this process faces significant bottlenecks: First, the investment costs for equipment and raw materials are extremely high. The investment for a single industrial-grade metal additive manufacturing equipment suitable for large-size combustion chambers can reach hundreds of thousands to tens of millions of yuan. The procurement cost of aerospace-grade spherical metal powder is 3 to 5 times that of traditional forged and rolled base materials, significantly increasing the production cost per unit. Second, controlling printing defects is difficult. The laser selective melting process is prone to defects such as internal micropores, incomplete interlayer fusion, printing stress cracks, and powder adhesion to the inner wall. The distribution of these defects varies depending on the specific process. The inherent limitations of 3D printing are significant. Firstly, the high rejection rate during subsequent non-destructive testing results in large variations in cooling heat exchange efficiency and structural strength between batches of combustion chambers, leading to a significant increase in test failure rates. Secondly, the excessively long component molding time, with single continuous printing of small and medium-sized regenerative cooling combustion chambers taking several days and high-thrust combustion chamber components taking one to two weeks, necessitates lengthy post-processing steps such as high-temperature stress-relief annealing, flow channel unblocking, and surface polishing. The overall delivery cycle cannot support the rapid iterative R&D and mass production needs of commercial aerospace, nor can it meet the demands of frequent test runs during the R&D process. Finally, 3D-printed combustion chambers are single, integral components; if localized ablation damage occurs on the inner wall, the entire unit must be scrapped, resulting in extremely high repair and reuse costs.

[0005] The traditional integral milling + multi-layer cylinder vacuum brazing forming process was a mature manufacturing route for early combustion chambers. The process involves more than ten steps, including large forging billets, integral deep groove milling of the cylinder wall, multi-layer cylinder assembly and positioning, multiple vacuum brazing processes, pressure leak testing, and precision machining. This process is lengthy and highly dependent on manual assembly and specialized large-scale processing equipment. This process has several shortcomings: First, the material utilization rate of integral forgings is low, and the superposition of multiple high-precision machining and brazing processes leads to high manufacturing costs per unit. Second, the processing cycle is as long as several months, which cannot match the production rhythm of frequent test runs and simultaneous delivery of multiple batches during the R&D phase. Third, the multi-layer cylinder relies on a large area of ​​brazing layer for sealing and connection; the brazing interface is prone to defects such as incomplete welds, microcracks, and flow channel blockage, making the combustion chamber highly susceptible to cooling medium leakage failure under high chamber pressure and high temperature conditions, resulting in a low product yield. Fourth, the brazing alloy bonding layer has weak resistance to thermal cycling shock; after multiple ignition tests and high / low temperature alternating loads, the brazing interface is prone to delamination and cracking, failing to meet the performance requirements of liquid rocket engines for repeated orbital insertions.

[0006] In addition, traditional rocket nozzle assembly generally adopts a high-temperature hot-fitting assembly process, which involves heating the outer shell to expand it before fitting it in, and then cooling and contracting it to achieve assembly. This process has obvious drawbacks: high-temperature heating can easily cause component oxidation, thermal deformation, and large residual assembly stress, resulting in uneven cooling channel gaps and poor structural dimensional accuracy. During engine service, it is prone to structural deformation, vibration loosening, and sealing failure, which seriously affects the working stability and service life of rocket engines.

[0007] In summary, existing integral brazing processes and integrated 3D printing processes cannot simultaneously meet the industry's demands for low cost, short manufacturing cycle, high structural reliability, good reusability, and batch performance consistency. There is an urgent need to propose a new regenerative cooling combustion chamber structure to overcome the various technical defects of existing manufacturing routes. Summary of the Invention

[0008] To overcome the shortcomings mentioned above, the present invention aims to provide a technical solution that can solve the above problems.

[0009] A rocket nozzle includes: a thrust chamber core, a segmented pressure-bearing outer shell, an upper cap plate, and a lower fastening ring; The thrust chamber core is a thin-walled Laval tube structure with multiple cooling ribs. The multiple cooling ribs are arranged in an axial array along the outer wall of the thrust chamber core, and a through-type regenerative cooling channel is formed between adjacent cooling ribs. The segmented pressure-bearing outer shell is composed of at least two segmented shells, which surround and cover the outside of the thrust chamber core. The segmented pressure-bearing outer shell and the thrust chamber core are assembled by low-temperature shrinkage cold installation interference fit, so that at least two segmented shells are tightly hugged to the outer side of the cooling rib. The lower end face of the upper cap plate is provided with an annular groove, the top of the segmented pressure-bearing shell is embedded in the annular groove, and the lower fastening ring radially locks the bottom of the segmented pressure-bearing shell. Preferably, the segmented pressure-bearing shell is composed of two symmetrically arranged segmented shells, and the vertical joint between the two segmented shells is sealed by sealing welding. Preferably, the rocket nozzle is assembled by a low-temperature shrinkage cold installation interference fit assembly method. The thrust chamber core is placed in a low-temperature environment of -200°C for rapid freezing in advance, and after it shrinks, it is taken out for assembly. The interference fit amount after assembly is 0.03 to 0.08 mm. Preferably, the thrust chamber core, the segmented pressure-bearing outer shell, the upper top cover plate, and the lower fastening ring are all made of stainless steel, and the thrust chamber core and the cooling ribs are an integrated stainless steel structure. Preferably, the annular groove covers and limits the top end of the segmented pressure-bearing shell; the lower fastening ring is an annular flange structure, which applies axial preload by bolt fastening to lock the radial clamping state and axial limiting of the segmented shell.

[0010] The present invention also proposes a method for manufacturing a rocket nozzle, which includes the following steps: S1. Separately process and prepare the thrust chamber core, segmented shell, upper cap plate and lower fastening ring made of stainless steel; S2. Place the processed thrust chamber core advance body in a low temperature environment of -200℃ for rapid freezing treatment, so that the entire thrust chamber core body shrinks at low temperature. During assembly, take out the thrust chamber core body after low temperature shrinkage and invert it. Use tooling to wrap and attach at least two segmented shells to the outer side of the cooling ribs on the outside of the thrust chamber core body. S3. Insert the lower fastening ring into the bottom of the segmented pressure-bearing shell, and apply radial locking force through bolts to lock the segmented shell in a radially clamped state. S4. After the thrust chamber core and segmented shell are assembled and locked in the inverted state, the whole is flipped and straightened in the right direction. The upper cover plate is closed from top to bottom, so that the top of the segmented shell is embedded and limited in the annular groove on the lower end face of the upper cover plate. The upper cover plate is locked and fixed by screws. S5. Seal the vertical joint between the segmented shells by welding, and perform precision machining on the overall shape of the nozzle to remove assembly allowance, thus obtaining the finished rocket nozzle. Preferably, the method for preparing the thrust chamber core is as follows: Stainless steel billets are selected and homogenized by forging to eliminate internal defects and obtain regular columnar pre-forged billets. The pre-forged billet is placed in the furnace and heated and fired. After being taken out, the pre-forged billet is hot-extruded using a special mold for Laval tube profiles to form a thrust chamber core blank with Laval tube shape, cooling ribs and regenerative cooling channels in one piece. Vacuum cryogenic stress relief was performed on the thrust chamber core blank. Precision grinding is performed on the outer surface of the cooling ribs to accurately control the assembly outer diameter; Preferably, the method for preparing the segmented shell is as follows: stainless steel sheet blanks are used, which are rolled and stamped into an arc-shaped structure, and the side butt bevels and cooling medium inlet and outlet are processed to precisely control the inner arc size of the shell.

[0011] Compared with the prior art, the advantages of the present invention are: This invention uses a forging and hot extrusion integrated molding process to replace 3D printing and deep groove milling processes. This invention has high raw material utilization, short processing cycle, and low production cost. In addition, its pipe wall has no micropores, the nozzle has strong stability, and with the vacuum low temperature stress relief process, residual stress in the molding is completely eliminated, ensuring the dimensional stability of the component. During assembly, only the vertical joints of the segmented shell are partially sealed by welding, without large-area welding operations, resulting in small welding deformation and high yield. It is suitable for the high-frequency test and mass production needs of commercial aerospace.

[0012] The thrust chamber core of this invention is a thin-walled Laval tube structure with multiple cooling ribs. This regenerative cooling channel achieves uniform heat exchange throughout the cooling medium. Its thin-walled structure not only significantly reduces the core matrix wall thickness and structural weight, but also shortens the heat exchange interval between the gas side and the cooling channel, greatly improving heat exchange efficiency. In addition, the axially uniformly distributed cooling ribs can significantly improve the structural rigidity and overall deformation resistance of the thin-walled thrust chamber core, effectively resisting high chamber pressure and high-speed gas impact loads, avoiding problems such as bulging, deformation, and vibration instability in the thin-walled rocket core, and significantly improving the overall structural strength and operational adaptability of the nozzle, enabling the rocket nozzle to withstand higher chamber pressure and higher heat flux.

[0013] This invention uses an upper cap plate and a lower fastening ring to lock the top and bottom of the rocket nozzle respectively. The upper and lower bidirectional mechanical locking achieves axial and radial full-dimensional locking and fixation of the nozzle, which further enhances the structural rigidity of the rocket nozzle and makes its seal more durable and reliable. This allows the rocket nozzle to further increase room temperature and chamber pressure, resulting in higher working efficiency.

[0014] The thrust chamber core adopts a low-temperature shrinkage cold installation interference fit assembly process to replace the traditional high-temperature hot fitting assembly. There is no high-temperature heating, no component oxidation thermal deformation, and no residual assembly stress throughout the process. It can stably achieve a precise interference fit of 0.03 to 0.08 mm. The segmented shell is evenly clamped and the cooling interlayer gap is consistent, resulting in uniform and stable heat exchange. This avoids vibration loosening and sealing failure from the root cause.

[0015] This invention only performs local sealing welding on the vertical short seams of the segmented shell, without large-area brazing operations, which greatly reduces welding deformation and welding defects; combined with vacuum low-temperature stress relief process, it completely eliminates residual stress in the forming process, resulting in excellent component dimensional stability, small performance dispersion of batch products, and high mass production reliability.

[0016] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0017] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0018] Figure 1 This is an exploded view of the structure of the present invention.

[0019] Figure 2 This is a cross-sectional view of the structure of the present invention.

[0020] Figure 3 This is a structural diagram of the upper capping plate of the present invention. Detailed Implementation

[0021] The technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0022] In the description of this invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0023] Furthermore, in the description of this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal connection of two components; they can refer to a wireless connection or a wired connection. Those skilled in the art can understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0024] Furthermore, the technical features involved in the different embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0025] Please see Figures 1-2 This invention provides a rocket nozzle, the main body of which consists of a thrust chamber core 1, a segmented pressure-bearing outer shell, an upper cap plate 3, and a lower fastening ring 4, all of which are made of stainless steel.

[0026] The thrust chamber core 1 is a thin-walled Laval tube structure with multiple cooling ribs 11. The multiple cooling ribs 11 are arranged in an axial array along the outer wall of the thrust chamber core 1, and a through-type regenerative cooling channel is formed between adjacent cooling ribs 11. This regenerative cooling channel achieves uniform heat exchange throughout the cooling medium. Its thin-walled structure not only significantly reduces the core matrix wall thickness and structural weight, but also shortens the heat exchange interval between the gas side and the cooling channel, greatly improving heat exchange efficiency. In addition, the axially uniformly distributed cooling ribs 11 can significantly improve the structural rigidity and overall deformation resistance of the thin-walled thrust chamber core 1, effectively resisting high chamber pressure and high-speed gas impact loads, avoiding problems such as bulging, deformation, and vibration instability of the thin-walled rocket core, and significantly improving the overall structural strength and operational adaptability of the nozzle, enabling the rocket nozzle to withstand higher chamber pressure and higher heat flux.

[0027] Combining the above technical solutions, such as Figures 2-3 As shown, the present invention also limits and locks the top and bottom of the rocket nozzle by using the annular groove 31 of the upper cap plate 3, thereby achieving axial and radial full-dimensional locking and fixing of the nozzle, further enhancing the structural rigidity of the rocket nozzle, making its sealing more durable and reliable, and enabling the rocket nozzle to further increase room temperature and chamber pressure, resulting in higher working efficiency.

[0028] The segmented pressure-bearing shell consists of two completely symmetrical segmented shells 2, which together enclose the outside of the thrust chamber core 1. The thrust chamber core 1 is tightly fitted to the outer side of the cooling rib 11 through a cryogenic freezing cold fitting, forming a sealed regenerative cooling interlayer between the shell and the core. The vertical joint gaps between the two segmented shells 2 are sealed by partial sealing welding, only sealing the gaps and not performing full welding, to minimize welding deformation and ensure the structural dimensional accuracy and the airtightness of the cooling interlayer.

[0029] The lower end face of the upper cap plate 3 has an annular groove 31. The top of the segmented pressure-bearing shell is integrally embedded in the annular groove 31 to achieve circumferential positioning and axial limitation. At the same time, it is fixed by screws to achieve top sealing and upper locking. The lower fastening ring 4 is an annular flange structure, which is fitted onto the bottom end of the segmented shell 2. Axial and radial preload is applied by bolt tightening to lock the segmented shell 2 into a clamped state. Together with the upper cap plate 3, it forms a two-way locking structure to ensure that the overall structure of the nozzle does not move or loosen.

[0030] In one embodiment, the rocket nozzle of the present invention is assembled by a cryogenic shrinkage cold installation interference fit assembly method. Through high-precision closed-loop machining and size-graded matching mode, combined with cryogenic cold installation process, a precise interference fit of 0.03 to 0.08 mm is stably achieved.

[0031] Specifically, each segmented shell 2 is formed using a high-precision mold. After forming, the inner arc dimension of each piece is inspected by an inner diameter precision measuring instrument, and the dimensions are graded and screened. At the same time, the outer diameter of the cooling rib 11 is also calibrated piece by piece using a three-coordinate measuring machine to strictly control the dimensional tolerance of each piece. During assembly, the thrust chamber core 1 is placed in a low temperature environment of -200℃ for rapid freezing beforehand. After it shrinks, it is taken out for assembly. After assembly and the thrust chamber core 1 returns to room temperature, the thrust chamber core 1 expands, so that the cooling rib 11 fits tightly against the inner wall of the segmented shell 2. The interference fit is 0.03~0.08mm, which can avoid assembly jamming and shell plastic deformation caused by excessive interference fit, and ensure that the segmented shell 2 has sufficient clamping force and no risk of loosening and leakage.

[0032] It should be noted that the interference fit amount mentioned above refers to the difference in size between the outer diameter of the outer side of the cooling rib 11 and the inner diameter of the inner arc of the segmented shell 2 under normal temperature conditions. This difference is the interference fit amount.

[0033] The present invention also proposes a method for manufacturing a rocket nozzle, which includes the following steps: S1. Prepare the thrust chamber core 1, segmented shell 2, upper top cover plate 3 and lower fastening ring 4 in advance after processing and inspection to ensure that all parts are dimensionally qualified and surface-free. S2. Place the thrust chamber core 1 in a low temperature environment of -200℃ for rapid freezing and heat preservation, so that the thrust chamber core 1 shrinks uniformly at low temperature and the outer diameter of the assembly is slightly reduced. Quickly take out the core after low temperature shrinkage and place it upside down. Use special tooling to accurately align and wrap the two symmetrical segmented shells 2, so that the inner arc of the segmented shells 2 is tightly attached to the outer side of the cooling rib 11. S3. Insert the lower fastening ring 4 into the bottom of the segmented pressure-bearing shell, tighten the fastening bolts evenly, apply radial locking force through the bolt pre-tightening force, lock the radial clamping state of the two segmented shells 2, and complete the initial assembly and fixation of the core and shell. S4. After the inverted assembly and locking are completed, the whole component is flipped upright and placed in the right direction. The upper cover plate 3 is taken out and installed from top to bottom, so that the top of the segmented shell 2 is accurately embedded in the annular groove 31 on the lower end face of the cover plate to achieve axial limiting and circumferential positioning. Then, the upper cover plate 3 is locked and fixed with screws to complete the locking and sealing of the upper structure. S5. Seal the vertical joint between the two segmented shells 2 by partial sealing weld. At the same time, seal the gap between the upper cap plate 3, the lower fastening ring 4 and the split shell by welding. After welding, perform precision machining on the overall shape of the nozzle and the mounting flange surface to remove assembly allowance and machining burrs, correct the overall dimensional accuracy, and finally obtain the finished rocket nozzle.

[0034] The preparation method of the thrust chamber core 1 is as follows: First, stainless steel billets are selected for homogenization forging. The forging process breaks down the original defects such as pores, looseness, and compositional segregation inside the billet, refines the metal grains, and obtains a columnar pre-forged billet with dense structure, uniform properties, and regular dimensions, ensuring the consistency of the performance of the subsequently formed components. The prepared pre-forged billet is placed in a heating furnace for heating and firing to improve the overall plasticity of the billet. Then, the high-temperature billet is taken out and hot extruded as a whole using a special Laval tube profile mold. The complete Laval tube aerodynamic profile, outer wall array cooling ribs 11 and thrust chamber core 1 billet with through-type regenerative cooling channel are formed simultaneously through a one-time extrusion process. The hot-extruded core blank is sent into a vacuum heat treatment furnace for vacuum low-temperature stress relief treatment. After being kept at a constant temperature within the non-phase transformation temperature range of stainless steel, it is slowly cooled in the furnace to completely eliminate the residual internal stress generated by hot extrusion plastic deformation and avoid subsequent processing springback, assembly deformation and cracking problems. Finally, the outer surface of the core cooling rib 11 is precision ground to accurately control the outer diameter of the cooling rib 11 assembly and strictly control the assembly tolerance, providing a reliable dimensional basis for subsequent low-temperature cold fitting and precise interference fit.

[0035] The preparation method of segmented shell 2 is as follows: stainless steel sheet blanks are used as raw materials, and special arc molds are used for bending or stamping to obtain an arc shell structure with regular curvature and uniform contour accuracy. After forming, the two sides of the shell are beveled, and the cooling medium inlet and outlet structures are precisely processed. The inner arc assembly dimensions of segmented shell 2 are strictly controlled to ensure that the multiple shells can be precisely matched with the outer side of the cooling rib 11 of the thrust chamber core 1 after being enclosed, so as to meet the preset interference fit requirements.

[0036] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above, and that the invention can be implemented in other specific forms without departing from the spirit or essential characteristics of the invention. Therefore, the embodiments should be considered in all respects as exemplary and non-limiting, and the scope of the invention is defined by the appended claims rather than the foregoing description. Thus, it is intended that all variations falling within the meaning and scope of equivalents of the claims be included within the present invention.

Claims

1. A rocket nozzle, comprising a thrust chamber core, a segmented pressure-bearing outer shell, an upper cap plate, and a lower fastening ring; characterized in that... : The thrust chamber core is a thin-walled Laval tube structure with multiple cooling ribs. The multiple cooling ribs are arranged in an axial array along the outer wall of the thrust chamber core, and a through-type regenerative cooling channel is formed between adjacent cooling ribs. The segmented pressure-bearing outer shell is composed of at least two segmented shells, which surround and cover the outside of the thrust chamber core. The segmented pressure-bearing outer shell and the thrust chamber core are assembled by low-temperature shrinkage cold installation interference fit, so that at least two segmented shells are tightly hugged to the outer side of the cooling rib. The lower end face of the upper cap plate is provided with an annular groove, the top of the segmented pressure-bearing shell is embedded in the annular groove, and the lower fastening ring radially locks the bottom end of the segmented pressure-bearing shell.

2. The rocket nozzle according to claim 1, characterized in that: The segmented pressure-bearing shell consists of two symmetrically arranged segmented shells, and the vertical joint between the two segmented shells is sealed by sealing welding.

3. The rocket nozzle according to claim 1, characterized in that: The rocket nozzle is assembled using a cryogenic shrinkage cold installation interference fit assembly method. The thrust chamber core is pre-frozen in a -200°C environment and then removed for assembly after shrinkage. The interference fit amount after assembly is 0.03 to 0.08 mm.

4. The rocket nozzle according to claim 1, characterized in that: The thrust chamber core, the segmented pressure-bearing outer shell, the upper cap plate, and the lower fastening ring are all made of stainless steel, and the thrust chamber core and the cooling ribs are an integrated stainless steel structure.

5. The rocket nozzle according to claim 1, characterized in that: The annular groove covers and limits the top of the segmented pressure-bearing shell; the lower fastening ring is an annular flange structure, which applies axial preload by bolt fastening to lock the radial clamping state and axial limit of the segmented shell.

6. A method for manufacturing a rocket nozzle, used to prepare the rocket nozzle according to any one of claims 1-5, characterized in that, Includes the following steps: S1. Separately process and prepare the thrust chamber core, segmented shell, upper cap plate and lower fastening ring made of stainless steel; S2. Place the processed thrust chamber core advance body in a low temperature environment of -200℃ for rapid freezing treatment, so that the entire thrust chamber core body shrinks at low temperature. During assembly, take out the thrust chamber core body after low temperature shrinkage and invert it. Use tooling to wrap and attach at least two segmented shells to the outer side of the cooling ribs on the outside of the thrust chamber core body. S3. Insert the lower fastening ring into the bottom of the segmented pressure-bearing shell, and apply radial locking force through bolts to lock the segmented shell in a radially clamped state. S4. After the thrust chamber core and segmented shell are assembled and locked in the inverted state, the whole is flipped and straightened in the right direction. The upper cover plate is closed from top to bottom, so that the top of the segmented shell is embedded and limited in the annular groove on the lower end face of the upper cover plate. The upper cover plate is locked and fixed by screws. S5. Seal the vertical joint between the segmented shells by welding, and perform precision machining on the overall shape of the nozzle to remove assembly allowance, thus obtaining the finished rocket nozzle.

7. The method for manufacturing a rocket nozzle according to claim 6, characterized in that: The method for preparing the thrust chamber core is as follows: Stainless steel billets are selected and homogenized by forging to eliminate internal defects and obtain regular columnar pre-forged billets. The pre-forged billet is placed in the furnace and heated and fired. After being taken out, the pre-forged billet is hot-extruded using a special mold for Laval tube profiles to form a thrust chamber core blank with Laval tube shape, cooling ribs and regenerative cooling channels in one piece. Vacuum cryogenic stress relief was performed on the thrust chamber core blank. The outer surface of the cooling ribs is precision ground to accurately control the assembly outer diameter.

8. The method for manufacturing a rocket nozzle according to claim 7, characterized in that, The method for preparing the segmented shell is as follows: stainless steel sheet blanks are used, which are rolled and stamped into an arc-shaped structure, and the side butt bevels and cooling medium inlet and outlet are processed to precisely control the inner arc size of the shell.

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