Liquid rocket thrust chamber and liquid rocket engine

Abstract

The present application provides a liquid rocket thrust chamber and a liquid rocket engine, relating to the technical field of liquid rocket engines. The thrust chamber comprises an inner wall and an outer wall, with a sandwich structure between the inner wall and the outer wall. A plurality of ribs are evenly arranged within the sandwich structure along the circumference of the thrust chamber, with adjacent ribs forming a cooling channel for the flow of a cooling medium. The thrust chamber comprises a combustion chamber, a throat, and a nozzle. The combustion chamber comprises a convergent section. The convergent section, the throat, and the nozzle are sequentially arranged along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channel, the cooling medium flows through the nozzle, the throat, and the convergent section in sequence. The ribs gradually decrease in width and height from the end of the nozzle away from the throat to the throat, and gradually increase in width and height from the throat to the end of the convergent section away from the throat. This arrangement can enhance the turbulence effect on the cooling medium, increase the heat exchange area between the cooling medium and the inner wall, and easily control the flow resistance loss of the cooling medium.

Description

Technical Field

[0001] The present application relates to the technical field of liquid rocket engines, and in particular to a liquid rocket thrust chamber and a liquid rocket engine. Background Art

[0002] The thrust chamber is the component of a liquid rocket engine that converts propellant energy and generates thrust. It consists of injectors, a combustion chamber, and a nozzle. During operation, the propellant is mixed and atomized by the injectors at the head of the thrust chamber. It is then heated and vaporized by the flames in the combustion chamber, where it burns intensely, transforming into high-temperature, high-pressure combustion gas. This combustion gas is then accelerated through the nozzle into a high-velocity stream and ejected backward, generating thrust on the engine and propelling the rocket forward.

[0003] During operation, the thrust chamber walls are subjected to intense heat loads from high-temperature combustion gases. To reduce this heat load and prevent structural damage to the thrust chamber walls due to high temperatures, regenerative cooling can be used to cool the thrust chamber.

[0004] The thrust chamber consists of an inner wall, an outer wall, and an intermediate layer. The intermediate layer is equipped with multiple ribs, with cooling channels between adjacent ribs, through which the propellant flows. Existing technology typically uses ribs with uniform cross-sections. The cooling channels between adjacent ribs only narrow and widen in the width direction, resulting in a relatively simple flow disturbance for the propellant, poor heat exchange, and difficulty in effectively cooling areas with high heat loads. Summary of the Invention

[0005] The present application provides a liquid rocket thrust chamber and a liquid rocket engine, which can improve the turbulence effect on the cooling medium, increase the heat exchange area between the cooling medium and the inner wall, and easily control the flow resistance loss of the cooling medium, thereby improving the cooling effect of the thrust chamber and the overall efficiency of the engine.

[0006] To achieve the above objectives, this application adopts the following technical solutions:

[0007] In a first aspect, the present application provides a liquid rocket thrust chamber, which has an inner wall and an outer wall, with a sandwich structure between the inner wall and the outer wall, and a plurality of ribs are evenly arranged along the circumference of the thrust chamber in the sandwich structure, and a cooling channel is formed between adjacent ribs, and the cooling channel is used for a cooling medium to flow; the thrust chamber includes a combustion chamber, a throat and a nozzle, and the combustion chamber has a convergent section, and the convergent section, throat and nozzle are distributed in sequence along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channel, the cooling medium flows through the nozzle, throat and convergent section in sequence; wherein, from the end of the nozzle away from the throat to the throat, the rib width and rib height of the ribs gradually decrease; from the throat to the end of the convergent section away from the throat, the rib width and rib height of the ribs gradually increase.

[0008] In some possible implementations, the rib is provided with at least one through hole at the throat location, and every two adjacent cooling channels are connected through the through hole. When the cooling medium flows through the throat location in the cooling channel, the cooling medium converges at the through hole.

[0009] In some possible implementations, the through holes on every two adjacent ribs among the plurality of ribs are connected along the circumference of the thrust chamber to form an annular passage in the sandwich structure.

[0010] In some possible implementations, the rib is provided with a plurality of through holes, and the plurality of through holes are distributed along the axial direction of the thrust chamber, so as to form a plurality of annular passages distributed along the axial direction of the thrust chamber in the sandwich structure.

[0011] In some possible implementations, the through-holes on the ribs have openings facing the inner wall. When the cooling medium flows in the cooling channel, the cooling medium flows through the inner wall positions corresponding to the openings of the through-holes.

[0012] In some possible implementations, the rib width and rib height at one end of the nozzle away from the throat are respectively the first rib width and the first rib height, the rib width and rib height at one end of the convergent section away from the throat are respectively the second rib width and the second rib height, and the rib width and rib height at the throat are respectively the third rib width and the third rib height;

[0013] Among them, the first rib width is greater than the second rib width, the second rib width is greater than the third rib width, the first rib height is greater than the second rib height, and the second rib height is greater than the third rib height.

[0014] In some possible implementations, the width of the side of the rib connected to the inner wall is K1, and the width of the side of the rib connected to the outer wall is K2, where K2>K1.

[0015] In some possible implementations, a connecting shell is provided on the outer wall of the nozzle at one end of the nozzle away from the throat, a cooling medium discharge cavity is provided between the connecting shell and the outer wall, and the cooling medium discharge cavity is communicated with the inlet of the cooling channel;

[0016] The connecting shell is provided with a discharge port connected to the cooling medium discharge cavity. The cross-sectional area of ​​the discharge port gradually decreases from the outside of the cooling medium discharge cavity to the inside of the cooling medium discharge cavity to increase the flow rate of the cooling medium discharged from the discharge port.

[0017] In some possible implementations, the combustion chamber further has a straight section, which is located at an end of the convergent section away from the throat;

[0018] The thrust chamber is also provided with a discharge portion, which is arranged at one end of the straight cylinder section away from the convergent section. The discharge portion is provided with a cooling medium discharge cavity, and the cooling medium discharge cavity is connected with the outlet of the cooling channel. The discharge portion is provided with a discharge outlet connected with the cooling medium discharge cavity. The cross-sectional area of ​​the discharge outlet gradually increases from the inside of the cooling medium discharge cavity to the outside of the cooling medium discharge cavity to slow down the flow rate of the cooling medium flowing out of the discharge outlet.

[0019] In a second aspect, the present application provides a liquid rocket engine, comprising the liquid rocket thrust chamber of the first aspect.

[0020] It can be seen from the above technical solution that this application has at least the following beneficial effects:

[0021] The liquid rocket thrust chamber provided by the present application has a rib width and a rib height that first gradually decrease and then gradually increase from the end of the nozzle away from the throat to the throat, and from the throat to the end of the convergent section of the combustion chamber away from the throat, so that the width and height of the cooling channel between adjacent ribs first gradually decrease and then gradually increase, so that the cooling medium flows in the cooling channel with a constantly changing flow cross-section, thereby improving the disturbing effect on the cooling medium, enhancing the degree of turbulence, improving the heat transfer coefficient between the cooling medium and the wall, improving the heat transfer effect, and enhancing the cooling effect of the thrust chamber; and compared with the uniform cross-section ribs in the prior art, the width of the cooling channel is increased, so that the contact area between the cooling medium and the inner wall is increased, thereby increasing the heat exchange area between the cooling medium and the inner wall, and improving the cooling effect.

[0022] Furthermore, the cooling channel corresponding to the nozzle is expanded, so that the flow resistance loss of the cooling medium flowing in from the cooling channel inlet is smaller; the rib height corresponding to the throat is smaller, the corresponding cooling channel is contracted, and the cooling medium flow rate is higher, which quickly takes away heat and achieves maximum cooling for the throat with the highest heat load; although the cooling medium has a large flow resistance loss in the cooling channel corresponding to the throat, the cooling medium has a small flow resistance loss in the cooling channel corresponding to the nozzle, so that the flow resistance loss of the cooling medium flowing from the nozzle to the throat in the cooling channel corresponding to the throat is easy to control within the required range, thereby making the cooling medium pressure and flow rate flowing into the cooling channel corresponding to the convergent section meet the requirements, ensuring the cooling effect on the convergent section with higher heat load, and realizing targeted cooling according to the difference in axial heat load of the thrust chamber to prevent local overheating, and reducing the pump power demand and improving the overall efficiency of the liquid rocket engine.

[0023] It should be understood that the description of technical features, technical solutions, beneficial effects or similar language in this application does not imply that all features and advantages can be realized in any single embodiment. On the contrary, it is understood that the description of a feature or beneficial effect means that a specific technical feature, technical solution or beneficial effect is included in at least one embodiment. Therefore, the description of a technical feature, technical solution or beneficial effect in this specification does not necessarily refer to the same embodiment. Furthermore, the technical features, technical solutions and beneficial effects described in the present embodiment can also be combined in any appropriate manner. Those skilled in the art will understand that the embodiment can be implemented without one or more specific technical features, technical solutions or beneficial effects of a specific embodiment. In other embodiments, additional technical features and beneficial effects can also be identified in specific embodiments that do not embody all embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 A schematic diagram of a liquid rocket thrust chamber provided in this application in a specific embodiment;

[0025] Figure 2 for Figure 1 Partial cross-sectional view;

[0026] Figure 3 for Figure 1 a cross-sectional view in one perspective;

[0027] Figure 4 for Figure 3 Enlarged view of position II in the middle;

[0028] Figure 5 for Figure 1 A cross-sectional view from another perspective;

[0029] Figure 6 for Figure 5 Enlarged view of point III in the middle;

[0030] Figure 7 for Figure 2 Enlarged view of point I in the middle.

[0031] Figure numbers: 1-inner wall; 2-outer wall; 21-connecting shell; 211-cooling medium discharge cavity; 3-rib; 31-through hole; 32-annular passage; 4-cooling channel; 5-nozzle; 6-combustion chamber; 61-convergent section; 62-straight cylinder section; 7-throat. DETAILED DESCRIPTION

[0032] The terms "first", "second" and "third" in this application specification and the accompanying drawings are used to distinguish different objects rather than to limit a specific order.

[0033] In the embodiments of this application, words such as "exemplary" or "for example" are used to indicate examples, illustrations, or descriptions. Any embodiment or design described as "exemplary" or "for example" in the embodiments of this application should not be interpreted as being preferred or advantageous over other embodiments or designs. Rather, the use of words such as "exemplary" or "for example" is intended to present the relevant concepts in a concrete manner.

[0034] To make the description of the following embodiments clear and concise, a brief introduction to the related technologies is first given:

[0035] A liquid rocket is a type of rocket that uses liquid propellant. It consists of multiple components, including the rocket body and engine. Liquid rockets mix and burn liquid fuel and oxidizer within the engine, generating high-temperature, high-pressure combustion gases that are ejected backward at high speed, generating forward thrust and propelling the rocket into space.

[0036] The liquid rocket engine is the core component of the liquid rocket, mainly composed of thrust chamber, propellant supply system and engine control system.

[0037] The thrust chamber is a key component of a liquid rocket engine. The propellant supply system transports liquid propellant to the thrust chamber, where it is mixed and burned in the combustion chamber of the thrust chamber to produce high-temperature and high-pressure combustion gas, which is ejected at high speed from the nozzle of the thrust chamber, thereby generating thrust.

[0038] The thrust chamber has an inner wall, an outer wall and an intermediate layer. A plurality of ribs are arranged in the intermediate layer. Each rib extends from one end of the combustion chamber to one end of the nozzle. There is a cooling channel between adjacent ribs.

[0039] During operation, the thrust chamber walls are subjected to intense heat loads from high-temperature combustion gases. To reduce this heat load and prevent structural damage to the thrust chamber walls due to high temperatures, regenerative cooling can be used to cool the thrust chamber.

[0040] Regenerative cooling involves transferring low-temperature propellant from one end of the thrust chamber nozzle into a cooling channel. The low-temperature propellant absorbs and removes heat from the walls, reducing the thrust chamber's thermal load. The heated propellant then flows out of the cooling channel and into the combustion chamber to participate in combustion. This method effectively cools the thrust chamber while utilizing the heat absorbed during the cooling process to preheat the propellant, improving its combustion efficiency.

[0041] The thrust chamber consists of a combustion chamber, a throat, and a nozzle. The combustion chamber has a straight section and a convergent section, arranged in this order along the thrust chamber's axis. The nozzle is the expansion section of the thrust chamber, with the convergent section and nozzle gradually converging from either side of the throat. Propellant burns in the straight section of the combustion chamber, producing high-temperature combustion gases. These gases flow through the convergent section, throat, and nozzle in sequence. The convergent section gradually increases the gas flow rate. The throat is the smallest part of the thrust chamber, where the gas accelerates rapidly. The nozzle further expands and accelerates the gas, achieving a higher injection velocity.

[0042] When the gas flows through the thrust chamber, the inner wall of the thrust chamber is directly subjected to the erosion of the high-temperature gas. When the propellant flows through the cooling channel, the propellant mainly absorbs the heat from the inner wall.

[0043] When the propellant burns in the straight section of the combustion chamber, the resulting high-temperature combustion gases come into contact with the corresponding inner walls at this location, undergoing intense heat exchange, increasing the heat load in this section. As the high-temperature combustion gases flow out of the straight section and through the convergent section, their velocity gradually increases. As they reach the throat, they accelerate rapidly there. Consequently, the heat load increases from the convergent section to the throat, reaching its highest point at the throat. From the throat to the nozzle, the gas expands at an accelerated rate, gradually decreasing its temperature and pressure. The intensity of the heat exchange with the corresponding inner walls also weakens, and the heat load gradually decreases from the throat to the nozzle.

[0044] In the prior art, uniform-section ribs are typically used. The spacing between adjacent uniform-section ribs decreases from the nozzle to the throat, and increases from the throat to the convergent section. Consequently, the cooling channels between adjacent uniform-section ribs in the thrust chamber narrow and widen only in the width direction, resulting in a relatively limited turbulence effect on the propellant. This, in turn, leads to poor heat transfer and cooling in the throat and convergent sections, where heat loads are high. Furthermore, the uniform-section ribs are the same size in both high- and low-heat-load areas, resulting in a certain amount of resource waste.

[0045] Moreover, in the cooling channel formed between adjacent ribs of equal cross-section, the pressure drop of the cooling medium at different positions in the cooling channel varies greatly. The pressure drop of the cooling medium is higher when it flows from the nozzle to the throat, resulting in the pressure and flow of the cooling medium flowing from the throat to the convergent section being difficult to meet the requirements, and the cooling effect on the inner wall corresponding to the convergent section is insufficient, which not only affects the overall cooling efficiency, but also increases unnecessary pump power requirements.

[0046] In view of this, an embodiment of the present application provides a liquid rocket thrust chamber, which can enable the cooling medium to flow in a cooling channel with a continuously changing flow cross-section, thereby improving the disturbing effect on the cooling medium, enhancing the degree of turbulence, improving the heat transfer coefficient between the cooling medium and the wall, improving the heat transfer effect, and enhancing the cooling effect of the thrust chamber; compared with the uniform cross-section ribs in the prior art, the width of the cooling channel is increased to increase the contact area between the cooling medium and the inner wall, thereby increasing the heat exchange area between the cooling medium and the inner wall and improving the cooling effect.

[0047] Compared with the equal-section ribs in the prior art, the cooling channel corresponding to the nozzle is enlarged, so that the flow resistance loss of the cooling medium flowing in from the cooling channel inlet is smaller, and the flow velocity of the cooling medium at the corresponding position of the throat is higher, which quickly takes away the heat and achieves maximum cooling of the throat with the highest heat load. Although the flow resistance loss of the cooling medium in the cooling channel corresponding to the throat is large, the flow resistance loss of the cooling medium in the cooling channel corresponding to the nozzle is small, so that the flow resistance loss of the cooling medium flowing from the nozzle to the throat in the cooling channel corresponding to the throat is easy to control within the required range, thereby making the pressure and flow of the cooling medium flowing into the cooling channel corresponding to the convergent section meet the requirements, ensuring the cooling effect of the convergent section with higher heat load, and realizing targeted cooling according to the difference in axial heat load of the thrust chamber to prevent local overheating, and reducing the pump power demand and improving the overall efficiency of the liquid rocket engine.

[0048] The following describes the liquid rocket thrust chamber provided in the embodiment of the present application with reference to the accompanying drawings:

[0049] like Figures 1 to 6 As shown, the thrust chamber has an inner wall 1 and an outer wall 2, with a sandwich structure between the inner wall 1 and the outer wall 2. A plurality of ribs 3 are evenly arranged in the sandwich structure along the circumference of the thrust chamber, and cooling channels 4 are formed between adjacent ribs 3 for the flow of cooling medium.

[0050] The thrust chamber includes a combustion chamber 6, a throat 7 and a nozzle 5. The combustion chamber 6 has a convergent section 61. The convergent section 61, the throat 7 and the nozzle 5 are distributed in sequence along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channel 4, the cooling medium flows through the nozzle 5, the throat 7 and the convergent section 61 in sequence.

[0051] Among them, the rib width and rib height of the rib 3 gradually decrease from the end of the nozzle 5 away from the throat 7 to the throat 7; and the rib width and rib height of the rib 3 gradually increase from the throat 7 to the end of the convergent section 61 away from the throat 7.

[0052] The thrust chamber includes a combustion chamber 6, a throat 7 and a nozzle 5 distributed along the axial direction. The combustion chamber 6, the throat 7 and the nozzle 5 can be integrally formed to ensure the overall strength of the thrust chamber.

[0053] The combustion chamber 6 has a straight section 62 and a convergent section 61. The propellant burns in the straight section 62 of the combustion chamber 6 to generate high-temperature combustion gas. The high-temperature combustion gas is ejected outward through the convergent section 61, the throat 7 and the nozzle 5 to generate propulsion.

[0054] like Figures 2 and 3 As shown, the thrust chamber gradually contracts from the convergent section 61 to the throat 7 , and gradually expands from the throat 7 to the nozzle 5 .

[0055] The inlet of cooling channel 4 is located at the end of nozzle 5 away from throat 7, and the outlet of cooling channel 4 is located at the end of straight section 62 of combustion chamber 6 away from the throat. Cooling medium flows into cooling channel 4 from the inlet, sequentially passing through nozzle 5, throat 7, convergent section 61, and straight section 62. The cooling medium absorbs and removes heat from the nozzle 5, throat 7, convergent section, and straight section 62, thereby cooling the thrust chamber.

[0056] The cooling medium may specifically be the propellant of a liquid rocket.

[0057] From the end of the nozzle 5 away from the throat 7 to the throat 7, the rib width and rib height of the ribs 3 gradually decrease, so that the width and height of the cooling channel 4 between adjacent ribs 3 gradually decrease. In the process of the cooling medium flowing from the end of the nozzle 5 away from the throat 7 to the throat 7, the flow rate of the cooling medium gradually increases. At the position of the throat 7, the flow rate of the cooling medium reaches the maximum to quickly take away the heat of the throat 7; from the throat 7 to the end of the convergent section 61 away from the throat 7, the rib width and rib height of the ribs 3 gradually increase, so that the width and height of the cooling channel 4 between adjacent ribs 3 gradually increase. In the process of the cooling medium flowing from the throat 7 to the end of the convergent section 61 away from the throat 7, the flow rate of the cooling medium gradually decreases from the maximum flow rate.

[0058] Specifically, the rib width and rib height of the rib 3 can change linearly from the end of the nozzle 5 away from the throat 7 to the throat 7; and the rib width and rib height of the rib 3 can change linearly from the throat 7 to the end of the convergent section 61 away from the throat 7.

[0059] In this embodiment, the rib width and rib height of the ribs 3 first gradually decrease and then gradually increase from the end of the nozzle 5 away from the throat 7 to the throat 7, and from the throat 7 to the end of the convergent section 61 away from the throat 7, so that the width and height of the cooling channel 4 between adjacent ribs 3 first gradually decrease and then gradually increase, so that the cooling medium flows in the cooling channel 4 with a continuously changing flow cross-section, thereby improving the disturbing effect on the cooling medium, enhancing the degree of turbulence, improving the heat transfer coefficient between the cooling medium and the wall, improving the heat transfer effect, and enhancing the cooling effect on the thrust chamber.

[0060] The rib width of the rib 3 gradually decreases from the end of the nozzle 5 away from the throat 7 to the throat 7, and from the end of the convergent section 61 away from the throat 7 to the throat 7. Compared with the uniform cross-section ribs in the prior art, the width of the cooling channel 4 is increased, so that the contact area between the cooling medium and the inner wall 1 is increased, thereby increasing the heat exchange area between the cooling medium and the inner wall 1 and improving the cooling effect.

[0061] Moreover, the rib height of the rib 3 corresponding to the nozzle 5 is greater than the rib height of the rib 3 corresponding to the throat 7, and the cooling channel 4 corresponding to the nozzle 5 is expanded, so that the flow resistance loss of the cooling medium flowing in from the inlet of the cooling channel 4 is smaller; the rib height corresponding to the throat 7 is smaller, and the corresponding cooling channel 4 is contracted, and the cooling medium flow rate is higher, which quickly takes away heat and achieves maximum cooling of the throat 7 with the highest heat load; although the cooling medium has a large flow resistance loss in the cooling channel 4 corresponding to the throat 7, the cooling medium has a small flow resistance loss in the cooling channel 4 corresponding to the nozzle 5, so that the flow resistance loss of the cooling medium flowing from the nozzle 5 to the throat 7 in the cooling channel 4 corresponding to the throat 7 is easy to control within the required range, thereby making the pressure and flow of the cooling medium flowing into the cooling channel 4 corresponding to the convergent section 61 meet the requirements, ensuring the cooling effect on the convergent section 61 with a higher heat load, and realizing targeted cooling according to the difference in axial heat load of the thrust chamber to prevent local overheating, and reducing the pump work demand and improving the overall efficiency of the liquid rocket engine.

[0062] In a specific embodiment, Figures 1-3 As shown, at the end of the nozzle 5 away from the throat 7, a connecting shell 21 is provided on the outer wall 2 of the nozzle 5, and a cooling medium discharge cavity 211 is provided between the connecting shell 21 and the outer wall 2. The cooling medium discharge cavity 211 is connected to the inlet of the cooling channel 4; a discharge inlet connected to the cooling medium discharge cavity 211 is provided on the connecting shell 21.

[0063] The discharge port is used to communicate with the propellant supply system. When the propellant supply system is working, the cryogenic propellant is discharged into the cooling medium discharge chamber 211 through the discharge port, and then flows into the cooling channel 4 through the inlet of the cooling channel 4.

[0064] Specifically, the cross-sectional area of ​​the discharge port gradually decreases from the outside of the cooling medium discharge chamber 211 to the inside of the cooling medium discharge chamber 211 to increase the flow rate of the cooling medium discharged from the discharge port into the cooling medium discharge chamber 211 .

[0065] In this embodiment, the cross section of the discharge port is tapered, and along the discharge direction of the cooling medium, the discharge port gradually contracts, so that the cooling medium enters the cooling channel 4 at a higher flow rate, thereby enhancing the initial heat exchange capacity.

[0066] The thrust chamber is also provided with a discharge portion, which is arranged at one end of the straight cylindrical section 62 of the combustion chamber 6 away from the convergent section 61. The discharge portion is provided with a cooling medium discharge cavity, and the cooling medium discharge cavity is connected with the outlet of the cooling channel 4. The discharge portion is provided with a discharge outlet connected with the cooling medium discharge cavity. From the inside of the cooling medium discharge cavity to the outside of the cooling medium discharge cavity, the cross-sectional area of ​​the discharge outlet gradually increases to slow down the flow rate of the cooling medium flowing out of the discharge outlet.

[0067] In this embodiment, the cross-section of the discharge port is conical, and the discharge port gradually expands along the discharge direction of the cooling medium, slowing down the flow rate of the cooling medium flowing out of the discharge port, reducing the pressure loss when the cooling medium flows out, and thereby reducing the pump work requirement, which is beneficial to improving the efficiency of the liquid rocket engine.

[0068] Specifically, the thrust chamber also includes a head and an injector. The head is arranged at one end of the straight cylindrical section 62 of the combustion chamber 6 away from the convergent section 61, and the injector is installed on the head; the discharge port of the discharge portion is connected with the injector, so that the heated propellant flowing out of the cooling channel 4 flows into the injector, and the injector injects the heated propellant into the straight cylindrical section 62 of the combustion chamber 6 at a certain speed and angle, which not only realizes effective cooling of the thrust chamber, but also utilizes the heat absorbed during the cooling process to preheat the propellant, thereby improving the combustion efficiency of the propellant.

[0069] In a specific embodiment, Figure 4 As shown, at the location of the throat 7, the rib 3 is provided with at least one through hole 31, and every two adjacent cooling channels 4 are connected through the through hole 31. When the cooling medium flows through the location of the throat 7 in the cooling channel 4, the cooling medium converges at the through hole 31.

[0070] The location of the throat 7 includes the throat 7 itself, the end of the convergent section 61 near the throat 7 and the starting part of the nozzle 5.

[0071] When the cooling medium flows through the throat 7 in the cooling channel 4, the cooling medium converges at the through hole 31, destroying the laminar flow state of the cooling medium, causing the cooling medium to be disturbed and mixed, forming a turbulent flow, and enhancing the heat exchange effect at the throat 7 where the heat load is the highest.

[0072] Specifically, if Figure 7 As shown, the through holes 31 on every two adjacent ribs 3 among the plurality of ribs 3 are connected along the circumferential direction of the thrust chamber to form an annular passage 32 in the sandwich structure.

[0073] In this embodiment, when the cooling medium flows through the throat 7 in the cooling channel 4, multiple streams of cooling medium intersect and impact each other in the annular passage 32, which will destroy the laminar flow state of the fluid to a greater extent, causing the fluid to be strongly disturbed and mixed, forming a strong turbulent flow, and further enhancing the heat exchange effect at the throat 7 where the heat load is the highest.

[0074] like Figure 4 、 Figure 7 As shown, at the location of the throat 7 , the rib 3 is provided with a plurality of through holes 31 , which are distributed along the axial direction of the thrust chamber to form a plurality of annular passages 32 distributed along the axial direction of the thrust chamber in the sandwich structure.

[0075] In this embodiment, when the cooling medium flows through the throat 7 in the cooling channel 4, multiple streams of cooling medium intersect and impact each other in multiple annular passages 32, forming a stronger turbulence, further enhancing the heat exchange effect at the throat 7 where the heat load is the highest.

[0076] Specifically, if Figure 4 As shown, the through hole 31 on the rib 3 has an opening, which faces the inner wall 1. When the cooling medium flows through the throat 7 in the cooling channel 4, the cooling medium flows through the position of the inner wall 1 corresponding to the opening of the through hole 31, so as to increase the heat exchange area between the cooling medium and the inner wall 1 at that position and improve the heat exchange effect.

[0077] In a specific embodiment, at the position where the throat 7 is located, the rib 3 is provided with a plurality of through holes 31, and the plurality of through holes 31 are distributed along the axial direction of the thrust chamber. The through holes 31 on every two adjacent ribs 3 among the plurality of ribs 3 are connected along the circumferential direction of the thrust chamber to form a plurality of annular passages 32 distributed along the axial direction of the nozzle 5 in the sandwich structure; the through holes 31 on the rib 3 have openings, and the openings face the inner wall 1. When the cooling medium flows through the position where the throat 7 is located in the cooling channel 4, the cooling medium converges at the through holes 31 and flows through the position of the inner wall 1 corresponding to the opening of the through holes 31.

[0078] In this embodiment, the heat exchange effect at the throat 7 where the heat load is the highest is enhanced, and the heat exchange area between the cooling medium and the inner wall 1 at that location is increased, thereby greatly improving the heat exchange effect. Compared with the uniform cross-section ribs in the prior art, the temperature at the throat 7 can be reduced by about 80K on the existing basis, or even lower.

[0079] In a specific embodiment, the rib width and rib height of the nozzle 5 at one end away from the throat 7 are respectively the first rib width and the first rib height, the rib width and rib height of the convergent section 61 at one end away from the throat 7 are respectively the second rib width and the second rib height, and the rib width and rib height of the throat 7 are respectively the third rib width and the third rib height; wherein, the first rib width is greater than the second rib width, the second rib width is greater than the third rib width, the first rib height is greater than the second rib height, and the second rib height is greater than the third rib height.

[0080] In this embodiment, the size of the ribs 3 corresponding to the nozzle 5 with low thermal load is larger than the size of the ribs 3 corresponding to the throat 7 with the highest thermal load and the convergent section 61 with higher thermal load. Compared with the uniform cross-section ribs in the prior art, this not only matches the axial thermal load difference of the thrust chamber, but also makes the material of the ribs 3 more reasonable.

[0081] Specifically, the first rib width is 1.4~3.5mm, the second rib width is 1.3~2.5mm, and the third rib width is 0.5~1.5mm; the first rib height is 1.3~3.5mm, the second rib height is 1.2~2.6mm, and the third rib height is 1~1.9mm.

[0082] For example, in a specific embodiment, the first rib width is 1.4 mm, the second rib width is 1.3 mm, and the third rib width is 0.5 mm; the first rib height is 1.3 mm, the second rib height is 1.2 mm, and the third rib height is 1 mm; that is, from the end of the nozzle 5 away from the throat 7 to the throat 7, the rib width of the rib 3 gradually decreases from 1.4 mm to 0.5 mm, and the rib height of the rib 3 gradually decreases from 1.3 mm to 1 mm; from the throat 7 to the end of the convergent section 61 away from the throat 7, the rib width of the rib 3 gradually increases from 0.5 mm to 1.3 mm, and the rib height of the rib 3 gradually increases from 1 mm to 1.2 mm.

[0083] In a specific embodiment, the width of the side where the rib 3 is connected to the inner wall 1 is K1, and the width of the side where the rib 3 is connected to the outer wall 2 is K2, K2>K1, so that the width of the cooling channel 4 between adjacent ribs 3 corresponding to the inner wall 1 is greater than the width of the side corresponding to the outer wall 2. When the cooling medium flows in the cooling channel 4, the contact area between the cooling medium and the inner wall 1 increases, and the heat exchange area between the cooling medium and the inner wall 1 increases, thereby improving the heat exchange effect on the inner wall 1.

[0084] Specifically, the cross section of the rib 3 is trapezoidal.

[0085] In a specific embodiment, along the axial direction of the thrust chamber, the rib 3 extends from one end of the nozzle 5 to one end of the convergent section 61 .

[0086] This arrangement helps enhance the structural strength of the thrust chamber, effectively transmitting and dissipating the forces under axial pressure and bending moment, and reducing deformation caused by pressure. Furthermore, the cooling medium can flow along the cooling channels 4 formed between adjacent ribs 3 and extending axially along the thrust chamber, helping to guide the flow of the cooling medium within the cooling channels 4 and improving cooling efficiency.

[0087] In another specific embodiment, the ribs 3 extend in a spiral shape from one end of the nozzle 5 to one end of the convergent section 61 , which can increase the degree of disturbance to the cooling medium, thereby improving the cooling effect.

[0088] In order to enhance heat exchange, spoiler columns are provided in the cooling channel 4 .

[0089] With such a configuration, the spoiler column can destroy the boundary layer of the cooling medium in the cooling channel 4, causing turbulence in the cooling medium, increasing the heat transfer coefficient between the cooling medium and the inner wall 1 of the cooling channel 4, thereby effectively improving the heat transfer effect. In this way, the heat of the inner wall 1 of the cooling channel 4 can be transferred to the cooling medium more quickly, thereby reducing the temperature of the inner wall 1.

[0090] Specifically, the manufacturing process of the thrust chamber is introduced as follows:

[0091] Material selection: The inner wall 1, outer wall 2 and ribs 3 of the thrust chamber are all made of alloy materials with high temperature resistance, high strength and good thermal conductivity, such as nickel-based high-temperature alloys, to meet the extreme environmental requirements when the liquid rocket engine is working.

[0092] Manufacturing Process: Advanced additive manufacturing technologies, such as selective laser melting, are employed to manufacture the thrust chamber's inner wall 1, outer wall 2, and ribs 3. This selective laser melting process precisely controls the complex shape and dimensional accuracy of the ribs 3, ensuring the structural quality. The specific process involves first applying alloy powder layer by layer onto a worktable based on the designed 3D model. A high-energy laser beam is then scanned along a predetermined trajectory, melting the powder layer by layer and solidifying it into a solid form. During the manufacturing process, process parameters such as laser power, scanning speed, and powder particle size are strictly controlled to ensure the desired structural dimensions.

[0093] Assembly and testing: Assemble the various components of the manufactured thrust chamber to ensure that the inlet and outlet connections of the cooling channel 4 and the supply system pipeline are sealed and leak-free. After assembly is complete, the thrust chamber can be subjected to a hot run test. During the hot run, the actual operating conditions of the rocket engine are simulated, and the cooling effect and performance of the structure are evaluated by measuring parameters such as the temperature distribution of the thrust chamber wall surface, the inlet and outlet temperature of the cooling medium, and the pressure. Based on the test results, the structural design and manufacturing process are optimized and improved to ensure that the cooling effect of the thrust chamber provided in the embodiment of the present application meets the actual use requirements of the entire engine system.

[0094] A second aspect of the embodiments of the present application provides a liquid rocket engine, comprising the liquid rocket thrust chamber described in any of the above embodiments.

[0095] In this embodiment, the cooling medium flows in the cooling channel 4 with a constantly changing flow cross-section of the thrust chamber, which improves the disturbing effect on the cooling medium, enhances the degree of turbulence, improves the heat transfer coefficient between the cooling medium and the wall, improves the heat transfer effect, enhances the cooling effect on the thrust chamber, and ensures the normal operation of the engine; from the end of the nozzle 5 away from the throat 7 to the throat 7, and from the end of the convergent section 61 away from the throat 7 to the throat 7, the rib width of the rib 3 gradually decreases, compared with the uniform cross-section ribs in the prior art, the width of the cooling channel 4 is increased, so that the contact area between the cooling medium and the inner wall 1 is increased, thereby increasing the heat exchange area between the cooling medium and the inner wall 1 and improving the cooling effect; moreover, the pressure drop difference of the cooling medium at different positions in the cooling channel 4 is reduced, so that the flow resistance loss of the cooling medium is easy to control, the flow requirements at different positions are met, the pump power requirement is reduced, and the overall efficiency of the liquid rocket engine is improved.

[0096] The above description is only a specific implementation method of the present application, but the protection scope of the present application is not limited thereto. Any changes or replacements within the technical scope disclosed in the present application should be included in the protection scope of the present application.

Claims

1. A liquid rocket thrust chamber, characterized in that: The thrust chamber has an inner wall and an outer wall, a sandwich structure is formed between the inner wall and the outer wall, a plurality of ribs are evenly arranged in the sandwich structure along the circumference of the thrust chamber, and cooling channels are formed between adjacent ribs, and the cooling channels are used for the flow of cooling medium; The thrust chamber includes a combustion chamber, a throat, and a nozzle. The combustion chamber has a convergent section. The convergent section, the throat, and the nozzle are sequentially distributed along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channel, the cooling medium flows through the nozzle, the throat, and the convergent section in sequence. Wherein, the rib width and rib height of the rib gradually decrease from the end of the nozzle away from the throat to the throat; From the throat to the end of the convergent section away from the throat, the rib width and rib height of the rib gradually increase; At the location of the throat, the rib is provided with at least one through-hole, and every two adjacent cooling channels are connected through the through-hole. When the cooling medium flows through the location of the throat in the cooling channel, the cooling medium converges at the through-hole. The through hole on the rib has an opening, and the opening faces the inner wall. When the cooling medium flows in the cooling channel, the cooling medium flows through the inner wall position corresponding to the opening of the through hole; The width of the rib at the side connected to the inner wall is K1, and the width of the rib at the side connected to the outer wall is K2, where K2>K1.

2. The liquid rocket thrust chamber according to claim 1, characterized in that: The through holes on every two adjacent ribs in the plurality of ribs are connected along the circumference of the thrust chamber to form an annular passage in the sandwich structure.

3. The liquid rocket thrust chamber according to claim 2, characterized in that: The rib is provided with a plurality of the through holes, and the plurality of the through holes are distributed along the axial direction of the thrust chamber to form a plurality of the annular passages distributed along the axial direction of the thrust chamber in the sandwich structure.

4. The liquid rocket thrust chamber according to claim 1, characterized in that: The rib width and rib height of the nozzle away from the throat are respectively the first rib width and the first rib height, the rib width and rib height of the convergent section away from the throat are respectively the second rib width and the second rib height, and the rib width and rib height of the throat are respectively the third rib width and the third rib height; The first rib width is greater than the second rib width, the second rib width is greater than the third rib width, the first rib height is greater than the second rib height, and the second rib height is greater than the third rib height.

5. The liquid rocket thrust chamber according to claim 1, characterized in that: At one end of the nozzle away from the throat, a connecting shell is provided on the outer wall of the nozzle, a cooling medium discharge cavity is provided between the connecting shell and the outer wall, and the cooling medium discharge cavity is communicated with the inlet of the cooling channel; The connecting shell is provided with a discharge port connected to the cooling medium discharge cavity. The cross-sectional area of ​​the discharge port gradually decreases from the outside of the cooling medium discharge cavity to the inside of the cooling medium discharge cavity to increase the flow rate of the cooling medium discharged from the discharge port.

6. The liquid rocket thrust chamber according to claim 1, characterized in that: The combustion chamber further comprises a straight section, the straight section being located at an end of the convergent section away from the throat; The thrust chamber is also provided with a discharge portion, which is arranged at one end of the straight cylinder section away from the convergent section. The discharge portion is provided with a cooling medium discharge cavity, and the cooling medium discharge cavity is connected with the outlet of the cooling channel. The discharge portion is provided with a discharge outlet connected with the cooling medium discharge cavity. The cross-sectional area of ​​the discharge outlet gradually increases from the inside of the cooling medium discharge cavity to the outside of the cooling medium discharge cavity to slow down the flow rate of the cooling medium flowing out of the discharge outlet.

7. A liquid rocket engine, characterized in that: A liquid rocket thrust chamber comprising the liquid rocket thrust chamber described in any one of claims 1-6.

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