Liquid rocket thrust chamber and liquid rocket engine

Abstract

The invention provides a liquid rocket thrust chamber and a liquid rocket engine, and relates to the technical field of liquid rocket engines, the thrust chamber is provided with an inner wall and an outer wall, an interlayer structure is arranged between the inner wall and the outer wall, a plurality of ribs are evenly arranged in the interlayer structure in the circumferential direction of the thrust chamber, and a cooling channel is formed between every two adjacent ribs; the cooling channel is used for allowing a cooling medium to flow; the thrust chamber comprises a combustion chamber, a throat part and a spray pipe, the combustion chamber is provided with a convergent section, the convergent section, the throat part and the spray pipe are sequentially distributed in the axial direction of the thrust chamber, and when a cooling medium flows in the cooling channel, the cooling medium sequentially flows through the spray pipe, the throat part and the convergent section; the rib width and the rib height of the rib are gradually reduced from one end, far away from the throat part, of the spray pipe to the throat part; the rib width and the rib height of the rib are gradually increased from the throat part to one end, far away from the throat part, of the convergence section. Through the arrangement, the turbulent flow effect on the cooling medium can be improved, the heat exchange area of the cooling medium and the inner wall is increased, and the flow resistance loss of the cooling medium is easy to control.

Description

Technical Field

[0001] This 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 a component in a liquid rocket engine that completes the energy conversion of the propellant and generates thrust. The thrust chamber of a liquid rocket engine includes an injector, a combustion chamber, and a nozzle. When the liquid rocket engine operates, the propellant is mixed and atomized through the injector at the head of the thrust chamber, then heated and vaporized by the flame in the combustion chamber and burns violently, turning into high-temperature and high-pressure gas in the combustion chamber. The gas is accelerated into a high-speed airflow through the nozzle and ejected backward, generating thrust acting on the engine and pushing the rocket forward.

[0003] During the working process, the wall of the thrust chamber bears the strong thermal load of the high-temperature gas. To reduce the thermal load of the thrust chamber and prevent the structural damage of the thrust chamber wall due to high temperature, the thrust chamber can be cooled by a regenerative cooling method.

[0004] The thrust chamber includes an inner wall, an outer wall, and an intermediate sandwich layer. A plurality of ribs are arranged in the intermediate sandwich layer. The cooling channels are formed between adjacent ribs, and the propellant flows in the cooling channels. In the prior art, equal-section ribs are usually adopted, and the cooling channels between adjacent equal-section ribs only narrow and expand in the width direction, resulting in a relatively single flow disturbance effect on the propellant, poor heat transfer effect, and it is difficult to achieve good cooling for areas with high thermal load. Summary of the Invention

[0005] This application provides a liquid rocket thrust chamber and a liquid rocket engine, which can improve the flow disturbance effect on the cooling medium, increase the heat transfer area between the cooling medium and the inner wall, the flow resistance loss of the cooling medium is easy to control, and the cooling effect on the thrust chamber and the overall efficiency of the engine are improved.

[0006] To achieve the above object, this application adopts the following technical solutions: In a first aspect, this application provides a liquid rocket thrust chamber. The thrust chamber has an inner wall and an outer wall, and there is a sandwich structure between the inner wall and the outer wall. A plurality of ribs are uniformly arranged along the circumferential direction of the thrust chamber in the sandwich structure. Cooling channels are formed between adjacent ribs for the cooling medium to flow through; the thrust chamber includes a combustion chamber, a throat, and a nozzle. The combustion chamber has a converging section, and the converging section, the throat, and the nozzle are arranged in sequence along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channels, the cooling medium flows through the nozzle, the throat, and the converging section in sequence; wherein, from the end of the nozzle far 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 converging section far from the throat, the rib width and rib height of the ribs gradually increase.

[0007] In some possible implementation manners, at the position where the throat is located, the rib is provided with at least one through hole, and every two adjacent cooling channels are communicated through the through hole. When the cooling medium flows through the position where the throat is located in the cooling channel, the cooling medium converges at the through hole.

[0008] In some possible implementation manners, the through holes on every two adjacent ribs among the multiple ribs are communicated along the circumferential direction of the thrust chamber, so as to form an annular passage in the sandwich structure.

[0009] In some possible implementation manners, the rib is provided with multiple through holes, and the multiple through holes are distributed along the axial direction of the thrust chamber, so as to form multiple annular passages distributed along the axial direction of the thrust chamber in the sandwich structure.

[0010] In some possible implementation manners, 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 position corresponding to the opening of the through hole on the inner wall.

[0011] In some possible implementation manners, the rib width and rib height at the end of the nozzle far from the throat are the first rib width and the first rib height respectively, the rib width and rib height at the end of the converging section far from the throat are the second rib width and the second rib height respectively, and the rib width and rib height at the throat are the third rib width and the third rib height respectively; 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.

[0012] In some possible implementation manners, the width of the side of the rib connected to the inner wall is K1, the width of the side of the rib connected to the outer wall is K2, and K2 > K1.

[0013] In some possible implementation manners, at the end of the nozzle far from the throat, a connecting shell is arranged on the outer wall of the nozzle, a cooling medium discharge cavity is arranged 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 communicated with the cooling medium discharge cavity. From the outside to the inside of the cooling medium discharge cavity, the cross-sectional area of the discharge port gradually decreases, so as to increase the flow rate of the cooling medium discharged from the discharge port.

[0014] In some possible implementation manners, the combustion chamber further has a straight cylinder section, and the straight cylinder section is located at the end of the converging section far from the throat; The thrust chamber is further provided with a discharge part, the discharge part is arranged at the end of the straight cylinder section far from the converging section, the discharge part is provided with a cooling medium discharge cavity, the cooling medium discharge cavity is communicated with the outlet of the cooling channel, the discharge part is provided with a discharge port communicated with the cooling medium discharge cavity, and from the inside to the outside of the cooling medium discharge cavity, the cross-sectional area of the discharge port gradually increases, so as to slow down the flow rate of the cooling medium flowing out from the discharge port.

[0015] In a second aspect, the present application provides a liquid rocket engine, including the liquid rocket thrust chamber in the first aspect.

[0016] As can be seen from the above technical solutions, the present application has at least the following beneficial effects: For the liquid rocket thrust chamber provided in the present application, from the end of the nozzle far away from the throat to the throat, and from the throat to the end of the converging section of the combustion chamber far away from the throat, the rib width and rib height of the ribs first gradually decrease and then gradually increase, so that the width and height of the cooling channels between adjacent ribs first gradually decrease and then gradually increase, enabling the cooling medium to flow in the cooling channels with a continuously changing flow cross-section, improving the turbulence effect on the cooling medium, enhancing the degree of turbulence, increasing the heat transfer coefficient between the cooling medium and the wall surface, improving the heat transfer effect, and enhancing the cooling effect on the thrust chamber; moreover, compared with the equal-section ribs in the prior art, the width of the cooling channels is increased, so that the contact area between the cooling medium and the inner wall is increased, thereby increasing the heat transfer area between the cooling medium and the inner wall and improving the cooling effect.

[0017] Furthermore, the cooling channels corresponding to the nozzle are enlarged, resulting in a smaller flow resistance loss of the cooling medium flowing in from the cooling channel inlet; the rib height corresponding to the throat is smaller, and the corresponding cooling channels contract, with a higher flow velocity of the cooling medium, which quickly takes away heat and achieves the maximum cooling of the throat with the highest heat load; although the flow resistance loss of the cooling medium in the cooling channels corresponding to the throat is relatively large, the flow resistance loss of the cooling medium in the cooling channels corresponding to the nozzle is relatively small, making it easy to control the flow resistance loss of the cooling medium flowing from the nozzle to the throat in the cooling channels corresponding to the throat within the required range, thereby ensuring that the pressure and flow rate of the cooling medium flowing into the cooling channels corresponding to the converging section meet the requirements, ensuring the cooling effect on the converging section with a relatively high heat load, achieving targeted cooling according to the axial heat load difference of the thrust chamber, preventing local overheating, and reducing the pump work requirement, improving the overall efficiency of the liquid rocket engine.

[0018] It should be understood that the description of technical features, technical solutions, beneficial effects or similar languages in the present application does not imply that all features and advantages can be achieved in any single embodiment. On the contrary, it can be understood that the description of features or beneficial effects means that specific technical features, technical solutions or beneficial effects are included in at least one embodiment. Therefore, the description of technical features, technical solutions or beneficial effects in this specification does not necessarily refer to the same embodiment. Furthermore, the technical features, technical solutions and beneficial effects described in this embodiment can be combined in any appropriate manner. Those skilled in the art will understand that an 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

[0019] Figure 1 Schematic diagram of the liquid rocket thrust chamber provided by the present application in a specific embodiment; Figure 2 is Figure 1 a partial cross-sectional view; Figure 3 is Figure 1 a cross-sectional view from a certain perspective; Figure 4 is Figure 3 an enlarged view of part II in; Figure 5 is Figure 1 a cross-sectional view from another perspective; Figure 6 is Figure 5 an enlarged view of part III in; Figure 7 is Figure 2 an enlarged view of part I in.

[0020] Reference numerals: 1 - inner wall; 2 - outer wall; 21 - connecting shell; 211 - cooling medium discharge chamber; 3 - rib; 31 - through hole; 32 - annular passage; 4 - cooling channel; 5 - nozzle; 6 - combustion chamber; 61 - converging section; 62 - straight section; 7 - throat. Detailed Description of the Embodiments

[0021] The terms "first", "second", "third", etc. in the description of the present application and the description of the drawings are used to distinguish different objects, rather than to limit a specific order.

[0022] In the embodiments of the present application, words such as "exemplary" or "for example" are used to represent examples, illustrations or explanations. Any embodiment or design described as "exemplary" or "for example" in the embodiments of the present application should not be construed as being more preferred or having more advantages than other embodiments or designs. Rather, the use of words such as "exemplary" or "for example" is intended to present related concepts in a specific manner.

[0023] For the sake of clear and concise description of the following embodiments, a brief introduction to the related art is given first: A liquid rocket is a rocket that uses liquid propellants. It consists of multiple parts such as a rocket body structure and an engine. A liquid rocket can mix and burn liquid fuel and oxidizer in the engine, generate high-temperature and high-pressure gas, eject it at high speed backward, generate forward thrust, and push the rocket into the air.

[0024] A liquid rocket engine is the core component of a liquid rocket, mainly consisting of a thrust chamber, a propellant supply system, an engine control system, etc.

[0025] The thrust chamber is a key component of a liquid rocket engine. The propellant supply system delivers liquid propellants to the thrust chamber, where they are mixed and burned in the combustion chamber of the thrust chamber to generate high-temperature and high-pressure gas, which is ejected at high speed from the nozzle of the thrust chamber, thereby generating thrust.

[0026] The thrust chamber has an inner wall, an outer wall, and an intermediate sandwich layer. A plurality of ribs are provided in the intermediate sandwich layer, and each rib extends from one end of the combustion chamber to one end of the nozzle. A cooling channel is formed between adjacent ribs.

[0027] During operation, the wall of the thrust chamber bears the intense heat load of the high-temperature gas. To reduce the heat load of the thrust chamber and prevent structural damage to the wall of the thrust chamber due to high temperature, the thrust chamber can be cooled by regenerative cooling.

[0028] The regenerative cooling method is to convey low-temperature propellants from one end of the nozzle of the thrust chamber to the cooling channel. The low-temperature propellants absorb and carry away the heat of the wall, reducing the heat load of the thrust chamber; the heated propellants flow out of the cooling channel and then flow into the combustion chamber to participate in combustion. In this way, not only is the effective cooling of the thrust chamber achieved, but also the heat absorbed during the cooling process is used to preheat the propellants, improving the combustion efficiency of the propellants.

[0029] The thrust chamber has a combustion chamber, a throat, and a nozzle. The combustion chamber has a straight section and a convergent section. Along the axis of the thrust chamber, the straight section, the convergent section, the throat, and the nozzle are arranged in sequence; the nozzle is the divergent section of the thrust chamber, and the convergent section and the nozzle gradually contract from both sides of the throat to the throat. The propellants burn in the straight section of the combustion chamber to generate high-temperature gas; the high-temperature gas will flow through the convergent section, the throat, and the nozzle in sequence. The convergent section increases the gas flow velocity gradually, and the throat is the part with the smallest cross-sectional area in the thrust chamber, where the gas accelerates sharply, and the nozzle further expands and accelerates the gas to obtain a higher ejection velocity.

[0030] During the process of the gas flowing through the thrust chamber, the inner wall of the thrust chamber directly bears the erosion of the high-temperature gas. When the propellants flow through the cooling channel, the propellants mainly absorb the heat of the inner wall.

[0031] When the propellants burn in the straight section of the combustion chamber, the generated high-temperature gas comes into contact with the corresponding inner wall at the position of the straight section of the combustion chamber and conducts intense heat exchange, causing the heat load of the straight section of the combustion chamber to increase. The high-temperature gas flows out of the straight section and through the convergent section, where the gas flow velocity gradually increases. When it reaches the throat, the gas accelerates sharply at the throat. Therefore, from the convergent section to the throat, the heat load gradually increases, and the heat load at the throat position is the highest. From the throat to the nozzle, the gas expands and accelerates, and the gas temperature and pressure gradually decrease, and the heat exchange intensity with the corresponding inner wall also gradually weakens. From the throat to the nozzle, the heat load gradually decreases.

[0032] In the prior art, equal cross-section ribs are usually adopted. In the direction from the nozzle to the throat, the distance between adjacent equal cross-section ribs gradually decreases, and in the direction from the throat to the convergent section, the distance between adjacent equal cross-section ribs gradually increases. Therefore, in the thrust chamber, the cooling channels between adjacent equal cross-section ribs only narrow and expand in the width direction, resulting in a relatively single flow disturbance effect on the propellant. As a result, in the throat and convergent section areas with high heat loads, the heat transfer effect of the propellant is poor, and the cooling effect in the corresponding areas is poor. Moreover, the sizes of the equal cross-section ribs are always the same in the areas with high heat loads and in the areas with low heat loads, which will cause a certain waste of resources.

[0033] Moreover, for the cooling channels formed between adjacent equal cross-section ribs, the pressure drop difference of the cooling medium at different positions in the cooling channels is relatively large. The pressure drop is relatively high when the cooling medium flows from the nozzle to the throat, resulting in the pressure and flow rate 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. This will not only affect the overall cooling efficiency but also increase the unnecessary pump work demand.

[0034] In view of this, the embodiment of the present application provides a liquid rocket thrust chamber, which can enable the cooling medium to flow in the cooling channels with continuously changing flow cross-sections, improving the flow disturbance effect on the cooling medium, enhancing the degree of turbulence, increasing the heat transfer coefficient between the cooling medium and the wall surface, improving the heat transfer effect, and enhancing the cooling effect on the thrust chamber. Compared with the equal cross-section ribs in the prior art, the width of the cooling channels is increased, so that the contact area between the cooling medium and the inner wall is increased, and then the heat transfer area between the cooling medium and the inner wall is increased, improving the cooling effect.

[0035] Compared with the equal cross-section ribs in the prior art, the cooling channels corresponding to the nozzle are enlarged, so that the flow resistance loss of the cooling medium flowing in from the cooling channel inlet is small, the flow velocity of the cooling medium at the corresponding position of the throat is high, and the heat is quickly carried away, achieving the maximum cooling of the throat with the highest heat load. Although the flow resistance loss of the cooling medium in the cooling channels corresponding to the throat is large, the flow resistance loss of the cooling medium in the cooling channels corresponding to the nozzle is small, making it easy to control the flow resistance loss of the cooling medium flowing from the nozzle to the throat in the cooling channels corresponding to the throat within the required range. Furthermore, the pressure and flow rate of the cooling medium flowing into the cooling channels corresponding to the convergent section meet the requirements, ensuring the cooling effect on the convergent section with a relatively high heat load, achieving targeted cooling according to the axial heat load difference of the thrust chamber, preventing local overheating, and reducing the pump work demand, improving the overall efficiency of the liquid rocket engine.

[0036] The following will describe the liquid rocket thrust chamber provided by the embodiment of the present application with reference to the accompanying drawings: As Figures 1 - 6As shown, this kind of thrust chamber has an inner wall 1 and an outer wall 2. There is a sandwich structure between the inner wall 1 and the outer wall 2. A plurality of ribs 3 are evenly arranged along the circumferential direction of the thrust chamber within the sandwich structure. Cooling channels 4 are formed between adjacent ribs 3, and the cooling channels 4 are used for the flow of a cooling medium.

[0037] 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 arranged in sequence along the axial direction of the thrust chamber. When the cooling medium flows in the cooling channels 4, the cooling medium flows through the nozzle 5, the throat 7, and the convergent section 61 in sequence.

[0038] Among them, from the end of the nozzle 5 far from the throat 7 to the throat 7, the rib width and rib height of the ribs 3 gradually decrease; from the throat 7 to the end of the convergent section 61 far from the throat 7, the rib width and rib height of the ribs 3 gradually increase.

[0039] 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.

[0040] 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 gas. The high-temperature gas is ejected outward through the convergent section 61, the throat 7, and the nozzle 5 to generate a propulsive force.

[0041] As Figures 2 - 3 shown, from the convergent section 61 to the throat 7, the thrust chamber gradually contracts, and from the throat 7 to the nozzle 5, the thrust chamber gradually expands.

[0042] The inlet of the cooling channel 4 is located at the end of the nozzle 5 far from the throat 7, and the outlet of the cooling channel 4 is located at the end of the straight section 62 of the combustion chamber 6 far from the throat. The cooling medium flows in from the inlet of the cooling channel 4 and flows through the nozzle 5, the throat 7, the convergent section 61, and the straight section 62 in sequence. The cooling medium absorbs and takes away the heat of the nozzle 5, the throat 7, the convergent section, and the straight section 62 to cool the thrust chamber.

[0043] The cooling medium can specifically be the propellant of a liquid rocket.

[0044] From the end of the nozzle 5 far away from the throat 7 to the throat 7, the rib width and rib height of the rib 3 gradually decrease, so that the width and height of the cooling channel 4 between adjacent ribs 3 gradually decrease. During the process that the cooling medium flows from the end of the nozzle 5 far away from the throat 7 to the throat 7, the flow velocity of the cooling medium gradually increases. At the position where the throat 7 is located, the flow velocity 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 far away from the throat 7, the rib width and rib height of the rib 3 gradually increase, so that the width and height of the cooling channel 4 between adjacent ribs 3 gradually increase. During the process that the cooling medium flows from the throat 7 to the end of the convergent section 61 far away from the throat 7, the flow velocity of the cooling medium gradually decreases from the maximum flow velocity.

[0045] Specifically, from the end of the nozzle 5 far away from the throat 7 to the throat 7, the rib width and rib height of the rib 3 can vary linearly; from the throat 7 to the end of the convergent section 61 far away from the throat 7, the rib width and rib height of the rib 3 can vary linearly.

[0046] In this embodiment, from the end of the nozzle 5 far away from the throat 7 to the throat 7, and from the throat 7 to the end of the convergent section 61 far away from the throat 7, the rib width and rib height of the rib 3 first gradually decrease and then gradually increase, so that the width and height of the cooling channel 4 between adjacent ribs 3 first gradually decrease and then gradually increase, enabling the cooling medium to flow in the cooling channel 4 with a continuously changing flow cross-section, improving the turbulence effect on the cooling medium, enhancing the degree of turbulence, increasing the heat transfer coefficient between the cooling medium and the wall surface, improving the heat transfer effect, and enhancing the cooling effect on the thrust chamber.

[0047] From the end of the nozzle 5 far away from the throat 7 to the throat 7, and from the end of the convergent section 61 far away from the throat 7 to the throat 7, the rib width of the rib 3 gradually decreases respectively. Compared with the equal-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, and further the heat transfer area between the cooling medium and the inner wall 1 is increased, improving the cooling effect.

[0048] Moreover, the rib height of the rib 3 corresponding to the nozzle 5 is greater than that of the rib 3 corresponding to the throat 7, and the cooling channel 4 corresponding to the nozzle 5 expands, so that the flow resistance loss of the cooling medium flowing into the cooling channel 4 from the inlet is small; the rib height corresponding to the throat 7 is small, and the corresponding cooling channel 4 contracts, and the flow velocity of the cooling medium is high, quickly taking away heat, achieving the maximum cooling of the throat 7 with the highest heat load; although the flow resistance loss of the cooling medium in the cooling channel 4 corresponding to the throat 7 is large, the flow resistance loss of the cooling medium in the cooling channel 4 corresponding to the nozzle 5 is small, 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 easily controlled within the required range, and further making the pressure and flow rate of the cooling medium flowing into the cooling channel 4 corresponding to the converging section 61 meet the requirements, ensuring the cooling effect on the converging section 61 with a relatively high heat load, achieving targeted cooling according to the axial heat load difference of the thrust chamber, preventing local overheating, reducing the pump power requirement, and improving the overall efficiency of the liquid rocket engine.

[0049] In a specific embodiment, as Figures 1 - 3 shown, at one end of the nozzle 5 far from the throat 7, a connecting housing 21 is provided on the outer wall 2 of the nozzle 5. A cooling medium discharge cavity 211 is provided between the connecting housing 21 and the outer wall 2, and the cooling medium discharge cavity 211 is communicated with the inlet of the cooling channel 4; a discharge port communicated with the cooling medium discharge cavity 211 is provided on the connecting housing 21.

[0050] The discharge port is used to communicate with the propellant supply system. When the propellant supply system works, cryogenic propellant is discharged into the cooling medium discharge cavity 211 through the discharge port, and then flows into the cooling channel 4 through the inlet of the cooling channel 4.

[0051] Specifically, in the direction from outside the cooling medium discharge cavity 211 to inside the cooling medium discharge cavity 211, the cross-sectional area of the discharge port gradually decreases to increase the flow velocity of the cooling medium discharged into the cooling medium discharge cavity 211 through the discharge port.

[0052] In this embodiment, the cross-section of the discharge port is conical, 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 relatively high flow velocity, enhancing the initial heat exchange capacity.

[0053] The thrust chamber is also provided with a discharge part, which is arranged at one end of the straight cylinder section 62 of the combustion chamber 6 far from the converging section 61. The discharge part is provided with a cooling medium discharge cavity, the cooling medium discharge cavity is communicated with the outlet of the cooling channel 4, and the discharge part is provided with a discharge port communicated with the cooling medium discharge cavity. In the direction from inside the cooling medium discharge cavity to outside the cooling medium discharge cavity, the cross-sectional area of the discharge port gradually increases to slow down the flow velocity of the cooling medium flowing out of the discharge port.

[0054] In this embodiment, the cross-section of the discharge port is conical. Along the discharge direction of the cooling medium, the discharge port gradually expands, 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 thus reducing the pump power requirement, which is beneficial to improving the efficiency of the liquid rocket engine.

[0055] Specifically, the thrust chamber further includes a head and an injector. The head is arranged at one end of the straight cylinder 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 part is communicated with the injector, so that the heated propellant flowing out of the cooling channel 4 flows into the injector. The injector injects the heated propellant into the straight cylinder section 62 of the combustion chamber 6 at a certain speed and angle, which not only realizes the effective cooling of the thrust chamber, but also utilizes the heat absorbed during the cooling process to preheat the propellant and improve the combustion efficiency of the propellant.

[0056] In a specific embodiment, as Figure 4 shown, at the position where the throat 7 is located, the rib 3 is provided with at least one through hole 31, and every two adjacent cooling channels 4 are communicated through the through hole 31. When the cooling medium flows through the position where the throat 7 is located in the cooling channel 4, the cooling media meet at the through hole 31.

[0057] The position where the throat 7 is located includes the throat 7 itself and the end of the convergent section 61 near the throat 7 and the starting part of the nozzle 5.

[0058] When the cooling medium flows through the position where the throat 7 is located in the cooling channel 4, the cooling media meet at the through hole 31, destroying the laminar flow state of the cooling medium, causing the cooling medium to produce disturbance and mixing, forming a turbulent flow, and strengthening the heat transfer effect at the position where the throat 7 with the highest heat load is located.

[0059] Specifically, as Figure 7 shown, the through holes 31 on every two adjacent ribs 3 among the multiple ribs 3 are communicated along the circumferential direction of the thrust chamber to form an annular passage 32 in the sandwich structure.

[0060] In this embodiment, when the cooling medium flows through the position where the throat 7 is located in the cooling channel 4, multiple cooling media meet and impact each other in the annular passage 32, which will more greatly destroy the laminar flow state of the fluid, cause the fluid to produce strong disturbance and mixing, form a strong turbulent flow, and further strengthen the heat transfer effect at the position where the throat 7 with the highest heat load is located.

[0061] As Figure 4 、 Figure 7 shown, at the position where the throat 7 is located, the rib 3 is provided with multiple through holes 31, and the multiple through holes 31 are distributed along the axial direction of the thrust chamber to form multiple annular passages 32 distributed along the axial direction of the thrust chamber in the sandwich structure.

[0062] In this embodiment, when the cooling medium flows through the position of the throat 7 in the cooling channel 4, multiple strands of the cooling medium intersect and impact with each other in the multiple annular passages 32, forming stronger turbulent flows, and further strengthening the heat exchange effect at the position of the throat 7 with the highest heat load.

[0063] Specifically, as Figure 4 shown, the through holes 31 on the rib 3 have openings facing the inner wall 1. When the cooling medium flows through the position of 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 this place and improve the heat exchange effect.

[0064] In a specific embodiment, at the position of the throat 7, the rib 3 is provided with multiple through holes 31. The multiple 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 multiple ribs 3 are connected in the circumferential direction of the thrust chamber, so as to form multiple 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 facing the inner wall 1. When the cooling medium flows through the position of the throat 7 in the cooling channel 4, the cooling medium intersects at the through holes 31 and flows through the position of the inner wall 1 corresponding to the opening of the through hole 31.

[0065] In this embodiment, not only the heat exchange effect at the position of the throat 7 with the highest heat load is strengthened, but also the heat exchange area between the cooling medium and the inner wall 1 at this place is increased, greatly improving the heat exchange effect. Compared with the equal-section ribs in the prior art, the temperature at the position of the throat 7 can be reduced by about 80K or even lower on the existing basis.

[0066] In a specific embodiment, the rib width and rib height at one end of the nozzle 5 far from the throat 7 are the first rib width and the first rib height respectively, the rib width and rib height at one end of the converging section 61 far from the throat 7 are the second rib width and the second rib height respectively, and the rib width and rib height of the throat 7 are the third rib width and the third rib height respectively; 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.

[0067] In this embodiment, the size of the rib 3 corresponding to the nozzle 5 with a low heat load is larger than the size of the rib 3 corresponding to the throat 7 with the highest heat load and the converging section 61 with a relatively high heat load. Compared with the equal-section ribs in the prior art, it not only matches the heat load difference in the axial direction of the thrust chamber, but also makes the material of the rib 3 more reasonable.

[0068] Specifically, the first rib width is 1.4 - 3.5 mm, the second rib width is 1.3 - 2.5 mm, and the third rib width is 0.5 - 1.5 mm; the first rib height is 1.3 - 3.5 mm, the second rib height is 1.2 - 2.6 mm, and the third rib height is 1 - 1.9 mm.

[0069] For example, in a specific embodiment, the width of the first rib is 1.4 mm, the width of the second rib is 1.3 mm, and the width of the third rib is 0.5 mm; the height of the first rib is 1.3 mm, the height of the second rib is 1.2 mm, and the height of the third rib is 1 mm. That is, from the end of the nozzle 5 far away from the throat 7 to the throat 7, the width of the rib 3 gradually decreases from 1.4 mm to 0.5 mm, and the 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 far away from the throat 7, the width of the rib 3 gradually increases from 0.5 mm to 1.3 mm, and the height of the rib 3 gradually increases from 1 mm to 1.2 mm.

[0070] In a specific embodiment, the width of the side of the rib 3 connected to the inner wall 1 is K1, and the width of the side of the rib 3 connected to the outer wall 2 is K2, where K2 > K1, such that the width of the cooling channel 4 corresponding to the inner wall 1 side between adjacent ribs 3 is greater than the width corresponding to the outer wall 2 side. When the cooling medium flows in the cooling channel 4, the contact area between the cooling medium and the inner wall 1 increases, the heat exchange area between the cooling medium and the inner wall 1 increases, and the heat exchange effect on the inner wall 1 is improved.

[0071] Specifically, the cross-section of the rib 3 is trapezoidal.

[0072] 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.

[0073] With such a setting, it is beneficial to enhance the structural strength of the thrust chamber. When bearing axial pressure and bending moment, it can effectively transfer and disperse the force, reducing the deformation caused by pressure. At the same time, the cooling medium can flow along the cooling channel 4 formed between adjacent ribs 3 and extending along the axial direction of the thrust chamber, which helps to guide the cooling medium to flow in the cooling channel 4 and improve the cooling efficiency.

[0074] In another specific embodiment, from one end of the nozzle 5 to one end of the convergent section 61, the rib 3 extends in a spiral shape, which can increase the degree of disturbance of the cooling medium, thereby improving the cooling effect.

[0075] In order to strengthen heat exchange, turbulators are provided in the cooling channel 4.

[0076] With such a setting, the turbulators can destroy the boundary layer of the cooling medium in the cooling channel 4, cause the cooling medium to generate turbulence, increase the heat transfer coefficient between the cooling medium and the inner wall 1 of the cooling channel 4, and thus effectively improve the heat exchange 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, reducing the temperature of the inner wall 1.

[0077] Specifically, the manufacturing process of the thrust chamber is introduced as follows: 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 superalloys, to meet the extreme environmental requirements during the operation of liquid rocket engines.

[0078] Manufacturing Process: Advanced additive manufacturing techniques are adopted. For example, the selective laser melting process can be used to manufacture the inner wall 1, outer wall 2, and ribs 3 of the thrust chamber. The selective laser melting process can precisely control the complex shape and dimensional accuracy of the ribs 3, ensuring the manufacturing quality of the structure. The specific process can be as follows: First, according to the designed 3D model, alloy powder is spread layer by layer on the workbench, and scanned by a high-energy laser beam along a predetermined trajectory, causing the powder to melt and solidify layer by layer to form a shape. During the manufacturing process, process parameters such as laser power, scanning speed, and powder particle size are strictly controlled to ensure the dimensional requirements of the formed structure.

[0079] Assembly and Testing: The components of the manufactured thrust chamber are assembled to ensure that the connections between the inlets and outlets of the cooling channels 4 and the supply system pipes are sealed without leakage. After assembly, a hot-fire test can be conducted on the thrust chamber. During the hot-fire test, the actual working conditions of the rocket engine are simulated, and parameters such as the temperature distribution on the wall of the thrust chamber, the inlet and outlet temperatures and pressures of the cooling medium are measured to evaluate the cooling effect and performance of the structure. According to 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 embodiments of the present application meets the actual usage requirements of the entire engine system.

[0080] In the second aspect of the embodiments of the present application, a liquid rocket engine is provided, including the liquid rocket thrust chamber described in any one of the above embodiments.

[0081] In this embodiment, the cooling medium flows in the cooling channels 4 with a continuously changing flow cross-section of the thrust chamber, which improves the turbulence effect on the cooling medium, enhances the degree of turbulence, increases the heat transfer coefficient between the cooling medium and the wall surface, 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 far away from the throat 7 to the throat 7, and from the end of the converging section 61 far away from the throat 7 to the throat 7, the rib widths of the ribs 3 gradually decrease respectively. Compared with the equal-section ribs in the prior art, the width of the cooling channels 4 is increased, so that the contact area between the cooling medium and the inner wall 1 is increased, and then the heat transfer area between the cooling medium and the inner wall 1 is increased, improving the cooling effect; moreover, the pressure drop difference of the cooling medium at different positions in the cooling channels 4 is reduced, making the flow resistance loss of the cooling medium easy to control, meeting the flow requirements at different positions, reducing the pump power requirement, and improving the overall efficiency of the liquid rocket engine.

[0082] As described above, it is only the specific implementation manner of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions within the technical scope disclosed in the present application shall be covered by 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, and there is a sandwich structure between the inner wall and the outer wall. A plurality of ribs are uniformly arranged in the circumferential direction of the thrust chamber within the sandwich structure, and cooling channels are formed between adjacent ribs for the flow of a cooling medium; the thrust chamber includes a combustion chamber, a throat and a nozzle. The combustion chamber has a converging section, and the converging 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 channels, the cooling medium sequentially flows through the nozzle, the throat and the converging section; 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 converging section away from the throat, the rib width and rib height of the ribs gradually increase.

2. The liquid rocket thrust chamber according to claim 1, wherein At the position where the throat is located, the rib is provided with at least one through hole, and every two adjacent cooling channels communicate with each other through the through hole. When the cooling medium flows through the position where the throat is located in the cooling channels, the cooling medium converges at the through hole.

3. The liquid rocket thrust chamber according to claim 2, characterized in that, The through holes on every two adjacent ribs among the plurality of ribs communicate with each other in the circumferential direction of the thrust chamber to form an annular passage in the sandwich structure.

4. The liquid rocket thrust chamber according to claim 3, characterized in that, The rib is provided with a plurality of the through holes, and the plurality of 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.

5. The liquid rocket thrust chamber according to claim 2, characterized in that, The through hole on the rib has an opening facing the inner wall. When the cooling medium flows in the cooling channel, the cooling medium flows through the position on the inner wall corresponding to the opening of the through hole.

6. The liquid rocket thrust chamber according to claim 1, characterized in that, The rib width and rib height at the end of the nozzle away from the throat are respectively a first rib width and a first rib height, the rib width and rib height at the end of the converging section away from the throat are respectively a second rib width and a second rib height, and the rib width and rib height of the throat are respectively a third rib width and a 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.

7. The liquid rocket thrust chamber according to claim 1, characterized in that, 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, and K2 > K1.

8. The liquid rocket thrust chamber according to claim 1, characterized in that, At the end of the nozzle away from the throat, a connection housing is provided on the outer wall of the nozzle. A cooling medium discharge cavity is provided between the connection housing and the outer wall, and the cooling medium discharge cavity communicates with the inlet of the cooling channel; The connection housing is provided with a discharge port communicating with the cooling medium discharge cavity. From the outside to the inside of the cooling medium discharge cavity, the cross-sectional area of the discharge port gradually decreases to increase the flow rate of the cooling medium discharged from the discharge port.

9. The liquid rocket thrust chamber according to claim 1, characterized in that The combustion chamber also has a straight cylinder section located at the end of the converging section away from the throat; The thrust chamber is further provided with a discharge part, the discharge part is arranged at one end of the straight cylinder section away from the converging section, the discharge part is provided with a cooling medium discharge cavity, the cooling medium discharge cavity is communicated with the outlet of the cooling channel, the discharge part is provided with a discharge port communicated with the cooling medium discharge cavity, and in the direction from inside the cooling medium discharge cavity to outside the cooling medium discharge cavity, the cross-sectional area of the discharge port gradually increases to slow down the flow rate of the cooling medium flowing out from the discharge port.

10. A liquid rocket engine, characterized in that, A liquid rocket thrust chamber according to any one of claims 1-9.

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