Fuel gas cavity for thrust chamber, engine thrust chamber, engine and liquid carrier rocket

Abstract

The embodiment of the invention provides a fuel gas cavity for a thrust chamber, an engine thrust chamber, an engine and a liquid carrier rocket. The fuel gas cavity comprises an annular cavity body (10), a heat exchanger (20) and a fuel gas ejection device (30). The annular cavity (10) is wound on the outer wall of the lower end of the expansion section (100), and a closed annular cavity channel is formed by the annular cavity (10) and the outer wall of the lower end of the expansion section (100); a gas inlet of the annular cavity (10) is connected to a turbine exhaust pipe, and turbine exhaust is used for receiving gas exhausted by turbines of other systems; the heat exchanger (20) is arranged in the annular cavity (10) along the annular cavity channel; and the fuel gas spraying device (30) is arranged on the fuel gas cavity close to the lower end of the expansion section (100). The heat exchanger is arranged along the annular cavity channel in the annular cavity, so that the height size of the general assembly layout of the engine is reduced, and the general assembly structure is compact.

Description

Technical Field

[0001] This invention relates to the field of rockets, specifically to a gas chamber for a thrust chamber, an engine thrust chamber, an engine, and a liquid-fueled launch vehicle. Background Technology

[0002] The thrust chamber of a liquid rocket engine includes an injector, a combustion chamber, and a nozzle. The nozzle comprises a contraction section, a throat, and an expansion section. The nozzle area ratio of the engine thrust chamber is typically between 40 and 200. The expansion section, with an area ratio of approximately 10, extends to the exit and its internal surface is designed in a unified manner with the preceding regenerative cooling section (i.e., the contraction section, throat, and expansion section excluding the extension section). Structurally, the expansion section with an area ratio of approximately 10 to the exit is considered as a separate component, referred to as the nozzle extension section. The inlet gas temperature of the nozzle extension section is above 2000K; therefore, structural cooling of the extension section inlet is a key focus of the product design.

[0003] In existing technologies, common cooling solutions for nozzle extension sections include regenerative cooling using a full-sandwich structure, cooling with special composite materials, or cooling with exhaust gas from turbines in other systems. Among these, the full-sandwich cooling structure is a traditional approach. Small nozzles typically utilize mature welding or 3D printing processes, but this results in relatively high nozzle weight and system supply requirements, leading to suboptimal engine performance parameters. Large nozzles made of composite materials such as silicon carbide and ceramics are not yet widely used. Cooling with exhaust gas from turbines in other systems is a common method for nozzle extension sections, offering two main advantages: first, it creates a film cooling protection for the nozzle extension, improving its reliability and reducing material requirements, thus lowering costs and weight; second, the fuel-rich nature of turbine exhaust gas increases the momentum of the nozzle exit gas, generating additional thrust and improving the engine's overall specific impulse.

[0004] In existing technologies, the structure for introducing turbine exhaust from other systems into the nozzle extension section mainly includes a cold regeneration cooling thrust chamber nozzle expansion section, an inlet flange, an inlet bend, an exhaust manifold, an exhaust nozzle, a cone, a nozzle flange, and a nozzle extension section. The turbine exhaust from other systems is first connected to a heat exchanger, then to the exhaust manifold, and then the turbine exhaust in the exhaust manifold is introduced into the nozzle extension section to cool the shell of the nozzle extension section.

[0005] In the process of developing this invention, the applicant discovered at least the following problems in the prior art:

[0006] Other systems require the turbine exhaust to be connected to a heat exchanger first and then to an exhaust manifold, which results in the use of more structural components, a larger footprint, increased engine height, and greater difficulty in engine assembly layout. Summary of the Invention

[0007] This invention provides a gas chamber for a thrust chamber, an engine thrust chamber, an engine, and a liquid launch vehicle, which can solve the technical problem in the prior art: "The turbine exhaust of other systems needs to be connected to a heat exchanger first and then to an exhaust collector, which makes the cooling method use more structural components, occupies a large area, increases the height of the engine, and increases the difficulty of the engine assembly layout."

[0008] To achieve the above objectives, in a first aspect, embodiments of the present invention provide a gas chamber for a thrust chamber, wherein the nozzle includes an expansion section and an extension section located downstream of the expansion section, the lower end of the expansion section is connected to the upper end of the extension section, and the gas chamber is wound around the lower end of the expansion section adjacent to the extension section.

[0009] The gas chamber includes an annular cavity, a heat exchanger, and a gas ejection device;

[0010] The annular cavity is wound around the outer wall of the lower end of the expansion section, forming a closed annular cavity with the outer wall of the lower end of the expansion section.

[0011] The annular cavity has a gas inlet, which is connected to a turbine exhaust pipe, and the turbine exhaust pipe is used to receive gas discharged from the turbines of other systems.

[0012] The heat exchanger is disposed along the annular cavity channel within the annular cavity.

[0013] The gas ejection device is located on the gas chamber near the lower end of the expansion section. The gas ejected from the gas ejection device flows in the same direction as the axis of the extension section and toward the inside of the extension section.

[0014] Secondly, embodiments of the present invention provide an engine thrust chamber, including a combustion chamber, a nozzle, and the aforementioned gas chamber for the thrust chamber. The nozzle includes an expansion section and an extension section located downstream of the expansion section. The lower end of the expansion section is connected to the upper end of the extension section. The gas chamber is wound around the lower end of the expansion section. The outer diameter of the cross-section of the gas chamber at the gas injection device is smaller than the inner diameter of the upper end of the extension section.

[0015] Thirdly, embodiments of the present invention provide an engine, including the aforementioned engine thrust chamber.

[0016] Fourthly, embodiments of the present invention provide a liquid-fueled launch vehicle, including the aforementioned engine.

[0017] The above technical solution has the following beneficial effects: In this embodiment of the invention, the exhaust gas from the turbine of other systems is introduced into the annular cavity through the turbine exhaust pipe. The temperature of the introduced gas is approximately 800-900K. Heat exchange is performed between the gas and the heat exchange medium in the heat exchanger, reducing the gas temperature to approximately 700K. The 700K gas is then injected into the extension section through the gas injection device, forming a gas film in the extension section. This gas film separates the exhaust gas from the inner wall of the extension section, thus providing cooling and protection for the extension section, ensuring its safe and reliable operation. Compared to connecting the gas to the heat exchanger first and then to the gas chamber, this embodiment of the invention arranges the heat exchanger along the annular channel within the annular cavity, resulting in a compact structure and significantly reduced space usage. This reduces the overall height of the engine assembly layout, making the overall assembly structure more compact and solving the problem of increasing the overall height of the engine assembly layout. Attached Figure Description

[0018] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0019] Figure 1 This is an isometric view of the gas chamber for the thrust chamber according to an embodiment of the present invention;

[0020] Figure 2 yes Figure 1 The corresponding top view;

[0021] Figure 3 yes Figure 1 The sectional view corresponding to the left view;

[0022] Figure 4 This is an isometric view (partial structure removed) of the gas chamber used in the thrust chamber according to an embodiment of the present invention.

[0023] The reference numerals in the attached figures are as follows:

[0024] 100. Expansion section; 10. Annular cavity; 20. Heat exchanger; 30. Gas injection device; 40. Inlet elbow; 50. Flow equalization plate; 60. First inlet; 70. Second inlet; 80. Intermediate partition plate;

[0025] 101. First gas collecting ring; 102. Second gas collecting ring; 301. Gas nozzle; 103. Heat exchange medium inlet; 104. Heat exchange medium outlet. Detailed Implementation

[0026] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts are within the scope of protection of the present invention.

[0027] like Figure 1 , Figure 2 and Figure 3 As shown, in conjunction with the present invention, a gas chamber for a thrust chamber includes an expansion section 100 and an extension section located downstream of the expansion section 100. The lower end of the expansion section 100 is connected to the upper end of the extension section, and the gas chamber is wound around the lower end of the expansion section 100 adjacent to the extension section.

[0028] The gas chamber includes an annular cavity 10, a heat exchanger 20, and a gas injection device 30. The annular cavity 10 is wound around the outer wall of the lower end of the expansion section 100, forming a closed annular channel with the outer wall of the lower end of the expansion section 100. The annular cavity 10 has a gas inlet connected to a turbine exhaust pipe, which is used to receive gas discharged from the turbines of other systems. The heat exchanger 20 is arranged within the annular cavity 10 along the annular channel. The gas injection device 30 is located on the gas chamber near the lower end of the expansion section 100, and the flow direction of the gas injected from the gas injection device 30 is consistent with the axial direction of the extension section and faces into the extension section. The other systems refer to the engine auxiliary circuit system consisting of a gas generator and a turbopump.

[0029] The thrust chamber is the component of a liquid rocket engine that generates thrust. It includes the injector, combustion chamber, and nozzle. In the combustion chamber, the propellant is converted into high-temperature, high-pressure combustion products, which are then discharged at high speed through the nozzle, generating thrust. The nozzle includes a contraction section, a throat, an expansion section, and an extension section.

[0030] In this embodiment of the invention, the exhaust gas from the turbine of another system is introduced into the annular cavity 10 through the turbine exhaust pipe. The temperature of the introduced gas is about 800-900K. The heat exchange medium in the heat exchanger 20 exchanges heat with the introduced gas, reducing the temperature of the gas to about 700K. The 700K gas is injected into the extension section through the gas injection device 30, forming a gas film in the extension section. The gas film separates the exhaust gas from the inner wall of the extension section from the exhaust gas entering the nozzle after being discharged from the engine, thus cooling and protecting the extension section and enabling it to operate safely and reliably.

[0031] Compared to connecting the heat exchanger first and then the combustion chamber, the embodiment of the present invention arranges the heat exchanger 20 along the annular channel inside the annular cavity 10, which is more compact and uses less space. This reduces the height of the engine's overall assembly layout and makes the overall assembly structure more compact, thus solving the problem of increasing the height of the engine's overall assembly layout.

[0032] Preferably, such as Figure 1 and Figure 2 As shown, the annular cavity 10 is an unclosed shell wound around the expansion section 100 along its extension direction; the unclosed portion of the shell on the annular cavity 10 is fastened to the outer wall of the lower end of the expansion section 100, forming a closed annular cavity with the outer wall of the lower end of the expansion section 100. The annular cavity 10 is a thin-walled shell wound around the expansion section 100 along its extension direction, which, compared to a fully closed shell wound around the expansion section 100, reduces the mass of the combustion chamber, thereby reducing the overall weight of the engine assembly. The annular cavity allows the exhaust gas from the turbine to form a circumferential flow.

[0033] The cross-section of the annular cavity 10 is a gradually decreasing cross-section from the gas inlet side to the opposite side of the gas inlet; after the gas discharged from the turbine of other systems enters the gas inlet of the annular cavity 10 through the turbine exhaust pipe, its flow rate on the gas inlet side will be greater than that on the opposite side of the gas inlet, ensuring that the flow velocity and static pressure at the two places are basically the same. Designing the cross-section of the annular cavity 10 to match the flow rate can reduce the mass of the annular cavity 10, thereby reducing the mass of the engine assembly.

[0034] Preferably, such as Figure 2 As shown, the gas chamber for the thrust chamber also includes an inlet elbow 40. The first end of the inlet elbow 40 is connected to the gas inlet, and the second end of the inlet elbow 40 is connected to the turbine exhaust pipe via a flange structure, thus smoothly connecting the turbine exhaust pipe to the annular cavity 10. The inlet elbow 40 can be integrally formed using 3D printing, which, compared to sheet metal forming, offers lower processing difficulty, higher precision, and a shorter processing cycle.

[0035] Preferably, such as Figure 4As shown, the gas chamber for the thrust chamber also includes a flow equalization plate 50 connected to the lowest end of the inlet elbow 40. The flow equalization plate 50 has multiple flow equalization holes spaced apart. If the flow equalization plate 50 is not provided, the gas introduced into the annular cavity 10 from the lowest end of the inlet elbow 40 will quickly and directly rush into the annular channel below the lowest end of the inlet elbow 40, greatly reducing its flow rate to other parts of the annular channel. Consequently, the flow rate of gas entering the gas ejection device 30 from other parts of the annular channel is greatly reduced, resulting in uneven gas flow from the gas ejection device 30 and uneven gas film formed by the gas entering the extension section, which has a poor cooling and protection effect on the extension section. In this application, a flow equalization plate 50 is provided at the lowest end of the inlet elbow 40. The flow equalization plate 50 has multiple flow equalization holes. These holes reduce the speed of the gas entering the inlet elbow 40, achieving axial throttling and guiding most of the inlet gas towards the two annular cavities for uniform distribution. This ensures that the gas flows evenly downwards, allowing it to reach the entire annular cavity. This guarantees that the gas flow rate from the gas ejection device 30 is essentially the same at every point, achieving uniform axial gas flow distribution along the extension section, i.e., even distribution on the inner wall of the extension section. Simulation calculations were used to verify the number and location of the flow equalization holes, achieving a relatively uniform gas flow distribution from the annular cavity into the gas nozzle 301. Therefore, in this embodiment of the invention, the optimal flow equalization hole parameters (number and location) are obtained.

[0036] Preferably, such as Figure 1 and Figure 2As shown, the inlet elbow 40 has a first inlet 60 and a second inlet 70 separated by a middle partition 80. The first inlet 60 and the second inlet 70 are located above the flow equalization plate 50 and are respectively connected to opposite sides of the annular cavity 10. The middle partition 80 is located above the flow equalization plate 50 and has a gap with the flow equalization plate 50. The annular cavity 10 includes at least a first gas collecting ring 101 and a second gas collecting ring 102. The first end face of the first gas collecting ring 101 and the first end face of the second gas collecting ring 102 are connected. The second end face of the first gas collecting ring 101 is connected to the first inlet 60, and the second end face of the second gas collecting ring 102 is connected to the second inlet 70. The gas entering the annular cavity 10 is diverted through the first inlet 60 and the second inlet 70, allowing the gas to flow along the annular cavity in two directions towards its distal end. Further diversion and equalization of the gas through multiple flow equalization holes spaced apart on the flow equalization plate 50 allows even more gas to flow along the annular cavity towards its distal end. In addition to the first gas collecting ring 101 and the second gas collecting ring 102, there can be multiple gas collecting rings, as long as the cross-section of the annular cavity 10 formed by connecting multiple gas collecting rings is a gradually decreasing cross-section from the gas inlet side to the opposite side of the gas inlet. Both the first gas collecting ring 101 and the second gas collecting ring 102 are 3D printed as a single unit, which is easier to process, has higher precision, and a shorter processing cycle compared to sheet metal forming.

[0037] Preferably, such as Figure 1 and Figure 3 As shown, the gas ejection device 30 includes a plurality of spaced gas nozzles 301. Each gas nozzle 301 may be a square orifice. The plurality of gas nozzles 301 form a ring-shaped structure. The gas ejected from the gas nozzles 301 flows in the same direction as the axial direction of the extension section and towards the inside of the extension section. The gas ejection device 30, i.e., the gas nozzles 301, is welded (e.g., by argon arc welding) to the outer wall of the expansion section and the annular cavity 10. The gas ejected from the plurality of annular gas nozzles 301 forms a ring. After being injected into the upper inlet of the extension section of the nozzle, it forms an annular gas film on the inner wall of the extension section, thus protecting the shell structure of the extension section.

[0038] Preferably, the axial cross-section of the gas nozzle 301 is a Laval nozzle profile. After the gas in the annular cavity enters the gas nozzle 301, it is first accelerated and depressurized, reaching the speed of sound at the throat of the gas nozzle 301. Then, it continues to be accelerated and depressurized in the expansion section of the gas nozzle 301. The gas is then injected from the outlet of the gas nozzle 301 into the upper inlet of the extension section of the nozzle, forming a low-temperature gas film to protect the shell structure of the extension section.

[0039] Preferably, such as Figure 1 As shown, the heat exchanger 20 includes a heat exchange medium inlet 103 and a heat exchange medium outlet 104 that penetrate the annular cavity 10.

[0040] The heat exchange medium inlet 103 is used to provide the heat exchanger 20 with a low temperature heat exchange medium, and the heat exchange medium outlet 104 is used to discharge the high temperature heat exchange medium after heat exchange with the gas, so as to ensure that the heat exchanger 20 has a low temperature heat exchange medium continuously, and realize the ability to continuously exchange heat and cool down the gas entering the annular cavity 10.

[0041] In conjunction with embodiments of the present invention, an engine thrust chamber is provided, including an injector, a combustion chamber, a nozzle, and a gas chamber for the thrust chamber as described above. The nozzle includes an expansion section 100 and an extension section located downstream of the expansion section 100. The lower end of the expansion section 100 is connected to the upper end of the extension section. The gas chamber is wound around the lower end of the expansion section 100. The outer diameter of the cross-section of the gas chamber at the gas injection device 30 is smaller than the inner diameter of the upper end of the extension section. The exhaust gas from other systems' turbines is introduced into the annular cavity 10 through the turbine exhaust pipe. The temperature of the introduced gas is about 800-900K. The heat exchange medium in the heat exchanger 20 exchanges heat with the introduced gas, reducing the gas temperature to about 700K. The 700K gas is then injected into the extension section through the gas injection device 30, forming a gas film in the extension section. This gas film separates the exhaust gas from the inner wall of the extension section, thus cooling and protecting the extension section and ensuring its safe and reliable operation.

[0042] In conjunction with embodiments of the present invention, an engine is provided, including the aforementioned engine thrust chamber.

[0043] In conjunction with embodiments of the present invention, a liquid-fueled launch vehicle is provided, including the aforementioned engine.

[0044] It should be understood that in the above detailed description, various features are combined together in a single embodiment to simplify this disclosure. This approach to disclosure should not be interpreted as reflecting an intention that embodiments of the claimed subject matter require more features than are explicitly stated in each claim. Rather, as reflected in the appended claims, the invention is presented with fewer features than all the features of a single disclosed embodiment. Therefore, the appended claims are hereby explicitly incorporated into the detailed description, wherein each claim stands alone as a preferred embodiment of the invention.

[0045] The disclosed embodiments have been described above to enable any person skilled in the art to implement or use the present invention. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein can be applied to other embodiments without departing from the spirit and scope of this disclosure. Therefore, this disclosure is not limited to the embodiments given herein, but is consistent with the broadest scope of the principles and novel features disclosed in this application.

[0046] The foregoing description includes examples of one or more embodiments. It is certainly impossible to describe all possible combinations of components or methods in order to describe the above embodiments, but those skilled in the art will recognize that further combinations and arrangements of the various embodiments are possible. Therefore, the embodiments described herein are intended to cover all such changes, modifications, and variations that fall within the scope of the appended claims. Furthermore, the term "comprising" as used in the specification or claims is interpreted in a manner similar to the term "including," as interpreted when used as a conjunction in the claims. Additionally, the use of any term "or" in the specification of the claims is intended to mean "non-exclusive or."

[0047] The specific implementation methods described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific implementation method of the present invention and is not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A gas chamber for a thrust chamber, the nozzle comprising an expansion section (100) and an extension section downstream of the expansion section (100), the lower end of the expansion section (100) being connected to the upper end of the extension section, the gas chamber being wound around the lower end of the expansion section (100) adjacent to the extension section; characterized in that, The gas chamber includes an annular cavity (10), a heat exchanger (20), and a gas ejection device (30); The annular cavity (10) is wound around the outer wall of the lower end of the expansion section (100), forming a closed annular cavity with the outer wall of the lower end of the expansion section (100). The annular cavity (10) has a gas inlet connected to a turbine exhaust pipe, which is used to receive gas discharged from the turbines of other systems. The heat exchanger (20) is disposed within the annular cavity (10) along the annular cavity channel; The gas ejection device (30) is located on the gas chamber near the lower end of the expansion section (100). The gas ejected from the gas ejection device (30) flows in the same direction as the axis of the extension section and toward the inside of the extension section.

2. The gas chamber for the thrust chamber according to claim 1, characterized in that, The annular cavity (10) is a shell that is not closed along its extension direction and is wrapped around the expansion section (100); Furthermore, the cross-section of the annular cavity (10) is a gradually decreasing cross-section from the gas inlet side to the opposite side of the gas inlet; The shell at the unsealed part of the annular cavity (10) is fastened to the outer wall of the lower end of the expansion section (100), forming a closed annular cavity with the outer wall of the lower end of the expansion section (100).

3. The gas chamber for the thrust chamber according to claim 1, characterized in that, It also includes an inlet elbow (40), the first end of which is connected to the gas inlet, and the second end of which is connected to the turbine exhaust pipe.

4. The gas chamber for the thrust chamber according to claim 3, characterized in that, It also includes a flow equalization plate (50) connected to the lowest end of the inlet elbow (40), the flow equalization plate (50) having a plurality of flow equalization holes spaced apart.

5. The gas chamber for the thrust chamber according to claim 4, characterized in that, The inlet elbow (40) has a first inlet (60) and a second inlet (70) separated by an intermediate partition (80), the intermediate partition (80) being located above the flow equalization plate (50) and having a gap between it and the flow equalization plate (50); The annular cavity (10) includes at least a first gas collecting ring (101) and a second gas collecting ring (102). The first end face of the first gas collecting ring (101) and the first end face of the second gas collecting ring (102) are connected. The second end face of the first gas collecting ring (101) is connected to the first inlet (60), and the second end face of the second gas collecting ring (102) is connected to the second inlet (70).

6. The gas chamber for the thrust chamber according to claim 1, characterized in that, The gas ejection device (30) includes a plurality of spaced gas nozzles (301), and the gas ejected by the gas nozzles (301) flows in the same direction as the axis of the extension section and toward the inside of the extension section.

7. The gas chamber for a thrust chamber according to claim 6, characterized in that, The axial section of the gas nozzle (301) is a Laval nozzle profile.

8. The gas chamber for the thrust chamber according to claim 1, characterized in that, The heat exchanger (20) includes a heat exchange medium inlet (103) and a heat exchange medium outlet (104) that penetrate the annular cavity (10).

9. An engine thrust chamber, characterized in that, The device includes an injector, a combustion chamber, a nozzle, and a gas chamber for a thrust chamber as described in any one of claims 1-8. The nozzle includes an expansion section (100) and an extension section located downstream of the expansion section (100). The lower end of the expansion section (100) is connected to the upper end of the extension section. The gas chamber is wound around the lower end of the expansion section (100). The outer diameter of the cross-section of the gas chamber at the gas ejection device (30) is smaller than the inner diameter of the upper end of the extension section.

10. An engine, characterized in that, Includes the engine thrust chamber as described in claim 9.

11. A liquid-fueled launch vehicle, characterized in that, Includes the engine as described in claim 10.

Images