Gas cavity for thrust chamber, engine thrust chamber, engine and liquid carrier rocket

Abstract

The embodiment of the present application provides a gas cavity for a thrust chamber, an engine thrust chamber, an engine and a liquid carrier rocket, the gas cavity comprises an annular cavity (10), a heat exchanger (20) and a gas ejection device (30); the annular cavity (10) is arranged on the outer wall of the lower end of the expansion section (100) and forms a closed annular cavity channel with the outer wall of the lower end of the expansion section (100); the gas inlet of the annular cavity (10) is connected to a turbine exhaust pipe, the turbine exhaust pipe is used for receiving the turbine exhaust gas of other systems; the heat exchanger (20) is arranged along the annular cavity channel in the annular cavity (10); the gas ejection device (30) is arranged on the gas cavity near 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 dimension of the overall layout of the engine is reduced, and the overall structure is compact.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of rockets, in particular to a gas cavity for a thrust chamber, an engine thrust chamber, an engine and a liquid carrier rocket. BACKGROUND

[0002] The engine thrust chamber of a liquid rocket includes an injector, a combustion chamber and a nozzle, the nozzle includes a convergent section, a throat section and a divergent section, the nozzle area ratio of the engine thrust chamber is usually 40-200, the inner profile of the divergent section with an area ratio of about 10 to the outlet is designed uniformly with the front regenerative cooling section (i.e. the convergent section, the throat section and the divergent section except the extended section), and the divergent section with an area ratio of about 10 to the outlet is considered as a component separately in structure and is called a nozzle extended section. The gas temperature at the inlet of the nozzle extended section is above 2000K, so the structural cooling of the inlet of the extended section is the focus of product design.

[0003] In the prior art, the nozzle extended section adopts full sandwich structure regenerative cooling, special composite material structure to achieve the cooling purpose or adopts gas cooling of turbine exhaust of other systems, which is the common cooling scheme at present. Among them, the full sandwich cooling structure is a traditional technical scheme, small-size nozzles generally adopt mature welding or 3D printing process, but it will cause relatively large mass of the nozzle and large system demand supply capacity, and the engine performance parameters are not optimal, and large-size nozzles of silicon carbide, ceramic and other composite materials have not entered the application stage. The gas cooling of turbine exhaust of other systems is a common cooling way for the structure of the nozzle extended section, which has two advantages: first, it forms a gas film cooling protection for the nozzle extended section, improves the working reliability of the nozzle extended section, reduces the material system requirement of the nozzle extended section, and reduces the cost and weight; second, the gas of turbine exhaust has rich combustion, which can increase the gas momentum at the nozzle outlet, generate additional thrust and improve the overall specific impulse of the engine.

[0004] In the prior art, the structure for introducing the turbine exhaust of other systems into the nozzle extended section mainly includes a cold regenerative cooling thrust chamber nozzle divergent section, an inlet flange, an inlet elbow, an exhaust collector, an exhaust nozzle, a cone cylinder, a nozzle flange and a nozzle extended section. The turbine exhaust of other systems is first connected to a heat exchanger, then connected to an exhaust collector, and then the turbine exhaust in the exhaust collector is introduced into the nozzle extended section to cool the shell of the nozzle extended section.

[0005] In the process of implementing the present application, the applicant found that at least the following problems exist in the prior art:

[0006] The turbine exhaust of other systems needs to be connected to a heat exchanger first and then connected to an exhaust collector, which makes the cooling way use more structural components, occupies a large area, increases the height size of the engine and increases the difficulty of the overall layout of the engine. SUMMARY

[0007] The embodiment of the present application provides a gas cavity for a thrust chamber, an engine thrust chamber, an engine and a liquid carrying rocket, and can solve the technical problem in the prior art that the turbine exhaust of other systems needs to be connected to a heat exchanger first and then connected to an exhaust collector, so that the cooling mode uses many structural components, occupies a large area, increases the height dimension of the engine, and increases the difficulty of the overall layout of the engine.

[0008] To achieve the above object, in a first aspect, the embodiment of the present application provides a gas cavity for a thrust chamber, a nozzle including an expansion section and an extension section located downstream of the expansion section, a lower end of the expansion section being connected to an upper end of the extension section, and the gas cavity being wound around a lower end of the expansion section adjacent to the extension section.

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

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

[0011] The annular cavity has a gas inlet connected to a turbine exhaust pipe for receiving the gas exhausted by the turbine of other systems.

[0012] The heat exchanger is arranged along the annular cavity in the annular cavity.

[0013] The gas ejection device is arranged on the gas cavity adjacent to the lower end of the expansion section, and the flow direction of the gas ejected from the gas ejection device is consistent with the axial direction of the extension section and is directed into the extension section.

[0014] In a second aspect, the embodiment of the present application provides an engine thrust chamber, including a combustion chamber, a nozzle and the aforementioned gas cavity for a thrust chamber, the nozzle including an expansion section and an extension section located downstream of the expansion section, a lower end of the expansion section being connected to an upper end of the extension section, the gas cavity being wound around the lower end of the expansion section, and a cross-sectional outer diameter of the gas cavity at the gas ejection device being smaller than an inner diameter of the upper end of the extension section.

[0015] In a third aspect, the embodiment of the present application provides an engine, including the aforementioned engine thrust chamber.

[0016] In a fourth aspect, the embodiment of the present application provides a liquid carrying rocket, including the aforementioned engine.

[0017] The technical scheme has the following beneficial effects: in the embodiment of the application, the turbine exhaust pipe is used to connect the turbine exhaust gas of other systems into the annular cavity, the temperature of the introduced gas is about 800-900K, the heat exchange medium in the heat exchanger exchanges heat with the introduced gas, and the temperature of the gas is reduced to about 700K; the 700K gas is sprayed into the extension section through the gas spraying device, the gas film is formed in the extension section, the gas film separates the engine exhaust gas entering the extension section of the nozzle from the inner wall of the extension section, and the purpose of cooling and protecting the extension section is achieved, so that the extension section can work safely and reliably. Compared with the prior art of connecting the heat exchanger and the gas cavity, the heat exchanger is arranged along the annular channel in the annular cavity in the embodiment of the application, the structure is compact, the space used is greatly reduced, so that the height size of the overall layout of the engine is reduced, the overall structure is compact, and the problem of increasing the height size of the overall layout of the engine is solved. BRIEF DESCRIPTION OF DRAWINGS

[0018] In order to more clearly illustrate the technical scheme in the embodiments of the application or the prior art, the drawings needed in the following embodiment or prior art description will be briefly introduced. Obviously, the drawings in the following description are only some embodiments of the application, and other drawings can be obtained by those skilled in the art without creative labor.

[0019] Figure 1 is the axial view of the gas cavity for the thrust chamber in the embodiment of the application;

[0020] Figure 2 is Figure 1 is the corresponding top view;

[0021] Figure 3 is Figure 1 is the corresponding sectional view of the left view of

[0022] Figure 4 is the axial view of the gas cavity for the thrust chamber in the embodiment of the application (part of the structure is removed).

[0023] The reference signs are as follows:

[0024] 100, expansion section; 10, annular cavity; 20, heat exchanger; 30, gas spraying device; 40, inlet elbow; 50, flow equalizing 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 DESCRIPTION

[0026] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative effort belong to the scope of the present application.

[0027] As shown in Figure 1 , Figure 2 and Figure 3 , a gas cavity for a thrust chamber according to the present application, a nozzle comprises a divergent section 100 and an extended section located downstream of the divergent section 100, a lower end of the divergent section 100 is connected to an upper end of the extended section, and the gas cavity is wound around the lower end of the divergent section 100 adjacent to the extended section;

[0028] The gas cavity comprises 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 divergent section 100 to form a closed annular cavity channel with the outer wall of the lower end of the divergent section 100. The annular cavity 10 has a gas inlet connected to a turbine exhaust pipe for receiving turbine exhaust gas from other systems. The heat exchanger 20 is arranged along the annular cavity channel in the annular cavity 10. The gas injection device 30 is arranged on the gas cavity adjacent to the lower end of the divergent section 100, and the flow direction of the gas injected from the gas injection device 30 is consistent with the axial direction of the extended section and is directed into the extended section. Wherein, the other system refers to a sub-system of an engine composed of a gas generator and a turbine pump.

[0029] The thrust chamber is a component of a liquid rocket engine for generating thrust, which comprises an injector, a combustion chamber and a nozzle. Propellant is converted into high-temperature and high-pressure combustion products in the combustion chamber, and is discharged at high speed through the nozzle to generate thrust. The nozzle comprises a convergent section, a throat, a divergent section and an extended section.

[0030] In the embodiments of the present application, turbine exhaust gas from other systems is introduced into the annular cavity 10 through a 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 to reduce the temperature of the gas to about 700K. The 700K gas is injected into the extended section through the gas injection device 30 to form a gas film in the extended section. The gas film separates the exhaust gas of the engine from the inner wall of the extended section after entering the extended section of the nozzle, thereby achieving the purpose of cooling and protecting the extended section, so that the extended section can work safely and reliably.

[0031] Compared with the prior art of connecting the heat exchanger and the gas cavity, the heat exchanger 20 is arranged along the annular cavity in the annular cavity 10, the structure is compact, the space used is greatly reduced, the height size of the overall layout of the engine is reduced, the overall structure is compact, and the problem of increasing the height size of the overall layout of the engine is solved.

[0032] Preferably, as shown in Figure 1 and Figure 2 The annular cavity 10 is a shell wound around the expansion section 100 in the extension direction and not closed; the shell not closed on the annular cavity 10 is buckled on the outer wall of the lower end of the expansion section 100, and forms 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 in the extension direction and not closed, compared with the shell wound around the expansion section 100 in the extension direction and fully closed, the mass of the gas cavity can be reduced. Further, the mass of the overall engine can be reduced. The annular cavity can realize the circumferential flow of the turbine exhaust gas.

[0033] The cross section of the annular cavity 10 is a gradually changing cross section gradually changing from the gas inlet side to the opposite side of the gas inlet; after the turbine exhaust gas discharged from the turbine exhaust pipe enters the gas inlet of the annular cavity 10, the flow rate at the gas inlet side is greater than that at the opposite side of the gas inlet, so as to ensure that the flow rate and static pressure at the two places are basically the same. The cross section of the annular cavity 10 is designed to match the flow rate, which can reduce the mass of the annular cavity 10, and further reduce the mass of the overall engine.

[0034] Preferably, as shown in Figure 2 The gas cavity for the thrust chamber further comprises 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 through a flange structure, so as to smoothly connect the turbine exhaust pipe to the annular cavity 10. The inlet elbow 40 can be integrally formed by 3D printing, which has small machining difficulty, high precision and short machining cycle compared with the sheet metal forming process of parts.

[0035] Preferably, as shown in Figure 4As shown, the gas cavity for the thrust chamber further comprises a flow uniformizing plate 50 connected to the lowermost end of the inlet elbow 40, and the flow uniformizing plate 50 is provided with a plurality of flow uniformizing holes arranged at intervals. If the flow uniformizing plate 50 is not arranged, the gas introduced into the annular cavity 10 from the lowermost end of the inlet elbow 40 will directly flow into the annular cavity below the lowermost end of the inlet elbow 40 at a high speed, and the flow of the gas to other places of the annular cavity is greatly reduced, and then the flow of the gas into the gas ejection device 30 at other places of the annular cavity is greatly reduced, which causes the flow of the gas ejected from the gas ejection device 30 to be uneven, and the gas film formed by the gas entering the extension section is uneven, and the cooling protection effect on the extension section is not good. In the present application, the flow uniformizing plate 50 is arranged at the lowermost end of the inlet elbow 40, and a plurality of flow uniformizing holes are arranged on the flow uniformizing plate 50. The flow uniformizing holes can slow down the gas of the inlet elbow 40 to realize axial throttling, and most of the inlet gas is guided to the two side annular cavities for uniform distribution, thereby playing a role in uniformly flowing downward, so that the gas can flow into the entire annular cavity, and the flow of the gas ejected from the gas ejection device 30 at each place is basically equivalent, and the flow of the gas along the axial direction of the extension section is uniformly distributed, i.e., the gas can be uniformly distributed on the inner wall side of the extension section. The number and arrangement position of the flow uniformizing holes are calculated and verified by simulation, and the flow distribution of the gas entering the gas nozzle 301 from the annular cavity is relatively uniform, so that the optimal flow uniformizing hole parameters (number and arrangement) are obtained in the embodiment of the present application.

[0036] Preferably, as Figure 1 and Figure 2As shown, the inlet elbow 40 has a first inlet 60 and a second inlet 70 obtained by separating the inlet elbow 40 by an intermediate partition plate 80, i.e. the first inlet 60 and the second inlet 70 are located above the flow equalizing plate 50 and respectively communicate with two sides of the annular cavity 10 in opposite directions, and the intermediate partition plate 80 is located above the flow equalizing plate 50 with a gap between the flow equalizing plate 50; the annular cavity 10 at least includes a first gas collecting ring 101 and a second gas collecting ring 102, the first end surface of the first gas collecting ring 101 and the first end surface of the second gas collecting ring 102 are connected, the second end surface of the first gas collecting ring 101 is connected to the first inlet 60, and the second end surface of the second gas collecting ring 102 is connected to the second inlet 70. The gas entering the annular cavity 10 is divided by the first inlet 60 and the second inlet 70, so that the gas flows in two directions along the annular cavity and to the distal end of the annular cavity, and further through the multiple flow equalizing holes arranged on the flow equalizing plate 50 to throttle and equalize the flow of the gas, so that more gas flows along the annular cavity and to the distal end of the annular cavity. Among them, 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 multiple gas collecting rings are connected to form an annular cavity 10 with a gradually changing cross section from the gas inlet side to the opposite side of the gas inlet. The first gas collecting ring 101 and the second gas collecting ring 102 are integrally formed by 3D printing, which has smaller processing difficulty, higher precision and shorter processing cycle compared with the process of part sheet metal forming.

[0037] Preferably, as shown in Figure 1 and Figure 3 The gas ejection device 30 includes multiple gas nozzles 301 arranged at intervals, the gas nozzles 301 can be square holes, and the multiple gas nozzles 301 form a ring mechanism, the flow direction of the gas ejected by the gas nozzles 301 is consistent with the axial direction of the extension section and is towards the extension section, and the gas ejection device 30, i.e. the gas nozzles 301, is welded (for example, argon arc welding process) on the outer wall of the expansion section and the annular cavity 10. The gas ejected by the multiple gas nozzles 301 in the form of a ring mechanism forms a ring, and after being injected into the upper end inlet of the extension section of the nozzle, a ring-shaped gas film is formed at the inner wall of the extension section of the nozzle to protect 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 decompressed, and the speed reaches the speed of sound at the throat of the gas nozzle 301, and then continues to accelerate and decompress in the expansion section of the gas nozzle 301. The gas is injected into the upper end inlet of the extension section of the nozzle from the outlet of the gas nozzle 301 to form a low-temperature gas film to protect the shell structure of the extension section.

[0039] Preferably, as shown in Figure 1 The heat exchanger 20 comprises a heat exchange medium inlet 103 and a heat exchange medium outlet 104 penetrating through the annular cavity 10.

[0040] The heat exchange medium inlet 103 is used to provide the heat exchanger 20 with 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 fuel gas, so as to ensure that the heat exchanger 20 continuously has low-temperature heat exchange medium, and heat exchange and temperature reduction of the fuel gas entering the annular cavity 10 can be realized.

[0041] In combination with the embodiments of the present application, an engine thrust chamber is provided, comprising an injector, a combustion chamber, a nozzle, and the fuel gas cavity for the thrust chamber as described above, the nozzle comprises 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 fuel gas cavity is wound around the lower end of the expansion section 100, and the cross-sectional outer diameter of the fuel gas cavity at the fuel gas ejection device 30 is smaller than the inner diameter of the upper end of the extension section. The fuel gas discharged by the turbine of other systems is introduced into the annular cavity 10 through the turbine exhaust pipe, the temperature of the introduced fuel gas is about 800-900K, the heat exchange medium in the heat exchanger 20 exchanges heat with the introduced fuel gas, and the temperature of the fuel gas is reduced to about 700K; the fuel gas at 700K is sprayed into the extension section through the fuel gas ejection device 30, and a gas film is formed in the extension section, which separates the fuel gas discharged by the engine and entering the extension section of the nozzle from the inner wall of the extension section, so as to achieve the purpose of cooling and protecting the extension section, and the extension section can work safely and reliably.

[0042] In combination with the embodiments of the present application, an engine is provided, comprising the engine thrust chamber as described above.

[0043] In combination with the embodiments of the present application, a liquid carrier rocket is provided, comprising the engine as described above.

[0044] It should be understood that in the foregoing detailed description, various features are combined in a single embodiment for simplicity. This disclosure does not reflect the intention that the claimed subject matter requires more features than expressly set forth in each claim. On the contrary, as reflected by the appended claims, the present application is susceptible to further combinations. Accordingly, the appended claims are hereby expressly incorporated into this detailed description, with each claim acting as a separate embodiment of the present application.

[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 embodiments described above further illustrate the purpose, technical solution, and beneficial effects of the present invention. It should be understood that the above description is only a specific embodiment 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 within the scope of protection of the present invention.

Claims

1. A gas cavity for a thrust chamber, the nozzle comprising a divergent section (100) and an extended section downstream of the divergent section (100), the lower end of the divergent section (100) being connected to the upper end of the extended section, the gas cavity being wound around the lower end of the divergent section (100) adjacent to the extended section, characterized in that, the gas cavity comprises 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 divergent section (100) to form a closed annular cavity channel with the outer wall of the lower end of the divergent section (100); the annular cavity (10) has a gas inlet connected to a turbine exhaust pipe for receiving turbine exhaust gas from other systems; the heat exchanger (20) is arranged along the annular cavity channel in the annular cavity (10); the gas injection device (30) is arranged on the gas cavity adjacent to the lower end of the divergent section (100), the flow direction of the gas injected from the gas injection device (30) being consistent with the axial direction of the extended section and towards the inside of the extended section.

2. The gas cavity for a thrust chamber according to claim 1, characterized in that, the annular cavity (10) is a shell that is not closed in the direction of its extension and is wound around the divergent section (100) for one turn; and the cross section of the annular cavity (10) is a gradually changing cross section that gradually decreases from the side of the gas inlet to the opposite side of the gas inlet; the shell that is not closed on the annular cavity (10) is fitted on the outer wall of the lower end of the divergent section (100) to form a closed annular cavity channel with the outer wall of the lower end of the divergent section (100).

3. The gas cavity for a thrust chamber according to claim 1, characterized in that, It further comprises an inlet elbow (40), the first end of the inlet elbow (40) being connected to the gas inlet, and the second end of the inlet elbow (40) being connected to the turbine exhaust pipe.

4. The gas cavity for a thrust chamber according to claim 3, characterized in that It further comprises a flow uniformizing plate (50) connected to the lowermost end of the inlet elbow (40), the flow uniformizing plate (50) being provided with a plurality of flow uniformizing holes arranged at intervals.

5. The gas cavity for a 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 plate (80), the intermediate partition plate (80) being located above the flow uniformizing plate (50) and having a gap with the flow uniformizing plate (50); The annular cavity (10) comprises 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) being connected to the first end face of the second gas collecting ring (102), the second end face of the first gas collecting ring (101) being connected to the first inlet (60), and the second end face of the second gas collecting ring (102) being connected to the second inlet (70).

6. The gas cavity for a thrust chamber of claim 1, wherein, The gas injection device (30) comprises a plurality of gas nozzles (301) arranged at intervals, the flow direction of the gas injected from the gas nozzles (301) being consistent with the axial direction of the extended section and towards the inside of the extended section.

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

8. The gas cavity for a thrust chamber of claim 1, wherein, The heat exchanger (20) comprises a heat exchange medium inlet (103) and a heat exchange medium outlet (104) penetrating through the annular cavity (10).

9. An engine thrust chamber characterized by, The gas cavity for the thrust chamber of any one of claims 1-8, wherein the gas cavity is wrapped around the lower end of the divergent section (100) and has a cross-sectional outer diameter at the gas ejection device (30) that is less than an inner diameter of the upper end of the extension section.

10. An engine characterized by, The engine thrust chamber of claim 9.

11. A liquid launch vehicle, characterized by, The engine of claim 10.

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