Turbine exhaust gas exhaust structure of a liquid rocket engine

Abstract

The present invention discloses a turbine exhaust gas exhaust structure for a liquid rocket engine, which includes a nozzle cascade, an exhaust part and a turbine disk. The gas in the combustion chamber is ejected through the nozzle cascade via a first channel, passes through a second channel on the turbine disk and then enters the guide vane cascade. While driving the turbine disk to rotate and output rotational power, it enters the nozzle through the guide vane cascade. The gas expands under the action of the guiding surface to generate thrust. After passing through the nozzle, the flow velocity increases sharply and the pressure decreases. The guiding surface on the exhaust part further increases the expansion area ratio of the nozzle, continues to accelerate the gas flowing out of the nozzle to increase the thrust, replaces the traditional exhaust pipe structure, simplifies the structural complexity, reduces the structural weight, can reduce the overall structural size and weight of the liquid rocket engine, helps to improve the thrust and specific impulse performance of the liquid rocket engine in different flight segments, and thus improves the carrying capacity of the rocket.

Description

Technical Field

[0001] This application relates to the technical field of liquid rocket engines, and particularly to a turbine exhaust gas exhaust structure of a liquid rocket engine. Background Art

[0002] The information provided in this section is for the purpose of generally presenting the background of the present disclosure. To the extent described in this section, the work of the currently named inventors and aspects that may not constitute prior art descriptions at the time of filing are neither expressly nor implicitly considered prior art to the present disclosure.

[0003] The turbine exhaust gas exhaust device is an important component of a gas generator cycle liquid rocket engine. Its function is to discharge the high-temperature gas that has done work on the turbine through the turbine exhaust gas pipe and generate a part of the thrust to reduce the specific impulse loss of the engine. Thrust and specific impulse are the main performance indicators of liquid rocket engines. Especially for vacuum engines, the vacuum specific impulse of the engine directly determines the carrying capacity of the rocket.

[0004] The traditional turbine exhaust gas pipe of a liquid rocket engine has a fixed expansion ratio and is often limited by the structural space, so the expansion ratio is relatively small. This results in a small thrust generated by the exhaust pipe and a variable specific impulse loss as the working altitude of the engine changes. Moreover, the structural dimensions are usually large and relatively heavy. These are all the deficiencies of the traditional turbine exhaust gas exhaust device of liquid rocket engines. Summary of the Invention

[0005] In view of the defects existing in the prior art, this application provides a turbine exhaust gas exhaust structure of a liquid rocket engine to solve the problems of large structure, heavy weight, and large specific impulse loss of the turbine exhaust gas exhaust device in the prior art.

[0006] The above object of this application is mainly achieved through the following technical solutions:

[0007] A turbine exhaust gas exhaust structure of a liquid rocket engine, comprising:

[0008] A nozzle cascade, the nozzle cascade is used to connect with a turbine inlet collector, and a plurality of first channels communicating with the turbine inlet collector are provided on the nozzle cascade;

[0009] An exhaust section, which is fixedly connected to the side of the nozzle cascade away from the turbine inlet collector. A plurality of nozzles are provided on the exhaust section. A plurality of guide vanes for guiding the gas to enter are provided on one side of the nozzles. The center line of the nozzle has an intersection with the center line of the exhaust section on the side away from the nozzle cascade and forms an angle, and the angle is an acute angle. A guiding surface is provided on the edge of the nozzle close to the center line of the exhaust section, and along the gas ejection direction, the distance between the guiding surface and the center line of the exhaust section shows a decreasing trend;

[0010] A turbine disk, which is rotatably arranged between the nozzle cascade and the exhaust section, and a plurality of second channels corresponding to the first channels are provided on the turbine disk.

[0011] Further, the turbine disk is coaxially arranged with the nozzle cascade.

[0012] Further, a stop block is sleeved on the turbine disk, and the stop block abuts between the turbine disk and the nozzle cascade.

[0013] Further, the guiding surface is a conical surface.

[0014] Further, the guiding surface is a parabolic surface.

[0015] Further, the exhaust section and the nozzle cascade are fixed by a detachable connecting member.

[0016] Further, the connecting member is a nut and a bolt that cooperate with each other.

[0017] Further, a sealing ring is provided between the exhaust section and the nozzle cascade.

[0018] Further, the nozzles are Laval nozzle structures arranged at intervals.

[0019] Further, the nozzles are annular nozzles arranged in the same layout as the exhaust section.

[0020] Compared with the prior art, the advantages of the present application are as follows:

[0021] In this application, a nozzle cascade is arranged on one side of the combustion chamber. A plurality of first channels communicating with the turbine intake collector are provided on the nozzle cascade. An exhaust part is arranged on the other side of the nozzle cascade. A plurality of nozzles are provided on the exhaust part. A plurality of guide vanes are arranged on one side of the nozzles. The center line of the nozzle and the center line of the exhaust part have an intersection on the side far from the nozzle cascade and form an included angle, and this included angle is an acute angle. A guiding surface is provided on the edge of the nozzle close to the center line of the exhaust part. And along the direction of the gas ejection, the distance between the guiding surface and the center line of the exhaust part shows a decreasing trend. A turbine disk is rotatably arranged between the nozzle cascade and the exhaust part. A plurality of second channels corresponding to the first channels are provided on the turbine disk. The gas in the turbine intake collector is ejected through the first channels, passes through the second channels on the turbine disk and then enters the guide vanes, drives the turbine disk to rotate and outputs rotational power, and at the same time passes through the guide vanes and enters the nozzles. Finally, the gas ejected from the nozzles continues to expand under the action of the guiding surface to generate thrust. The flow velocity of the gas increases sharply and the pressure decreases after passing through the nozzles. The guiding surface on the exhaust part further increases the expansion area ratio of the nozzles, and further accelerates the gas flowing out of the nozzles. Thus, the exhaust expansion ratio can be greatly increased in a smaller space, and partial thrust is additionally generated on the end surface of the exhaust part far from the nozzle cascade side, thereby greatly improving the thrust, replacing the traditional exhaust pipe structure, simplifying the structural complexity, reducing the structural weight, reducing the overall structural size and weight of the liquid rocket engine, and helping to improve the thrust and specific impulse performance of the liquid rocket engine in different flight segments, thereby improving the carrying capacity of the rocket. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the following will briefly introduce the drawings required for the description of the embodiments. Obviously, the drawings in the following description are only some embodiments of the present application. For those of ordinary skill in the art, without creative efforts, other drawings can be obtained based on these drawings.

[0023] Figure 1 It is a sectional view of the turbine exhaust gas exhaust structure provided by the embodiment of the present application;

[0024] Figure 2 It is a side view of the turbine exhaust gas exhaust structure provided by the embodiment of the present application;

[0025] In the figure: 1. Nozzle cascade; 11. First channel; 2. Exhaust part; 21. Nozzle; 22. Guide vane; 23. Guiding surface; 3. Turbine disk; 31. Second channel; 4. Stopper; 51. Nut; 52. Bolt; 6. Sealing ring. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0026] The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments. It should be noted here that the description of these embodiments is used to help understand the present invention, but does not constitute a limitation on the present invention. The specific structural and functional details disclosed herein are only used to describe the exemplary embodiments of the present invention. However, the present invention can be embodied in many alternative forms and should not be construed as limited to the embodiments described herein.

[0027] As Figure 1-2 shown, a turbine exhaust gas exhaust structure of a liquid rocket engine can be applied to liquid rocket engines using propellants such as liquid oxygen-liquid hydrogen, liquid oxygen-kerosene, or liquid oxygen-methane. It includes a nozzle vane cascade 1, an exhaust section 2, and a turbine disk 3, where:

[0028] The nozzle vane cascade 1 is used to connect to the turbine inlet collector, and a plurality of first channels 11 communicating with the turbine inlet collector are provided on the nozzle vane cascade 1. The first channels 11 of the nozzle vane cascade 1 can be formed by a plurality of sheet-like structures. When the gas in the turbine inlet collector ejects, the first channels 11 guide the gas to be discharged.

[0029] As Figure 1 shown, the exhaust section 2 is fixedly connected to the side of the nozzle vane cascade 1 away from the turbine inlet collector. A plurality of nozzles 21 are provided on the exhaust section 2. A plurality of guide vane cascades 22 for guiding the gas to enter are provided on one side of the nozzles 21. The center line of the nozzles 21 and the center line of the exhaust section 2 have an intersection and form an angle on the side away from the nozzle vane cascade 1. The angle is an acute angle. That is to say, the gas entering the nozzles 21 is ejected at a certain angle under the guidance of the arrangement direction of the nozzles 21, and the ejection direction of the gas in the nozzle forms an acute angle with the overall center line of the exhaust section 2. The direction of the gas ejection is restricted by the intersection of the center line of the nozzles 21 and the center line of the exhaust section 2 on the side away from the nozzle vane cascade 1. Relative to the exhaust section 2, the gas ejected from the nozzles 21 is in a converging state with respect to the center line of the exhaust section 2. As Figure 1 shown, the center line of the nozzles 21 and the center line of the exhaust section 2 have an intersection and form an angle on the side away from the nozzle vane cascade 1, which is the angle formed by L1 and L2 in the figure on the side away from the turbine inlet collector.

[0030] One side edge of the nozzle 21 close to the center line of the exhaust part 2 is provided with a guiding surface 23, and along the gas ejection direction, the distance between the guiding surface 23 and the center line of the exhaust part 2 shows a decreasing trend. On the premise that the gas ejected from the nozzle 21 is in a converging state with respect to the center line of the exhaust part 2, during the expansion process of the gas ejected from the nozzle 21, a pushing force is formed on the guiding surface 23. For the nozzle 21 on the exhaust part 2, only one side is provided with the guiding surface 23, and the other side of the nozzle 21 is directly the flight environment. The pressure of the gas flowing out of the exhaust part 2 is always the same as the ambient atmospheric pressure. Therefore, the specific impulse loss of the exhaust part 2 does not change with the change of the engine working altitude, which is beneficial to improving the thrust and specific impulse performance of the liquid rocket engine in different flight segments and also improving the flight reliability.

[0031] The turbine disk 3 is rotatably arranged between the nozzle cascade 1 and the exhaust part 2, and a plurality of second channels 31 corresponding to the first channels 11 are arranged on the turbine disk 3. After the gas passes through the first channels 11, it is ejected through the second channels 31, and when the gas passes through the second channels 31, the gas drives the turbine disk 3 to rotate. The turbine disk 3 can output rotational power and can be used as a power source for the work of other mechanisms.

[0032] The working principle of this embodiment is as follows: A nozzle cascade 1 is arranged on one side of the turbine intake collector. A plurality of first channels 11 communicating with the turbine intake collector are provided on the nozzle cascade 1. An exhaust part 2 is arranged on the other side of the nozzle cascade 1. A plurality of nozzles 21 are provided on the exhaust part 2. A plurality of guide vanes 22 are arranged on one side of the nozzles 21. The center line of the nozzles 21 and the center line of the exhaust part 2 have an intersection on the side far from the nozzle cascade 1 and form an angle, and this angle is an acute angle. A guiding surface 23 is provided on the edge of the side of the nozzles 21 close to the center line of the exhaust part 2. Along the direction of the gas ejection, the distance between the guiding surface 23 and the center line of the exhaust part 2 shows a decreasing trend. A turbine disk 3 is rotatably arranged between the nozzle cascade 1 and the exhaust part 2. A plurality of second channels 31 corresponding to the first channels 11 are provided on the turbine disk 3. The gas in the turbine intake collector is ejected through the first channels 11, passes through the second channels 31 on the turbine disk 3 and then enters the guide vanes 22, drives the turbine disk 3 to rotate and outputs rotational power, and at the same time passes through the guide vanes 22 and enters the nozzles 21. Finally, the gas after ejecting from the nozzles 21 generates thrust under the action of the guiding surface 23. The flow velocity of the gas increases sharply and the pressure decreases after passing through the nozzles 21. The guiding surface 23 on the exhaust part 2 further increases the expansion area ratio of the nozzles 21 and further accelerates the gas flowing out of the nozzles 21. Thus, the exhaust expansion ratio can be greatly increased in a relatively small space, and additional thrust is generated at the end face on the side of the exhaust part far from the nozzle cascade 1, thereby greatly improving the thrust, replacing the traditional exhaust pipe structure, simplifying the structural complexity, reducing the structural weight, reducing the overall structural size and weight of the liquid rocket engine, and contributing to improving the thrust and specific impulse performance of the liquid rocket engine in different flight segments, thereby improving the carrying capacity of the rocket.

[0033] Further, on the basis of the above embodiment, the turbine disk 3 and the nozzle cascade 1 are coaxially arranged. The coaxial arrangement of the turbine disk 3 and the nozzle cascade 1 maintains the accuracy of the correspondence between the first channels 11 and the second channels 31, which can improve the smoothness of the gas passing through and the utilization effect of the gas, and avoid unnecessary work waste.

[0034] Further, on the basis of the above embodiment, a stop block 4 is sleeved on the turbine disk 3. The stop block 4 abuts between the turbine disk 3 and the nozzle cascade 1. The setting of the stop block 4 can keep the center line of the turbine disk 3 always coincident with the nozzle cascade 1 when the turbine disk 3 rotates, avoiding the vibration generated during the rotation of the turbine disk 3 and the impact with the nozzle cascade 1, and improving the stability during the overall operation.

[0035] Further, on the basis of the above embodiment, the guiding surface 23 is a conical surface, which maintains the convenience of processing the guiding surface 23, and at the same time, the gas ejected from the nozzles 21 can quickly form a pushing force on the guiding surface 23 during the expansion process.

[0036] Further, on the basis of the above embodiments, the guiding surface 23 is a parabolic surface. During the expansion process of the fuel gas ejected from the nozzle 21, a pushing force is formed on the guiding surface 23. Thanks to the guiding surface 23 of the parabolic surface, the pushing effect of the fuel gas in the area of the guiding surface 23 is effectively improved, and the thrust generated by the expansion of the fuel gas is utilized to a greater extent, thereby improving the thrust and specific impulse performance.

[0037] Further, on the basis of the above embodiments, the end face of the exhaust part 2 on the side far from the nozzle cascade 1 is a flat end face. Thanks to the flat end face of the exhaust part 2, additional thrust is generated by the fuel gas flowing out of the guiding surface 23 on the flat end face.

[0038] Further, on the basis of the above embodiments, the exhaust part 2 and the nozzle cascade 1 are fixed by a detachable connecting member, which is convenient for disassembly and assembly and for the components to be replaced and maintained relatively independently.

[0039] Further, on the basis of the above embodiments, the connecting member is a nut 51 and a bolt 52 that cooperate with each other, which not only maintains the convenience of detachability but also can form a stable assembly connection relationship.

[0040] Further, on the basis of the above embodiments, a sealing ring 6 is provided between the exhaust part 2 and the nozzle cascade 1. The fuel gas generated by combustion can be guided more completely through the exhaust part 2, avoiding gas leakage caused by gaps under different jitter working conditions, and thus avoiding the loss of thrust and specific impulse performance.

[0041] Further, on the basis of the above embodiments, the nozzle 21 is a Laaval nozzle structure arranged at intervals. The nozzles 21 are arranged in multiple and at intervals, and are set as Laaval nozzle structures. The Laaval nozzle structure has a narrow throat in the middle, and expands outward from small to large after the narrow throat, which can change the speed of the fuel gas due to the change of the spray cross-sectional area, so that the fuel gas changes from subsonic speed to sonic speed and is accelerated to supersonic speed to obtain better thrust and specific impulse performance.

[0042] As Figure 2 shown, further, on the basis of the above embodiments, the nozzle 21 is an annular nozzle 21 arranged in the same way as the exhaust part 2, which is directly set as a continuous annular structure, can obtain a lighter overall mass, and the cross-sectional channel of the annular nozzle 21 can also be set in the shape of a Laaval nozzle.

[0043] It should be understood that the terms "first", "second", etc. are only used for descriptive distinction and should not be construed as indicating or implying relative importance. Although the terms "first", "second", etc. may be used herein to describe various units, these units should not be limited by these terms. These terms are only used to distinguish one unit from another. For example, the first unit may be referred to as the second unit, and similarly, the second unit may be referred to as the first unit, without departing from the scope of the exemplary embodiments of the present invention.

[0044] It should be understood that the term "and / or" herein is merely a description of the association relationship of associated objects, indicating that three relationships may exist. For example, A and / or B may represent: A exists alone, B exists alone, and both A and B exist simultaneously. The term " / and" herein describes another association object relationship, indicating that two relationships may exist. For example, A / and B may represent: A exists alone, and both A and B exist. Additionally, the character " / " herein generally represents that the associated objects before and after are in an "or" relationship.

[0045] It should be understood that in the description of the present invention, the orientation or positional relationship indicated by the terms "upper", "vertical", "inner", "outer", etc. is the orientation or positional relationship in which the disclosed product is customarily placed during use, or the orientation or positional relationship commonly understood by those skilled in the art. It is only for the convenience of describing the present invention and simplifying the description, and does not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on the present invention.

[0046] In the description of the present invention, it should also be noted that unless otherwise clearly specified and defined, the terms "arranged", "installed", "connected" should be understood in a broad sense. For example, it may be a fixed connection, a detachable connection, or an integral connection; it may be directly connected, or indirectly connected through an intermediate medium, and it may be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0047] The terms used herein are only for describing specific embodiments and are not intended to limit the exemplary embodiments of the present invention. As used herein, the singular forms "a", "an", and "the" are intended to include the plural forms unless the context clearly indicates otherwise. It should also be understood that the terms "comprise", "comprises", "include", and / or "includes" when used herein specify the existence of the stated features, integers, steps, operations, units, and / or components, and do not exclude the existence or addition of one or more other features, quantities, steps, operations, units, components, and / or their combinations.

[0048] Specific details are provided in the following description to facilitate a complete understanding of the example embodiments. However, those of ordinary skill in the art should understand that the example embodiments may be implemented without these specific details. In other embodiments, well-known processes, structures, and techniques may not be shown in unnecessary detail to avoid obscuring the example embodiments.

[0049] The above are only specific implementation manners of the present application, enabling those skilled in the art to understand or implement the present application. Various modifications to these embodiments will be obvious to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the present application. Therefore, the present application will not be limited to these embodiments shown herein, but rather should be accorded the widest scope consistent with the principles and novel features claimed herein.

[0050] It should be noted that the information disclosed in the above background art section is only used to enhance the understanding of the background of the present disclosure, and thus may include information that does not constitute the prior art known to those of ordinary skill in the art.

Claims

1. A turbine exhaust structure for a liquid rocket engine, characterized in that It includes: A nozzle cascade for connecting with a turbine inlet collector, and a plurality of first channels communicating with the turbine inlet collector are provided on the nozzle cascade; An exhaust part fixedly connected to the side of the nozzle cascade away from the turbine inlet collector. A plurality of nozzles are provided on the exhaust part. A plurality of guide vanes for guiding gas to enter are provided on one side of the nozzles. The center line of the nozzles and the center line of the exhaust part have an intersection and form an angle on the side away from the nozzle cascade. The angle is an acute angle. A guiding surface is provided on the side edge of the nozzle close to the center line of the exhaust part. Along the gas ejection direction, the distance between the guiding surface and the center line of the exhaust part shows a decreasing trend. The end surface of the exhaust part on the side away from the nozzle cascade is set as a flat end surface; A turbine disk rotatably arranged between the nozzle cascade and the exhaust part, and a plurality of second channels corresponding to the first channels are provided on the turbine disk.

2. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, characterized in that: The turbine disk is coaxially arranged with the nozzle cascade.

3. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 2, characterized in that: A stop block is sleeved on the turbine disk, and the stop block abuts between the turbine disk and the nozzle cascade.

4. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, wherein: The guiding surface is a conical surface.

5. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, characterized in that: The guiding surface is a parabolic surface.

6. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, characterized in that: The exhaust part and the nozzle cascade are fixed through a detachable connecting member.

7. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 6, characterized in that: The connecting member is a nut and a bolt that cooperate with each other.

8. The exhaust structure of the turbine waste gas of the liquid rocket engine according to claim 1, wherein: A sealing ring is provided between the exhaust part and the nozzle cascade.

9. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, wherein: The nozzles are Laval nozzle structures arranged at intervals.

10. The turbine exhaust gas exhaust structure of the liquid rocket engine according to claim 1, characterized in that: The nozzles are annular nozzles arranged in the same layout as the exhaust part.

Images