阅读说明：本技术 一种火箭发动机涡轮泵的密封结构 (Sealing structure of turbopump of rocket engine ) 是由 李晓俊 林言丕 朱祖超 李林敏 于 2021-09-15 设计创作，主要内容包括：本发明公开了一种火箭发动机涡轮泵的密封结构,其包括离心叶轮(4),离心叶轮包括后盘(41)、前盘、多个叶片,第一密封环与第二密封环之间形成密封副；其特征在于：在后盘(41)的外侧且位于第一密封环(42)的径向内端设置有多个辅助叶片(43、44),多个辅助叶片沿周向均匀分布,且在径向上,多个辅助叶片与第一密封环之间形成回流空间,在后盘(41)上且位于第一密封环(42)的径向内端设置有多个回流孔(45),回流空间与回流孔连通。本发明将一部分的从密封副泄漏的泄漏流引导至回流空间,并从回流空间引导至回流孔,以实现回流,从而减少泄漏流向机械密封处的泄漏,从而提高涡轮泵的整体性能和效率。(The invention discloses a sealing structure of a turbopump of a rocket engine, which comprises a centrifugal impeller (4), wherein the centrifugal impeller comprises a rear disc (41), a front disc and a plurality of blades, and a sealing pair is formed between a first sealing ring and a second sealing ring; the method is characterized in that: a plurality of auxiliary blades (43 and 44) are arranged on the outer side of the rear disc (41) and on the radial inner end of the first sealing ring (42), the auxiliary blades are evenly distributed along the circumferential direction, a backflow space is formed between the auxiliary blades and the first sealing ring in the radial direction, a plurality of backflow holes (45) are formed in the rear disc (41) and on the radial inner end of the first sealing ring (42), and the backflow space is communicated with the backflow holes. The invention guides part of leakage flow leaked from the sealing pair to the backflow space and guides the leakage flow from the backflow space to the backflow hole to realize backflow, thereby reducing the leakage flow to the mechanical seal, and improving the overall performance and efficiency of the turbine pump.)

1. A sealing structure of a turbopump of a rocket engine comprises a first shell (1), a second shell (2), a third shell (3), a first centrifugal impeller (4), a first spiral inducer (5), a common shaft (6), a second centrifugal impeller (7), a second spiral inducer (8), a mechanical seal (9) and an inlet flow channel, wherein one end of the first shell is connected with the second shell through a connecting piece, the other end of the first shell is connected with the third shell through a connecting piece, the upstream end of the first centrifugal impeller is provided with the first spiral inducer, the first spiral inducer is adjacent to the inlet flow channel, the upstream end of the second centrifugal impeller is provided with the second spiral inducer, the first centrifugal impeller, the first spiral inducer, the second centrifugal impeller and the second spiral inducer are respectively arranged on the common shaft, the mechanical seal is arranged on the periphery of the common shaft in the first shell, the first centrifugal impeller and the second centrifugal impeller are arranged back to back relative to the mechanical seal; the first centrifugal impeller (4) and/or the second centrifugal impeller (7) comprise a rear disc (41), a front disc and a plurality of blades, the blades are circumferentially distributed between the rear disc and the front disc, a first sealing ring (42) is arranged on the outer side of the rear disc and at the radial outer end of the rear disc, a second sealing ring (11) is arranged at one end of the first shell, and a sealing pair is formed between the first sealing ring and the second sealing ring;

the method is characterized in that: a plurality of auxiliary blades (43 and 44) are arranged on the outer side of the rear disc (41) and on the radial inner end of the first sealing ring (42), the auxiliary blades are evenly distributed along the circumferential direction, a backflow space is formed between the auxiliary blades and the first sealing ring in the radial direction, a plurality of backflow holes (45) are formed in the rear disc (41) and on the radial inner end of the first sealing ring (42), and the backflow space is communicated with the backflow holes.

2. A sealing structure of a turbopump of a rocket engine according to claim 1, wherein said auxiliary vanes (43, 44) comprise a plurality of first auxiliary vanes (43) and second auxiliary vanes (44), the heights of the first auxiliary vanes and the second auxiliary vanes are different in the axial direction, and the height of the second auxiliary vanes is 0.3 to 0.7 times the height of the first auxiliary vanes.

3. A sealing structure of a turbopump of a rocket engine according to claim 2, wherein a plurality of first auxiliary vanes (43) and second auxiliary vanes (44) are alternately arranged at intervals in a circumferential direction, and one or two second auxiliary vanes are arranged between every two first auxiliary vanes.

4. A sealing structure of a turbopump of a rocket engine according to claim 3, wherein a radially outer end of said first auxiliary vane (43) has a first arc-shaped portion (46) which is a substantially semicircular arc-shaped portion.

5. A rocket engine turbopump sealing structure as claimed in claim 4, wherein the radially outer end of said second auxiliary vane (44) has a second arcuate portion (47) which is a substantially 1/4 circular arcuate portion.

6. A rocket engine turbopump sealing structure as claimed in claim 5, wherein the radially inner end of said first sealing ring (42) has a third arcuate portion (48) which is a substantially 1/4 circular arcuate portion, the radially inner end of the first sealing ring has an inclined surface, and the third arcuate portion is connected to the inclined surface.

7. A rocket engine turbopump sealing structure as claimed in claim 6, characterized in that in an axial cross-sectional view, the first curved portion (46), the second curved portion (47) and the third curved portion (48) substantially constitute a portion 3/4 of a circle, and the return space is a substantially circular return space.

8. A rocket engine turbopump sealing structure as recited in claim 6, wherein said first auxiliary vane (43) has a setting angle of 30-75 °, said second auxiliary vane (44) has a setting angle of 30-75 °, said first auxiliary vane is a two-dimensional vane or a three-dimensional vane, and said second auxiliary vane is a two-dimensional vane or a three-dimensional vane.

Technical Field

The invention relates to the technical field of turbopumps of rocket engines, in particular to a sealing structure of a turbopump of a rocket engine.

Background

The turbopump of the rocket engine mainly comprises an inducer, a centrifugal impeller, a mechanical seal, a bearing, a shafting supporting system, a shell and the like. However, the existing sealing structure of the turbine pump has the problems of large leakage and incapability of backflow.

Disclosure of Invention

The invention aims to overcome the defects in the prior art, and provides a sealing structure of a turbopump of a rocket engine, wherein a backflow space/backflow channel is formed between a plurality of auxiliary blades and a first sealing ring in the radial direction through the design of the auxiliary blades, a plurality of backflow holes are formed in the rear disc and positioned at the inner end of the first sealing ring in the radial direction, and the backflow space/backflow channel is communicated with the backflow holes and used for guiding part/most of leakage flow leaked from a sealing pair to the backflow space and guiding the leakage flow from the backflow space to the backflow holes to realize backflow, so that the leakage of the leakage flow to a mechanical sealing position is reduced, the sealing effect is better than that of the traditional labyrinth sealing, and the overall performance and the efficiency of the turbopump are improved. Through the design in circular shape backward flow space, can promote to leak the interior runner of flow direction backward flow space to guide to the backward flow hole from the backward flow space, in order to realize the backward flow, thereby reduce the leakage that leaks flow direction to mechanical seal department, thereby improve turbopump's wholeness ability and efficiency.

In order to achieve the purpose, the invention adopts the technical scheme that:

a sealing structure of a turbopump of a rocket engine comprises a first shell (1), a second shell (2), a third shell (3), a first centrifugal impeller (4), a first spiral inducer (5), a common shaft (6), a second centrifugal impeller (7), a second spiral inducer (8), a mechanical seal (9) and an inlet flow channel, wherein one end of the first shell is connected with the second shell through a connecting piece, the other end of the first shell is connected with the third shell through a connecting piece, the upstream end of the first centrifugal impeller is provided with the first spiral inducer, the first spiral inducer is adjacent to the inlet flow channel, the upstream end of the second centrifugal impeller is provided with the second spiral inducer, the first centrifugal impeller, the first spiral inducer, the second centrifugal impeller and the second spiral inducer are respectively arranged on the common shaft, the mechanical seal is arranged on the periphery of the common shaft in the first shell, the first centrifugal impeller and the second centrifugal impeller are arranged back to back relative to the mechanical seal; the first centrifugal impeller (4) and/or the second centrifugal impeller (7) comprise a rear disc (41), a front disc and a plurality of blades, the blades are circumferentially distributed between the rear disc and the front disc, a first sealing ring (42) is arranged on the outer side of the rear disc and at the radial outer end of the rear disc, a second sealing ring (11) is arranged at one end of the first shell, and a sealing pair is formed between the first sealing ring and the second sealing ring; the method is characterized in that: a plurality of auxiliary blades (43 and 44) are arranged on the outer side of the rear disc (41) and on the radial inner end of the first sealing ring (42), the auxiliary blades are evenly distributed along the circumferential direction, a backflow space is formed between the auxiliary blades and the first sealing ring in the radial direction, a plurality of backflow holes (45) are formed in the rear disc (41) and on the radial inner end of the first sealing ring (42), and the backflow space is communicated with the backflow holes.

Further, the auxiliary blades (43, 44) comprise a plurality of first auxiliary blades (43) and second auxiliary blades (44), the heights of the first auxiliary blades and the second auxiliary blades are different in the axial direction, and the height of the second auxiliary blades is 0.3-0.7 times that of the first auxiliary blades.

Further, in the circumferential direction, a plurality of first auxiliary blades (43) and second auxiliary blades (44) are alternately arranged at intervals, and one or two second auxiliary blades are arranged between every two first auxiliary blades.

Further, the radially outer end of the first auxiliary vane (43) has a first arc-shaped portion (46) which is a substantially semicircular arc-shaped portion.

Further, the radially outer end of the second auxiliary vane (44) has a second arc-shaped portion (47) which is a circular arc-shaped portion of substantially 1/4.

Further, the radially inner end of the first seal ring (42) has a third arcuate portion (48) which is a substantially 1/4 circular arcuate portion, the radially inner end of the first seal ring has an inclined surface, and the third arcuate portion is connected to the inclined surface.

Further, in an axial cross-sectional view, the first arc-shaped portion (46), the second arc-shaped portion (47), and the third arc-shaped portion (48) substantially form a portion 3/4 of a circle, and the return space is a substantially circular return space.

Further, the installation angle of the first auxiliary blade (43) is 30-75 degrees, the installation angle of the second auxiliary blade (44) is 30-75 degrees, the first auxiliary blade is a two-dimensional blade or a three-dimensional blade, and the second auxiliary blade is a two-dimensional blade or a three-dimensional blade.

According to the sealing structure of the rocket engine turbopump, the auxiliary blades are designed, so that a backflow space/backflow channel is formed between the auxiliary blades and the first sealing ring in the radial direction, the backflow holes are formed in the rear disc and located at the inner end of the first sealing ring in the radial direction, and the backflow space/backflow channel is communicated with the backflow holes and used for guiding part/most of leakage flow leaked from the sealing pair to the backflow space and guiding the leakage flow from the backflow space to the backflow holes to achieve backflow, so that leakage flowing to a mechanical sealing position is reduced, the sealing structure is better than that of a traditional labyrinth seal, and the overall performance and efficiency of the turbopump are improved. Through the design in circular shape backward flow space, can promote to leak the interior runner of flow direction backward flow space to guide to the backward flow hole from the backward flow space, in order to realize the backward flow, thereby reduce the leakage that leaks flow direction to mechanical seal department, thereby improve turbopump's wholeness ability and efficiency.

Drawings

FIG. 1 is a schematic view of a turbopump of a rocket engine according to the present invention;

FIG. 2 is a schematic structural view of a sealing structure of a turbopump of the rocket engine according to the present invention;

fig. 3 is a structural schematic view (side view) of a sealing structure of a turbopump of a rocket engine according to the present invention.

In the figure: the centrifugal impeller comprises a first housing 1, a second housing 2, a third housing 3, a first centrifugal impeller 4, a first spiral inducer 5, a common shaft 6, a second centrifugal impeller 7, a second spiral inducer 8, a mechanical seal 9, a second seal ring 11, a rear disc 41, a first seal ring 42, a first auxiliary blade 43, a second auxiliary blade 44, a backflow hole 45, a first arc-shaped portion 46, a second arc-shaped portion 47, and a third arc-shaped portion 48.

Detailed Description

In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.

The present invention will be described in further detail with reference to the accompanying drawings.

As shown in fig. 1-3, a sealing structure of a turbopump of a rocket engine comprises a first casing 1, a second casing 2, a third casing 3, a first centrifugal impeller 4, a first spiral inducer 5, a common shaft 6, a second centrifugal impeller 7, a second spiral inducer 8, a mechanical seal 9, and an inlet flow channel, wherein one end of the first casing 1 is connected with the second casing 2 through a connecting piece, the other end of the first casing is connected with the third casing 3 through a connecting piece, the upstream end of the first centrifugal impeller 4 is provided with the first spiral inducer 5, the first spiral inducer 5 is adjacent to the inlet flow channel, the upstream end of the second centrifugal impeller 7 is provided with the second spiral inducer 8, the first centrifugal impeller 4, the first spiral inducer 5, the second centrifugal impeller 7, and the second spiral inducer 8 are respectively installed on the common shaft 6, the mechanical seal 9 is installed in the first casing 1 and located at the periphery of the common shaft 6, the first centrifugal impeller 4 and the second centrifugal impeller 7 are arranged back to back with respect to the mechanical seal 9, the first pump with the first centrifugal impeller 4 is used for pumping low-temperature methane (such as-160-.

The first centrifugal impeller 4 and/or the second centrifugal impeller 7 comprise a rear disc 41, a front disc, a plurality of blades circumferentially distributed between the rear disc 41 and the front disc, a first sealing ring 42 is arranged at the outer side of the rear disc 41 and at the radial outer end of the rear disc 41, a second sealing ring 11 is arranged at one end of the first shell 1, a sealing pair is formed between the first sealing ring 42 and the second sealing ring 11, a plurality of auxiliary blades (43, 44) are arranged at the outer side of the rear disc 41 and at the radial inner end of the first sealing ring 42, the plurality of auxiliary blades (43, 44) are uniformly distributed along the circumferential direction, and, in the radial direction, a return space/return channel is formed between the plurality of auxiliary vanes and the first seal ring 42, a plurality of return holes 45 are provided in the rear disc 41 at the radially inner end of the first seal ring 42, the return space/return channel communicating with the return holes 45.

Further, the auxiliary blade (43, 44) includes a plurality of first auxiliary blades 43 and second auxiliary blades 44, and the heights of the first auxiliary blades 43 and the second auxiliary blades 44 are different in the axial direction, and specifically, the height of the second auxiliary blades 44 is 0.4-0.6 times the height of the first auxiliary blades 43.

In the circumferential direction, a plurality of first auxiliary blades 43 and second auxiliary blades 44 are alternately arranged at intervals, and specifically, one second auxiliary blade 44 is arranged between every two first auxiliary blades 43.

According to the sealing structure of the turbopump of the rocket engine, through the design of the auxiliary blades (43 and 44), a backflow space/backflow channel is formed between the auxiliary blades and the first sealing ring 42 in the radial direction, the backflow hole 45 is formed in the rear disc 41 and is positioned at the inner end of the first sealing ring 42 in the radial direction, and the backflow space/backflow channel is communicated with the backflow hole 45 and is used for guiding part/most of leakage flow leaked from the sealing pairs (42 and 11) to the backflow space and guiding the leakage flow from the backflow space to the backflow hole 45 to realize backflow, so that the leakage flow to the mechanical seal 9 is reduced, the sealing effect is better than that of the traditional labyrinth seal, and the overall performance and the efficiency of the turbopump are improved.

As shown in fig. 2-3, further, the radially outer end of the first auxiliary vane 43 has a first arc-shaped portion 46, and the first arc-shaped portion 46 is a substantially semicircular arc-shaped portion. The radially outer end of the second auxiliary vane 44 has a second arc-shaped portion 47, and the second arc-shaped portion 47 is a circular arc-shaped portion of substantially 1/4.

Further, the radially inner end of the first seal ring 42 has a third arcuate portion 48, the third arcuate portion 48 being a substantially 1/4 circular arcuate portion, the radially inner end of the first seal ring 42 having an inclined surface, the third arcuate portion 48 being connected to the inclined surface.

In an axial cross-sectional view, the first arc-shaped portion 46, the second arc-shaped portion 47 and the third arc-shaped portion 48 substantially form a circular 3/4 portion, and the return flow space is a substantially circular return flow space.

According to the sealing structure of the turbopump of the rocket engine, leakage can be promoted to flow to the flow channel in the backflow space through the design of the circular backflow space, and the leakage is guided to the backflow hole 45 from the backflow space to realize backflow, so that the leakage of the leakage flowing to the mechanical seal 9 is reduced, and the overall performance and efficiency of the turbopump are improved.

Further, the installation angle of the first auxiliary blade 43 is 40 to 75 °, the installation angle of the second auxiliary blade 44 is 40 to 75 °, the first auxiliary blade 43 is a two-dimensional blade or a three-dimensional blade, and the second auxiliary blade 44 is a two-dimensional blade or a three-dimensional blade.

According to the sealing structure of the turbopump of the rocket engine, through the design of the auxiliary blades (43 and 44), a backflow space/backflow channel is formed between the auxiliary blades and the first sealing ring 42 in the radial direction, the backflow hole 45 is formed in the rear disc 41 and is positioned at the inner end of the first sealing ring 42 in the radial direction, and the backflow space/backflow channel is communicated with the backflow hole 45 and is used for guiding part/most of leakage flow leaked from the sealing pairs (42 and 11) to the backflow space and guiding the leakage flow from the backflow space to the backflow hole 45 to realize backflow, so that the leakage flow to the mechanical seal 9 is reduced, the sealing effect is better than that of the traditional labyrinth seal, and the overall performance and the efficiency of the turbopump are improved. Through the design of circular shape backward flow space, can promote to leak the interior runner of flow direction backward flow space to guide to backward flow hole 45 from the backward flow space, in order to realize the backward flow, thereby reduce the leakage that leaks flow direction to mechanical seal 9 department, thereby improve turbopump's wholeness ability and efficiency.

The above-described embodiments are illustrative of the present invention and not restrictive, it being understood that various changes, modifications, substitutions and alterations can be made herein without departing from the principles and spirit of the invention, the scope of which is defined by the appended claims and their equivalents.