CN113775560B - A sealing structure for a rocket engine turbopump - Google Patents

Abstract

The invention discloses a sealing structure of a rocket engine turbopump, which includes a centrifugal impeller (4). The centrifugal impeller includes a rear disk (41), a front disk, and a plurality of blades. A first sealing ring and a second sealing ring form a sealing structure. Sealing pair; characterized in that: a plurality of auxiliary blades (43, 44) are provided outside the rear disk (41) and located at the radial inner end of the first sealing ring (42), and the plurality of auxiliary blades are evenly distributed along the circumferential direction. , and in the radial direction, a backflow space is formed between the plurality of auxiliary blades and the first sealing ring, and a plurality of backflow holes (45) are provided on the rear disk (41) and located at the radial inner end of the first sealing ring (42) ), the return space is connected with the return hole. The invention guides a part of the leakage flow from the seal pair to the return space, and from the return space to the return hole to realize the return flow, thereby reducing the leakage flow to the mechanical seal, thereby improving the overall performance and efficiency of the turbine pump.

Description

A sealing structure for a rocket engine turbopump

Technical field

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

Background technique

The turbopump of a rocket engine is mainly composed of an inducer, a centrifugal impeller, a mechanical seal, a bearing, a shafting support system and a casing. However, the sealing structure of the existing turbine pump has the problem of large leakage and inability to flow back.

Contents of the invention

The purpose of the present invention is to overcome the deficiencies in the prior art and provide a sealing structure for a rocket engine turbopump. Through the design of the auxiliary blades, a backflow space is formed between a plurality of auxiliary blades and the first sealing ring in the radial direction. /Return channel, a plurality of return holes are provided on the rear plate and located at the radial inner end of the first sealing ring. The return space/return channel is connected with the return hole and is used to absorb part/most of the leakage from the sealing pair. The flow is guided to the return space and from the return space to the return hole to achieve return flow, thereby reducing leakage flow to the mechanical seal, which is better than the traditional labyrinth seal, thus improving the overall performance and efficiency of the turbine pump. Through the design of the circular return space, the leakage flow can be promoted to the flow channel in the return space, and guided from the return space to the return hole to realize the return flow, thereby reducing the leakage flow to the mechanical seal, thereby improving the overall performance of the turbine pump and efficiency.

In order to achieve the above objects, the technical solutions adopted by the present invention are:

A sealing structure of a rocket engine turbopump, which includes a first housing (1), a second housing (2), a third housing (3), a first centrifugal impeller (4), a first spiral inducer ( 5), common shaft (6), second centrifugal impeller (7), second spiral inducer (8), mechanical seal (9), inlet flow channel, one end of the first housing is connected to the second housing through a connector The other end is connected to the third housing through a connecting piece. The upstream end of the first centrifugal impeller is provided with a first spiral inducer. The first spiral inducer is adjacent to the inlet flow channel. The upstream end of the second centrifugal impeller is A second spiral inducer is provided at the end. The first centrifugal impeller, the first spiral inducer, the second centrifugal impeller, and the second spiral inducer are respectively installed on the common shaft. A mechanical device is installed in the first housing and on the outer periphery of the common shaft. Seal, the first centrifugal impeller and the second centrifugal impeller are arranged back-to-back with respect to the mechanical seal; the first centrifugal impeller (4) and/or the second centrifugal impeller (7) include a rear disk (41), a front disk, a plurality of blades, and a plurality of The blades are distributed circumferentially between the rear disk and the front disk. A first sealing ring (42) is provided outside the rear disk and at the radial outer end of the rear disk. A second sealing ring (42) is provided at one end of the first housing. The sealing ring (11) forms a sealing pair between the first sealing ring and the second sealing ring; it is characterized in that: there are multiple plurality of sealing rings located on the outside of the rear disk (41) and at the radial inner end of the first sealing ring (42). A plurality of auxiliary blades (43, 44) are evenly distributed along the circumferential direction, and in the radial direction, a return space is formed between the plurality of auxiliary blades and the first sealing ring, which is on the rear disk (41) and located on the first sealing ring. The radial inner end of the sealing ring (42) is provided with a plurality of return holes (45), and the return space is connected with the return holes.

Further, the auxiliary blades (43, 44) include a plurality of first auxiliary blades (43) and second auxiliary blades (44). In the axial direction, the heights of the first auxiliary blades and the second auxiliary blades are different, And the height of the second auxiliary blade is 0.3-0.7 times the height of the first auxiliary blade.

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

Further, the radial outer end of the first auxiliary blade (43) has a first arc portion (46), and the first arc portion is generally a semicircular arc portion.

Further, the radial outer end of the second auxiliary blade (44) has a second arc portion (47), and the second arc portion is generally a 1/4 circular arc portion.

Further, the radially inner end of the first sealing ring (42) has a third arcuate portion (48). The third arcuate portion is generally a 1/4 circular arcuate portion, and the diameter of the first sealing ring (42) is The inward end has an inclined surface, and the third arc portion is connected to the inclined surface.

Further, in the axial cross-sectional view, the first arcuate portion (46), the second arcuate portion (47), and the third arcuate portion (48) generally constitute a 3/4 part of a circle, and the return space is approximately Above is a circular return space.

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

The sealing structure of a rocket engine turbopump of the present invention, through the design of the auxiliary blades, forms a return space/return channel between a plurality of auxiliary blades and the first sealing ring in the radial direction, on the rear disk and located on the first sealing ring. The radial inner end of the sealing ring is provided with multiple return holes. The return space/reflow channel is connected with the return holes and is used to guide part/most of the leakage flow leaking from the sealing pair to the return space, and from the return space to The return hole is used to achieve return flow, thereby reducing leakage flow to the mechanical seal, which is better than the traditional labyrinth seal, thus improving the overall performance and efficiency of the turbine pump. Through the design of the circular return space, the leakage flow can be promoted to the flow channel in the return space, and guided from the return space to the return hole to realize the return flow, thereby reducing the leakage flow to the mechanical seal, thus improving the overall performance and performance of the turbine pump. efficiency.

Description of drawings

Figure 1 is a schematic structural diagram of a rocket engine turbopump of the present invention;

Figure 2 is a schematic structural diagram of the sealing structure of the rocket engine turbopump of the present invention;

Figure 3 is a schematic structural diagram (side view) of the sealing structure of the rocket engine turbopump of the present invention.

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

Detailed ways

In order to make the purpose, 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 in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments These are some embodiments of the present invention, rather than all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without making creative efforts fall within the scope of protection of the present invention.

The present invention will be further described in detail below in conjunction with the accompanying drawings.

As shown in Figures 1-3, a sealing structure of a rocket engine turbopump includes a first housing 1, a second housing 2, a third housing 3, a first centrifugal impeller 4, and a first spiral inducer 5 , common shaft 6, second centrifugal impeller 7, second spiral inducer 8, mechanical seal 9, inlet flow channel, one end of the first housing 1 is connected to the second housing 2 through a connecting piece, and the other end is connected through a connecting piece Connected to the third housing 3, a first spiral inducer 5 is provided at the upstream end of the first centrifugal impeller 4. The first spiral inducer 5 is adjacent to the inlet flow channel. The upstream end of the second centrifugal impeller 7 is provided with a first spiral inducer 5. 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, inside the first housing 1 and located on the common shaft. A mechanical seal 9 is installed on the outer periphery of 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 to pump low-temperature methane (such as -160-170°C ) or low-temperature liquid oxygen (such as -180-185°C), the second pump with the second centrifugal impeller 7 is used to pump low-temperature methane (such as -160-170°C) or low-temperature liquid oxygen (such as -180-185°C) .

The first centrifugal impeller 4 and/or the second centrifugal impeller 7 includes a rear disk 41 , a front disk, and a plurality of blades. The plurality of blades are circumferentially distributed between the rear disk 41 and the front disk, outside the rear disk 41 and located A first sealing ring 42 is provided at the radially outer end of the rear disk 41, and a second sealing ring 11 is provided at one end of the first housing 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 provided outside the rear disk 41 and at the radially inner end of the first sealing ring 42. The plurality of auxiliary blades (43, 44) are evenly distributed along the circumferential direction, and in the radial direction, A return flow space/reflow channel is formed between the plurality of auxiliary blades and the first sealing ring 42. A plurality of return flow holes 45 are provided on the rear disk 41 and located at the radial inner end of the first sealing ring 42. The return flow space/reflow channel is connected with The return hole 45 is connected.

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

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

The sealing structure of a rocket engine turbopump of the present invention uses the design of the auxiliary blades (43, 44) to form a return space/reflow channel between the plurality of auxiliary blades and the first sealing ring 42 in the radial direction. A plurality of return holes 45 are provided on the disk 41 and located at the radial inner end of the first sealing ring 42. The return space/reflow channel is connected with the return holes 45 for transferring part/most of the components from the sealing pair (42, 11) The leakage flow is guided to the return space, and from the return space to the return hole 45 to achieve return flow, thereby reducing the leakage flow to the mechanical seal 9, which is better than the traditional labyrinth seal, thereby improving the overall performance of the turbine pump and efficiency.

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

Further, the radially inner end of the first sealing ring 42 has a third arc portion 48. The third arc portion 48 is generally a 1/4 circular arc portion. The radially inner end of the first sealing ring 42 has a The inclined surface, the third arc portion 48 is connected to the inclined surface.

In the axial cross-sectional view, the first arcuate portion 46 , the second arcuate portion 47 , and the third arcuate portion 48 generally form a 3/4 portion of a circle, and the return flow space is generally a circular return space.

The sealing structure of a rocket engine turbine pump of the present invention can promote the leakage flow to the flow channel in the return space through the design of the circular return space, and guide it from the return space to the return hole 45 to realize the return flow, thereby reducing the leakage flow direction. Leakage at 9 points of the mechanical seal, thereby improving the overall performance and efficiency of the turbine pump.

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

The sealing structure of a rocket engine turbopump of the present invention uses the design of the auxiliary blades (43, 44) to form a return space/reflow channel between the plurality of auxiliary blades and the first sealing ring 42 in the radial direction. A plurality of return holes 45 are provided on the disk 41 and located at the radial inner end of the first sealing ring 42. The return space/reflow channel is connected with the return holes 45 for transferring part/most of the components from the sealing pair (42, 11) The leakage flow is guided to the return space, and from the return space to the return hole 45 to achieve return flow, thereby reducing the leakage flow to the mechanical seal 9, which is better than the traditional labyrinth seal, thereby improving the overall performance of the turbine pump and efficiency. Through the design of the circular return space, the leakage flow can be promoted to the flow channel in the return space, and guided from the return space to the return hole 45 to realize the return flow, thereby reducing the leakage flow to the mechanical seal 9, thereby improving the overall performance of the turbine pump Performance and efficiency.

The above-mentioned embodiments are illustrative of the present invention, not limitations of the present invention. It can be understood that various changes, modifications, substitutions and modifications can be made to these embodiments without departing from the principles and spirit of the present invention. The protection scope of the present invention as defined by the appended claims and their equivalents.

Claims (7)

1. The utility model provides a rocket engine turbopump's seal structure, it includes first casing (1), second casing (2), third casing (3), first centrifugal impeller (4), first spiral inducer (5), public axle (6), second centrifugal impeller (7), second spiral inducer (8), mechanical seal (9), import runner, first casing's one end is connected with second casing through the connecting piece, the other end is connected with third casing through the connecting piece, first centrifugal impeller's upstream end is provided with first spiral inducer, first spiral inducer is adjacent with import runner, second centrifugal impeller's upstream end is provided with second spiral inducer, first centrifugal impeller, first spiral inducer, second centrifugal impeller, second spiral inducer are installed on public axle respectively, mechanical seal is installed to the periphery that just is located public axle in the first casing, first centrifugal impeller and second centrifugal impeller are about mechanical seal back to back; the first centrifugal impeller (4) and/or the second centrifugal impeller (7) comprises a rear disc (41), a front disc and a plurality of blades, wherein the blades are circumferentially distributed between the rear disc and the front disc, a first sealing ring (42) is arranged at 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, 44) are arranged on the outer side of the rear disc (41) and positioned at the radial inner end of the first sealing ring (42), the plurality of auxiliary blades are uniformly distributed along the circumferential direction, a backflow space is formed between the plurality of auxiliary blades and the first sealing ring in the radial direction, a plurality of backflow holes (45) are arranged on the rear disc (41) and positioned at the radial inner end of the first sealing ring (42), and the backflow space is communicated with the backflow holes; 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 of the height of the first auxiliary blades.

2. A sealing structure of a turbopump of a rocket engine according to claim 1, wherein a plurality of first auxiliary vanes (43) and second auxiliary vanes (44) are alternately arranged at intervals in the circumferential direction, and one or two second auxiliary vanes are provided between each two first auxiliary vanes.

3. A sealing arrangement for a turbopump of a rocket engine according to claim 2, wherein the radially outer end of the first auxiliary vane (43) has a first arcuate portion (46) which is a semi-circular arcuate portion.

4. A sealing arrangement for a turbopump of a rocket engine according to claim 3, characterized in that the radially outer end of the second auxiliary vane (44) has a second arcuate portion (47) which is a 1/4 circular arcuate portion.

5. A sealing arrangement for a turbopump of a rocket engine in accordance with claim 4, wherein the radially inner end of the first sealing ring (42) has a third arcuate portion (48) which is a 1/4 circular arcuate portion, the radially inner end of the first sealing ring having an inclined surface, the third arcuate portion being connected to the inclined surface.

6. A sealing arrangement for a turbopump of a rocket engine according to claim 5, characterized in that, in an axial sectional view, the first arcuate portion (46), the second arcuate portion (47), the third arcuate portion (48) substantially form a 3/4 part of a circle, and the return space is a circular return space.

7. A sealing structure of a turbopump of a rocket engine according to claim 5, characterized in that the first auxiliary vane (43) has a mounting angle of 30-75 ° and the second auxiliary vane (44) has a mounting angle of 30-75 °, the first auxiliary vane being a two-dimensional vane or a three-dimensional vane, and the second auxiliary vane being a two-dimensional vane or a three-dimensional vane.