CN113775560B - Sealing structure of rocket engine turbine pump - Google Patents
Sealing structure of rocket engine turbine pump Download PDFInfo
- Publication number
- CN113775560B CN113775560B CN202111083022.5A CN202111083022A CN113775560B CN 113775560 B CN113775560 B CN 113775560B CN 202111083022 A CN202111083022 A CN 202111083022A CN 113775560 B CN113775560 B CN 113775560B
- Authority
- CN
- China
- Prior art keywords
- auxiliary
- backflow
- centrifugal impeller
- sealing ring
- arcuate portion
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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Links
- 238000007789 sealing Methods 0.000 title claims abstract description 60
- 238000000034 method Methods 0.000 claims abstract description 3
- 239000000411 inducer Substances 0.000 claims description 24
- 238000011144 upstream manufacturing Methods 0.000 claims description 6
- 230000008676 import Effects 0.000 claims description 4
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 description 4
- 230000000694 effects Effects 0.000 description 3
- MYMOFIZGZYHOMD-UHFFFAOYSA-N Dioxygen Chemical compound O=O MYMOFIZGZYHOMD-UHFFFAOYSA-N 0.000 description 2
- 238000009434 installation Methods 0.000 description 2
- 238000005086 pumping Methods 0.000 description 2
- 230000004075 alteration Effects 0.000 description 1
- 230000007547 defect Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/08—Sealings
- F04D29/083—Sealings especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/08—Sealings
- F04D29/10—Shaft sealings
- F04D29/102—Shaft sealings especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/26—Rotors specially for elastic fluids
- F04D29/28—Rotors specially for elastic fluids for centrifugal or helico-centrifugal pumps for radial-flow or helico-centrifugal pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
- F04D29/4206—Casings; Connections of working fluid for radial or helico-centrifugal pumps especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
- F04D29/44—Fluid-guiding means, e.g. diffusers
- F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The invention discloses a sealing structure of a rocket engine turbopump, 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, 44) are arranged on the outer side of the rear disc (41) and located at the radial inner end of the first sealing ring (42), the auxiliary blades are uniformly distributed along the circumferential direction, a backflow space is formed between the auxiliary blades and the first sealing ring in the radial direction, and a plurality of backflow holes (45) are formed in the rear disc (41) and located at the radial inner end of the first sealing ring (42) and are communicated with the backflow holes. The invention guides a part of leakage flow leaking from the sealing pair to the backflow space and guides the leakage flow from the backflow space to the backflow hole so as to realize backflow, thereby reducing leakage flowing to the mechanical sealing part, and improving the overall performance and efficiency of the turbine pump.
Description
Technical Field
The invention relates to the technical field of turbopumps of rocket engines, in particular to a sealing structure of a rocket engine turbopump.
Background
The turbine pump of rocket engine is mainly composed of inducer, centrifugal impeller, mechanical seal, bearing, shafting supporting system and shell. However, the sealing structure of the conventional 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 rocket engine turbine pump, wherein a plurality of auxiliary blades and a first sealing ring form a backflow space/backflow channel in the radial direction through the design of the auxiliary blades, a plurality of backflow holes are formed in a rear disc and positioned at the radial inner end of the first sealing ring, and the backflow space/backflow channel is communicated with the backflow holes and used for guiding a part/most of leakage flow leaking from a sealing pair to the backflow space and guiding the leakage flow from the backflow space to the backflow holes so as to realize backflow, thereby reducing leakage flowing to a mechanical sealing part. Through the design in circular backflow space, can promote leakage flow direction backward flow space internal flow passageway to guide to the backward flow hole from the backward flow space, in order to realize the backward flow, thereby reduce leakage flow direction mechanical seal department's leakage, thereby improve turbo pump's wholeness ability and efficiency.
In order to achieve the above purpose, the technical scheme adopted by the invention is as follows:
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 located at the radial inner end of the first sealing ring (42), the auxiliary blades are uniformly distributed along the circumferential direction, a backflow space is formed between the auxiliary blades and the first sealing ring in the radial direction, and a plurality of backflow holes (45) are formed in the rear disc (41) and located at the radial inner end of the first sealing ring (42) and are communicated with the backflow holes.
Further, the auxiliary blades (43, 44) comprise a plurality of first auxiliary blades (43), second auxiliary blades (44), the heights of the first auxiliary blades and the second auxiliary blades are unequal in the axial direction, and the heights of the second auxiliary blades are 0.3-0.7 times of the heights 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 arcuate portion (46) which is substantially a semicircular arcuate portion.
Further, the radially outer end of the second auxiliary blade (44) has a second arcuate portion (47) which is substantially a 1/4 circular arcuate portion.
Further, the radially inner end of the first seal ring (42) has a third arcuate portion (48) which is substantially a 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 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 space is a generally 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 turbine pump, through the design of the auxiliary blades, a backflow space/backflow channel is formed between the plurality of auxiliary blades and the first sealing ring in the radial direction, a plurality of backflow holes are formed in the rear disc and located at the radial inner end of the first sealing ring, the backflow space/backflow channel is communicated with the backflow holes and used for guiding a part/most of leakage flow leaking from the sealing pair to the backflow space and guiding the leakage flow to the backflow holes from the backflow space so as to realize backflow, and therefore leakage flowing to a mechanical sealing position is reduced, compared with a traditional labyrinth sealing effect, the overall performance and efficiency of the turbine pump are improved. Through the design in circular backflow space, can promote leakage flow direction backward flow space internal flow passageway to guide to the backward flow hole from the backward flow space, in order to realize the backward flow, thereby reduce leakage flow direction mechanical seal department's leakage, thereby improve turbo pump's wholeness ability and efficiency.
Drawings
FIG. 1 is a schematic view of a rocket engine turbopump of the present invention;
FIG. 2 is a schematic view of the seal structure of the rocket engine turbopump of the present invention;
fig. 3 is a schematic view (side view) of a seal structure of a turbopump of a rocket engine according to the present invention.
In the figure: the centrifugal impeller 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, a second seal ring 11, a rear disk 41, a first seal ring 42, a first auxiliary blade 43, a second auxiliary blade 44, a backflow hole 45, a first arc-shaped part 46, a second arc-shaped part 47 and a third arc-shaped part 48.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present invention more apparent, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and it is apparent that the described embodiments are some embodiments of the present invention, but not all embodiments of the present invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
The invention is described in further detail below with reference to the accompanying drawings.
As shown in fig. 1-3, a sealing structure of a rocket engine turbopump 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 passage, wherein one end of the first shell 1 is connected with the second shell 2 through a connecting piece, the other end is connected with the third shell 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 passage, 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 arranged on the common shaft 6, a mechanical seal 9 is arranged in the first shell 1 and positioned on the periphery of the common shaft 6, the first centrifugal impeller 4 and the second centrifugal impeller 7 are arranged back to back relative to the mechanical seal 9, the first pump with the first centrifugal impeller 4 is used for pumping low-temperature methane (such as-160-170 ℃) or low-temperature liquid oxygen (such as-180-185 ℃), and the second pump with the second centrifugal impeller 7 is used for pumping low-temperature methane (such as-160-170 ℃) or low-temperature liquid oxygen (such as-180-185 ℃).
The first centrifugal impeller 4 and/or the second centrifugal impeller 7 includes a rear disk 41, a front disk, a plurality of blades circumferentially distributed between the rear disk 41 and the front disk, a first seal ring 42 is provided on the outside of the rear disk 41 and at the radially outer end of the rear disk 41, a second seal ring 11 is provided on one end of the first casing 1, a seal pair is formed between the first seal ring 42 and the second seal ring 11, a plurality of auxiliary blades (43, 44) are provided on the outside of the rear disk 41 and at the radially inner end of the first seal ring 42, the plurality of auxiliary blades (43, 44) are uniformly distributed in the circumferential direction, and a backflow space/backflow passage is formed between the plurality of auxiliary blades and the first seal ring 42 in the radial direction, a plurality of backflow holes 45 are provided on the rear disk 41 and at the radially inner end of the first seal ring 42, and the backflow space/backflow passage communicates with the backflow holes 45.
Further, the auxiliary blades (43, 44) include a plurality of first auxiliary blades 43, 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, specifically, the height of the second auxiliary blades 44 is 0.4 to 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, specifically, one second auxiliary blade 44 is provided between every two first auxiliary blades 43.
According to the sealing structure of the rocket engine turbopump, through the design of the auxiliary blades (43, 44), a backflow space/backflow channel is formed between the auxiliary blades and the first sealing ring 42 in the radial direction, a plurality of backflow holes 45 are formed in the rear disc 41 and located at the radial inner end of the first sealing ring 42, the backflow space/backflow channel is communicated with the backflow holes 45 and is used for guiding a part/most of leakage flow leaking from the sealing pairs (42, 11) to the backflow space and guiding the leakage flow from the backflow space to the backflow holes 45 so as to realize backflow, and therefore leakage flowing to the mechanical seal 9 is reduced, and compared with the traditional labyrinth sealing effect, the overall performance and 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 arcuate portion 46, the first arcuate portion 46 being a generally semicircular arcuate portion. The radially outer end of the second auxiliary vane 44 has a second arcuate portion 47, and the second arcuate portion 47 is a substantially 1/4 circular arcuate portion.
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 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 space is a generally circular return space.
According to the sealing structure of the rocket engine turbopump, disclosed by the invention, leakage can be promoted to flow to the inner flow channel of the backflow space through the design of the round backflow space, and is guided to the backflow hole 45 from the backflow space so as to realize backflow, so that leakage of the leakage to the mechanical seal 9 is reduced, and the overall performance and efficiency of the turbopump are improved.
Further, the first auxiliary blade 43 has a mounting angle of 40 to 75 °, the second auxiliary blade 44 has a mounting angle of 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 rocket engine turbopump, through the design of the auxiliary blades (43, 44), a backflow space/backflow channel is formed between the auxiliary blades and the first sealing ring 42 in the radial direction, a plurality of backflow holes 45 are formed in the rear disc 41 and located at the radial inner end of the first sealing ring 42, the backflow space/backflow channel is communicated with the backflow holes 45 and is used for guiding a part/most of leakage flow leaking from the sealing pairs (42, 11) to the backflow space and guiding the leakage flow from the backflow space to the backflow holes 45 so as to realize backflow, and therefore leakage flowing to the mechanical seal 9 is reduced, and compared with the traditional labyrinth sealing effect, the overall performance and efficiency of the turbopump are improved. By the design of the circular backflow space, the leakage flow direction can be promoted to flow into the backflow space inner flow channel, and the leakage flow is guided to the backflow hole 45 from the backflow space, so that backflow is realized, leakage of the leakage flow direction to the mechanical seal 9 is reduced, and the overall performance and efficiency of the turbine pump are improved.
The above-described embodiments are illustrative of the present invention and are not intended to be limiting, and it is to be understood that various changes, modifications, substitutions and alterations can be made to these embodiments without departing from the spirit and 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.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111083022.5A CN113775560B (en) | 2021-09-15 | 2021-09-15 | Sealing structure of rocket engine turbine pump |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111083022.5A CN113775560B (en) | 2021-09-15 | 2021-09-15 | Sealing structure of rocket engine turbine pump |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113775560A CN113775560A (en) | 2021-12-10 |
| CN113775560B true CN113775560B (en) | 2023-12-19 |
Family
ID=78844386
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111083022.5A Active CN113775560B (en) | 2021-09-15 | 2021-09-15 | Sealing structure of rocket engine turbine pump |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113775560B (en) |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5156534A (en) * | 1990-09-04 | 1992-10-20 | United Technologies Corporation | Rotary machine having back to back turbines |
| JP2009024572A (en) * | 2007-07-19 | 2009-02-05 | Hitachi Appliances Inc | Electric blower and vacuum cleaner equipped with the same |
| CN102589057A (en) * | 2011-01-10 | 2012-07-18 | Lg电子株式会社 | Outdoor unit for air conditioner |
| JP2013189916A (en) * | 2012-03-13 | 2013-09-26 | Mitsubishi Heavy Ind Ltd | Pump, and mechanism for suppressing interference of pump leakage flow |
| CN204783705U (en) * | 2015-08-06 | 2015-11-18 | 安徽卧龙泵阀股份有限公司 | There is not magnetic force actuation of leakage self priming pump |
| CN112012957A (en) * | 2020-09-24 | 2020-12-01 | 江西省子轩科技有限公司 | A compressor for industrial production |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6439843B1 (en) * | 2000-11-16 | 2002-08-27 | Ametek, Inc. | Motor/fan assembly having a radial diffuser bypass |
-
2021
- 2021-09-15 CN CN202111083022.5A patent/CN113775560B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5156534A (en) * | 1990-09-04 | 1992-10-20 | United Technologies Corporation | Rotary machine having back to back turbines |
| JP2009024572A (en) * | 2007-07-19 | 2009-02-05 | Hitachi Appliances Inc | Electric blower and vacuum cleaner equipped with the same |
| CN102589057A (en) * | 2011-01-10 | 2012-07-18 | Lg电子株式会社 | Outdoor unit for air conditioner |
| JP2013189916A (en) * | 2012-03-13 | 2013-09-26 | Mitsubishi Heavy Ind Ltd | Pump, and mechanism for suppressing interference of pump leakage flow |
| CN204783705U (en) * | 2015-08-06 | 2015-11-18 | 安徽卧龙泵阀股份有限公司 | There is not magnetic force actuation of leakage self priming pump |
| CN112012957A (en) * | 2020-09-24 | 2020-12-01 | 江西省子轩科技有限公司 | A compressor for industrial production |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113775560A (en) | 2021-12-10 |
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