CN114651132A - Bleed air device in a turbojet compressor rotor - Google Patents
Bleed air device in a turbojet compressor rotor Download PDFInfo
- Publication number
- CN114651132A CN114651132A CN202080062706.4A CN202080062706A CN114651132A CN 114651132 A CN114651132 A CN 114651132A CN 202080062706 A CN202080062706 A CN 202080062706A CN 114651132 A CN114651132 A CN 114651132A
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- Prior art keywords
- bleed
- bleed air
- air
- angle
- axis
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
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- 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
- F04D29/58—Cooling; Heating; Diminishing heat transfer
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Physics & Mathematics (AREA)
- Thermal Sciences (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
Bleed air arrangement in a turbojet compressor rotor, consisting of a bleed air duct for guiding cooling air into the turbine and a bleed air groove, wherein the bleed air duct has a lip, an upper chamfer and a lower chamfer being formed at an angle γ and an angle δ, respectively, relative to the engine axis and equal to 30 ° … 60 °. The front impeller and the rear impeller are provided with thrust shoulders, and the air guide pipe is arranged on the lip along the radial direction. The bleed slot bleed air provided in the flange of the rear impeller is rectangular with rounded corners, wherein the ratio of the slot length L to the slot width E is equal to 2 … 2.5.5 and the slot is configured such that the angle between the slot axis and the engine axis is α and the angle between the slot axis and the bleed duct axis is β, wherein the ratio of the angle α to the angle β is equal to 1 … 2. The bleed ducts are mounted in upper and lower openings of an intermediate ring which is located on a lip of the bleed duct between the thrust shoulders of the front and rear impellers. The result is improved efficiency, manufacturability, and fatigue life.
Description
Technical Field
The present invention relates to aircraft engine manufacturing, and in particular to a high-pressure compressor rotor for a turbojet engine.
Background
Gas turbine engine compressor rotors with dual bleed air ducts attached to tailored flanges are known (U.S. Pat. No.5275534, IPC F01D11/00, F01D5/06, F02C7/18, published on 1, 4, 1994). Disadvantages of the known construction are the short length of the bleed air duct, the complicated flange connection and the presence of the collar.
The closest in technical substance to the claimed invention is the gas turbine engine compressor rotor (russian patent No.2386864, IPC F04D 29/32, published on 20.4.2010) used as a prototype comprising a front and rear impeller, a cooling bleed air device inside the rotor consisting of a bleed air duct, a bleed air duct which introduces cooling air into the turbine; the air guide pipe is internally provided with a lip part, an upper inclined cutting part and a lower inclined cutting part; thrust shoulders are arranged on the front impeller and the rear impeller; the bleed air duct is mounted on the thrust shoulder in a radial direction and the bleed air groove is provided in the duct for bleeding air from the pipeline.
The disadvantages of the known structure used as prototype are as follows: the low efficiency of the bleed air results from the loss of high air pressure diverted through the air inlet slots in the rim junction of the impeller over the tubes used to bleed air from the pipeline and the stress points for mounting the bleed air tubes in the openings in the upper and lower lips of the disk, and the poor manufacturability of the structure from having to form these lips for the connection of the bleed air tubes.
A technical problem that cannot be solved using prototypes, but only by implementing the claimed invention, is the low efficiency of the bleed air arrangement due to the high air pressure losses on the tubes used for bleed air from the pipeline during the flow diversion through the bleed air duct.
Disclosure of Invention
The technical problem solved by the claimed invention is to improve the operating efficiency of bleed air devices in the turbojet compressor rotor by reducing the turbine cooling air pressure losses.
This technical problem is solved in the following manner: the air-entraining device of the turbojet engine compressor rotor consists of an air-entraining groove and an air-entraining pipe for introducing cooling air into the turbine; wherein the bleed pipe has a lip, an upper chamfer and a lower chamfer; the front impeller and the rear impeller are provided with thrust shoulders; according to the invention, the bleed air groove is formed in the rear impeller flange as a rectangle with rounded corners, the ratio of the groove length L to the groove width E is equal to 2.. 2.5, and the angle between the groove axis and the engine axis is alpha, and the angle between the groove axis and the bleed air pipe axis is beta, wherein the ratio of the angle alpha to the angle beta is equal to 1.. 2, the bleed air pipe has an upper chamfer and a lower chamfer formed respectively at an angle gamma and delta relative to the engine axis and equal to 30 °.60 °, the bleed air pipe being mounted in the upper and lower openings of an intermediate ring located on the lip of the bleed air pipe between the thrust shoulders of the front and rear impellers.
In the claimed invention, compared to the prototype, a rectangular bleed air duct with rounded corners is used, in which the ratio of the duct length L to the duct width E is equal to 2.. 2.5 and is placed at an angle α between the duct axis and the engine axis and an angle β between the duct axis and the bleed air duct axis, in which the ratio of angle α to angle β is equal to 1.. 2, the bleed air duct having upper and lower chamfers formed at angles γ and δ, respectively, equal to 30 °.60 ° relative to the engine axis, the bleed air duct being mounted in upper and lower openings of an intermediate ring for fastening the bleed air duct in the compressor rotor, allowing to increase the operating efficiency of the bleed air apparatus in the high-pressure compressor due to the reduced pressure loss of air sucked into the compressor rotor for turbine cooling and to eliminate stress points in the rotor disc caused by the bleed air apparatus placement.
A reduction of the ratio of the slot length L to the slot width E below 2 leads to an increased flow velocity in the slot and to an increased air pressure loss in the bleed air device. Increasing the ratio of the slot length L to the slot width E above 2.5 will impair the fatigue life of the disc at the bleed slot location.
A smaller or larger range of the angles α to β and of the angles γ, δ of 30 °.60 ° exacerbates the intake air flow of the bleed air device and influences the parameters of the sucked-in air in terms of flow, temperature and pressure.
In the claimed invention, the intermediate ring for fastening the bleed air duct in the compressor rotor makes it possible, in comparison with a prototype, to eliminate stress points in the disk, to improve its manufacturability and to improve the operating efficiency of the bleed air arrangement in the turbojet compressor rotor.
Drawings
Fig. 1 shows a longitudinal section of a turbojet high-pressure compressor rotor.
Figure 2 is an enlarged view of the bleed air arrangement.
Figure 3 shows an air entrainment tank.
Figure 4 shows an air entrainment tank.
Fig. 5 shows the intermediate ring (bottom view).
Fig. 6 shows an intermediate ring (top view).
Detailed Description
The turbojet high-pressure compressor rotor (fig. 1) comprises a flanged impeller 2 with a pipeline bleed slot 9 contoured to match the departure angle of the impeller 1, an intermediate ring 10, a bleed duct 11, an impeller 3, a labyrinth 4 downstream of the high-pressure compressor, and a duct 8 (fig. 1 and 2). The bleed ducts 11 are mounted in a radial direction with respect to the axis K of the engine. The intermediate ring 10 is mounted between the thrust shoulder 12 of the impeller 1 and the thrust shoulder 13 of the impeller 2 with an axial and radial press fit. The bleed ducts 11 are mounted on a thrust shoulder 12 and a thrust shoulder 13 using an intermediate ring 10. The bleed duct 11 is press-fitted into the intermediate ring 10 and secured in radial displacement by a lip 14 on the bleed duct 11. The bleed air pipe 11 and the pipeline bleed air groove 9 are located in the same plane. The bleed air duct 11 has a chamfer 15 on the upper end for the intake of air sucked in from the compressor system (not shown in the figures) and a chamfer 16 on the lower end for guiding the air between the hubs 5, 6 and 7 of the impellers 2 and 3 and the labyrinth 4 and duct 8 downstream of the high-pressure compressor and into the turbine (not shown in the figures). The provision of the bleed grooves 9 in the flange of the impeller 2 rectangular and with rounded corners 17 at the corners and positioned at the angle α between the axis "pi" of the grooves 9 and the axis "K" of the engine and at the angle β between the axis "pi" of the bleed grooves 9 and the axis "tt" of the bleed duct 11, and the provision of the bleed duct 11 with the undercut 16 and the upper chamfer 15, respectively formed at the angles γ and δ, relative to the axis "K" of the engine, equal to 30 °.60 °, reduces the pressure loss of the air sucked into the compressor rotor for the turbine cooling and thus improves the operating efficiency of the bleed apparatus.
The ratio of the length L to the width E of the bleed air channel 9 in the range of 2.. 2.5 and the angle α between the axis pi of the bleed air channel 9 and the axis tt of the bleed air duct 11 in the range of 1.. 2 and the angles γ and δ between the upper and lower bevels 14 and 15 of the bleed air duct 11 and the axis "k" of the engine in the range of 30.. 60 ° are selected on the basis of the conditions of the parameters of the turbine cooling air required for maximum operating efficiency of the bleed air arrangement. The device operates in the following manner. During engine operation, air is drawn from the compressor system (not numbered) through the bleed air slots 9 and is first directed to the upper chamfer 15 of the bleed air duct 11, then to the lower chamfer 16 via the duct 11, and then to the hubs 5, 6, 7 of the impellers 2, 3 and to the labyrinth 4 downstream of the high pressure compressor and below the duct 8 to provide a cavity for suction into the rotor (not shown) towards the turbine (not shown) for cooling the turbine.
The bleed air arrangement in the high-pressure compressor rotor of the claimed structure has been implemented in a baseline high-pressure compressor configuration of a gas turbine engine, based on successful test results in a pilot gas generator.
The claimed invention, having the above-mentioned distinctive features, in combination with the known features, thus allows to improve the operating efficiency of the bleed air device in exchange for reducing the pressure loss of the air sucked into the high-pressure compressor rotor of the turbojet for turbine cooling.
Claims (1)
1. A bleed air arrangement in a turbojet compressor rotor comprising an impeller and a flanged impeller, the impeller consisting of a bleed air duct, a bleed air slot, for directing cooling air into the turbine, wherein the bleed air duct is formed at an end with a lip, an upper chamfer and a lower chamfer; the impeller and the impeller are provided with a thrust shoulder and a thrust shoulder, the air-entraining pipe is arranged on the thrust shoulder and the thrust shoulder along the radial direction, the air-entraining groove is positioned above the air-entraining pipe, and the air-entraining device is characterized by further comprising an intermediate ring which is positioned on a lip of the air-entraining pipe between the thrust shoulder and the thrust shoulder of the impeller and the impeller, and the air-entraining pipe is arranged in the intermediate ring; wherein the bleed air groove is formed in a flange of the impeller and is rectangular in shape with rounded corners at the corners, the ratio of the bleed air groove length L to the groove width E being equal to 2.. 2.5, and an angle α between the axis of the bleed air groove and the axis of the engine, and an angle β between the axis of the bleed air groove and the axis of the bleed air duct, wherein the ratio of the angle α to the angle β is equal to 1.. 2, the upper and lower chamfer of the bleed air duct being formed at an angle γ and an angle δ, respectively, equal to 30 °.60 ° relative to the axis of the engine.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| RU2019128008 | 2019-09-05 | ||
| RU2019128008A RU2728550C1 (en) | 2019-09-05 | 2019-09-05 | Air bleeder in rotor of turbojet compressor |
| PCT/RU2020/000454 WO2021045645A1 (en) | 2019-09-05 | 2020-08-27 | Device for bleeding air in the rotor of a turbojet engine compressor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN114651132A true CN114651132A (en) | 2022-06-21 |
| CN114651132B CN114651132B (en) | 2023-07-18 |
Family
ID=70478572
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202080062706.4A Active CN114651132B (en) | 2019-09-05 | 2020-08-27 | Bleed air device in a turbojet compressor rotor |
Country Status (3)
| Country | Link |
|---|---|
| CN (1) | CN114651132B (en) |
| RU (1) | RU2728550C1 (en) |
| WO (1) | WO2021045645A1 (en) |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20080112793A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Interstage cooled turbine engine |
| RU2386864C1 (en) * | 2008-10-27 | 2010-04-20 | Открытое акционерное общество "Авиадвигатель" | Gas turbine engine compressor rotor |
| RU2010125468A (en) * | 2010-06-21 | 2011-12-27 | Открытое акционерное общество "Авиадвигатель" (RU) | GAS TURBINE ENGINE COMPRESSOR ROTOR |
| CN203097955U (en) * | 2012-12-24 | 2013-07-31 | 中航商用航空发动机有限责任公司 | Air guiding assembly of gas turbine engine |
| US20140096536A1 (en) * | 2012-10-08 | 2014-04-10 | United Technologies Corporation | Bleed air slot |
Family Cites Families (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2834758B1 (en) * | 2002-01-17 | 2004-04-02 | Snecma Moteurs | DEVICE FOR STRAIGHTENING THE SUPPLY AIR OF A CENTRIPETE SAMPLING IN A COMPRESSOR |
| DE10310815A1 (en) * | 2003-03-12 | 2004-09-23 | Rolls-Royce Deutschland Ltd & Co Kg | Vortex rectifier in tubular design with retaining ring |
| RU189794U1 (en) * | 2017-08-29 | 2019-06-04 | Акционерное общество "Объединенная двигателестроительная корпорация" (АО "ОДК") | ROTOR COMPRESSOR GAS TURBINE ENGINE |
-
2019
- 2019-09-05 RU RU2019128008A patent/RU2728550C1/en active
-
2020
- 2020-08-27 CN CN202080062706.4A patent/CN114651132B/en active Active
- 2020-08-27 WO PCT/RU2020/000454 patent/WO2021045645A1/en active Application Filing
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20080112793A1 (en) * | 2006-11-10 | 2008-05-15 | General Electric Company | Interstage cooled turbine engine |
| RU2386864C1 (en) * | 2008-10-27 | 2010-04-20 | Открытое акционерное общество "Авиадвигатель" | Gas turbine engine compressor rotor |
| RU2010125468A (en) * | 2010-06-21 | 2011-12-27 | Открытое акционерное общество "Авиадвигатель" (RU) | GAS TURBINE ENGINE COMPRESSOR ROTOR |
| US20140096536A1 (en) * | 2012-10-08 | 2014-04-10 | United Technologies Corporation | Bleed air slot |
| CN203097955U (en) * | 2012-12-24 | 2013-07-31 | 中航商用航空发动机有限责任公司 | Air guiding assembly of gas turbine engine |
Also Published As
| Publication number | Publication date |
|---|---|
| WO2021045645A1 (en) | 2021-03-11 |
| CN114651132B (en) | 2023-07-18 |
| RU2019128008A3 (en) | 2020-03-17 |
| RU2019128008A (en) | 2020-03-18 |
| RU2728550C1 (en) | 2020-07-31 |
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