CN114651132B - Bleed air device in a turbojet compressor rotor - Google Patents
Bleed air device in a turbojet compressor rotor Download PDFInfo
- Publication number
- CN114651132B CN114651132B CN202080062706.4A CN202080062706A CN114651132B CN 114651132 B CN114651132 B CN 114651132B CN 202080062706 A CN202080062706 A CN 202080062706A CN 114651132 B CN114651132 B CN 114651132B
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- CN
- China
- Prior art keywords
- bleed
- bleed air
- angle
- slot
- axis
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- 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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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
-
- 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
Landscapes
- 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 device in a turbojet compressor rotor, consisting of bleed air slots and bleed air ducts for guiding cooling air into a turbine, wherein the bleed air ducts have lips, forming an upper chamfer and a lower chamfer at an angle γ and an angle δ, respectively, with respect to the engine axis and equal to 30 °.60 °. The front impeller and the rear impeller are provided with thrust shoulders, and the bleed air pipe is arranged on the lip in 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 slot length L to slot width E is equal to 2..2.5, 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 air tube axis is β, wherein the ratio of the angle α to the angle β is equal to 1..2. The bleed air pipe is mounted in the upper and lower openings of the intermediate ring, and the intermediate ring is positioned on the lip of the bleed air pipe between the front and rear impeller thrust shoulders. 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 double bleed ducts attached to a tailored flange are known (U.S. Pat. No.5275534, IPC F01D11/00, F01D5/06, F02C7/18, published 1/4 1994). A disadvantage of the known construction is the short length of the bleed air duct, the complex flange connection and the presence of the retaining ring.
In technology essentially closest to the claimed invention is the gas turbine engine compressor rotor (russian patent No.2386864, IPC F04D 29/32, published 4/20 2010) used as a prototype comprising front and rear impellers, a cooled bleed air device inside the rotor consisting of bleed air slots, bleed air ducts for introducing cooled air into the turbine; the inside of the air guide pipe is provided with a lip part, an upper beveling part and a lower beveling part; the front impeller and the rear impeller are provided with thrust shoulders; the bleed air duct is mounted on the thrust shoulder in a radial direction, and bleed air grooves are provided on the duct for bleed air from the pipeline.
The drawbacks of the known structure used as prototype are as follows: the bleed efficiency is low due to the high air pressure loss on the pipes for bleed air from the pipeline and the stress points in the openings for mounting the bleed air pipes in the upper and lower lips of the disk, and the manufacturability of the structure is poor due to the necessity to form these lips for the connection of the bleed air pipes.
The technical problem that could not be solved using the prototype, only by implementing the claimed invention, is that the efficiency of the bleed air device is low due to the high air pressure loss on the pipes used for bleed air from the pipeline during the flow diversion through the bleed air channel.
Disclosure of Invention
The technical problem addressed by the claimed invention is to increase the operating efficiency of bleed air devices in turbojet compressor rotors by reducing turbine cooling air pressure losses.
The technical problem is solved in the following way: the bleed air device of the turbojet compressor rotor is composed of bleed air tanks, bleed air pipes for introducing cooling air into the turbine; wherein the bleed duct has a lip, an upper chamfer and a lower chamfer; thrust shoulders are arranged on the front impeller and the rear impeller; the bleed duct is mounted on the thrust shoulder in the radial direction, a bleed slot is provided above the duct for bleed air from the pipeline, according to the invention the bleed slot is formed in the rear impeller flange as a rectangle with rounded corners, the ratio of slot length L to slot width E is equal to 2.5, and 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.
In the claimed invention, compared to the prototype, rectangular bleed grooves with rounded corners are used, wherein the ratio of the groove length L to the groove width E is equal to 2..2.5 and is placed with an angle α between the groove axis and the engine axis and an angle β between the groove axis and the bleed tube axis, wherein the ratio of the angle α to the angle β is equal to 1..2, the bleed tube having upper and lower bevelled portions formed with angles γ and δ with respect to the engine axis and equal to 30..60 °, respectively, the bleed tube being mounted in upper and lower openings of an intermediate ring for fastening the bleed tube in the compressor rotor, allowing to increase the working efficiency of the bleed device in the high pressure compressor and to eliminate stress points in the rotor disc caused by the placement of the bleed device, due to reduced pressure losses of air for turbine cooling in the compressor rotor.
The reduction of the ratio of the slot length L to the slot width E below 2 causes an increase in the flow velocity in the slot and an increase in the air pressure loss in the bleed air device. Increasing the ratio of slot length L to slot width E above 2.5 can compromise the disk fatigue life of the bleed slot location.
The ratio of the angle α to the angle β and the 30 °.60 ° range of angles γ, δ becomes smaller or larger, which aggravates the intake flow of the bleed air device and affects the parameters of the sucked air in terms of flow, temperature and pressure.
In the claimed invention, the intermediate ring for the bleed air duct fastened in the compressor rotor allows to eliminate stress points in the disk, to improve its manufacturability and to improve the working efficiency of the bleed air device in the turbojet compressor rotor, compared to the prototype.
Drawings
Fig. 1 shows a longitudinal section of a turbojet high-pressure compressor rotor.
Fig. 2 is an enlarged view of the bleed air device.
Figure 3 shows a bleed slot.
Figure 4 shows a bleed slot.
Fig. 5 shows the intermediate ring (bottom view).
Fig. 6 shows the intermediate ring (top view).
Detailed Description
The turbojet high pressure compressor rotor (fig. 1) comprises a flanged impeller 2 with a line bleed slot 9 with a profile matching the departure angle of the impeller 1, an intermediate ring 10, a bleed air duct 11, the impeller 3, a labyrinth 4 downstream of the high pressure compressor and a duct 8 (fig. 1 and 2). The bleed duct 11 is mounted in a radial direction with respect to the axis K of the engine. The intermediate ring 10 is mounted with an axial and radial press fit between the thrust shoulder 12 of the impeller 1 and the thrust shoulder 13 of the impeller 2. The bleed ducts 11 are mounted on the thrust shoulders 12 and 13 using intermediate rings 10. The gas-guide tube 11 is press-fitted into the intermediate ring 10 and is fixed in radial displacement by a lip 14 on the gas-guide tube 11. The bleed air pipe 11 is located in the same plane as the pipeline bleed air groove 9. The air induction pipe 11 has a chamfer 15 on the upper end for the intake of air drawn in from the compressor system (not shown in the figures) and a chamfer 16 on the lower end for guiding air between the hubs 5, 6 and 7 of the impeller 2 and 3 and the labyrinth 4 and pipe 8 downstream of the high pressure compressor and into the turbine (not shown in the figures). The bleed slot 9 in the flange of the impeller 2 is arranged rectangular and has rounded corners 17 at the corners and is positioned such that the angle alpha between the axis of the slot 9 pi and the axis of the engine K and the angle beta between the axis of the bleed slot 9 pi and the axis of the bleed duct 11 t, and the bleed duct 11 is provided with a lower chamfer 16 and an upper chamfer 15 formed respectively at angles gamma and delta with respect to the axis of the engine of about 30 deg. 60 deg., reducing the pressure loss of the air sucked into the compressor rotor for turbine cooling, thus improving the working efficiency of the bleed device.
The ratio of the length L to the width E of the bleed slot 9 in the range 2.5 and the angle α between the axis ii of the bleed slot 9 and the axis t of the bleed duct 11 in the range 1.2 and the angles γ and δ between the upper and lower bevelled portions 14 and 15 of the bleed duct 11 and the axis k of the engine in the range 30.60 are chosen on the basis of the conditions of the parameters of the turbine cooling air required for the highest operating efficiency of the bleed air device. The device works as follows. During operation of the engine, air is drawn from the compressor system (not numbered) via the bleed slot 9 and is first directed to the upper chamfer 15 of the bleed duct 11, then to the lower chamfer 16 via the duct 11, then to the hubs 5, 6, 7 of the impellers 2, 3 and beneath the labyrinth 4 and duct 8 downstream of the high pressure compressor to provide a suction for entering the rotor (not shown) towards the turbine (not shown) for cooling the turbine.
Based on successful test results within the pilot gas generator, bleed air devices in the high pressure compressor rotor of the claimed structure have been implemented in the baseline high pressure compressor structure of the gas turbine engine.
The claimed invention with the above-mentioned distinguishing features, in combination with the known features, thus allows to increase the operating efficiency of the bleed air device in exchange for reducing the pressure loss of the air for turbine cooling in the turbojet engine high-pressure compressor rotor.
Claims (1)
1. A bleed air device in a turbojet compressor rotor, comprising a first impeller (1) and a second impeller (2), the second impeller (2) having a flange, a bleed air groove (9), a bleed air duct (11) for guiding cooling air into the turbine, wherein the bleed air duct (11) is formed with a lip (14) and an upper chamfer (15) and a lower chamfer (16) at both ends, respectively; -a first thrust shoulder (12) is provided on the first impeller (1), -a second thrust shoulder (13) is provided on the second impeller (2), -the bleed air duct (11) is mounted on the first and second thrust shoulders (12, 13) in a radial direction, -the bleed air slot (9) is located above the bleed air duct (11), characterized in that the bleed air device further comprises an intermediate ring (10), the intermediate ring (10) being located between the first and second thrust shoulders (12, 13) and on a lip (14) of the bleed air duct (11), and-the bleed air duct (11) is mounted in the intermediate ring; wherein the bleed slot (9) is formed in the flange of the second impeller (2) and has a rectangular shape with rounded corners (17) at the corners, the ratio of the bleed slot length L to the slot width E being in the range of 2-2.5 and an angle α between the axis of the bleed slot (9) and the axis (K) of the engine and an angle β between the axis (ii) of the bleed slot and the axis (T) of the bleed tube (11), wherein the ratio of the angle α to the angle β is in the range of 1-2, the upper chamfer (15) and the lower chamfer (16) of the bleed tube (11) being formed at an angle γ and an angle δ, respectively, in the range of 30 ° -60 ° with respect to the axis (K) of the engine.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| RU2019128008A RU2728550C1 (en) | 2019-09-05 | 2019-09-05 | Air bleeder in rotor of turbojet compressor |
| RU2019128008 | 2019-09-05 | ||
| 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 CN114651132A (en) | 2022-06-21 |
| CN114651132B true 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) |
Family Cites Families (8)
| 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 |
| US7870742B2 (en) * | 2006-11-10 | 2011-01-18 | General Electric Company | Interstage cooled turbine engine |
| RU2386864C1 (en) * | 2008-10-27 | 2010-04-20 | Открытое акционерное общество "Авиадвигатель" | Gas turbine engine compressor rotor |
| RU2451840C2 (en) * | 2010-06-21 | 2012-05-27 | Открытое акционерное общество "Авиадвигатель" | Compressor rotor of gas-turbine engine |
| US9677472B2 (en) * | 2012-10-08 | 2017-06-13 | 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 |
| 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 WO PCT/RU2020/000454 patent/WO2021045645A1/en active Application Filing
- 2020-08-27 CN CN202080062706.4A patent/CN114651132B/en active Active
Also Published As
| Publication number | Publication date |
|---|---|
| CN114651132A (en) | 2022-06-21 |
| RU2728550C1 (en) | 2020-07-31 |
| RU2019128008A3 (en) | 2020-03-17 |
| RU2019128008A (en) | 2020-03-18 |
| WO2021045645A1 (en) | 2021-03-11 |
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