CN109630209A - A kind of band is prewhirled the turbine disk chamber seal structure of bleed - Google Patents
A kind of band is prewhirled the turbine disk chamber seal structure of bleed Download PDFInfo
- Publication number
- CN109630209A CN109630209A CN201811503703.0A CN201811503703A CN109630209A CN 109630209 A CN109630209 A CN 109630209A CN 201811503703 A CN201811503703 A CN 201811503703A CN 109630209 A CN109630209 A CN 109630209A
- Authority
- CN
- China
- Prior art keywords
- bleed
- seal structure
- chamber seal
- disk chamber
- turbine
- 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.)
- Pending
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Classifications
-
- 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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
-
- 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
Abstract
It prewhirls the turbine disk chamber seal structure of bleed the invention discloses a kind of band, several bleed flow paths of prewhirling are provided on inside support ring 1, bleed flow path of prewhirling successively is made of bleed hole 2, air collecting chamber 3 and venthole 4, wherein, bleed hole 2 is several circumferentially uniformly distributed circular holes, is set on the end face of inside support ring air inlet outer ring；Air collecting chamber 3 is cavity structure；The angular range of venthole 4 and engine centerline is 65 to 80 degree.After this patent disk chamber seal structure, movable vane root works in design conditions, and movable vane channel secondary flow loss substantially reduces.Compared with existing disk chamber seal structure, turbine efficiency can be made to improve nearly 1% after taking this patent.This patent structure is simple, can largely borrow existing disk chamber seal structure, and economy and realizability are good.
Description
Technical field:
This patent is related to gas-turbine unit technical field, is that turbine part disk chamber therein obturages knot specifically
Structure, and in particular to a kind of band is prewhirled the turbine disk chamber seal structure of bleed.
Background technique
The working principle of gas-turbine unit is generated after being mixed and burned using the compressed air and fuel of compressor
High-temperature high-pressure fuel gas drives turbine high speed rotation, and turbine drives compressor further through turbine wheel shaft, to form continuous running.Whirlpool
Wheel is made of the movable vane (forming rotor part) of static guide vane (forming stator part) and rotation.For ensure turbine rotor can
By operating, avoids turning touching mill between stator, turning all the presence of certain gap between the stator turbine disk.Guide vane and movable vane root it
Between axially and radially gap presence, may cause turbine channel high-temperature fuel gas by the gap and invade turbine disk chamber, to make
It is excessively high at turbine disk temperature and influence the work safety and service life of the turbine disk.Therefore high pressure usually is introduced in turbine disk chamber
The cold air of blower outlet prevents combustion gas from invading turbine disk chamber, while also cooling down to the turbine disk as gas is obturaged.Although
Introducing obturage cold air can prevent mainstream combustion gas invade ablation disk chamber, but due to enter mainstream channel obturage cold air flow velocity compared with
Low, air-flow prewhirl angle is smaller (exporting prewhirl angle compared to guide vane root), and obturaging cold air can fire obturaging seam outlet with mainstream
Gas carries out strong blending, and loss is caused to increase, and turbine efficiency reduces.As secondary speed, turbine inlet temperature and efficiency are wanted
The raising asked turns to obturage the decline of turbine performance caused by the low flow velocity and low prewhirl angle of cold air between stator increasingly apparent.Therefore,
Disk chamber seal structure needing to turn stator to turbine optimizes, to meet the requirement of turbine performance raising.
A kind of existing typical turbomachine disk chamber seal structure is as shown in Figure 1, the structure is connected by A1 inside support ring, A2 bolt
It connects, A3 front flow guiding plate, A4 front apron, A5 front apron air inlet, A6 obturage comb tooth, prop up in A7 honeycomb, A8 preswirl nozzle, A9
Pushing out ring air inlet composition.Wherein A1 inside support ring is connect by snap ring with high-pressure turbine guide vane, using bolt and A5 front flow guiding plate
Connection, A4 front apron are connect using fast fastening structure with turbine rotor blade.The working principle of the turbine disk chamber seal structure is: from
The cold air introduced before compressor enters air collecting chamber by the A9 inside support ring air inlet on A1 inside support ring, by A8 preswirl nozzle
Decompression accelerate, be divided into three strands of cold air；Blast of cold air directly passes through A5 front apron air inlet, and to enter movable vane air collecting chamber cooling dynamic
Leaf；Second strand of cold air passes downwardly through that A6 obturages comb tooth and the gap of obturaging of A7 honeycomb composition enters the cooling whirlpool in turbine disk root
Wheel disc；Third stock cold air obturages the gap of obturaging of comb tooth and A7 honeycomb composition upwardly through A6 at two, and by turning between stator
Seam of obturaging enter mainstream channel, realize cooling to guide vane and movable vane root, and prevent the invasion of mainstream high-temperature fuel gas.
The defect of existing turbine disk chamber seal structure is: in Fig. 1, third stock cold air obturages comb tooth and honeycomb group by two-stage
At seal structure after, the pitot loss of cold air is larger, and the prewhirl angle of cooling rates and cold air is all lower, this just with guide vane root
There are great differences for portion's mainstream combustion gas, and radial direction can be had with the high-velocity fluid of mainstream by causing to obturage after cold air enters mainstream channel
With circumferential blending, and the mainstream gas into movable vane root is made very big negative angle of attack state occur, causes the work of movable vane root non-
Design point operating condition, this will increase the risk of movable vane root air-flow separation and enhances the intensity of channel secondary flow loss, finally makes
The flow losses of turbine channel increase, and turbine efficiency reduces, performance decline.For high-efficiency turbine, this influence is especially bright
It is aobvious.
Summary of the invention:
Goal of the invention:
The shortcomings that for presently used turbine disk chamber seal structure, this patent propose a kind of band and prewhirl the novel whirlpool of bleed
Wheel disc chamber seal structure, the purpose is to increase obturage seam outlet cold air speed and prewhirl angle, thus reduce obturage cold air with
The blending of mainstream improves the efficiency of turbine.
Technical solution:
A kind of band is prewhirled the turbine disk chamber seal structure of bleed, which is characterized in that is provided with several on inside support ring 1
It prewhirls bleed flow path, bleed flow path of prewhirling successively is made of bleed hole 2, air collecting chamber 3 and venthole 4, wherein bleed hole 2 is week
To several uniformly distributed circular holes, it is set on the end face of inside support ring air inlet outer ring；Air collecting chamber 3 is cavity structure；Venthole 4
Angular range with engine centerline is 65 to 80 degree.
3 cross-sectional area of air collecting chamber is constantly reduced from bleed hole 2 on the direction of venthole 4.
It is interconnected between the air collecting chamber 3 of several bleed flow paths of prewhirling.
The angle of the venthole 4 and engine centerline is 75 degree.
Beneficial effect
A) using prewhirl bleed flow path after, obturage cold air outlet air velocity and prewhirl angle increase, can reduce with
Guide vane root exports the speed triangle difference of mainstream combustion gas, guarantees the work of movable vane root in design conditions.
B) using after this patent disk chamber seal structure, movable vane root works in design conditions, movable vane channel secondary flow loss
Substantially reduce.Compared with existing disk chamber seal structure, turbine efficiency can be made to improve nearly 1% after taking this patent.
C) this patent structure is simple, can largely borrow existing disk chamber seal structure, and economy and realizability are good.
Detailed description of the invention
Fig. 1 is existing disk chamber seal structure figure.
The disk chamber seal structure figure of Fig. 2 this patent.
The structural representation of Fig. 3 this patent inside support ring 1 and bleed flow circuit diagram.
Fig. 4 a and Fig. 4 b are that current velocity Huo existing speed triangle and this patent speed triangle compare.
Wherein, inside support ring 1, bleed hole 2, air collecting chamber 3 and venthole 4；S- guide vane, R- movable vane, V- mainstream absolute velocity,
U- rotation speed, W- mainstream relative velocity, Vseal- obturage absolute velocity, and Wseal- obturages relative velocity.
Specific embodiment
Technical solution of the present invention is described in detail with reference to the accompanying drawings of the specification.
The novel turbine disk chamber seal structure that this patent proposes is as shown in Figure 2.This patent remains existing disk chamber and obturages knot
A2 in structure is bolted, A3 front flow guiding plate, A4 front apron, A5 front apron air inlet, A6 obturage comb tooth, A7 honeycomb, A8
Preswirl nozzle and A9 inside support ring air inlet, and the basic structure of A1 inside support ring is remained, it is only pre- in A7 honeycomb and A8
Rotation upper end of nozzle increases bleed flow path.As shown in Fig. 2, the difference of new building chamber seal structure and existing structure is only that band is prewhirled
The inside support ring 1 of bleed.Band prewhirl bleed inside support ring 1 detailed construction it is as shown in Figure 3.It is increased prewhirl bleed flow path by
Bleed hole 2, air collecting chamber 3 and venthole 4 form.Wherein, bleed hole 2 is several circumferentially uniformly distributed circular holes；Air collecting chamber 3 is sky
Chamber, and cross-sectional area from bleed hole 2 on the direction of venthole 4 constantly reduce is formed shrink channel, and difference air collecting chamber 3 it
Between can be connected to, or be tapered annular entirety cavity；Venthole 4 is circumferential uniformly several inclined holes, with engine centerline at 75 °
Angle, cross-sectional view is as shown in the B-B of Fig. 3.
The turbine disk chamber seal structure working principle of this patent is: the compressor cold air of introducing divides two-way to flow into seal pan
Chamber, first via cold air are entered by the inside support ring bleed hole A9, and flow path is consistent with the existing disk chamber seal structure of description, preceding
Two strands of cold air cool down movable vane and the turbine disk respectively, and third stock cold air cools down guide vane and movable vane root and prevents mainstream high-temperature fuel gas
Invasion；Second road cold air by bleed hole 2 enter air collecting chamber 3, by shrinkage type channel realize accelerate, then by with engine centerline
4 high velocity jet of venthole of 75 ° of angles goes out, and realizes the acceleration of air-flow and increases prewhirl angle.The high speed cold that the strand is prewhirled
Gas cold air of prewhirling low with the third stock low speed in flow path one is blended, make outflow turn stator obturage seam obturage cold air flow velocity and
Prewhirl angle all increases, and is close to or up to the flow velocity and prewhirl angle of the outlet of guide vane root, reduces the blending of the two, makes movable vane root
The air-flow of portion's import reaches movable vane design point, to reduce secondary flow loss, reaches and increases turbine acting ability, improve turbine
The purpose of efficiency.This patent largely remains existing turbine disk chamber seal structure, and the prewhirl inside support ring of bleed of 1 band is adopted
It is processed with no surplus precision casting technology, guarantees to obtain accurately bleed hole, air collecting chamber and air outlet hole structure, remaining part uses
Existing manufacturing process completes the process.
After flow path using bleed of prewhirling, the air velocity and prewhirl angle for obturaging cold air outlet increase, and can reduce and lead
Blade root exports the speed triangle difference of mainstream combustion gas, guarantees the work of movable vane root in design conditions, such as Fig. 4 a and Fig. 4 b institute
Show.Existing disk chamber seal structure obturage cold air outlet absolute velocity size and Orientation all with guide vane outlet mainstream it is widely different,
In the case where movable vane revolving speed is certain, lead to the relative velocity difference for obturaging cold air relative velocity Yu mainstream combustion gas of movable vane import
Also big, to influence movable vane operating condition.After this patent disk chamber seal structure, cold air outlet absolute velocity size and Orientation is obturaged
With mainstream combustion gas difference very little, thus guarantee obturage cold air with respect to movable vane import relative velocity size and Orientation and mainstream difference
Very little can reduce mixing loss in this way, guarantee the work of movable vane root in design conditions.
Claims (4)
- The turbine disk chamber seal structure of bleed 1. a kind of band is prewhirled, which is characterized in that be provided with several on inside support ring (1) It prewhirls bleed flow path, bleed flow path of prewhirling successively is made of bleed hole (2), air collecting chamber (3) and venthole (4), wherein bleed hole (2) it is several circumferentially uniformly distributed circular holes, is set on the end face of inside support ring air inlet outer ring；Air collecting chamber (3) is cavity knot Structure；The angular range of venthole (4) and engine centerline is 65 to 80 degree.
- The turbine disk chamber seal structure of bleed 2. a kind of band according to claim 1 is prewhirled, which is characterized in that the air collecting chamber (3) cross-sectional area is constantly reduced from bleed hole (2) on the direction of venthole (4).
- The turbine disk chamber seal structure of bleed 3. a kind of band according to claim 1 is prewhirled, which is characterized in that described several Prewhirl bleed flow path air collecting chamber (3) between interconnect.
- 4. prewhirling the turbine disk chamber seal structure of bleed according to a kind of band of claim 1, which is characterized in that the venthole (4) Angle with engine centerline is 75 degree.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811503703.0A CN109630209A (en) | 2018-12-10 | 2018-12-10 | A kind of band is prewhirled the turbine disk chamber seal structure of bleed |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201811503703.0A CN109630209A (en) | 2018-12-10 | 2018-12-10 | A kind of band is prewhirled the turbine disk chamber seal structure of bleed |
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| CN109630209A true CN109630209A (en) | 2019-04-16 |
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| CN201811503703.0A Pending CN109630209A (en) | 2018-12-10 | 2018-12-10 | A kind of band is prewhirled the turbine disk chamber seal structure of bleed |
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Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111441828A (en) * | 2020-03-12 | 2020-07-24 | 中国科学院工程热物理研究所 | Engine turbine disc cavity structure with prewhirl nozzle and flow guide disc |
| CN112483193A (en) * | 2020-11-27 | 2021-03-12 | 北京化工大学 | Turbine damping disc edge structure capable of inhibiting Helmholtz resonance gas invasion |
| CN112855283A (en) * | 2021-01-11 | 2021-05-28 | 中国科学院工程热物理研究所 | Engine prerotation system capable of improving receiving hole flow coefficient |
| CN114961871A (en) * | 2021-02-24 | 2022-08-30 | 中国航发商用航空发动机有限责任公司 | Turbine and aircraft engine |
| CN114961869A (en) * | 2021-02-24 | 2022-08-30 | 中国航发商用航空发动机有限责任公司 | Rim sealing system and aeroengine |
| CN115853598A (en) * | 2022-11-29 | 2023-03-28 | 中国航空发动机研究院 | Turbine blade air conditioning supercharging impeller with axial air intake and pre-rotation supercharging air supply system |
| CN116220913A (en) * | 2023-05-08 | 2023-06-06 | 中国航发四川燃气涡轮研究院 | Low-loss engine pre-rotation air supply system |
| CN116537895A (en) * | 2023-07-04 | 2023-08-04 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
| CN116822081A (en) * | 2023-06-26 | 2023-09-29 | 中国航发沈阳发动机研究所 | Intermediate stage bleed air design method for high-load fan |
| CN117090643A (en) * | 2023-10-20 | 2023-11-21 | 中国航发沈阳发动机研究所 | Turbine rotor blade air feed structure of forced cooling |
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| EP1988260A2 (en) * | 2007-05-01 | 2008-11-05 | General Electric Company | Method and system for regulating a cooling fluid within a turbomachine in real time |
| US20120027575A1 (en) * | 2010-07-29 | 2012-02-02 | Rolls-Royce Plc | Labyrinth seal |
| CN105264175A (en) * | 2013-03-01 | 2016-01-20 | 西门子能源公司 | Active bypass flow control for a seal in a gas turbine engine |
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Cited By (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111441828A (en) * | 2020-03-12 | 2020-07-24 | 中国科学院工程热物理研究所 | Engine turbine disc cavity structure with prewhirl nozzle and flow guide disc |
| CN111441828B (en) * | 2020-03-12 | 2022-09-16 | 中国科学院工程热物理研究所 | Engine turbine disc cavity structure with prewhirl nozzle and flow guide disc |
| CN112483193A (en) * | 2020-11-27 | 2021-03-12 | 北京化工大学 | Turbine damping disc edge structure capable of inhibiting Helmholtz resonance gas invasion |
| CN112855283A (en) * | 2021-01-11 | 2021-05-28 | 中国科学院工程热物理研究所 | Engine prerotation system capable of improving receiving hole flow coefficient |
| CN112855283B (en) * | 2021-01-11 | 2022-05-20 | 中国科学院工程热物理研究所 | Engine prerotation system capable of improving receiving hole flow coefficient |
| CN114961871A (en) * | 2021-02-24 | 2022-08-30 | 中国航发商用航空发动机有限责任公司 | Turbine and aircraft engine |
| CN114961869A (en) * | 2021-02-24 | 2022-08-30 | 中国航发商用航空发动机有限责任公司 | Rim sealing system and aeroengine |
| CN115853598B (en) * | 2022-11-29 | 2023-09-22 | 中国航空发动机研究院 | Turbine blade cold air supercharging impeller for axial air intake and pre-rotation supercharging air supply system |
| CN115853598A (en) * | 2022-11-29 | 2023-03-28 | 中国航空发动机研究院 | Turbine blade air conditioning supercharging impeller with axial air intake and pre-rotation supercharging air supply system |
| CN116220913A (en) * | 2023-05-08 | 2023-06-06 | 中国航发四川燃气涡轮研究院 | Low-loss engine pre-rotation air supply system |
| CN116220913B (en) * | 2023-05-08 | 2023-08-18 | 中国航发四川燃气涡轮研究院 | Low-loss engine pre-rotation air supply system |
| CN116822081A (en) * | 2023-06-26 | 2023-09-29 | 中国航发沈阳发动机研究所 | Intermediate stage bleed air design method for high-load fan |
| CN116822081B (en) * | 2023-06-26 | 2024-04-09 | 中国航发沈阳发动机研究所 | Intermediate stage bleed air design method for high-load fan |
| CN116537895A (en) * | 2023-07-04 | 2023-08-04 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
| CN116537895B (en) * | 2023-07-04 | 2023-09-15 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
| CN117090643A (en) * | 2023-10-20 | 2023-11-21 | 中国航发沈阳发动机研究所 | Turbine rotor blade air feed structure of forced cooling |
| CN117090643B (en) * | 2023-10-20 | 2024-01-02 | 中国航发沈阳发动机研究所 | Turbine rotor blade air feed structure of forced cooling |
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Application publication date: 20190416 |