CN103003529A - Method and system for controlling the clearance at the blade tips of a turbine rotor - Google Patents
Method and system for controlling the clearance at the blade tips of a turbine rotor Download PDFInfo
- Publication number
- CN103003529A CN103003529A CN2011800275441A CN201180027544A CN103003529A CN 103003529 A CN103003529 A CN 103003529A CN 2011800275441 A CN2011800275441 A CN 2011800275441A CN 201180027544 A CN201180027544 A CN 201180027544A CN 103003529 A CN103003529 A CN 103003529A
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- Prior art keywords
- valve
- stage
- turbine
- engine
- air
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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
- 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
-
- 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
- F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
- F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
- F01D11/20—Actively adjusting tip-clearance
- F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
Abstract
The invention relates to a method for controlling the clearance (38) between the tips of mobile blades of a turbine rotor of an aeroplane gas-turbine engine and a turbine shroud of an outer casing surrounding the blades, the method consisting of controlling, according to the operating speed of the engine, a valve arranged in an air conduit opening at a compressor stage of the engine and leading into a control housing arranged around the outer surface of the turbine shroud and supplied with air coming only from said compressor stage. The valve is opened in order to cool the turbine shroud during a high-speed operating phase (TO+CL) which corresponds to the take-off and climb phases of an aeroplane propelled by the engine and during a nominal-speed phase (CR) following the high-speed phase and corresponding to the cruise phase of the aeroplane. The invention also relates to a system for implementing such a method.
Description
Technical field
The present invention relates to the common field for the turbine wheel of aircraft gas-turbine engine.It more specifically relates on the one hand, and the tip of the moving blade of turbine rotor and on the other hand, surrounds the control in the gap between the turbine cover of shell of blade.
Background technique
In order to improve the performance of turbine, will be present in turbine bucket tip and reduce to as far as possible minimum around the gap between the cover of blade is a known practice.This blade tip clearance depends on the dimensional changes between rotating part (consisting of dish and the blade of vane rotor) and the standing part (shell comprises as its a part of turbine cover).These dimensional changes are all because of thermal source (interrelating with the temperature variation of blade, dish and housing) and mechanical sources (concrete centrifugal force with being applied on the turbine rotor interrelates).
For making this gap minimum, relying on active control system is known practice.These systems turn round by guiding on the outer surface of turbine cover from the cool air of the fan of a compressor and/or turbine engine usually.The cool air that is sent on the outer surface of turbine cover has the effect of the outer surface of cooling turbine cover, to limit its thermal expansion.Such ACTIVE CONTROL is controlled by the full powers control system (or FADEC) of for example turbine engine, and is determined by its different running levels.
Document EP 1,860,281 have described an example of active control system, wherein from air cooling turbine cover in the cruising flight phase process of turbine engine fan.Yet, such system has many shortcomings, for example it takies than large space in the cabin of turbine engine, strongly relies on the effect of its Aerodynamic Heating condition within being present in nacelle, and the performance loss that interrelates with outflow from the air-flow of the fan that does not participate in providing thrust.
Two different phases that another active control system is included in the compressor of turbine engine flow out air, and the transmission of regulating the stream of each these outflow, guide to the temperature of the mixture on the turbine cover outer surface with control.Although such system is effective, its valve that shows employing complexity and large volume is to regulate the shortcoming of cooling blast.Particularly, for the situation that is applied to less turbine engine, using such valve be not very desirable aspect quality (mass) and the cost.
Summary of the invention
Therefore, main purpose of the present invention is to overcome above-mentioned shortcoming, and a kind of minimum ACTIVE CONTROL scheme that requires aspect quality and cost is provided.
The method in the gap between the tip of the moving blade of the turbine rotor of this purpose by being used for control aircraft gas-turbine engine and the turbine cover of blade shell on every side realizes, the method comprises the running speed according to engine, control is arranged in the compressor stage of leading to this engine and introduces the valve of an air duct that is positioned near the control room the turbine cover outer surface, and described control room is supplied with only from the air of described compressor stage.According to the present invention, corresponding to the aircraft that is advanced by engine take off and the phase process that runs up of ramp-up period in and in corresponding to the rated velocity phase process behind the high speed stage of the cruising phase of described aircraft, this valve is opened with the turbine cover of cooled enclosure.
Relatively, the invention provides a kind of vane tip of the turbine rotor for control aircraft gas-turbine engine and around the system in the gap between the turbine cover of the frame of blade, this system comprises an air duct, this air duct is designed to open in the compressor stage of engine, and lead to a control room, the outer surface around turbine cover is orientated in this control room as, and be supplied the air that is flowed out by described compressor stage with only, one is arranged in the valve of described air duct, with a circuit, this circuit can be controlled described valve, with corresponding to by aircraft that engine was advanced take off and the phase process that runs up of ramp-up period in and in corresponding to the rated velocity phase process behind the high speed stage of the cruising phase of described aircraft, open it.
At high speed stage, it means the speed stage greater than the rated velocity stage of turbine engine.In an airplane turbine engine, the rated velocity stage is the flight cruise stage, will select this stage in the most of the time of flight, and high speed stage is the stage that is higher than this flight cruise stage, is used in particular for taking off and ramp-up period of aircraft.
Unusual part of the present invention particularly in, it uses an independent air tap at the compressor place, it ensures enough pressure differences and is sent to turbine cover (this control room only show single with unique air supplies) to guarantee cool air.In addition, this air that flows out at the compressor place only is sent in this control room, is not supplied to any other parts of this engine.And, when this valve is closed, there is not air really to flow out from compressor, this limits the loss of pressure head in it.Tracheae in the engine and aeroembolism can be reduced to minimum by this way, and use the simplest possible valve (aspect structure and control).The low-cost control system that consequently has less quality.
Preferably, this valve cuts out after the stage and in corresponding to the flight idle phase process near the stage before the aircraft landing in rated velocity.
Equally preferably, this valve cuts out before the stage and in the ground idling phase process corresponding to the aircraft taxi stage before taking off in rated velocity.
This idling stage of turbo machine is one to be lower than the level in turbo machine rated velocity stage.In aircraft gas-turbine engine, idling stage thereby be the stage that is lower than the flight cruise stage.
Advantageously, the air outer surface that is sent to turbine cover reduces in the transfer process between the stage gradually at high speed stage and rated velocity.In the situation of variable position valve, this kind that air transmits decrescence can obtain by closing gradually this valve.In the situation of on-off valve, what this air transmitted decrescence can obtain by the opening and closing stage that changes this valve.
The present invention also provides a kind of aircraft gas-turbine engine with clearance control system that the front limits.
Description of drawings
With reference to accompanying drawing, will present by following description other features and advantages of the present invention, it not is determinate embodiments of the invention that described accompanying drawing illustrates.Wherein:
Fig. 1 is the schematic longitdinal cross-section diagram that is equipped with according to the gas-turbine aeronautical engine of control system of the present invention;
Fig. 2 is the enlarged view of engine among Fig. 1, specifically shows its high-pressure turbine;
Fig. 3 shows a suite line, and described plotted curve is illustrated in the corresponding variation that of running level in the gas-turbine aeronautical engine changes rotor and the radial dimension of stator; And
Fig. 4 A-4C demonstration expression is used for the curve according to the example of the control of control system one embodiment's of the present invention on-off valve.
Embodiment
Fig. 1 schematically shows the biaxial type turbojet engine 10 of this bypass, and the present invention is applied to the type especially.Certainly, the present invention also is not limited to the gas-turbine aeronautical engine of this special type.
As everyone knows, this turbojet engine 10 with longitudinal axis X-X specifically comprises a fan 12, and this fan is delivered to airflow in sprue 14 neutralization and the coaxial secondary fluid course 16 of this sprue.Be from upstream to the downstream along the air current flow direction of passing sprue 14, this sprue 14 comprises low pressure compressor 18, high pressure compressor 20, firing chamber 22, high-pressure turbine 24 and low-pressure turbine 26.
Show more accurately that in Fig. 2 the high-pressure turbine of turbojet engine comprises a rotor, this rotor comprises dish 28, at this dish 28 a plurality of moving blades 30 is installed, and described blade is arranged in sprue 14.This rotor is surrounded by turbine shroud 32, and this turbine shroud 32 comprises turbine cover 34, and turbine cover 34 is carried by outer turbine shroud 36 by mounting bracket 37.
According to the present invention, a system is provided, it can come control gap 38 by the inner diameter that reduces outer turbine shroud 36 in controlled mode.
For this reason, a control room 40 is arranged on around the turbine shroud 36.This chamber utilizes an air duct 42 to receive cool air, and this air duct 42 at its upstream end (for example utilization ventilating hole known so not shown in the diagram itself) leads in the passage of the main flow at a level place of high pressure compressor 20.Particularly, this control room is only supplied with air by this single tap at the compressor place (other air-sources of this chamber of unavailability).
The cool air of circulation in air duct 42 (utilizing a plurality of through holes on control room 40 walls for example) flows out on the outer turbine shroud 36 fully, makes its cooling, thereby reduces its inner diameter.Particularly, the air that flows out in the high pressure compressor level does not offer any other parts except this control room.
As shown in fig. 1, valve 44 is arranged in the air duct 42.This valve is by 46 controls of full powers control system (or FADEC) of the turbojet engine of the running level that depends on this turbojet engine.
By the valve 44 of control as the function of the different mission phases of aircraft, thus the inner diameter of turbine shroud 36 can in this task process, change outside-also so change gap between the tip of blade 30 of the inner diameter of described turbine cover 34-control turbine cover and High Pressure Turbine Rotor.
Fig. 3 shows the variation in the typical mission process intermediate gap 38 of aircraft that obtains by control system according to the present invention and method.
In this figure, shown different curves, it is the rotational speed of the high-pressure shaft of curve 100 these turbojet engines of diagram, the outer diameter of curve 200 diagram High Pressure Turbine Rotors (dish 28 and blade 30), curve 300 illustrates the inner diameter of the stator (outer turbine shroud 36 and turbine cover 23) of the high-pressure turbine of being controlled by control system according to the present invention, and curve 300a (dotted line) diagram is without the inner diameter of the stator under the control.
These different curves show according to the different phase of the running of the turbojet engine of expression one typical mission, described different phase is namely: ground idling stage GI (corresponding to the coast period of the front aircraft that takes off), follow by high speed stage TO+CL (corresponding to taking off and ramp-up period of aircraft), follow by rated velocity stage CR (corresponding to the cruising phase of aircraft), follow by flight idle stage F I (corresponding to approaching of aircraft before landing), following by deboost stage REV (corresponding to the braking at ground plane), is another ground idling stage GI subsequently.
Shown in curve 100, it should be noted that high speed stage TO+CL occurs under than the higher speed of the rated velocity (CR stage) of turbojet engine.The idling stage (ground and flight) occurs under than the lower speed of the rated velocity of turbojet engine, and flight idle stage F I has the speed of the speed that is lower than equally ground idling stage GI.Should also be noted that rated velocity stage CR is used in the major part process of this task.
As follows to the control of valve 44 according to the present invention:
-in ground idling stage GI process, this valve cuts out, and the inner diameter of stator roughly remains unchanged.In the translate phase process between stage, this valve still cuts out, free expansion under the impact of the hot air of this stator in the passage of main flow in GI stage and TO+CL.In this same transitions phase process, it should be noted that under the impact of centrifugal force, rotor begins mechanically to expand.
-in high speed TO+CL phase process, valve 44 is opened, and this cools off this stator and therefore reduces its inner diameter.Described gap is less, is significantly reduced with the situation contrast that lacks control.Consequently significantly increase in this stage performance.Should notice more accurately that the opening narrow through occuring later of this valve is namely in case occur when arriving transition point between the thermal expansion of mechanical swelling stage of rotor and rotor.
-in rated velocity stage CR process, valve 44 stays open to cool off this stator, thus obtain a less gap, and this performance for engine is beneficial.
It should be noted that terminal in the TO+CL stage, in the process of rated velocity stage CR conversion, air reduces gradually to the transmission of stator.Should also be noted that in the CR phase process this identical air transmits and depends on flying height and can be greater or lesser.The distinct methods that the transmission of acquisition air reduces will be described in conjunction with Fig. 4 hereinafter in more detail.
-in flight idle stage F I process, valve 44 cuts out again, so that freely expand under the impact of the hot air of stator in flowing in the passage of main flow.Before aircraft landing near in the phase process, open in this gap, to prepare to require the aircraft fortuitous event of (thereby recovering at a high speed) that again takes off.
-last, in the process of deboost stage REV and ground idling stage GI, valve 44 keeps cutting out.
Can use different valve arrangements, to realize such gap control.This valve 44 can be to transmit controlled type (under FADEC control), and this air that is conducive to control to stator transmits, and is particularly terminal and at CR in the stage in the TO+CL stage.
Yet, be in the reason of cost and reliability, it is favourable adopting the valve of dibit pattern.In order to obtain to change the opening and closing stage of this valve to adopting air that such valve carries out towards the adjusting of stator transmission.
Fig. 4 A shows the different transmission that the control of available this on-off valve type obtains with 4C.Shown square-wave signal in these figure, its y coordinate represents the position (the 0=valve is opened, and the 1=valve cuts out) of ripple, and abscissa represents time t.Curve C a-Cc represents to depend on the different opening times of valve and transmitted by the average air that this valve is supplied: valve (respectively opening the cycle) is opened longlyer, the average air transmission higher (and on the contrary) of then being supplied by this valve.
In this way, be appreciated that on the one hand that by the frequency of opening of operating valve, on the other hand, the wheel by operating valve turns the opening/closing ratio, can obtain air towards the variation of the average transmission of stator.
Different dibit pattern valve arrangements is well known to those skilled in the art, does not therefore describe at this.Preferably, can select electrically-controlled valve, it will be maintained in its closed position (thereby guaranteeing that in the situation that control was lost efficacy valve keeps cutting out) in situation that lacks electric power supply.
Claims (10)
1. the method in the gap (38) between the turbine cover (34) of the tip of the moving blade (30) of a turbine rotor that is used for control aircraft gas-turbine engine and blade shell (36) on every side, the method comprises the running speed according to engine, control is arranged in compressor (20) level of leading to this engine and the valve (44) of introducing an air duct (42) that is arranged near the control room (40) of turbine cover outer surface, described control room is supplied with only from the air of described compressor stage, it is characterized in that, corresponding to the aircraft that is advanced by described engine take off and the phase process that runs up of ramp-up period in and in corresponding to the rated velocity phase process behind the high speed stage of the cruising phase of described aircraft, this valve is opened to cool off the turbine cover (34) of described shell (36).
2. the method for claim 1, wherein said valve cuts out after the stage and in corresponding to the flight idle phase process near the stage before the aircraft landing in described rated velocity.
3. method as claimed in claim 1 or 2, wherein said valve cuts out before the stage and in the ground idling phase process corresponding to the coast period before taking off in described rated velocity.
4. such as any one described method among the claim 1-3, its Air reduces gradually towards being transmitted in high speed stage and the transfer process of rated velocity between the stage of described turbine cover outer surface.
5. method as claimed in claim 4, wherein said valve is the adjustable position valve, described transfer process Air towards the transmission of this turbine cover outer surface reduce gradually obtain by closing gradually described valve.
6. method as claimed in claim 4, wherein said valve is an on-off valve, described transfer process Air towards the transmission of this turbine cover outer surface reduce gradually obtain by the opening and closing stage that changes this valve.
7. the system in the gap (38) between the turbine cover (34) of the tip of the moving blade (30) of a turbine rotor that is used for control aircraft gas-turbine engine and the frame (36) that centers on described blade, this system comprises:
One air duct (42), compressor (20) level that this air duct is designed at this engine is opened, and lead to a control room (40), this control room is around the outer surface of described turbine cover, and is designed to be supplied with only from the air of described compressor stage;
One is arranged in the valve (44) of described air duct; With
One circuit, this circuit can be controlled described valve, with corresponding to by aircraft that this engine was advanced take off and the high speed stage process of ramp-up period in and in corresponding to the rated velocity phase process behind the high speed stage of the cruising phase of described aircraft, this valve is opened.
8. system as claimed in claim 7, wherein said valve is the adjustable position valve.
9. system as claimed in claim 7, wherein said valve is on-off valve.
10. one kind comprises the aircraft gas-turbine engine such as any one described clearance control system among the claim 7-9.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| FR1054366 | 2010-06-03 | ||
| FR1054366A FR2960905B1 (en) | 2010-06-03 | 2010-06-03 | METHOD AND SYSTEM FOR CONTROLLING TURBINE ROTOR BLACK SUMP |
| PCT/FR2011/051261 WO2011151602A1 (en) | 2010-06-03 | 2011-06-01 | Method and system for controlling the clearance at the blade tips of a turbine rotor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103003529A true CN103003529A (en) | 2013-03-27 |
| CN103003529B CN103003529B (en) | 2015-09-30 |
Family
ID=43471088
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201180027544.1A Expired - Fee Related CN103003529B (en) | 2010-06-03 | 2011-06-01 | For controlling the method and system in the gap at the vane tip place of turbine rotor |
Country Status (8)
| Country | Link |
|---|---|
| US (1) | US20130177414A1 (en) |
| EP (1) | EP2576994A1 (en) |
| CN (1) | CN103003529B (en) |
| BR (1) | BR112012030635A2 (en) |
| CA (1) | CA2801193A1 (en) |
| FR (1) | FR2960905B1 (en) |
| RU (1) | RU2566510C2 (en) |
| WO (1) | WO2011151602A1 (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104963729A (en) * | 2015-07-09 | 2015-10-07 | 中国航空工业集团公司沈阳发动机设计研究所 | Heavy-duty gas turbine high-vortex tip clearance control structure |
| CN107120146A (en) * | 2016-02-25 | 2017-09-01 | 通用电气公司 | Active HPC clearance controls |
| CN110318823A (en) * | 2019-07-10 | 2019-10-11 | 中国航发沈阳发动机研究所 | Active clearance control method and device |
| CN111255572A (en) * | 2018-12-03 | 2020-06-09 | 劳斯莱斯有限公司 | Method and apparatus for controlling at least a portion of a start or re-ignition process of a gas turbine engine |
| CN112081664A (en) * | 2019-06-12 | 2020-12-15 | 劳斯莱斯有限公司 | Improving the deceleration of a gas turbine |
| CN112211731A (en) * | 2019-07-12 | 2021-01-12 | 劳斯莱斯有限公司 | Gas turbine engine generator |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2997443B1 (en) | 2012-10-31 | 2015-05-15 | Snecma | CONTROL UNIT AND METHOD FOR CONTROLLING THE AUBES TOP SET |
| US9266618B2 (en) * | 2013-11-18 | 2016-02-23 | Honeywell International Inc. | Gas turbine engine turbine blade tip active clearance control system and method |
| US10344614B2 (en) | 2016-04-12 | 2019-07-09 | United Technologies Corporation | Active clearance control for a turbine and case |
| FR3105980B1 (en) * | 2020-01-08 | 2022-01-07 | Safran Aircraft Engines | METHOD AND CONTROL UNIT FOR CONTROLLING THE GAME OF A HIGH PRESSURE TURBINE FOR REDUCING THE EGT OVERRIDE EFFECT |
| US11788425B2 (en) * | 2021-11-05 | 2023-10-17 | General Electric Company | Gas turbine engine with clearance control system |
| US11808157B1 (en) * | 2022-07-13 | 2023-11-07 | General Electric Company | Variable flowpath casings for blade tip clearance control |
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| GB2388407A (en) * | 2002-05-10 | 2003-11-12 | Rolls Royce Plc | Gas turbine blade tip clearance control structure |
| US20050109016A1 (en) * | 2003-11-21 | 2005-05-26 | Richard Ullyott | Turbine tip clearance control system |
| CN1664318A (en) * | 2004-03-04 | 2005-09-07 | Snecma发动机公司 | Axial maintenance device to support the strut of a stator ring of the high-pressure turbine of a turbomachine |
| EP1798381A2 (en) * | 2005-12-16 | 2007-06-20 | General Electric Company | Thermal control of gas turbine engine rings for active clearance control |
| EP2025878A2 (en) * | 2007-08-03 | 2009-02-18 | General Electric Company | Aircraft gas turbine engine blade tip clearance control |
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| RU2006593C1 (en) * | 1991-07-01 | 1994-01-30 | Иван Анатольевич Черняев | Method of control of radial clearance between rotor blade tips and housing of turbomachine of gas-turbine engine |
| RU2175410C1 (en) * | 2000-04-18 | 2001-10-27 | Открытое акционерное общество "Авиадвигатель" | Gas-turbine engine compressor |
| GB2363864B (en) * | 2000-06-23 | 2004-08-18 | Rolls Royce Plc | A control arrangement |
| US7431557B2 (en) * | 2006-05-25 | 2008-10-07 | General Electric Company | Compensating for blade tip clearance deterioration in active clearance control |
| US7686569B2 (en) * | 2006-12-04 | 2010-03-30 | Siemens Energy, Inc. | Blade clearance system for a turbine engine |
| US20090053042A1 (en) * | 2007-08-22 | 2009-02-26 | General Electric Company | Method and apparatus for clearance control of turbine blade tip |
-
2010
- 2010-06-03 FR FR1054366A patent/FR2960905B1/en active Active
-
2011
- 2011-06-01 RU RU2012157775/06A patent/RU2566510C2/en not_active IP Right Cessation
- 2011-06-01 BR BR112012030635A patent/BR112012030635A2/en not_active IP Right Cessation
- 2011-06-01 EP EP11728349.9A patent/EP2576994A1/en not_active Withdrawn
- 2011-06-01 CN CN201180027544.1A patent/CN103003529B/en not_active Expired - Fee Related
- 2011-06-01 CA CA2801193A patent/CA2801193A1/en not_active Abandoned
- 2011-06-01 US US13/701,700 patent/US20130177414A1/en not_active Abandoned
- 2011-06-01 WO PCT/FR2011/051261 patent/WO2011151602A1/en active Application Filing
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2388407A (en) * | 2002-05-10 | 2003-11-12 | Rolls Royce Plc | Gas turbine blade tip clearance control structure |
| US20050109016A1 (en) * | 2003-11-21 | 2005-05-26 | Richard Ullyott | Turbine tip clearance control system |
| CN1664318A (en) * | 2004-03-04 | 2005-09-07 | Snecma发动机公司 | Axial maintenance device to support the strut of a stator ring of the high-pressure turbine of a turbomachine |
| EP1798381A2 (en) * | 2005-12-16 | 2007-06-20 | General Electric Company | Thermal control of gas turbine engine rings for active clearance control |
| EP2025878A2 (en) * | 2007-08-03 | 2009-02-18 | General Electric Company | Aircraft gas turbine engine blade tip clearance control |
Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104963729A (en) * | 2015-07-09 | 2015-10-07 | 中国航空工业集团公司沈阳发动机设计研究所 | Heavy-duty gas turbine high-vortex tip clearance control structure |
| CN107120146A (en) * | 2016-02-25 | 2017-09-01 | 通用电气公司 | Active HPC clearance controls |
| CN111255572A (en) * | 2018-12-03 | 2020-06-09 | 劳斯莱斯有限公司 | Method and apparatus for controlling at least a portion of a start or re-ignition process of a gas turbine engine |
| CN111255572B (en) * | 2018-12-03 | 2023-08-11 | 劳斯莱斯有限公司 | Method and apparatus for controlling at least a portion of a start-up or re-ignition process of a gas turbine engine |
| CN112081664A (en) * | 2019-06-12 | 2020-12-15 | 劳斯莱斯有限公司 | Improving the deceleration of a gas turbine |
| CN112081664B (en) * | 2019-06-12 | 2023-05-05 | 劳斯莱斯有限公司 | Improving deceleration of gas turbines |
| CN110318823A (en) * | 2019-07-10 | 2019-10-11 | 中国航发沈阳发动机研究所 | Active clearance control method and device |
| CN112211731A (en) * | 2019-07-12 | 2021-01-12 | 劳斯莱斯有限公司 | Gas turbine engine generator |
Also Published As
| Publication number | Publication date |
|---|---|
| RU2012157775A (en) | 2014-07-20 |
| FR2960905A1 (en) | 2011-12-09 |
| FR2960905B1 (en) | 2014-05-09 |
| WO2011151602A1 (en) | 2011-12-08 |
| BR112012030635A2 (en) | 2016-08-16 |
| EP2576994A1 (en) | 2013-04-10 |
| US20130177414A1 (en) | 2013-07-11 |
| CN103003529B (en) | 2015-09-30 |
| RU2566510C2 (en) | 2015-10-27 |
| CA2801193A1 (en) | 2011-12-08 |
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