CN103003529B - For controlling the method and system in the gap at the vane tip place of turbine rotor - Google Patents
For controlling the method and system in the gap at the vane tip place of turbine rotor Download PDFInfo
- Publication number
- CN103003529B CN103003529B CN201180027544.1A CN201180027544A CN103003529B CN 103003529 B CN103003529 B CN 103003529B CN 201180027544 A CN201180027544 A CN 201180027544A CN 103003529 B CN103003529 B CN 103003529B
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- Prior art keywords
- valve
- stage
- turbine
- engine
- air
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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.)
- Expired - Fee Related
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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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Control Of Turbines (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
The present invention relates to a kind of moving blade for controlling aircraft gas-turbine engine tip and around the frame of described blade turbine cover between the system in gap (38), the method comprises: according to the running speed of engine, control to be positioned at the level of the compressor leading to this engine, and introduce and be arranged in turbine cover proximity and be supplied with only from the valve of the air duct in the control room of the air of described compressor stage.This valve correspond to advanced by described engine aircraft take off and ramp-up period (TO+CL) process that runs up the stage in and correspond to described aircraft cruising phase high speed stage after rated velocity stage (CR) process in, this valve is opened to cool described turbine cover.The invention still further relates to a kind of system implementing the method.
Description
Technical field
The present invention relates to the usual field of the turbine wheel for aircraft gas-turbine engine.It more specifically relates on the one hand, the tip of the moving blade of turbine rotor, with another aspect, and the control in the gap between the turbine cover of the shell of encirclement blade.
Background technique
In order to improve the performance of turbine, by be present in turbine bucket tip and around blade cover between to reduce to minimum be as far as possible a known practice in gap.This blade tip clearance depends on the dimensional changes between rotating part (forming dish and the blade of vane rotor) with standing part (shell comprises the turbine cover as its part).These dimensional changes are all because of thermal source (being associated with the temperature variation of blade, dish and housing) and mechanical sources (being specifically associated with the centrifugal force be applied on turbine rotor).
For making this gap minimum, be known practice by means of active control system.These systems are usually by the outer surface of the cool air of the fan from a compressor and/or turbine engine being guided to turbine cover operates.The cool air be 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 controlled by the full powers control system (or FADEC) of such as turbine engine, and determined by its different operational level.
Document EP 1,860,281 describes an example of active control system, wherein from air cooling turbine cover in cruising flight phase process of turbine engine air blower.But, such system has many shortcomings, such as it takies larger space in the cabin of turbine engine, strong depend-ence its in the effect being present in the Aerodynamic Heating condition in nacelle, and with from the performance loss not participating in providing the outflow of the air-flow of the fan of thrust to be associated.
Two different phases that another active control system is included in the compressor of turbine engine flow out air, and regulate the transmission of each stream that these flow out, to control the temperature of the mixture guided on turbine cover outer surface.Although such system is effective, it shows and adopts the valve of complicated and large volume to regulate the shortcoming of cooling blast.Particularly, for the situation being applied to less turbine engine, such valve is used not to be very desirable in quality (mass) and cost.
Summary of the invention
Therefore, main purpose of the present invention overcomes above-mentioned shortcoming, provides a kind of in quality and cost, require minimum ACTIVE CONTROL scheme.
This object is realized by the method in the gap between the turbine cover of the shell around the tip of the moving blade of the turbine rotor for controlling aircraft gas-turbine engine and blade, the method comprises the running speed according to engine, control to be arranged in the compressor stage of leading to this engine and introduce the valve being positioned at an air duct in a control room of turbine cover proximity, described control room is supplied with only from the air of described compressor stage.According to the present invention, correspond to advanced by engine aircraft take off and ramp-up period the phase process that runs up in and correspond to described aircraft cruising phase high speed stage after rated velocity phase process in, this valve is opened with the turbine cover of cooled enclosure.
Relatively, the invention provides a kind of turbine rotor for controlling aircraft gas-turbine engines vane tip and around the frame of blade turbine cover between the system in gap, 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 with the air only flowed out by described compressor stage, one valve being arranged in described air duct, with a circuit, this circuit can control described valve, with correspond to advanced by engine aircraft take off and ramp-up period the phase process that runs up in and correspond to described aircraft cruising phase high speed stage after rated velocity phase process in open it.
At high speed stage, it means the speed stage in the rated velocity stage being greater than turbine engine.In an airplane turbine engine, the rated velocity stage is the flight cruise stage, and in the most of the time of flying by this stage of selection, and high speed stage is the stage 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 compressor place, and it ensures enough pressure differences to guarantee that cool air is sent to turbine cover (this control room only show single with unique air supplies).In addition, this air flowed out at compressor place is only sent in this control room, is not supplied to any other parts of this engine.And when the valve is closed, do not have air really to flow out from compressor, this limits the loss of pressure head in it.Tracheae in engine and aeroembolism can be reduced to minimum by this way, and use possible the simplest valve (in structure and control).Consequently there is the low cost control system of less quality.
Preferably, closing close in the flight idle phase process in stage before this valve also corresponds to aircraft landing after the rated velocity stage.
Equally preferably, close in the ground idling phase process of this valve before the rated velocity stage and corresponding to the aircraft taxi stage before taking off.
This idling stage of turbo machine is one lower than the level in turbo machine rated velocity stage.In aircraft gas-turbine engines, the idling stage because of but lower than stage in flight cruise stage.
Advantageously, the outer surface that air is sent to turbine cover reduces gradually in transfer process between high speed stage and rated velocity stage.When variable position valve, this kind that air transmits decrescence obtains by closing this valve gradually.When on-off valve, the opening and closing stage decrescence by changing this valve that this air transmits obtains.
The present invention also provide a kind of have before the aircraft gas-turbine engines of clearance control system that limits.
Accompanying drawing explanation
With reference to accompanying drawing, will be presented by following description other features and advantages of the present invention, described accompanying drawing illustrates not determinate embodiments of the invention.Wherein:
Fig. 1 is the schematic longitudinal section figure of the gas-turbine aeronautical engine be equipped with according to control system of the present invention;
Fig. 2 is the enlarged view of engine in Fig. 1, concrete its high-pressure turbine of display;
Fig. 3 shows a suite line, and described plotted curve illustrates that a change rotor of operational level in gas-turbine aeronautical engine changes with the corresponding of radial dimension of stator; And
Fig. 4 A-4C shows the curve representing and be used for according to the example of the control of the on-off valve in control system one embodiment of the present invention.
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 is also not limited thereto the gas-turbine aeronautical engine of special type.
As everyone knows, this turbojet engine 10 with longitudinal axis X-X specifically comprises a fan 12, and airflow is delivered to sprue 14 and neutralized in the secondary fluid course 16 coaxial with this sprue by this fan.Be from upstream to downstream along the air current flow direction through 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 in fig. 2, the high-pressure turbine of turbojet engine comprises a rotor, and this rotor comprises dish 28, and this dish 28 is provided with multiple moving blade 30, and described blade is arranged in sprue 14.This rotor surrounded by turbine shroud 32, and this turbine shroud 32 comprises turbine cover 34, and turbine cover 34 carried by outer turbine shroud 36 by mounting bracket 37.
Turbine cover 34 can be formed by multiple adjacent joint.In inner side, it is equipped with the layer 34a of high-abrasive material, and around the blade 30 of this rotor, leaves the gap 38 of the most advanced and sophisticated 30a with them.
According to the present invention, provide a system, it carrys out control gap 38 by the inner diameter reducing outer turbine shroud 36 in a controlled manner.
For this reason, a control room 40 is arranged on around turbine shroud 36.This room utilizes an air duct 42 to receive cool air, and this air duct 42 at its upstream end (ventilating hole that such as utilization itself is known therefore not shown in the diagram) leads in the passage of the main flow at a level place at high pressure compressor 20.Particularly, this control room is supplied by means of only this single tap at compressor place (other air-sources of this room of unavailability) with air.
In air duct 42, the cool air (utilizing the multiple through holes on such as control room 40 wall) of circulation flows out on outer turbine shroud 36 completely, makes it cool, thus reduces its inner diameter.Particularly, the air flowed out in high pressure compressor level is not supplied to any other parts except this control room.
As shown in fig. 1, valve 44 is arranged in air duct 42.This valve controlled by the full powers control system (or FADEC) 46 of turbojet engine of the operational level depending on this turbojet engine.
By controlling the valve 44 of the function as the different mission phase of aircraft, can change the inner diameter-of outer turbine shroud 36 and therefore change the inner diameter-of described turbine cover 34 thus gap between the tip of the blade 30 of control turbine cover and High Pressure Turbine Rotor in this task process.
The change of the typical mission process intermediate gap 38 at aircraft that Fig. 3 display is obtained by control system according to the present invention and method.
Show different curves in this figure, namely curve 100 illustrates the rotational speed of the high-pressure shaft of this turbojet engine, curve 200 illustrates the outer diameter of High Pressure Turbine Rotor (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 controlled by control system according to the present invention, and curve 300a (dotted line) diagram is without the inner diameter of the stator under 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 (coast period corresponding to aircraft before taking off), follow by high speed stage TO+CL (taking off and ramp-up period corresponding to aircraft), follow by rated velocity stage CR (cruising phase corresponding to aircraft), follow by flight idle stage FI (corresponding to the close of front aircraft of landing), follow by deboost stage REV (corresponding to the braking at ground plane), another ground idling stage GI subsequently.
As shown in curve 100, it should be noted that, high speed stage TO+CL occurs under the speed that the rated velocity (CR stage) than turbojet engine is higher.The idling stage (ground and flight) occurs under the speed lower than the rated velocity of turbojet engine, and flight idle stage FI has equally lower than the speed of the speed of 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 according to the control of the present invention to valve 44:
-in ground idling stage GI process, this valve cuts out, and the inner diameter of stator roughly remains unchanged.In translate phase process between GI stage and TO+CL 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 this same transitions phase process, it should be noted that, under the influence of the centrifugal force, rotor starts mechanically to expand.
-in high speed TO+CL phase process, valve 44 is opened, and this cools this stator and therefore reduces its inner diameter.Described gap is less, and situation about controlling with shortage contrasts and is significantly reduced.Consequently significantly increase in this stage performance.Should notice more accurately, this valve open at narrow point through occurring later, namely occur once the transition point arrived between mechanical swelling stage of rotor and the thermal expansion of rotor.
-in rated velocity stage CR process, valve 44 stays open to cool this stator, thus obtains a less gap, and this performance for engine is beneficial to.
It should be noted that, at TO+CL stage end, in the process changed to rated velocity stage CR, air reduces gradually to the transmission of stator.Should also be noted that in CR phase process, this identical air transmission is depended on flying height and can be greater or lesser.The distinct methods obtaining air transmission reduction will describe by composition graphs 4 hereinafter in more detail.
-in flight idle stage FI process, valve 44 cuts out again, freely expands under making the impact of the hot air of stator in the passage flowing in main flow.Before aircraft landing close in phase process, this gap is opened, with the fortuitous event preparing to require that aircraft takes off again (thus recovering at a high speed).
-last, in the process of deboost stage REV and ground idling stage GI, valve 44 keeps cutting out.
Different valve arrangements can be used, to realize such gap control.This valve 44 can be transmit controlled type (under FADEC controls), and this is conducive to controlling the air transmission to stator, particularly at TO+CL stage end with in the CR stage.
But be in the reason of cost and reliability, the valve adopting dibit pattern is favourable.In order to obtain the adjustment transmitted towards stator the air adopting such valve to carry out, the opening and closing stage of this valve can be changed.
The difference transmission that the control that Fig. 4 A and 4C shows available this on-off valve type obtains.Show square-wave signal in these figures, its y coordinate represents the position (0=valve is opened, and 1=valve cuts out) of ripple, abscissa representing time t.Curve C a-Cc represents the different opening time depending on valve and the average air supplied by this valve transmission: valve (respectively opening the cycle) is opened longer, then the average air supplied by this valve transmission higher (and on the contrary).
In this way, be appreciated that on the one hand, open frequency by operating valve, on the other hand, turn opening/closing ratio by the wheel of operating valve, the change of air towards the average transmission of stator can be obtained.
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 when lacking electric power supply (thus guaranteeing when controlling inefficacy that valve keeps cutting out).
Claims (10)
1. one kind for control the shell (36) around the tip of the moving blade (30) of the turbine rotor of aircraft gas-turbine engine and blade turbine cover (34) between the method in gap (38), the method comprises the running speed according to engine, control to be arranged in compressor (20) level of leading to this engine and also introduce the valve (44) being arranged in an air duct (42) in a control room (40) of turbine cover proximity, described control room is supplied with only from the air of described compressor stage, it is characterized in that, correspond to advanced by described engine aircraft take off and ramp-up period the phase process that runs up in and correspond to described aircraft cruising phase high speed stage after rated velocity phase process in, this valve is opened the turbine cover (34) cooling described shell (36).
2. the method for claim 1, closing close in the flight idle phase process in stage before wherein said valve also corresponds to aircraft landing after the described rated velocity stage.
3. method as claimed in claim 1 or 2, closes in the ground idling phase process of wherein said valve before the described rated velocity stage and corresponding to the coast period before taking off.
4. method as claimed in claim 1 or 2, wherein air reduces gradually in the transfer process being transmitted between high speed stage and rated velocity stage of described turbine cover outer surface.
5. method as claimed in claim 4, wherein said valve is adjustable position valve, and in described transfer process, air is closed described valve gradually towards being reduced by gradually of transmission of this turbine cover outer surface and being obtained.
6. method as claimed in claim 4, wherein said valve is an on-off valve, and in described transfer process, air obtained towards the opening and closing stage changing this valve that is reduced by gradually of the transmission of this turbine cover outer surface.
7. for control the moving blade (30) of the turbine rotor of aircraft gas-turbine engine tip and around the frame (36) of described blade turbine cover (34) between the system in gap (38), this system comprises:
One air duct (42), this air duct is designed to open in compressor (20) level of this engine, and lead to a control room (40), this control room around the outer surface of described turbine cover, and is designed to be supplied with only from the air of described compressor stage;
One valve (44) being arranged in described air duct; With
One circuit, this circuit can control described valve, with correspond to advanced by this engine aircraft take off and ramp-up period high speed stage process in and correspond to described aircraft cruising phase high speed stage after rated velocity phase process in this valve is opened.
8. system as claimed in claim 7, wherein said valve is 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 as the system in claim 7-9 as described in any one.
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 CN103003529A (en) | 2013-03-27 |
| CN103003529B true 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) |
Families Citing this family (14)
| 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 |
| 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 |
| US10138752B2 (en) * | 2016-02-25 | 2018-11-27 | General Electric Company | Active HPC clearance control |
| US10344614B2 (en) | 2016-04-12 | 2019-07-09 | United Technologies Corporation | Active clearance control for a turbine and case |
| GB201819695D0 (en) * | 2018-12-03 | 2019-01-16 | Rolls Royce Plc | Methods and apparatus for controlling at least part of a start-up or re-light process of a gas turbine engine |
| GB2584693A (en) * | 2019-06-12 | 2020-12-16 | Rolls Royce Plc | Improving deceleration of a gas turbine |
| CN110318823B (en) * | 2019-07-10 | 2022-07-15 | 中国航发沈阳发动机研究所 | Active clearance control method and device |
| GB201910008D0 (en) * | 2019-07-12 | 2019-08-28 | Rolls Royce Plc | Gas turbine engine electrical generator |
| 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 |
| US11982189B2 (en) | 2021-06-04 | 2024-05-14 | Rtx Corporation | Warm start control of an active clearance control for a gas turbine engine |
| US11788425B2 (en) * | 2021-11-05 | 2023-10-17 | General Electric Company | Gas turbine engine with clearance control system |
| US12012859B2 (en) | 2022-07-11 | 2024-06-18 | General Electric Company | Variable flowpath casings for blade tip clearance control |
| US11808157B1 (en) * | 2022-07-13 | 2023-11-07 | General Electric Company | Variable flowpath casings for blade tip clearance control |
Citations (4)
| 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 |
| 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 |
Family Cites Families (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| US20050109016A1 (en) * | 2003-11-21 | 2005-05-26 | Richard Ullyott | Turbine tip clearance control system |
| 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 CN CN201180027544.1A patent/CN103003529B/en not_active Expired - Fee Related
- 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
- 2011-06-01 EP EP11728349.9A patent/EP2576994A1/en not_active Withdrawn
- 2011-06-01 RU RU2012157775/06A patent/RU2566510C2/en not_active IP Right Cessation
- 2011-06-01 CA CA2801193A patent/CA2801193A1/en not_active Abandoned
- 2011-06-01 BR BR112012030635A patent/BR112012030635A2/en not_active IP Right Cessation
Patent Citations (4)
| 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 |
| 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 |
Also Published As
| Publication number | Publication date |
|---|---|
| FR2960905B1 (en) | 2014-05-09 |
| EP2576994A1 (en) | 2013-04-10 |
| FR2960905A1 (en) | 2011-12-09 |
| RU2012157775A (en) | 2014-07-20 |
| US20130177414A1 (en) | 2013-07-11 |
| WO2011151602A1 (en) | 2011-12-08 |
| BR112012030635A2 (en) | 2016-08-16 |
| CN103003529A (en) | 2013-03-27 |
| RU2566510C2 (en) | 2015-10-27 |
| CA2801193A1 (en) | 2011-12-08 |
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