GB2129880A - Gas turbine rotor tip clearance control apparatus - Google Patents
Gas turbine rotor tip clearance control apparatus Download PDFInfo
- Publication number
- GB2129880A GB2129880A GB08231924A GB8231924A GB2129880A GB 2129880 A GB2129880 A GB 2129880A GB 08231924 A GB08231924 A GB 08231924A GB 8231924 A GB8231924 A GB 8231924A GB 2129880 A GB2129880 A GB 2129880A
- Authority
- GB
- United Kingdom
- Prior art keywords
- shroud ring
- control apparatus
- ring
- rotor
- tip clearance
- 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.)
- Withdrawn
Links
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
- F01D11/08—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
- F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A stationary shroud structure 26 surrounding a turbine rotor 22 comprises a ring 30 having a frusto conical inner surface 32 defining a small running clearance 29 with blades 24. The ring 30 is supported from a structure 42, 46 by flexible springs 38, 40 which permit axial movement of the ring from an initial position. In operation a pressure difference is formed axially across the ring 30 and a gas bearing is formed between the ring and the radial extremities of the blades 24. The pressure difference produces a first force on the ring 30 acting in a downstream direction; and the air bearing produces a second force on the ring acting in an upstream direction. Axial movement of the shroud ring from the initial position produces a third spring force on the ring acting towards the equilibrium position. The axial position of the ring 30 is determined by the resultant of the first, second and third forces and controls the running clearance. The gas bearing may be improved by the supply of air from each blade interior 58 to a cavity 52 at the blade tip. Winglets may also be provided at the tip. The ring 30 may include a heat pipe or thermosyphon to reduce thermal gradients. <IMAGE>
Description
SPECIFICATION
A tip clearance control apparatus for a rotor
The present invention relates to a shroud for a rotor assembly of a gas turbine engine, and is concerned with controlling a tip clearance between the shroud and the radial extremities of blades on the rotor.
It is desirable for gas turbine engines to be as efficient as possible, particularly in the case of gas turbine engines for use in propelling aircraft, in order to reduce fuel consumption and so that greater loads may be carried.
One way of increasing the efficiency of a gas turbine engine is to maintain the closest possible clearance between the radial extremities of the rotor blades and their encircling shrouds, because any gases which pass through the clearance gap represents a loss of energy to the rotor and hence a loss of efficiency to the gas turbine engine.
In the turbine region of the gas turbine engine the high speeds and temperatures involved lead to considerable thermal and centrifugal loads on the rotor. During operation of an aero engine there are many changes in engine speed which alter the centrifugal and thermal loads of the rotor, these changes cause the rotor to expand or contract and as a consequence the clearance between the radial extremities of the rotor blades and the shroud will change.
It is desirable to find a method of maintaining as small a clearance as possible between the shroud and the blade extremities. Several methods of controlling the clearance have been devised, for example various metals have been selected with which to make the shroud and blades because their thermal properties assist in matching their radial expansion/contraction at different engine operating conditions. Another method has been to direct variable temperature air onto a shroud support structure, so that the expansion/contraction rate of the shroud can be controlled in an effort to match the expansion/ contraction rate of the rotor at different engine operating conditions. A further method has been to move a shroud structure with a frustoconical inner surface axially to vary the clearance in a predetermined manner by using actuators, as in our previous patent application GB 2042646A.
The use of air to cool the shroud in order to control the clearance between the shroud and the radial extremities of the blades represents a loss of efficiency of the gas turbine engine as this air has been tapped from the compressor.
The use of the actuating means to move the shroud with a frustoconical inner surface is mechanically complex and a means to measure the clearance is required, both of which add weight to the engine.
Accordingly the present invention provides a rotor tip clearance control apparatus for a gas turbine engine comprising a shroud for use in conjunction with a plurality of blades arranged circumferentially on the rotor, the shroud comprising a continuous shroud ring having a frustoconical inner surface, the shroud ring being positioned coaxially around said rotor and said inner surface of said shroud ring being spaced radially from the radial extremities of the blades by a small clearance, the shroud ring being supported from a structure by at least one flexible structure, the flexible structure permitting axial movement of the shroud ring from an initial position, in operation a pressure difference being formed axially across the shroud ring and a gasbearing being formed between the inner surface of the shroud ring and the radial extremities of the blades, the pressure difference producing a first force on said shroud ring acting in a downstream direction, the gas bearing producing a second force on said shroud ring acting in an upstream direction, axial movement of the shroud ring from the initial position producing a third spring force on said shroud ring acting towards the initial position, the axial position of the ring being determined be the resultant of the first, second and third forces, the axial position of the shroud ring controlling the clearance between the inner surface of the shroud ring and the radial extremities of the blades.
There may be two flexible structures secured at the axial ends of the shroud ring, the flexible structures may be springs and the springs may be circular and have a generally U-shaped crosssection, the arms of the U may extend in a radial direction.
In operation the shroud ring may move axially in a downstream direction from the initial position producing the third spring force on said shroud ring acting in an upstream direction towards the initial position.
In operation the first, second and third forces acting on the shroud ring may be selected so that the resulting axial position of the shroud ring gives an optimum clearance between the inner surface of the shroud ring and the radial extremities of the blades when the rotor is operating at cruise conditions.
In operation if there is a change in the clearance between the inner surface of the shroud ring and the radial extremities of the blades there may be a change in the second force which will cause the resultant of the first, second and third forces to move the shroud ring axially to oppose the change in the clearance.
In operation if there is a change in rotor speed there may be a change in the first force which will cause the resultant of the first, second and third forces to move the shroud axially.
The shroud ring having an upstream and a downstream face, the areas of the upstream and downstream faces may be preselected to give a preselected first force acting on the shroud ring when the rotor is operating at cruise conditions.
The springs may be preselected to give a selected third force acting on the shroud ring when the rotor is operating at cruise conditions.
The shroud ring may be relatively massive in order to withstand thermal loads applied to the shroud ring during operation of the rotor.
The shroud ring may be hollow, and may contain a heat pipe or a thermosyphon.
The shroud ring may be solid.
The blades may be unshrouded, and the radial extremities may have cavities or circumferentially extending winglets.
The blades may be shrouded.
The present invention will be more fully described by way of example and witch reference to the accompanying drawings in which: Figure 1 is a partly broken away view of a gas turbine engine showing a turbine shroud structure according to the present invention.
Figure 2 is an enlarged section through the turbine area of figure 1 showing the turbine shroud structure and associated turbine blade.
Figure 3 is an enlarged section through the turbine area of figure 1 showing the turbine shroud structure and an alternative turbine blade.
Figure 4 is a section along line X-X in figure 3.
Figure 1 shows a gas turbine engine 10 which comprises in flow series a fan 12, a compressor section 14, a combustion section 16, a turbine section 1 8 and an exhaust nozzle 20. The turbine section 18 comprises a number of rotors 22 and stators 28, and a plurality of turbine blades 24 are arranged circumferentially on each rotor 22, and each turbine blade extends radially from the corresponding rotor. A shroud structure 26 is positioned coaxially around each rotor 22 and is spaced radially from the radial extremities 25 of the turbine blades 24 by a small clearance 29 which may be seen more clearly in figure 2.
In operation air enters the gas turbine engine 10 and flows through and is compressed by the fan 12 and the compressor section 14. The compressed air flows into the combustion section 1 6 where fuel is burnt in the compressed air to produce hot gases. The hot gases produced by the combustion of the fuel flow from the combustion section 1 6 through the turbine section 1 8 and the exhaust nozzle 20 to atmosphere. The hot gases drive the turbine rotors 22 which in turn drive the fan 12 and the compressor section 14 via shafts (not shown).
Figure 2 shows in greater detail part of the turbine section 18. A single rotor 22 only is shown with one of the turbine blades 24, which in this example are unshrouded. The rotor 22 is positioned downstream of the stators which in this case are nozzle guide vanes 28 which direct the hot gases onto the turbine blades 24. At their outer radial extremities 25 the turbine blades 24 run very close to a shroud structure 26. The shroud structure 26 comprises a continuous ring 30 which has a frustoconical inner surface 32, an upstream surface 34 and a downstream surface 36. The ring 30 in this case is hollow, but could be a solid ring. The frustoconical inner surface 32 of the hollow ring 30 is spaced from the outer radial extremities of the turbine blades 25 by the small clearance 29.The hollow ring 30 is supported from static structures 42 and 46 by circular springs 38 and 40 respectively which have a generally U-shaped cross-section and which permit axial movement of the hollow ring 30. The springs 38 and 40 are secured at the axial ends of the shroud ring 30, and the arms of the U extend radially. Spring 38 is secured to the static structure 42 by bolts 44 which also secure the nozzle guide vanes 28, and the static structure 46 is secured to casings 48 and 50. Initially, when the rotors of the gas turbine engine are at rest, the hollow ring 30 is in an initial position, and the springs 38 and 40 are neither in compression nor in expansion.
In operation the hot gases from the combustion system 1 6 are directed onto the turbine blades 24 at the optimum angle by the nozzle guide vanes 28. The turbine blades 24 extract energy from the hot gases as they flow between the turbine blades, and the energy is used to drive the fan or compressor section as mentioned above. The small clearance 29 between the outer radial extremities 25 of the turbine blades 24 and the frustoconical inner surface 32 of the hollow ring 30 is very important to the overall efficiency of the turbine rotor 22, and it must be maintained as small as possible but must not be allowed to close up completely otherwise the blades or the hollow ring will be damaged.
When the engine is running a pressure difference is formed axially across the turbine rotor 22, the associated turbine blades 24 and the hollow ring 30, due to the expansion of the hot gases through the turbine rotors 22. The upstream surface 34 of the hollow ring 30 is preselected to be greater than the downstream surface 36, and the pressure acting on the upstream surface 34 is greater than the pressure acting on the downstream surface 36, so there is a first resultant force on the hollow ring 30 due to the pressure difference acting in a downstream direction. This first force causes the hollow ring 30 to move axially in a downstream direction, and this results in a closing of the clearance 29 between the hollow ring 30 and the extremities 25 of the turbine blades 24.This movement aiso results in the generation of a gas bearing between the hollow ring 30 and the radial extremities 25 of the turbine blades 24. The gas bearing produces a second force which opposes the closing of the clearance 29 i.e. it acts axially in an upstream direction on the hollow ring 30.
Movement of the hollow ring 30 axially from the initial position produces a third force, a restoring force, due to the springs 38 and 40, acting towards the initial position i.e. in general acting in an upstream direction. The hollow ring 30 finally takes up an axial position which is determined by the resultant of the first, second and third forces, the axial position of the hollow ring controlling the clearance between the inner surface 32 of the hollow ring 30 and the radial extremities of the turbine blades 24.
The shroud structure 26 may be arranged so that, when the gas turbine engine is at rest, the hollow ring 30 is in its initial position as previously mentioned, and this position is upstream of the positions the hollow ring 30 may take up when the gas turbine is in operation. The spring strengths and the areas of the upstream and downstream faces must be chosen so that the first force will balance the second and third forces acting on the hollow ring 30 to give the optimum clearance 29 between the hollow ring 30 and the outer radial extremities 25 of the turbine blades 24 when the gas turbine engine 10 is operating at cruise conditions.
The gas bearing may be improved by using turbine blades 24 which have a cavity 52 at their outer radial extremities 25 as shown in figure 3.
The cavity 52 may be supplied with air from the interior 58 of the turbine blade 24 through apertures 56 in a wall 54 recessed from the outer radial extremity 25 of the turbine blade 24. The air supplied to the cavity 52 from the interior 58 of the turbine blade may have been used to cool the turbine blade 24. The turbine blades 24 may also have winglets 60 and 62 which extend circumferentially from their outer radial extremities 25 as seen in figure 4.
The winglets 60 and 62 will also increase the gas bearing effect between the radial extremities 25 of the turbine blades 24 and the inner surface 32 of the hollow ring 30.
The use of the first, second and third forces acting on the hollow ring to position the hollow ring axially, so as to controFthe clearance between the radial extremities of the turbine blades and the hollow ring, also allows the clearance to be self adjusting. For example, when the engine is operating at cruise conditions, the clearance is at its optimum and the first force is balanced by the second and third forces on the hollow ring. If the clearance increases for some reason, the gas bearing effect and the resulting second force will decrease, this will result in an inbalance between the first, second and third forces with the first force overpowering the sum of the second and third forces and producing a resultant force on the hollow ring acting in a downstream direction.The resultant force will act until the axial movement of the hollow ring reduces the clearance i.e.
increases the second force, and increases the spring force i.e. the third force, sufficiently for the second and third forces to once again balance the first force.
Similarly if the clearance decreases for some reason, the gas bearing effect will increase, and there will be an inbalance between the first second and third forces with the first force being overpowered by the sum of the second and third forces producing a resultant force on the hollow ring acting in an upstream direction. The resultant force will act until the axial movement of the hollow ring increases the clearance i.e. decreases the second force, and decreases the spring force i.e. the third force, sufficiently for the first force to once again balance the second and third forces.
Changes in engine speed will cause the pressure difference across the turbine blades and hollow ring to change which in turn will result in a change in the first force. This change will be followed by corresponding changes in the second and third forces following axial movement of the hollow ring to a new position.
It may be desirable to generate a pressure difference axially across the hollow ring by another method, for example, bleed air from the compressor, or other suitable position, could be supplied to one face of the hollow ring.
The hollow ring may be constructed so that it is quite massive so that it can withstand the thermal loads applied to it during the operation of the engine, and the hollow ring could be provided with a heat pipe or thermosyphon to minimise thermal gradients.
Although the figures and description refer to the use of a shroud structure with a frustoconical inner surface in cooperation with unshrouded turbine blades it may be possible to apply the principle to shrouded turbine blades. The air bearing could be formed by the discharging of cooling air from the interior of the blade through apertures in the blade shroud.
Claims (20)
1. A rotor tip clearance control apparatus for a gas turbine engine comprising a shroud for use in conjunction with a plurality of blades arranged circumferentially on a rotor, the shroud comprising a continuous shroud ring having a frusto conical inner surface, the shroud ring being positioned coaxially around said rotor and said inner surface of said shroud ring being spaced radially from the radial extremities of the blades by a small clearance, the shroud ring being supported from a structure by at least one flexible structure, the flexible structure permitting axial movement of the shroud ring from an initial position, in operation a pressure difference being formed axially across the shroud ring and a gas bearing being formed between the inner surface of the shroud ring and the radial extremities of the blades, the pressure difference producing a first force on said shroud ring acting in a downstream direction, the gas bearing producing a second force on said shroud ring acting in an upstream direction, axial movement of the shroud ring from the initial position producing a third spring force on said shroud ring acting towards the initial position, the axial position of the shroud ring being determined by the resultant of the first, second and third forces, the axial position of the shroud ring controlling the clearance between the inner surface of the shroud ring and the radial extremities of the blades.
2. A rotor tip clearance control apparatus as claimed in claim 1 in which there are two flexible structures secured at the axial ends of the shroud ring.
3. A rotor tip clearance control apparatus as claimed in claim 2 in which the flexible structures are springs, the springs being circular and having a generally U-shaped cross-section, the arms of the
U extending in a radial direction.
4. A rotor tip clearance control apparatus as claimed in claim 3 in which in operation the shroud ring moves axially in a downstream direction from the initial position producing the third spring force on said shroud ring acting in an upstream direction towards the initial position.
5. A rotor tip clearance control apparatus as claimed in any of claims 1 to 4 in which in operation the first, second and third forces acting on the shroud ring are selected so that the resulting axial position of the shroud ring gives an optimum clearance between the inner surface of the shroud ring and the radial extremities of the blades when the rotor is operating at cruise conditions.
6. A rotor tip clearance control apparatus as claimed in claim 5 in which in operation if there is a change in the clearance between the inner surface of the shroud ring and the radial extremities of the blades there is a corresponding change in the second force which will cause the resultant of the first, second and third forces to move the shroud ring axially to oppose the change in clearance.
7. A rotor tip clearance control apparatus as claimed in claim 5 in which in operation if there is a change in the speed of the rotor there is a corresponding change in the first force which will cause the resultant of the first, second and third forces to move the shroud ring axially.
8. A rotor tip clearance control apparatus as claimed in claim 5 in which the shroud ring has an upstream and a downstream face, the areas of the upstream and downstream faces of the shroud ring being preselected to give a selected first force acting on the shroud ring when the rotor is operating at cruise conditions.
9. A rotor tip clearance control apparatus as claimed in claim 5 in which the springs are preselected to give a selected third force acting on the shroud ring when the rotor is operating at cruise conditions.
10. A rotor tip clearance control apparatus as claimed in any of claims 1 to 9 in which the shroud ring is relatively massive in order to minimise the effects of thermal loads applied to the shroud ring during operation of the rotor.
11. A rotor tip clearance control apparatus as claimed in claim 10 in which the shroud ring is hollow.
12. A rotor tip clearance control apparatus as claimed in claim 11 in which the hollow shroud ring contains a heat pipe or a thermosyphon.
13. A rotor tip clearance control apparatus as claimed in claim 10 in which the shroud ring is solid.
14. A rotor tip clearance control apparatus as claimed in any of claims 1 to 1 3 in which the blades are unshrouded blades.
1 5. A rotor tip clearance control apparatus as claimed in claim 14 in which the radial extremities of the unshrouded blades have cavities.
1 6. A rotor tip clearance control apparatus as claimed in claim 14 or in claim 1 5 in which the radial extremities of the unshrouded blades have winglets extending circumferentially therefrom.
1 7. A rotor tip clearance control apparatus as claimed in any of claims 1 to 13 in which the blades are shrouded blades.
1 8. A rotor tip clearance control apparatus as claimed in any of the preceding claims in which the rotor is a turbine or a compressor rotor.
1 9. A rotor tip clearance control apparatus substantially as herein described and with reference to the accompanying drawings.
20. A gas turbine engine comprising a rotor tip clearance control apparatus as claimed in any of the preceding claims.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08231924A GB2129880A (en) | 1982-11-09 | 1982-11-09 | Gas turbine rotor tip clearance control apparatus |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB08231924A GB2129880A (en) | 1982-11-09 | 1982-11-09 | Gas turbine rotor tip clearance control apparatus |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| GB2129880A true GB2129880A (en) | 1984-05-23 |
Family
ID=10534128
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08231924A Withdrawn GB2129880A (en) | 1982-11-09 | 1982-11-09 | Gas turbine rotor tip clearance control apparatus |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2129880A (en) |
Cited By (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2206651A (en) * | 1987-07-01 | 1989-01-11 | Rolls Royce Plc | Turbine blade shroud structure |
| US5456576A (en) * | 1994-08-31 | 1995-10-10 | United Technologies Corporation | Dynamic control of tip clearance |
| WO2001044624A1 (en) * | 1999-12-14 | 2001-06-21 | Pratt & Whitney Canada Corp. | Split ring for tip clearance control |
| EP1691054A1 (en) * | 2005-02-12 | 2006-08-16 | Hubert Antoine | Gas Turbine |
| US7195452B2 (en) | 2004-09-27 | 2007-03-27 | Honeywell International, Inc. | Compliant mounting system for turbine shrouds |
| US7246994B2 (en) | 2004-05-27 | 2007-07-24 | Rolls-Royce Plc | Spacing arrangement |
| US7594792B2 (en) * | 2005-04-27 | 2009-09-29 | Snecma | Sealing device for a chamber of a turbomachine, and aircraft engine equipped with said sealing device |
| CN102046926A (en) * | 2008-05-28 | 2011-05-04 | 斯奈克玛公司 | High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box |
| EP2302167A3 (en) * | 2009-09-28 | 2013-03-13 | Rolls-Royce plc | A gas turbine sealing component |
| US8894349B2 (en) | 2009-08-20 | 2014-11-25 | Rolls-Royce Plc | Turbomachine casing assembly |
| US8944756B2 (en) | 2011-07-15 | 2015-02-03 | United Technologies Corporation | Blade outer air seal assembly |
| US9593589B2 (en) | 2014-02-28 | 2017-03-14 | General Electric Company | System and method for thrust bearing actuation to actively control clearance in a turbo machine |
| EP3299584A1 (en) * | 2016-09-23 | 2018-03-28 | Rolls-Royce plc | Gas turbine engine |
| GB2614760A (en) * | 2022-01-13 | 2023-07-19 | Rolls Royce Plc | Turbine for gas turbine engine |
-
1982
- 1982-11-09 GB GB08231924A patent/GB2129880A/en not_active Withdrawn
Cited By (18)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2206651A (en) * | 1987-07-01 | 1989-01-11 | Rolls Royce Plc | Turbine blade shroud structure |
| GB2206651B (en) * | 1987-07-01 | 1991-05-08 | Rolls Royce Plc | Turbine blade shroud structure |
| US5456576A (en) * | 1994-08-31 | 1995-10-10 | United Technologies Corporation | Dynamic control of tip clearance |
| WO1996007018A1 (en) * | 1994-08-31 | 1996-03-07 | United Technologies Corporation | Dynamic control of tip clearance |
| WO2001044624A1 (en) * | 1999-12-14 | 2001-06-21 | Pratt & Whitney Canada Corp. | Split ring for tip clearance control |
| US6368054B1 (en) | 1999-12-14 | 2002-04-09 | Pratt & Whitney Canada Corp. | Split ring for tip clearance control |
| US7246994B2 (en) | 2004-05-27 | 2007-07-24 | Rolls-Royce Plc | Spacing arrangement |
| US7195452B2 (en) | 2004-09-27 | 2007-03-27 | Honeywell International, Inc. | Compliant mounting system for turbine shrouds |
| EP1691054A1 (en) * | 2005-02-12 | 2006-08-16 | Hubert Antoine | Gas Turbine |
| US7594792B2 (en) * | 2005-04-27 | 2009-09-29 | Snecma | Sealing device for a chamber of a turbomachine, and aircraft engine equipped with said sealing device |
| CN102046926A (en) * | 2008-05-28 | 2011-05-04 | 斯奈克玛公司 | High pressure turbine of a turbomachine with improved assembly of the mobile blade radial clearance control box |
| US8894349B2 (en) | 2009-08-20 | 2014-11-25 | Rolls-Royce Plc | Turbomachine casing assembly |
| EP2302167A3 (en) * | 2009-09-28 | 2013-03-13 | Rolls-Royce plc | A gas turbine sealing component |
| US8727709B2 (en) | 2009-09-28 | 2014-05-20 | Rolls-Royce Plc | Casing component |
| US8944756B2 (en) | 2011-07-15 | 2015-02-03 | United Technologies Corporation | Blade outer air seal assembly |
| US9593589B2 (en) | 2014-02-28 | 2017-03-14 | General Electric Company | System and method for thrust bearing actuation to actively control clearance in a turbo machine |
| EP3299584A1 (en) * | 2016-09-23 | 2018-03-28 | Rolls-Royce plc | Gas turbine engine |
| GB2614760A (en) * | 2022-01-13 | 2023-07-19 | Rolls Royce Plc | Turbine for gas turbine engine |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |