US20080138197A1 - Transition duct for a gas turbine engine - Google Patents

Transition duct for a gas turbine engine Download PDF

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Publication number
US20080138197A1
US20080138197A1 US11/984,976 US98497607A US2008138197A1 US 20080138197 A1 US20080138197 A1 US 20080138197A1 US 98497607 A US98497607 A US 98497607A US 2008138197 A1 US2008138197 A1 US 2008138197A1
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United States
Prior art keywords
fairing
duct
perturbations
transition duct
end wall
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.)
Abandoned
Application number
US11/984,976
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English (en)
Inventor
Matthew J. Green
Neil W. Harvey
Edward M. J. Naylor
Cecilia Ortiz-Duenas
Howard P. Hodson
Robert J. Miller
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
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Rolls Royce PLC
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Assigned to ROLLS-ROYCE PLC reassignment ROLLS-ROYCE PLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GREEN, MATTHEW JAMES, HARVEY, NEIL WILLIAM, ORTIZ-DUENAS, CECILIA, HODSON, HOWARD PETER, MILLER, ROBERT JOHN, NAYLOR, EDWARD MICHAEL JOHN
Publication of US20080138197A1 publication Critical patent/US20080138197A1/en
Abandoned legal-status Critical Current

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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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • 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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/126—Baffles or ribs
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to transition ducts in gas turbine engines.
  • the low-pressure (LP) system has a lower rotational speed and larger radius than the high-pressure (HP) core system.
  • HP high-pressure
  • intermediate “S” shaped transition ducts are needed to connect the compressor or turbine of the high-radius LP system with the corresponding compressor or turbine of the low-radius HP system.
  • These intermediate ducts often carry loads, support bearings and have thick structural struts passing through them, making them large, heavy and expensive structures of considerable complexity. Improving these ducts can lead to significant benefits both in the weight and in the performance of the engine.
  • curvature of the duct is made more pronounced, the required change of radius can be accomplished with a shorter duct and the whole engine can be made shorter and lighter.
  • curvature exacerbates the undesirable aerodynamic effects in the duct. This may worsen the aerodynamic performance of a downstream compressor or turbine, and may cause instability in either an upstream or a downstream compressor.
  • transition duct for a gas turbine engine as claimed in claim 1 .
  • FIG. 1 is a perspective view of part of a hub end wall of a known transition duct containing a fairing, showing contours of static pressures near to the wall;
  • FIG. 2 is a sectional side view of the duct of FIG. 1 , showing contours of velocity close to the fairing surface;
  • FIG. 3 is a schematic plan view of a fairing in a transition duct, showing the changes to the hub end wall in a first embodiment of the invention
  • FIG. 4 shows the total pressure contours in axial planes close to the fairing leading edge ( FIG. 4( a )) and trailing edge ( FIG. 4( b )) for the embodiment of FIG. 3 ;
  • FIG. 5 is a perspective view of part of a hub end wall of a transition duct containing a fairing, showing contours of static pressures near to the wall in a known duct ( FIG. 5( a )) and in a duct according to the invention ( FIG. 5( b ));
  • FIGS. 6( a ) and 6 ( b ) are sectional side views of the ducts of FIGS. 5( a ) and 5 ( b ), showing contours of velocity close to the fairing surface;
  • FIG. 7 shows contours of total pressure in the axial plane downstream of a fairing in a transition duct, without ( FIG. 7( a )) and with ( FIG. 7( b )) non-axisymmetric end wall profiling;
  • FIG. 8 is a graph of static pressure against fairing surface perimeter for the boundary layer air flow in the embodiment of FIG. 9 ;
  • FIG. 9 is a schematic plan view of a fairing in a transition duct, showing the changes to the hub end wall in a second embodiment of the invention.
  • a typical duct used between compressors today has an area ratio (A in /A out ) of 1.0-2.5 and an aspect ratio ( ⁇ R/length) of 0.3-0.6.
  • the ducts used in current engines may in fact achieve only minimal diffusion. This is a very conservative choice in order to be absolutely sure that there is no risk of separations in the duct.
  • a conservative duct leads to significant compromises in both upstream and downstream compressors. Rear stages of the upstream compressor have to be designed at a non-optimum radius, and the design flow coefficient of the downstream compressor is limited by the exit Mach number of the upstream compressor and the choice of a conservative area ratio through the duct. There are inevitable cost, weight and fuel-burn penalties associated with such compromises, and these will be present for the complete life cycle of the product. If it were possible to use more pronounced curvatures in duct designs and design tools, much better optimization of the whole compression system would be possible.
  • Transition ducts between compressors should ideally diffuse the flow through them in order to minimise the flow velocity into the downstream compressor. That is to say, the cross-sectional area of the duct should ideally increase in the flow direction.
  • struts When structural struts are present they will typically have a form of aerodynamic fairing around them to present an acceptably smooth surface to the gas flow.
  • the fairing is symmetrical in shape around the camber line of the strut, and both its surfaces are largely convex in plan view.
  • the strut and fairing will not be separate. Rather, the strut itself will be aerodynamically shaped.
  • fairing will be used for such an aerodynamic fairing, or for a fairing-shaped strut.
  • the fairing will also be so aligned. However, if the inlet flow is swirling (has a tangential velocity component) then the fairings will be staggered to be aligned to this flow.
  • FIGS. 1 and 2 illustrate the problems with a diffusing duct with a (thick) fairing 12 in it. In this case the flow into the duct has no tangential component.
  • FIG. 1 shows a plan view of contours of static pressure on the hub (radially inner) end wall 14 .
  • the circumferential direction is shown by the arrow 16 , and the direction of the inlet gas flow by the arrow 18 .
  • the lowest values of static pressure are seen over the portion 20 of the end wall that has convex curvature, and it can be seen that low static pressure persists in the region 22 near to the fairing, even as the static pressure further away from the fairing (and at the same axial position) increases. It is these regions of low static pressure that give rise to the flow separations. Further downstream, the static pressure rises in the region 24 of the end wall that has concave curvature, and as the cross-sectional area of the duct increases.
  • FIG. 2 shows, in sectional side view, contours of flow velocity close to the fairing 12 of FIG. 1 .
  • the direction of the inlet gas flow is shown by the arrow 18 , and the radially outward direction by the arrow 26 .
  • the inlet 23 to the transition duct is therefore at a greater radius than the outlet 25 .
  • the velocities shown are just outside the aerofoil boundary layer, but the effects of the end wall boundary layer can be seen. Note also that the static pressure and fluid velocity are in an inverse relationship.
  • FIG. 2 the curvature of the hub end wall 14 can be clearly seen.
  • the curvature In the front part 28 of the transition duct the curvature is convex. This accelerates the flow, raising velocities and reducing the static pressure. This reduction in static pressure can be seen in FIG. 1 (region 20 ).
  • the hub end wall curvature is concave. This will have the effect of diffusing the flow. Because the duct is reducing in radius, from inlet to exit, it will be understood that this arrangement, with first convex and then concave curvature on the hub end wall, will generally be unavoidable. The flow field on the hub end wall is a particular problem.
  • the early convex curvature raises the local flow velocity above the value that would be expected from the area change alone, and it may in fact be significantly higher than the inlet value. From this local peak in velocity the flow must then diffuse to the exit of the duct. Thus the maximum (local) diffusion may be much higher than that obtained from simply considering the ratio of exit to inlet flow velocities.
  • region 34 shows the increased flow velocity arising from the blockage and convex surface curvature of the fairing 12 .
  • the aerodynamic conditions may be so adverse that flow separation occurs. This is shown in FIG. 2 where a significant region of reverse flow 32 occurs towards the downstream end of the fairing 12 surface near the hub. This significantly increases the losses in the duct. Note that this region of reverse flow typically does not extend across the whole circumferential width of the passage, but is confined to the region near both surfaces of the fairing 12 .
  • the location of the maximum thickness of the fairing 12 is typically in its forward (upstream) part. This will usually correspond to the location of the maximum thickness of the strut within it. By having the maximum thickness in the forward part of the strut the diffusion gradient along the surface of the fairing is reduced and the losses in the boundary layer (away from the end walls) are minimised.
  • Three-dimensional shaping of vanes in gas turbine engines is a well-known technique, used to improve aerodynamic efficiency and suppress end wall flow separation.
  • Examples of such shaping are the application of lean (often “compound lean”), sweep and dihedral.
  • lean often “compound lean”
  • sweep and dihedral are the applications of lean (often “compound lean”), sweep and dihedral.
  • the strut will typically be a major load-bearing part of the gas turbine engine and the loads must be carried in a largely radial manner through it, so it is not possible to apply any significant three-dimensional shaping to such components.
  • Boundary layer bleed from the strut and/or the end walls, is another known means of preventing separation in strongly diffusing flows.
  • this is rarely if ever applied to practical gas turbine engines, because usually the benefit in aerodynamic efficiency will be lost due to the cost and weight of the bleed system.
  • the action of bleeding off the flow and either dumping it overboard or re-introducing it back into the gas flow in some other part of the engine will typically incur as much extra loss elsewhere as was gained in the duct.
  • FIG. 3 illustrates a first embodiment of the invention, and shows schematically profiling of the hub end wall around a fairing 12 in a transition duct.
  • the strut, and the fairing around it will be aligned substantially radially. They may be leant in an axial sense but they will not usually be leant tangentially.
  • the first compressor will be at a higher radius than the second and rotating more slowly than it.
  • this transition duct will typically have a reducing radius through it.
  • each perturbation 36 , 38 will generally take the form of a protruding “blister” or a recessed “hollow” in the axisymmetric annulus profile.
  • the perturbations 36 , 38 will be described in more detail in due course.
  • FIG. 4 illustrates the total pressure contours in axial planes close to the fairing 12 leading edge 46 ( FIG. 4( a )) and trailing edge 48 ( FIG. 4( b )).
  • the positions of the perturbations 36 , 38 are also shown in FIGS. 4( a ) and 4 ( b ) respectively.
  • FIG. 5 illustrates the effect of the hub end wall profiling of FIG. 3 , showing contours of static pressures near to the wall in a known duct ( FIG. 5( a )) and in a duct according to the invention with hub end wall profiling ( FIG. 5( b )).
  • FIG. 6 shows contours of velocity close to the vane fairing surface in a known duct ( FIG. 6( a )) and in a duct according to the invention with hub end wall profiling ( FIG. 6( b )).
  • the perturbation 36 is symmetrical about the fairing camber line 40 . Noting that this is a hub end wall, the radius (from the engine centreline) is reduced at the camber line.
  • the perturbation 36 has its lowest point at the camber line, and therefore defines a “hollow” in the end wall. Away from the camber line, the amplitude of the perturbation reduces steadily until the end wall radius (from the engine centreline) is the same as that of the axisymmetric annulus. This occurs at the position indicated by the notional line 37 , which defines the circumferential extent of the perturbation.
  • the amplitude of the perturbation also reduces in an axially upstream direction from its maximum, to blend into the axisymmetric annulus at the appropriate point on the notional line 37 .
  • the perturbation 38 has the opposite arrangement to the perturbation 36 . Its highest point (i.e. the point of maximum radius from the engine centreline) is at the camber line 40 , and it therefore defines a “blister” in the end wall. Away from the camber line, the amplitude of the perturbation reduces steadily until the end wall radius (from the engine centreline) is the same as that of the axisymmetric annulus. This occurs at the position indicated by the notional line 39 , which defines the circumferential extent of the perturbation. The amplitude of the perturbation also reduces in an axially downstream direction from its maximum, to blend into the axisymmetric annulus at the appropriate point on the notional line 39 .
  • the perturbations will have a circumferential extent, on each side of the fairing, typically about 50% of the fairing pitch at each location.
  • the range of useful values will lie between 5% and 100% of the pitch.
  • the circumferential extent of a perturbation is defined as described in the discussion of FIG. 3 above, particularly in terms of the notional lines shown as 37 and 39 in that drawing. For example, if two adjacent perturbations each had a circumferential extent of 50% of pitch, then one or both of their notional lines 37 , 39 would be coincident.
  • the fairing arrangement is not periodically uniform. It might happen, for example, that only every alternate fairing has the profiling applied. In this case the perturbation around one fairing could extend beyond the mid-pitch position, so that its circumferential extent would be greater than 50%, and it may extend as far as the next fairing, in which case its circumferential extent would be 100%.
  • the perturbations at different axial locations may have different maximum amplitudes.
  • the maximum amplitude will be 7% of the chord, but may lie in the range 2% to 15% of chord depending on the details of the flow conditions. (Higher speed flows require smaller amplitudes to achieve the same aerodynamic effects as lower speed flows).
  • a smooth end wall shape is obtained by applying a spline in the streamwise direction through circumferentially corresponding points at these axial locations, to blend the perturbations smoothly into the axisymmetric end wall shape so as to present a smooth surface to the gas flow.
  • FIG. 7 shows contours of total pressure in an axial plane at the exit from the duct, for a known duct ( FIG. 7( a )) and for a duct according to the invention with non-axisymmetric perturbations ( FIG. 7( b )).
  • the areas of greatest loss are shown by the regions 72 in FIG. 7( a ) and by the region 74 in FIG. 7( b ), from which it can be seen that the extent and depth of the losses have been significantly reduced. Reductions in loss of 20% may be achieved, with a corresponding improvement in the aerodynamic efficiency.
  • Careful design of the area ruling should enable the flow near the fairing to experience the same diffusion (from inlet to exit) that the mid-passage region experiences. This embodiment is illustrated in FIGS. 8 and 9 .
  • FIG. 8 illustrates the typical static pressure distribution along one surface of the fairing where it intersects with the axisymmetric end wall—in the particular case where there is no net diffusion through the duct.
  • Static pressure, P is plotted against fairing surface perimeter, S.
  • a second embodiment of the invention in which the end wall profiling is as shown in FIG. 9 .
  • the amplitudes of the perturbations, and their axial locations, are defined such that they compensate for the reduction in the cross-sectional area of the passage due to the blockage of the fairing.
  • This will be known as “non-axisymmetric end wall area ruling”.
  • there will be a perturbation 92 in the leading edge region such that the radius of the end wall with respect to the engine centreline is increased at the intersection with the camber line 40 .
  • This will locally reduce the flow area, raising flow velocities and reducing the static pressure. This will compensate for the increase in static pressure where the flow stagnates at the leading edge.
  • the maximum amplitude of the perturbation may be at the leading edge or upstream of it.
  • a smooth shape is obtained for the end wall by applying a spline fit in the streamwise direction through circumferentially corresponding points at different axial locations, to blend the perturbations smoothly into the axisymmetric end wall shape.
  • the end wall profiling is defined by perturbations at, at least, two axial locations.
  • the end wall profiling may be applied to either (radially inner or radially outer) end wall of the transition duct. Typically for a transition duct with reducing radius it will be most effective if applied to the hub end wall.
  • the annulus must locally be raised or lowered, which will alter the flow area. Conversely, the area ruling cannot be implemented without changing the local surface curvature.
  • An optimum aerodynamic design for the annulus wall may be obtained by the designing the duct without a fairing present to achieve maximum diffusion along the annulus wall while avoiding (2-D) boundary layer separation.
  • the fairing is then replaced and non axisymmetric end wall profiling applied to restore the end wall static pressure field to be as close as possible to what it was in the absence of the fairing.
  • the fairings are not arranged uniformly in the circumferential direction.
  • the dimensions of the perturbations may be based on the chord and pitch local to each fairing.

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  • Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US11/984,976 2006-12-05 2007-11-26 Transition duct for a gas turbine engine Abandoned US20080138197A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GB0624294.5 2006-12-05
GBGB0624294.5A GB0624294D0 (en) 2006-12-05 2006-12-05 A transition duct for a gas turbine engine

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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150016983A1 (en) * 2013-03-14 2015-01-15 Rolls-Royce Corporation Subsonic shock strut
US9222437B2 (en) 2012-09-21 2015-12-29 General Electric Company Transition duct for use in a turbine engine and method of assembly
EP2233726A3 (fr) * 2009-03-10 2016-04-06 Rolls-Royce Deutschland Ltd & Co KG Canal de dérivation d'un turboréacteur
US9951633B2 (en) 2014-02-13 2018-04-24 United Technologies Corporation Reduced length transition ducts
US11028707B2 (en) * 2017-01-19 2021-06-08 Gkn Aerospace Sweden Ab Zoned surface roughness
US11242770B2 (en) 2020-04-02 2022-02-08 General Electric Company Turbine center frame and method
US11434832B2 (en) 2019-05-23 2022-09-06 Rolls-Royce Plc Geared gas turbine engine

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6035946B2 (ja) 2012-07-26 2016-11-30 株式会社Ihi エンジンダクト及び航空機エンジン
GB201512516D0 (en) * 2015-07-17 2015-08-19 Rolls Royce Plc A gas turbine engine

Citations (5)

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Publication number Priority date Publication date Assignee Title
US4677828A (en) * 1983-06-16 1987-07-07 United Technologies Corporation Circumferentially area ruled duct
US6283713B1 (en) * 1998-10-30 2001-09-04 Rolls-Royce Plc Bladed ducting for turbomachinery
US6494044B1 (en) * 1999-11-19 2002-12-17 General Electric Company Aerodynamic devices for enhancing sidepanel cooling on an impingement cooled transition duct and related method
US20060024158A1 (en) * 2004-07-28 2006-02-02 Martin Hoeger Flow structure for a gas turbine
US20060051200A1 (en) * 2004-09-03 2006-03-09 Martin Hoeger Flow structure for a gas turbine

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JPH06257597A (ja) * 1993-03-02 1994-09-13 Jisedai Koukuuki Kiban Gijutsu Kenkyusho:Kk 軸流圧縮機の翼列構造
JP5124276B2 (ja) * 2004-10-07 2013-01-23 ボルボ エアロ コーポレイション ガスタービン中間構造および該中間構造を含むガスタービンエンジン
GB0518628D0 (en) * 2005-09-13 2005-10-19 Rolls Royce Plc Axial compressor blading

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4677828A (en) * 1983-06-16 1987-07-07 United Technologies Corporation Circumferentially area ruled duct
US6283713B1 (en) * 1998-10-30 2001-09-04 Rolls-Royce Plc Bladed ducting for turbomachinery
US6494044B1 (en) * 1999-11-19 2002-12-17 General Electric Company Aerodynamic devices for enhancing sidepanel cooling on an impingement cooled transition duct and related method
US20060024158A1 (en) * 2004-07-28 2006-02-02 Martin Hoeger Flow structure for a gas turbine
US20060051200A1 (en) * 2004-09-03 2006-03-09 Martin Hoeger Flow structure for a gas turbine

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2233726A3 (fr) * 2009-03-10 2016-04-06 Rolls-Royce Deutschland Ltd & Co KG Canal de dérivation d'un turboréacteur
US9222437B2 (en) 2012-09-21 2015-12-29 General Electric Company Transition duct for use in a turbine engine and method of assembly
US20150016983A1 (en) * 2013-03-14 2015-01-15 Rolls-Royce Corporation Subsonic shock strut
US10309236B2 (en) * 2013-03-14 2019-06-04 Rolls-Royce Corporation Subsonic shock strut
US9951633B2 (en) 2014-02-13 2018-04-24 United Technologies Corporation Reduced length transition ducts
US11028707B2 (en) * 2017-01-19 2021-06-08 Gkn Aerospace Sweden Ab Zoned surface roughness
US11434832B2 (en) 2019-05-23 2022-09-06 Rolls-Royce Plc Geared gas turbine engine
US11761384B2 (en) 2019-05-23 2023-09-19 Rolls-Royce Plc Geared gas turbine engine
US11994075B2 (en) 2019-05-23 2024-05-28 Rolls-Royce Plc Geared gas turbine engine
US12140084B2 (en) 2019-05-23 2024-11-12 Rolls-Royce Plc Geared gas turbine engine
US11242770B2 (en) 2020-04-02 2022-02-08 General Electric Company Turbine center frame and method

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EP1936120A2 (fr) 2008-06-25
EP1936120A3 (fr) 2011-07-20
GB0624294D0 (en) 2007-01-10

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