US20070147989A1 - Fan or compressor casing - Google Patents

Fan or compressor casing Download PDF

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Publication number
US20070147989A1
US20070147989A1 US11604703 US60470306A US2007147989A1 US 20070147989 A1 US20070147989 A1 US 20070147989A1 US 11604703 US11604703 US 11604703 US 60470306 A US60470306 A US 60470306A US 2007147989 A1 US2007147989 A1 US 2007147989A1
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US
Grant status
Application
Patent type
Prior art keywords
casing
fan
grooves
helical groove
blades
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
US11604703
Inventor
Larry Collins
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
Original Assignee
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

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Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/40Casings; Connections of working fluid
    • F04D29/52Casings; Connections of working fluid for axial pumps
    • F04D29/522Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
    • F04D29/526Details of the casing section radially opposing blade tips
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/66Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/685Inducing localised fluid recirculation in the stator-rotor interface
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/10Stators
    • F05D2240/11Shroud seal segments
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/10Stators
    • F05D2240/14Casings or housings protecting or supporting assemblies within
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/25Three-dimensional helical

Abstract

The surge margin of a fan or compressor stage may be increased by radial, circular grooves in the casing (2) encircling the rotor stage. The present invention provides improved aerodynamic performance by replacing circumferential grooves by multiple helical grooves 18. Preferably the grooves (18) have multiple start and end positions (20,22) spaced apart around the casing (2) and at least the downstream ends (22) of the grooves (18) are tapered to aid expulsion of particles in the grooves (18).

Description

  • The invention relates to a fan or compressor casing and more particularly to a modified fan or compressor casing for having improved surge margin performance.
  • The invention may be utilised in a fan section at the front of a gas turbine engine, or in an axial flow compressor section of the engine, or in an independent rotary fan arrangement for example of the kind used as a lift fan for a vertical/short take off aircraft. Such arrangements have in common a rotary stage enclosed within a generally cylindrical casing. The aerodynamic performance of such arrangements and the phenomenon of surge are well understood. It is known to increase their surge margin by adopting certain modifications to the casing wall in the vicinity of the rotary stage that allow limited re-circulation of air over the tips of the blades.
  • One such modification for improving the surge margin of a rotary stage of this kind uses a number of circumferential grooves formed in the casing wall. In operation, as the stage rotates a certain amount of air passing through the stage is spilled from the blade tips into the grooves and can travel around the groove according to pressure distribution therein. The drawback in such an arrangement is that foreign objects are also centrifuged into the grooves and can accumulate therein leading to impaired performance of the modification. This is undesirable as it leads inevitably to a reduction of the available surge margin in the rotary stage.
  • The present invention solves the above-mentioned drawback by avoiding the accumulation of foreign objects in the aforementioned tip treatment grooves.
  • According to the present invention a gas turbine engine driven fan or compressor stage casing comprising a generally cylindrical casing portion encircling a fan or compressor rotor having a multiplicity of blades spaced apart around the rotor, the casing portion having formed in its inner surface over the tips of the blades at least one helical groove characterised in that the at least one helical groove has a gradually reducing depth in the downstream direction.
  • In one form of the invention the at least one helical groove extends for a plurality of turns around the inner surface of the casing, and in addition to the depth of groove or grooves being tapered towards the downstream direction, the grooved portion may also extend beyond the trailing edges of the rotor blades. The groove or grooves may also be tapered towards the forward end and the grooved portion may extend ahead of the leading edges of the rotor blades.
  • The invention will now be described by way of example only with reference to the accompanying drawings, in which:
  • FIG. 1 shows a section through part of a known axial flow compressor fan casing incorporating a known surge margin treatment and a blade tip portion of an associated rotary fan;
  • FIG. 2 shows a corresponding view through an axial flow compressor fan casing provided with a surge margin treatment in accordance with the present invention;
  • FIG. 3 shows a surge margin treatment according to the present invention in which the treatment grooves taper at their upstream and downstream ends; and
  • FIG. 4 shows a surge margin treatment similar to that of FIG. 3 un which the treatment grooves extend beyond the leading and trailing edges of the blades.
  • Referring to FIG. 1 there is shown a generally cylindrical portion 2 of an axial flow compressor fan casing, in this case the fan casing forms part of lift fan arrangement for a V/STOL aircraft but it could equally form part of the intake to a gas turbine propulsion engine. At either end the casing 2 incorporates integral annular flanges 4,6 by means of which the casing may be attached to adjacent lift fan or engine sections in axial flow series as appropriate. A tip portion of a blade forming part of an associated rotary fan is shown at 8 to illustrate the close spacing between the two parts. The radially outer tip surfaces 10 of the blades 8 are positioned close to the inner surface 12 of the casing 2. In operation, the exact size of the tip clearance gap varies with rotor speed mainly due to elongation of the blade radius as a result mainly of the effect of centrifugal forces and blade untwist.
  • Under certain operating conditions the compressor, or fan, working line may be forced beyond the stability line leading to a surge. The stage is designed such that in normal operating range a surge margin is maintained between the working and stability lines sufficient to avoid surge. The situations that produce surge and the transient operating conditions during a surge event are well known, and it is not intended to analyse the several possible types of surge in depth here. However, to summarise briefly, during a surge event, the inlet mass flow of the rotary stage stalls or varies with time as the flow rate oscillates and may even become negative. Certain design steps may be taken to avoid or alleviate the conditions that give rise to a stall. One such step involves the design of the casing wall in the vicinity of the blade tips.
  • It is known to modify the compressor casing wall to increase the surge margin, as illustrated in FIG. 1 by providing a plurality of substantially radial, circular grooves 14 a, 14 b . . . 14 j in the surface 12 of cylindrical wall 2 in the vicinity of the blade tips 10. Fluid pressure in these grooves 14 a-j tends to follow pressure at the blade tips 10 so that local variations around the stage circumference, as occur under stall conditions, result in circumferential fluid flow in some at least of the grooves 14 a-j and has the effect of evening out pressure variations or filling-in localised low-pressure regions.
  • The possibility of stall conditions occurring in the axial flow through the compressor is therefore reduced thus increasing the surge margin. For ease of construction and the possibility of repair by replacement the grooves 14 a-j are cut into a demountable, liner section 16 secured in a cut-out formed in the inner surface 12 of the casing section 2. The liner section 16 may be made from different material compared to the remainder of the casing 2. It may be made in a number of segments and secured to the casing by appropriate means.
  • However, such an arrangement suffers from a drawback in that foreign objects of size equal to or smaller than the width of the grooves 14 may become trapped in the grooves. Air borne foreign objects will inevitably enter the air flow through the fan or compressor stage, it is simply impractical to try to exclude them without severely impeding the intake fluid flow. A proportion of these objects will be thrown into the tip treatment grooves 14 a-j and will become trapped, unable to escape until the rotor has stopped or nearly stopped. Some may even become permanently trapped. A sufficient quantity of trapped particles will reduce the efficiency of the grooves and may abrade the blade tips 10 thereby reducing stage efficiency.
  • In the improvements illustrated in the accompanying drawings, like parts carry like references. Referring first to FIG. 2 the circular grooves 14 a-j of the prior art (FIG. 1) are replaced by at least one or a single helical groove 18 cut into the liner section 16. The helix or helical groove 18 is handed to suit the direction of rotation of the rotor and blades 8. Air entrained in the groove and foreign particles are forced to travel along the helix from their point of entry and are expelled at the end of the groove. The direction of the helix may be chosen so that particles are transported either upstream or downstream relative to their entry point into the groove.
  • There may be a plurality of such helical grooves 18 arranged to commence at a like number of different points 20 spaced apart around the circumference of the casing. In the case of a single helical groove 18 particles or foreign objects will tend to be expelled from the groove in one circumferential position, basically at the downstream end 22 of the groove but factors such as orientation may be an influence. It is considered undesirable to concentrate potentially abrasive particles in one position as it could result in increased localised abrasion of components downstream of the ejection point against which the debris is incident. Therefore it is preferred to employ a plurality of helical grooves 18 in the liner 16. Each of these helical grooves has the same number of turns and is of the same length, and the start points 20 or upstream ends are equidistantly spaced apart around the circumference of the liner 16. The exit or downstream points 22 are correspondingly spaced apart around the circumference of the liner surface 16. Thus the helix exit points 22 are not concentrated at one location on the circumference and localised abrasion is reduced. Furthermore it is preferred that the termination of each groove 18 is such that foreign particles are deflected away from adjacent, downstream components.
  • Also it is preferred that the depth of the helical grooves 18 is progressively reduced towards the chosen exit point 22 at least to reduce the chance that particles may become trapped at an abrupt end as in the prior art. It is envisaged that normally, that is in most cases only the depth of a portion of a helical grove 18 towards the downstream end 22 would be formed with a progressively reducing depth, as illustrated in FIG. 2. However, the upstream ends 20 of the or each helix or helical groove 18 may be tapered in similar fashion, that is the helical groove 18 may be formed with a progressively decreasing depth in an upstream direction, as illustrated in FIG. 3.
  • Multiple helical grooves 18 are more efficient at expelling particles, as the pitch of each helix 18 increases in proportion to the number of grooves, and less abrasion takes place. With less abrasion, or erosion, an optimum design geometry is maintained for a longer period so that the level of aerodynamic performance is preserved for a longer period of time compared to the existing prior art arrangement.
  • The grooves 18 serve a dual purpose. In addition to the grooves 18 functioning to collect and transport through the stage ingested foreign object particles, their primary function is to provide a re-circulation path in the vicinity of the tips of the rotor blades 8 to increase the stability and surge margin of the rotor stage in known fashion. Therefore, as is well known in that particular art the start 20 of the re-circulation grooves 18 may extend upstream of the leading edges 24 of the blades 8, and the exit 22 of the grooves 18 may extend downstream of the trailing edges 26 of the blades 8, as illustrated in FIG. 4 of the drawings.

Claims (6)

  1. 1. A gas turbine engine driven fan or compressor stage casing comprising a generally cylindrical casing portion (2) encircling a fan or compressor rotor having a multiplicity of blades (8) spaced apart around the rotor, the casing portion (2) having formed in its inner surface over the tips (10) of the blades (8) at least one helical groove (18) characterised in that the at least one helical groove (18) has a gradually reducing depth in the downstream direction.
  2. 2. A casing for a fan or compressor stage as claimed in claim 1 characterised in that the at least one helical groove (18) extends for a plurality of turns around the inner surface of the casing (2).
  3. 3. A casing for a fan or compressor stage as claimed in claim 1 characterised in that the casing (2) and the at least one helical groove (18) extends downstream of the trailing edge (26) of the tips of the blades (8).
  4. 4. A casing for a fan or compressor stage as claimed in claim 3 characterised in that the at least one helical groove (18) has a gradually reducing depth in the upstream direction.
  5. 5. A casing for a fan or compressor stage as claimed in claim 4 characterised in that the at least one helical groove (18) extends upstream of the leading edge (24) of the blades (8).
  6. 6. A casing for a fan or compressor stage as claimed in claim 2 characterised in that the casing (2) and the at least one helical groove (18) extends downstream of the trailing edge (26) of the tips of the blades (8).
US11604703 2005-12-22 2006-11-28 Fan or compressor casing Abandoned US20070147989A1 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
GB0526011.2 2005-12-22
GB0526011A GB0526011D0 (en) 2005-12-22 2005-12-22 Fan or compressor casing

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US20070147989A1 true true US20070147989A1 (en) 2007-06-28

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EP (1) EP1801361A1 (en)
GB (1) GB0526011D0 (en)

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090116188A1 (en) * 2007-11-05 2009-05-07 Microsoft Corporation Liquid resistant a/c adaptor
US20100232943A1 (en) * 2009-03-15 2010-09-16 Ward Thomas W Buried casing treatment strip for a gas turbine engine
US20110085896A1 (en) * 2008-03-28 2011-04-14 Snecma Casing for a moving-blade wheel of turbomachine
US20120201671A1 (en) * 2011-02-03 2012-08-09 Rolls-Royce Plc turbomachine comprising an annular casing and a bladed rotor
US8602720B2 (en) 2010-06-22 2013-12-10 Honeywell International Inc. Compressors with casing treatments in gas turbine engines
US20150037142A1 (en) * 2012-03-15 2015-02-05 Snecma Casing for turbomachine blish and turbomachine equipped with said casing
US20150132121A1 (en) * 2013-11-14 2015-05-14 Hon Hai Precision Industry Co., Ltd. Fan
US20150226078A1 (en) * 2012-09-25 2015-08-13 Snecma Turbine engine casing and rotor wheel
US20160010475A1 (en) * 2013-03-12 2016-01-14 United Technologies Corporation Cantilever stator with vortex initiation feature
US20160230776A1 (en) * 2015-02-10 2016-08-11 United Technologies Corporation Optimized circumferential groove casing treatment for axial compressors
US20160326899A1 (en) * 2015-05-08 2016-11-10 United Technologies Corporation Intermittent grooved soft abradable material to reduce blade tip temperature

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9726043B2 (en) * 2011-12-15 2017-08-08 General Electric Company Mounting apparatus for low-ductility turbine shroud
EP2687684A1 (en) * 2012-07-17 2014-01-22 MTU Aero Engines GmbH Abradable coating with spiral grooves in a turbomachine
FR2994718B1 (en) * 2012-08-27 2017-04-21 Snecma Carter treatments arasants housing
US9874104B2 (en) 2015-02-27 2018-01-23 General Electric Company Method and system for a ceramic matrix composite shroud hanger assembly

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US3261228A (en) * 1964-04-02 1966-07-19 United Aircraft Corp Disk fragment energy absorption and containment means
US4086022A (en) * 1975-09-25 1978-04-25 Rolls-Royce Limited Gas turbine engine with improved compressor casing for permitting higher air flow and pressure ratios before surge
US4197052A (en) * 1977-10-11 1980-04-08 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Safety device for an axially rotating machine
US5137419A (en) * 1984-06-19 1992-08-11 Rolls-Royce Plc Axial flow compressor surge margin improvement
US6290458B1 (en) * 1999-09-20 2001-09-18 Hitachi, Ltd. Turbo machines
US6302643B1 (en) * 1999-04-26 2001-10-16 Hitachi, Ltd. Turbo machines
US6499940B2 (en) * 2001-03-19 2002-12-31 Williams International Co., L.L.C. Compressor casing for a gas turbine engine
US6832890B2 (en) * 2002-07-20 2004-12-21 Rolls Royce Plc Gas turbine engine casing and rotor blade arrangement

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FR2558900B1 (en) * 1984-02-01 1988-05-27 Snecma Device for peripheral sealing of axial compressor blading
US6234747B1 (en) * 1999-11-15 2001-05-22 General Electric Company Rub resistant compressor stage

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Publication number Priority date Publication date Assignee Title
US3261228A (en) * 1964-04-02 1966-07-19 United Aircraft Corp Disk fragment energy absorption and containment means
US4086022A (en) * 1975-09-25 1978-04-25 Rolls-Royce Limited Gas turbine engine with improved compressor casing for permitting higher air flow and pressure ratios before surge
US4197052A (en) * 1977-10-11 1980-04-08 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Safety device for an axially rotating machine
US5137419A (en) * 1984-06-19 1992-08-11 Rolls-Royce Plc Axial flow compressor surge margin improvement
US6302643B1 (en) * 1999-04-26 2001-10-16 Hitachi, Ltd. Turbo machines
US6290458B1 (en) * 1999-09-20 2001-09-18 Hitachi, Ltd. Turbo machines
US6435819B2 (en) * 1999-09-20 2002-08-20 Hitachi, Ltd. Turbo machines
US20020182069A1 (en) * 1999-09-20 2002-12-05 Kouichi Irie Turbo machines
US6499940B2 (en) * 2001-03-19 2002-12-31 Williams International Co., L.L.C. Compressor casing for a gas turbine engine
US6832890B2 (en) * 2002-07-20 2004-12-21 Rolls Royce Plc Gas turbine engine casing and rotor blade arrangement

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7679906B2 (en) 2007-11-05 2010-03-16 Microsoft Corporation Liquid resistant A/C adaptor
US20090116188A1 (en) * 2007-11-05 2009-05-07 Microsoft Corporation Liquid resistant a/c adaptor
US20110085896A1 (en) * 2008-03-28 2011-04-14 Snecma Casing for a moving-blade wheel of turbomachine
US8777558B2 (en) * 2008-03-28 2014-07-15 Snecma Casing for a moving-blade wheel of turbomachine
US8177494B2 (en) * 2009-03-15 2012-05-15 United Technologies Corporation Buried casing treatment strip for a gas turbine engine
US20100232943A1 (en) * 2009-03-15 2010-09-16 Ward Thomas W Buried casing treatment strip for a gas turbine engine
US8602720B2 (en) 2010-06-22 2013-12-10 Honeywell International Inc. Compressors with casing treatments in gas turbine engines
US20120201671A1 (en) * 2011-02-03 2012-08-09 Rolls-Royce Plc turbomachine comprising an annular casing and a bladed rotor
US9004859B2 (en) * 2011-02-03 2015-04-14 Rolls-Royce Plc Turbomachine comprising an annular casing and a bladed rotor
EP2484913A3 (en) * 2011-02-03 2018-04-11 Rolls-Royce plc A turbomachine comprising an annular casing and a bladed rotor
US20150037142A1 (en) * 2012-03-15 2015-02-05 Snecma Casing for turbomachine blish and turbomachine equipped with said casing
US9651060B2 (en) * 2012-03-15 2017-05-16 Snecma Casing for turbomachine blisk and turbomachine equipped with said casing
US20150226078A1 (en) * 2012-09-25 2015-08-13 Snecma Turbine engine casing and rotor wheel
US9982554B2 (en) * 2012-09-25 2018-05-29 Snecma Turbine engine casing and rotor wheel
US20160010475A1 (en) * 2013-03-12 2016-01-14 United Technologies Corporation Cantilever stator with vortex initiation feature
US20150132121A1 (en) * 2013-11-14 2015-05-14 Hon Hai Precision Industry Co., Ltd. Fan
US20160230776A1 (en) * 2015-02-10 2016-08-11 United Technologies Corporation Optimized circumferential groove casing treatment for axial compressors
US10066640B2 (en) * 2015-02-10 2018-09-04 United Technologies Corporation Optimized circumferential groove casing treatment for axial compressors
US20160326899A1 (en) * 2015-05-08 2016-11-10 United Technologies Corporation Intermittent grooved soft abradable material to reduce blade tip temperature
US9951642B2 (en) * 2015-05-08 2018-04-24 United Technologies Corporation Intermittent grooved soft abradable material to reduce blade tip temperature

Also Published As

Publication number Publication date Type
EP1801361A1 (en) 2007-06-27 application
GB0526011D0 (en) 2006-02-01 grant

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Owner name: ROLLS-ROYCE PLC, GREAT BRITAIN

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:COLLINS, LARRY;REEL/FRAME:018622/0602

Effective date: 20060919