US20100251688A1 - Containment system for a gas turbine engine - Google Patents
Containment system for a gas turbine engine Download PDFInfo
- Publication number
- US20100251688A1 US20100251688A1 US11/651,553 US65155307A US2010251688A1 US 20100251688 A1 US20100251688 A1 US 20100251688A1 US 65155307 A US65155307 A US 65155307A US 2010251688 A1 US2010251688 A1 US 2010251688A1
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- Prior art keywords
- engine
- force absorbing
- casing
- aircraft according
- arrangement
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- 230000004323 axial length Effects 0.000 claims abstract description 7
- 239000011358 absorbing material Substances 0.000 claims description 44
- NJPPVKZQTLUDBO-UHFFFAOYSA-N novaluron Chemical compound C1=C(Cl)C(OC(F)(F)C(OC(F)(F)F)F)=CC=C1NC(=O)NC(=O)C1=C(F)C=CC=C1F NJPPVKZQTLUDBO-UHFFFAOYSA-N 0.000 claims description 28
- 239000012634 fragment Substances 0.000 claims description 7
- 239000000463 material Substances 0.000 claims description 7
- 238000005728 strengthening Methods 0.000 claims description 7
- 229920003235 aromatic polyamide Polymers 0.000 claims description 4
- 238000011144 upstream manufacturing Methods 0.000 claims description 4
- MHSKRLJMQQNJNC-UHFFFAOYSA-N terephthalamide Chemical compound NC(=O)C1=CC=C(C(N)=O)C=C1 MHSKRLJMQQNJNC-UHFFFAOYSA-N 0.000 claims description 3
- 238000012423 maintenance Methods 0.000 description 4
- 229920000271 Kevlar® Polymers 0.000 description 2
- 239000004761 kevlar Substances 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 239000004952 Polyamide Substances 0.000 description 1
- 239000004760 aramid Substances 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000009172 bursting Effects 0.000 description 1
- 210000003746 feather Anatomy 0.000 description 1
- 229920002647 polyamide Polymers 0.000 description 1
- 230000001141 propulsive effect Effects 0.000 description 1
Images
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
- F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
- F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
- F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/522—Casings; Connections of working fluid for axial pumps especially adapted for elastic fluid pumps
- F04D29/526—Details of the casing section radially opposing blade tips
-
- 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
- F05D2250/60—Structure; Surface texture
- F05D2250/61—Structure; Surface texture corrugated
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/40—Organic materials
- F05D2300/43—Synthetic polymers, e.g. plastics; Rubber
- F05D2300/433—Polyamides, e.g. NYLON
-
- 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
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/603—Composites; e.g. fibre-reinforced
Definitions
- This invention relates to containment systems for gas turbine engines.
- a containment system for a gas turbine engine comprising a force absorbing arrangement for absorbing the force exerted thereon by an ejected portion of a failed component of the engine, wherein the force absorbing arrangement circumferentially surrounds at least a major proportion of the axial length of the engine.
- the aforesaid component may comprise a high energy component.
- the component may comprise a disk.
- the component may comprise a compressor or a turbine disk.
- the force absorbing arrangement extends substantially the axial length of the engine.
- the force absorbing arrangement comprises a force absorbing material.
- a suitable material is a polyamide, such as an aramid, typically a para-aramid.
- the force absorbing material may comprise poly-paraphenylene terephthalamide.
- An example of such a material is sold under the registered trade mark KEVLAR.
- the engine may include rotary driving components.
- a casing may surround the rotary driving components.
- the rotary driving components may comprise a fan, a compressor arrangement, and a turbine arrangement.
- the casing may comprise radially inner and outer skins.
- the force absorbing arrangement may be provided on the casing, preferably between the inner and outer skins of the casing.
- the casing may comprise a nacelle for surrounding the engine.
- the force absorbing arrangement may be provided as a layer on one of the inner and/or outer skins. Desirably, the force absorbing arrangement is provided on the inner skin.
- the force absorbing arrangement may comprise a corrugated outer skin of the casing.
- a plurality of support members may be provided within the casing.
- Each support member may extend circumferentially around the casing, and may extend from the inner skin to the outer skin.
- the support members provide strengthening between the inner and outer skins.
- Each support member may constitute a bulkhead extending from the inner skin to the outer skin of the casing.
- the support members may be spaced from one another axially along the casing.
- the support members may divide the casing into a plurality of compartments.
- a pedestal may extend from the casing to mount the engine on an aircraft.
- support members extend to the pedestal.
- each support member transmits loads thereon to the pedestal.
- the provision of the support members provides the advantage in an embodiment of the invention that a load exerted on the force absorbing arrangement is constrained by the support members to prevent excessive deformation of the casing and of force absorbing material.
- the force absorbing arrangement may comprise a portion of the force absorbing material surrounding the fan, and a further portion of the force absorbing material circumferentially surrounding the remainder of the rotary components.
- the force absorbing material may comprise a first portion of the force absorbing material extending from a region surrounding the fan to a region surrounding the intermediate pressure compressor of the engine.
- the force absorbing arrangement may comprise a second portion of the force absorbing material extending from a region surrounding the high pressure compressor to a region surrounding or downstream of the low pressure turbine.
- the force absorbing material may be wrapped around the engine with appropriate tension to allow the force absorbing material to deform radially outwardly within the casing when absorbing the force of the ejected portion of the failed component.
- the outer skin of the casing may be formed of a corrugated material to allow radially outward deformation thereof if the force absorbing material impacts thereon on radially outwardly formation of the force absorbing material.
- the engine may be secured to the case by strengthening arrangements. Each strengthening arrangement may comprise an A frame extending from the engine to the casing.
- the force absorbing portion may comprise a corrugated outer skin.
- an engine arrangement comprising a gas turbine engine having a containment system as described above: and a pedestal to support the engine on an aircraft.
- the containment system includes a casing extending circumferentially around rotary driving components of the engine, as a plurality of support members extending circumferentially around the casing wherein each support member extends to the casing.
- a casing extending circumferentially around rotary driving components of the engine, as a plurality of support members extending circumferentially around the casing wherein each support member extends to the casing.
- the engine arrangement may include a chamber externally of the engine to include ancillary components of the engine.
- the ancillary components may include a power off take arrangement to extract power from a main shaft of the engine.
- the power off take arrangement may include an off take shaft extending from the aforesaid main shaft of the engine to the chamber, and a gear arrangement to transmit rotary power from the shaft to further components.
- FIG. 1A is a sectional side view of a three shaft gas turbine engine, with a nacelle surrounding the engine;
- FIG. 1B is a sectional side view of a two shaft gas turbine engine, with a nacelle surrounding the engine;
- FIG. 2 is a part sectional view of the nacelle shown in FIG. 1 supported by a pedestal;
- FIG. 3 is a view of the region marked III in FIG. 2 ;
- FIG. 4 is a perspective view of the nacelle shown in FIG. 2 indicating support members within the nacelle;
- FIG. 5 is a front sectional view of an aircraft fuselage incorporating two engines.
- FIG. 6 is a sectional view of an alternative embodiment of a containment system used in a gas turbine engine.
- FIG. 1A there is shown a sectional side view of a three shaft gas turbine engine 10 A housed in a nacelle 12 .
- the engine 10 A comprises a fan 14 and an engine core 16 .
- a bypass duct 18 is defined between the nacelle 12 and the engine core.
- the engine core 16 comprises an intermediate compressor 20 , a high pressure compressor 22 , a combustor 24 and high, intermediate and low pressure turbines 26 , 28 , 30 respectively.
- FIG. 1B shows a sectional side view of a two shaft gas turbine engine 10 B housed in a nacelle 12 .
- the engine 10 A comprises an engine core 16 .
- a bypass duct 18 is defined between the engine core 16 and the nacelle 12 .
- the engine core 16 comprises a high pressure compressor 22 , a combustor 24 , and high and low turbines 26 , 30 respectively.
- Some of the air driven by the fan 14 passes along the bypass duct 18 to be exhausted through an exhaust nozzle 32 to provide a propulsive force.
- the remainder of the air driven by the fan 14 passes through the engine core 16 to the exhaust via the exhaust nozzle 32 .
- the fan 14 , the compressors 23 , 24 and the turbines 26 , 28 , 30 are mounted on rotating discs.
- the nacelle 12 is provided with a force absorbing material 34 provided around the fan 14 and the engine core 16 .
- the force absorbing material 34 is provided to absorb the force from a fragment of a failed disc, and to contain the fragment within the nacelle 10 .
- the nacelle 12 is formed of inner and outer skins 36 , 38 , and the force absorbing material is provided as a layer between the inner and outer skins 36 , 38 on the inner skin 36 .
- the force absorbing material 34 is in the form of poly-paraphenylene terephthalamide.
- a suitable such material is sold under the registered trade mark KEVLAR.
- the force absorbing material is wrapped around the inner skin 36 of the nacelle 12 with appropriate low tension to allow it to deform radially outwardly in the event of impact thereon by a failed fragment of a rotating component such as a compressor or turbine disc.
- the force absorbing material 34 extends substantially the axial length of the engine, to the region upstream of the fan 14 to a region downstream of the low pressure turbine 30 .
- the rotating components of the engines 10 A and 10 B are circumferentially surrounded by the force absorbing material 34 .
- a central region 35 A of the nacelle 12 is not covered by the force absorbing material 34 , because this region is strengthened by struts 40 .
- the central region 35 A of the nacelle 12 generally circumferentially surrounds an intermediate case 41 A from which the struts 40 extend.
- the intermediate case 41 A is provided between the intermediate pressure compressor 20 and the high pressure compressor or 22 .
- a central region 35 B of the nacelle 12 is not covered by the force absorbing material 34 , because this region is strengthened by struts 40 .
- the central region 35 B of the nacelle 12 generally circumferentially surrounds an intermediate case 41 B from which the struts 40 extend.
- the intermediate case 41 B is provided generally upstream of the high pressure compressor 22 and downstream of the fan 14 .
- FIG. 1 shows an off take shaft 42 extending from a main shaft of the engine 10 to the off take gearbox.
- the off-take shaft 42 extends through one of the struts 40 .
- FIG. 2 shows a part sectional perspective view of the nacelle 12 of the gas turbine engine.
- the intermediate case 41 A or 41 B is shown, together with the struts 40 in the form of A frames 44 extending from the intermediate case 41 A or 41 B to the nacelle 12 . It is in the regions where the A frames 44 engage the inner skin 36 of the nacelle 12 that if desired, the nacelle can be devoid of the force absorbing material. However, as shown in FIG. 2 , the force absorbing material can extend without a gap along the inner skin 36 of the nacelle 12 .
- the inner and outer skins 36 , 38 define between them a circumferentially extending space 46 , and within the space 46 are provided a plurality of support members in the form of annular bulkheads 48 , 50 , 52 , 53 each of which extends around the nacelle 12 .
- the bulkheads 48 , 50 , 52 , 53 provide support and strengthening to the nacelle 12 , in that in the event of the force absorbing material 34 being radially outwardly deformed by an impact thereon, the bulkheads 48 , 50 , 52 , 53 resist the deformation of the inner skin to maintain the integrity of the nacelle 12 .
- FIG. 3 which is a close up of the region marked III in FIG. 2 it can be seen that the inner skin 36 comprises sound absorbing material 54 and the bulkhead 48 extends there from radially outwardly, and against which the force absorbing material 34 abuts.
- the nacelle 12 is supported by a pedestal 56 which provides a means of attachment of the engine to the fuselage of an aircraft.
- FIG. 4 which shows a perspective view of the nacelle 12 shown in FIG. 2 , but with the inner and outer skins removed.
- the bulkheads 48 , 50 , 52 , 53 are shown, and it can be seen that the bulkheads 48 , 50 , 52 , 53 extend outwardly by means of extensions 48 A, 50 A, 52 A, 53 A of each respective bulkhead 48 , 50 , 52 , 53 to the pedestal 56 , whereby loads on the bulkheads 48 , 50 , 52 are transmitted to the pedestal 56 , thereby supporting and strengthening the nacelle 12 .
- FIG. 5 there is shown a diagrammatic cross sectional front view of a fuselage 58 of an aircraft in which a pair of engines, 10 X, 10 Y are mounted thereon extending upwardly from the fuselage.
- the off take shaft 42 extends radially outwardly from each engine through the pedestal 56 to a chamber 60 within the fuselage 58 of the aircraft.
- Each of the off take shafts 42 extends to a respective gearbox 61 , upon which are mounted ancillary components 61 A, such as generators, pumps and the like, which allows power to be transferred from the engine 10 X, 10 Y to the necessary regions of the aircraft, for example, to provide electrical power.
- ancillary components 61 A such as generators, pumps and the like
- a trap door 62 provides an opening 63 to allow access from the chamber 60 to the engines 10 X, 10 Y.
- the trapdoor 62 can be connected by a hinge 64 to the fuselage 58 .
- the trap door 62 is connected by releasable hinges 64 to each opposite side of the opening 63 , thus allowing it to be opened in either direction, as shown in solid lines and in broken lines in FIG. 55 .
- the left hand hinge 64 is released to allow the trap door 64 to be swung outwardly towards the engine 10 Y to the position shown in R solid lines in FIG. 5 . In this position, the trapdoor 62 can be used as a work surface by the maintenance worker.
- the trapdoor 62 can be closed and the left hand hinge re-engaged with the trap door 62 .
- the right hand hinge can then be released to allow the trapdoor 62 to be swung outwardly towards the engine 10 X to the position shown in broken lines in FIG. 5 . In this position, the trap door 62 can be used as a work surface, as before.
- Each of the engines 10 a , 10 b can be located in its correct position on the fuselage 58 by dowels 64 provided at opposite end regions of the pedestal 56 at the interface between the pedestal 56 and the fuselage 58 .
- the dowels 64 extend into suitable recesses in the respective pedestal 56 of each of the engines 10 a , 10 b.
- FIG. 6 there is shown a diagrammatic cross sectional front view of a further embodiment of the invention, in which the force absorbing arrangement comprises a corrugated outer skin 138 of the nacelle 12 .
- the force absorbing arrangement comprises a corrugated outer skin 138 of the nacelle 12 .
- the provision of the corrugations in the outer skin 138 of the nacelle allows the outer skin to stretch radially outwardly in the event of impact thereon by a fragment of a failed component.
- the corrugated outer skin 138 can be used in conjunction with a force absorbing material 34 provided as a layer on the inner skin 36 between the inner and outer skins 36 , 138 .
- the force absorbing material 34 is shown in broken lines in FIG. 6 .
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Abstract
Description
- This invention relates to containment systems for gas turbine engines.
- The mounting of gas turbine engines above the fuselage of an aircraft has advantages with respect to noise reduction. However, there may be some issues concerning rotor failures that need to be addressed.
- According to one aspect of this invention, there is provided a containment system for a gas turbine engine, the containment system comprising a force absorbing arrangement for absorbing the force exerted thereon by an ejected portion of a failed component of the engine, wherein the force absorbing arrangement circumferentially surrounds at least a major proportion of the axial length of the engine.
- The aforesaid component may comprise a high energy component. The component may comprise a disk. The component may comprise a compressor or a turbine disk.
- Preferably, the force absorbing arrangement extends substantially the axial length of the engine.
- Preferably, the force absorbing arrangement comprises a force absorbing material. A suitable material is a polyamide, such as an aramid, typically a para-aramid. In one embodiment, the force absorbing material may comprise poly-paraphenylene terephthalamide. An example of such a material is sold under the registered trade mark KEVLAR.
- The engine may include rotary driving components. A casing may surround the rotary driving components. The rotary driving components may comprise a fan, a compressor arrangement, and a turbine arrangement.
- The casing may comprise radially inner and outer skins. The force absorbing arrangement may be provided on the casing, preferably between the inner and outer skins of the casing. In one embodiment, the casing may comprise a nacelle for surrounding the engine. The force absorbing arrangement may be provided as a layer on one of the inner and/or outer skins. Desirably, the force absorbing arrangement is provided on the inner skin.
- Alternatively or in addition, the force absorbing arrangement may comprise a corrugated outer skin of the casing.
- A plurality of support members may be provided within the casing. Each support member may extend circumferentially around the casing, and may extend from the inner skin to the outer skin. In a preferred embodiment, the support members provide strengthening between the inner and outer skins. Each support member may constitute a bulkhead extending from the inner skin to the outer skin of the casing.
- The support members may be spaced from one another axially along the casing. The support members may divide the casing into a plurality of compartments. A pedestal may extend from the casing to mount the engine on an aircraft. In one embodiment, support members extend to the pedestal. Thus, in this embodiment, each support member transmits loads thereon to the pedestal.
- The provision of the support members provides the advantage in an embodiment of the invention that a load exerted on the force absorbing arrangement is constrained by the support members to prevent excessive deformation of the casing and of force absorbing material.
- Where the engine includes a fan, the force absorbing arrangement may comprise a portion of the force absorbing material surrounding the fan, and a further portion of the force absorbing material circumferentially surrounding the remainder of the rotary components.
- Where the gas turbine engine includes an intermediate case, arranged between the high pressure and intermediate pressure compressor, the force absorbing material may comprise a first portion of the force absorbing material extending from a region surrounding the fan to a region surrounding the intermediate pressure compressor of the engine. The force absorbing arrangement may comprise a second portion of the force absorbing material extending from a region surrounding the high pressure compressor to a region surrounding or downstream of the low pressure turbine.
- The force absorbing material may be wrapped around the engine with appropriate tension to allow the force absorbing material to deform radially outwardly within the casing when absorbing the force of the ejected portion of the failed component. The outer skin of the casing may be formed of a corrugated material to allow radially outward deformation thereof if the force absorbing material impacts thereon on radially outwardly formation of the force absorbing material. The engine may be secured to the case by strengthening arrangements. Each strengthening arrangement may comprise an A frame extending from the engine to the casing. In one embodiment, the force absorbing portion may comprise a corrugated outer skin.
- According to another aspect of this invention, there is provided an engine arrangement comprising a gas turbine engine having a containment system as described above: and a pedestal to support the engine on an aircraft.
- Preferably, the containment system includes a casing extending circumferentially around rotary driving components of the engine, as a plurality of support members extending circumferentially around the casing wherein each support member extends to the casing. Thus, loads of support members can be transmitted to the pedestal.
- The engine arrangement may include a chamber externally of the engine to include ancillary components of the engine. The ancillary components may include a power off take arrangement to extract power from a main shaft of the engine. The power off take arrangement may include an off take shaft extending from the aforesaid main shaft of the engine to the chamber, and a gear arrangement to transmit rotary power from the shaft to further components.
- Embodiments of the invention will now be described by way of example only, with reference to the accompanying drawings, in which:
-
FIG. 1A is a sectional side view of a three shaft gas turbine engine, with a nacelle surrounding the engine; -
FIG. 1B is a sectional side view of a two shaft gas turbine engine, with a nacelle surrounding the engine; -
FIG. 2 is a part sectional view of the nacelle shown inFIG. 1 supported by a pedestal; -
FIG. 3 is a view of the region marked III inFIG. 2 ; -
FIG. 4 is a perspective view of the nacelle shown inFIG. 2 indicating support members within the nacelle; -
FIG. 5 is a front sectional view of an aircraft fuselage incorporating two engines; and -
FIG. 6 is a sectional view of an alternative embodiment of a containment system used in a gas turbine engine. - Referring to
FIG. 1A , there is shown a sectional side view of a three shaftgas turbine engine 10A housed in anacelle 12. Theengine 10A comprises afan 14 and anengine core 16. Abypass duct 18 is defined between thenacelle 12 and the engine core. Theengine core 16 comprises anintermediate compressor 20, ahigh pressure compressor 22, acombustor 24 and high, intermediate andlow pressure turbines -
FIG. 1B shows a sectional side view of a two shaftgas turbine engine 10B housed in anacelle 12. Theengine 10A comprises anengine core 16. Abypass duct 18 is defined between theengine core 16 and thenacelle 12. Theengine core 16 comprises ahigh pressure compressor 22, acombustor 24, and high andlow turbines - Some of the air driven by the
fan 14 passes along thebypass duct 18 to be exhausted through anexhaust nozzle 32 to provide a propulsive force. The remainder of the air driven by thefan 14 passes through theengine core 16 to the exhaust via theexhaust nozzle 32. - The
fan 14, thecompressors 23, 24 and theturbines - In each embodiment shown in
FIGS. 1A and 1B , thenacelle 12 is provided with aforce absorbing material 34 provided around thefan 14 and theengine core 16. Theforce absorbing material 34 is provided to absorb the force from a fragment of a failed disc, and to contain the fragment within the nacelle 10. - The
nacelle 12 is formed of inner andouter skins outer skins inner skin 36. - The
force absorbing material 34 is in the form of poly-paraphenylene terephthalamide. A suitable such material is sold under the registered trade mark KEVLAR. - The force absorbing material is wrapped around the
inner skin 36 of thenacelle 12 with appropriate low tension to allow it to deform radially outwardly in the event of impact thereon by a failed fragment of a rotating component such as a compressor or turbine disc. - The
force absorbing material 34 extends substantially the axial length of the engine, to the region upstream of thefan 14 to a region downstream of thelow pressure turbine 30. Thus, the rotating components of theengines force absorbing material 34. - In the embodiment shown in
FIG. 1A , acentral region 35A of thenacelle 12 is not covered by theforce absorbing material 34, because this region is strengthened bystruts 40. Thecentral region 35A of thenacelle 12 generally circumferentially surrounds anintermediate case 41A from which thestruts 40 extend. In the threeshaft engine 10A shown inFIG. 1A , theintermediate case 41A is provided between theintermediate pressure compressor 20 and the high pressure compressor or 22. - In the embodiment shown in
FIG. 1B , acentral region 35B of thenacelle 12 is not covered by theforce absorbing material 34, because this region is strengthened bystruts 40. Thecentral region 35B of thenacelle 12 generally circumferentially surrounds anintermediate case 41B from which thestruts 40 extend. InFIG. 13 theintermediate case 41B is provided generally upstream of thehigh pressure compressor 22 and downstream of thefan 14. - In the above described embodiment, the ancillary components of the engine, such as the power take off gearbox are held within a chamber in the aircraft fuselage, and this is described in more detail below with reference to
FIG. 5 .FIG. 1 shows anoff take shaft 42 extending from a main shaft of the engine 10 to the off take gearbox. The off-take shaft 42 extends through one of thestruts 40. - Referring to
FIGS. 2 and 3 ,FIG. 2 shows a part sectional perspective view of thenacelle 12 of the gas turbine engine. Theintermediate case struts 40 in the form of A frames 44 extending from theintermediate case nacelle 12. It is in the regions where the A frames 44 engage theinner skin 36 of thenacelle 12 that if desired, the nacelle can be devoid of the force absorbing material. However, as shown inFIG. 2 , the force absorbing material can extend without a gap along theinner skin 36 of thenacelle 12. - The inner and
outer skins circumferentially extending space 46, and within thespace 46 are provided a plurality of support members in the form ofannular bulkheads nacelle 12. Thebulkheads nacelle 12, in that in the event of theforce absorbing material 34 being radially outwardly deformed by an impact thereon, thebulkheads nacelle 12. - Referring to
FIG. 3 , which is a close up of the region marked III inFIG. 2 it can be seen that theinner skin 36 comprisessound absorbing material 54 and thebulkhead 48 extends there from radially outwardly, and against which theforce absorbing material 34 abuts. - The
nacelle 12 is supported by apedestal 56 which provides a means of attachment of the engine to the fuselage of an aircraft. - Referring to
FIG. 4 , which shows a perspective view of thenacelle 12 shown inFIG. 2 , but with the inner and outer skins removed. InFIG. 4 , thebulkheads bulkheads extensions respective bulkhead pedestal 56, whereby loads on thebulkheads pedestal 56, thereby supporting and strengthening thenacelle 12. - Referring to
FIG. 5 , there is shown a diagrammatic cross sectional front view of a fuselage 58 of an aircraft in which a pair of engines, 10X, 10Y are mounted thereon extending upwardly from the fuselage. As can be seen, the off takeshaft 42 extends radially outwardly from each engine through thepedestal 56 to a chamber 60 within the fuselage 58 of the aircraft. Each of the off takeshafts 42 extends to a respective gearbox 61, upon which are mounted ancillary components 61A, such as generators, pumps and the like, which allows power to be transferred from the engine 10X, 10Y to the necessary regions of the aircraft, for example, to provide electrical power. The arrangement of such ancillary components in the chamber 60 provides the advantage that it is no longer necessary for a maintenance worker to carry out work on the ancilliary components 61A within the engine. - A trap door 62 provides an opening 63 to allow access from the chamber 60 to the engines 10X, 10Y. The trapdoor 62 can be connected by a hinge 64 to the fuselage 58.
- The trap door 62 is connected by releasable hinges 64 to each opposite side of the opening 63, thus allowing it to be opened in either direction, as shown in solid lines and in broken lines in
FIG. 55 . - When it is desired to carry out maintenance on the right hand engine 10Y (as shown in
FIG. 5 ), the left hand hinge 64 is released to allow the trap door 64 to be swung outwardly towards the engine 10Y to the position shown in R solid lines inFIG. 5 . In this position, the trapdoor 62 can be used as a work surface by the maintenance worker. - If the maintenance worker desires to work on the left hand engine 10X, the trapdoor 62 can be closed and the left hand hinge re-engaged with the trap door 62. The right hand hinge can then be released to allow the trapdoor 62 to be swung outwardly towards the engine 10X to the position shown in broken lines in
FIG. 5 . In this position, the trap door 62 can be used as a work surface, as before. - Each of the engines 10 a, 10 b can be located in its correct position on the fuselage 58 by dowels 64 provided at opposite end regions of the
pedestal 56 at the interface between thepedestal 56 and the fuselage 58. The dowels 64 extend into suitable recesses in therespective pedestal 56 of each of the engines 10 a, 10 b. - Referring to
FIG. 6 , there is shown a diagrammatic cross sectional front view of a further embodiment of the invention, in which the force absorbing arrangement comprises a corrugatedouter skin 138 of thenacelle 12. The provision of the corrugations in theouter skin 138 of the nacelle allows the outer skin to stretch radially outwardly in the event of impact thereon by a fragment of a failed component. It will be appreciated that the corrugatedouter skin 138 can be used in conjunction with aforce absorbing material 34 provided as a layer on theinner skin 36 between the inner andouter skins force absorbing material 34 is shown in broken lines inFIG. 6 . - There is thus described a simple and effective way to contain fragments of failed components of a gas turbine engine to prevent the failed components bursting from the engine and causing damage elsewhere, for example to the fuselage, wings or tail fins of the aeroplane, or to the adjacent other engine. A further modification is that other force absorbing materials could be used, for example metallic or non-metallic structures, knitted structures, feather type armour, or bulk head containment. In the further modification, a tie member could extend between the engines to share the load.
- Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.
Claims (26)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0602362A GB2434837B (en) | 2006-02-07 | 2006-02-07 | A containment system for a gas turbine engine |
GB0602362.6 | 2006-02-07 |
Publications (2)
Publication Number | Publication Date |
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US7806364B1 US7806364B1 (en) | 2010-10-05 |
US20100251688A1 true US20100251688A1 (en) | 2010-10-07 |
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Application Number | Title | Priority Date | Filing Date |
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US11/651,553 Active 2029-01-28 US7806364B1 (en) | 2006-02-07 | 2007-01-10 | Containment system for a gas turbine engine |
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US (1) | US7806364B1 (en) |
GB (1) | GB2434837B (en) |
Families Citing this family (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2459646B (en) * | 2008-04-28 | 2011-03-30 | Rolls Royce Plc | A fan assembly |
US8092169B2 (en) * | 2008-09-16 | 2012-01-10 | United Technologies Corporation | Integrated inlet fan case |
US8827629B2 (en) * | 2011-02-10 | 2014-09-09 | United Technologies Corporation | Case with ballistic liner |
FR2961483B1 (en) * | 2010-06-18 | 2013-01-18 | Aircelle Sa | AIRCRAFT TURBOKIN ENGINE BLOWER HOUSING |
US9534505B2 (en) | 2012-07-23 | 2017-01-03 | United Technologies Corporation | Integrated nacelle inlet and metallic fan containment case |
FR2994942B1 (en) * | 2012-09-06 | 2015-08-07 | Airbus Operations Sas | LATERAL PROPULSIVE ASSEMBLY FOR AIRCRAFT COMPRISING A SUPPORT ARM OF A TURBOMOTEUR. |
US9376935B2 (en) | 2012-12-18 | 2016-06-28 | Pratt & Whitney Canada Corp. | Gas turbine engine mounting ring |
US10550718B2 (en) | 2017-03-31 | 2020-02-04 | The Boeing Company | Gas turbine engine fan blade containment systems |
US10487684B2 (en) | 2017-03-31 | 2019-11-26 | The Boeing Company | Gas turbine engine fan blade containment systems |
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US2608056A (en) * | 1950-11-07 | 1952-08-26 | A V Roe Canada Ltd | Power take-off from the forward end of aircraft propulsive power units |
US3093361A (en) * | 1958-07-07 | 1963-06-11 | Bristol Siddeley Engines Ltd | Engines |
US3540682A (en) * | 1964-12-02 | 1970-11-17 | Gen Electric | Turbofan type engine frame and support system |
US3735946A (en) * | 1970-07-09 | 1973-05-29 | Rolls Royce | Aircraft engine mountings |
US3974313A (en) * | 1974-08-22 | 1976-08-10 | The Boeing Company | Projectile energy absorbing protective barrier |
US5694765A (en) * | 1993-07-06 | 1997-12-09 | Rolls-Royce Plc | Shaft power transfer in gas turbine engines with machines operable as generators or motors |
US6575694B1 (en) * | 2000-08-11 | 2003-06-10 | Rolls-Royce Plc | Gas turbine engine blade containment assembly |
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Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB2262313B (en) * | 1991-12-14 | 1994-09-21 | Rolls Royce Plc | Aerofoil blade containment |
GB2281941B (en) * | 1993-09-15 | 1996-05-08 | Rolls Royce Plc | Containment structure |
GB2288639B (en) * | 1994-04-20 | 1998-10-21 | Rolls Royce Plc | Ducted fan gas turbine engine nacelle assembly |
GB2365925B (en) * | 2000-08-11 | 2005-02-23 | Rolls Royce Plc | A gas turbine engine blade containment assembly |
-
2006
- 2006-02-07 GB GB0602362A patent/GB2434837B/en active Active
-
2007
- 2007-01-10 US US11/651,553 patent/US7806364B1/en active Active
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2608056A (en) * | 1950-11-07 | 1952-08-26 | A V Roe Canada Ltd | Power take-off from the forward end of aircraft propulsive power units |
US3093361A (en) * | 1958-07-07 | 1963-06-11 | Bristol Siddeley Engines Ltd | Engines |
US3540682A (en) * | 1964-12-02 | 1970-11-17 | Gen Electric | Turbofan type engine frame and support system |
US3735946A (en) * | 1970-07-09 | 1973-05-29 | Rolls Royce | Aircraft engine mountings |
US3974313A (en) * | 1974-08-22 | 1976-08-10 | The Boeing Company | Projectile energy absorbing protective barrier |
US5694765A (en) * | 1993-07-06 | 1997-12-09 | Rolls-Royce Plc | Shaft power transfer in gas turbine engines with machines operable as generators or motors |
US6575694B1 (en) * | 2000-08-11 | 2003-06-10 | Rolls-Royce Plc | Gas turbine engine blade containment assembly |
Also Published As
Publication number | Publication date |
---|---|
GB2434837B (en) | 2008-04-09 |
US7806364B1 (en) | 2010-10-05 |
GB0602362D0 (en) | 2006-03-15 |
GB2434837A (en) | 2007-08-08 |
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