GB2365926A - Gas turbine engine blade containment with spirally arranged corrugated sheet material - Google Patents
Gas turbine engine blade containment with spirally arranged corrugated sheet material Download PDFInfo
- Publication number
- GB2365926A GB2365926A GB0019803A GB0019803A GB2365926A GB 2365926 A GB2365926 A GB 2365926A GB 0019803 A GB0019803 A GB 0019803A GB 0019803 A GB0019803 A GB 0019803A GB 2365926 A GB2365926 A GB 2365926A
- Authority
- GB
- United Kingdom
- Prior art keywords
- gas turbine
- turbine engine
- blade containment
- containment assembly
- corrugations
- 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.)
- Granted
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- 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
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- 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
- F05D2250/10—Two-dimensional
- F05D2250/15—Two-dimensional spiral
-
- 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
- F05D2250/20—Three-dimensional
- F05D2250/25—Three-dimensional helical
-
- 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
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
An arrangement for containment of detached portions of fan, compressor or turbine blades in a gas turbine engine, comprises at least one corrugated sheet metal spiral 54,56. The corrugations extend with axial and/or circumferential components. One embodiment has two sheets spirally wound with corrugations arranged at opposite angles a , a 2. Other embodiments (fig 12-14) has corrugations extending purely circumferentially or axially. The spirals may surround a generally annular or frustoconical thin metallic inner casing 52, be joined to the casing and flanges 40 at their axial ends, have apertures 63 for acoustic attenuation, foam filling or be surrounded by strong fibrous material 64. The flange 40 may have a fence or hook (45, fig 8) to prevent forward movement of the blade, or a purely circumferential corrugation (59, fig 10) upstream of the rings. Suitable materials for the sheets are disclosed, including steel.
Description
<Desc/Clms Page number 1>
A GAS TURBINE ENGINE BLADE CONTAINMENT ASSEMBLY The present invention relates to gas turbine engine casings, particularly gas turbine engine fan casings and turbine casings, more particularly to an improved blade containment assembly for use within or forming a part of the gas turbine engine casing.
Turbofan gas turbine engines for powering aircraft conventionally comprise a core engine, which drives a fan. The fan comprises a number of radially extending fan blades mounted on a fan rotor enclosed by a generally cylindrical fan casing. The core engine comprises one or more turbines, each one of which comprises a number of radially extending turbine blades enclosed by a cylindrical, or frustoconical, casing.
There is a remote possibility with such engines that part, or all, of a fan blade, or a turbine blade, could become detached from the remainder of the fan or turbine. In the case of a fan blade becoming detached this may occur as the result off for example, the turbofan gas turbine engine ingesting a bird or other foreign object.
The use of containment rings for turbofan gas turbine engine casings is well known. It is known to provide generally cylindrical, or frustoconical, relatively thick metallic containment rings. It is known to provide generally cylindrical, or frustoconical, locally thickened, isogrid, metallic containment rings. It is known to provide strong fibrous material wound around relatively thin metallic casinqs or around the above mentioned containment casings. In the event that a blade becomes detached it passes through the casing and is contained by the fibrous material.
However, the relatively thick containment casings are relatively heavy, the relatively thin casings enclosed by the fibrous material are lighter but are more expensive to manufacture. The relatively thick casings with fibrous material are both heavier and more expensive to manufacture.
<Desc/Clms Page number 2>
Accordingly the present invention seeks to provide a novel gas turbine engine casing which overcomes the above mentioned problems.
Accordingly the present invention provides a gas turbine engine blade containment assembly comprising a generally cylindrical, or frustoconical, casing, the casing being arranged in operation to surround a rotor carrying a plurality of radially extending rotor blades, at least one corrugated metal sheet wound into a spiral surrounding the casing, wherein the corrugations of the at least one corrugated metal sheet wound into a spiral extend with axial and/or circumferential components.
The casing may be a fan casing or a turbine casing. Preferably the corrugations are equi-spaced.
The corrugations in the at least one corrugated metal sheet wound into a spiral may extend with purely circumferential components.
The corrugations in the at least one corrugated metal sheet wound into a spiral may extend with purely axial components. Preferably the corrugations in the at least one corrugated metal sheet wound into a spiral may extend with both axial and circumferential components.
The casing may comprise a single metal sheet wound into a spiral.
Preferably the casing comprises a plurality of metal sheets, each of which is wound into a spiral and each one of which is corrugated.
Preferably the corrugations in different metal sheets are arranged to extend at different angles.
The corrugations in a first metal sheet may be arranged to extend with purely circumferential components and the corrugations in a second metal sheet are arranged to extend with purely axial components.
The corrugations in a first metal sheet may be arranged to extend with purely circumferential components and the
<Desc/Clms Page number 3>
corrugations in a second metal sheet are arranged to extend with both axial and circumferential components.
Preferably the corrugations in a first metal sheet are arranged to extend with both axial and circumferential components and the corrugations in a second metal sheet are arranged to extend with both axial and circumferential components.
The at least one metal sheet wound into a spiral may be provided with apertures therethrough to attenuate noise.
The plurality of metal sheets wound into spirals define spaces therebetween, the spaces may be filled with a energy absorbing material to increase the blade containment capability of the casing.
The at least one metal sheet wound into a spiral defines spaces therebetween, the spaces may be filled with a energy absorbing material to increase the blade containment capability of the casing.
The at least one metal sheet may be formed from titanium, an alloy of titanium, aluminium or steel.
The present invention will be more fully described by way of example with reference to the accompanying drawings in which:- Figure 1 is a partially cut away view of a gas turbine engine having a fan blade containment assembly according to the present invention.
Figure 2 is an enlarged view of the fan blade containment assembly shown in figure 1.
Figure 3 is a further enlarged view of the fan blade containment assembly shown in figure 2.
Figure 4 is a cross-sectional view in the direction of arrows A-A in figure 3.
Figure 5 is a view in the direction of arrow B in figure 3.
Figure 6 is a cut away view in the direction of arrow C in figure 3.
<Desc/Clms Page number 4>
Figure 7 is a cut away perspective view of the fan blade containment assembly shown in figure 3.
Figure 8 is a cross-sectional view of an alternative attachment of the fan blade containment assembly to the fan casing. Figure 9 is a cross-sectional view of a further attachment of the fan blade containment assembly to the fan casing.
Figure 10 is a cut away perspective view of a single sheet fan blade containment assembly according to the present invention.
Figure 11 is an alternative view in the direction of arrow B in figure 3.
Figure 12 is an enlarged view of an alternative fan blade containment assembly shown in figure 1.
Figure 13 is a further enlarged view of the fan blade containment assembly shown in figure 12.
Figure 14 is a view in the direction of arrow D in figure 13.
A turbofan gas turbine engine 10, as shown in figure 1, comprises in flow series an intake 12, a fan section 14, a compressor section 16, a combustor section 18, a turbine section 20 and an exhaust 22. The turbine section 20 comprises one or more turbines arranged to drive one or more compressors in the compressor section 16 via shafts. The turbine section 20 also comprises a turbine to drive the fan section 14 via a shaft. The fan section 14 comprises a fan duct 24 defined partially by a fan casing 26. The f an duct 24 has an outlet 28 at its axially downstream end. The fan casing 26 is secured to the core engine casing 36 by a plurality of radially extending fan outlet guide vanes 30. The fan casing 26 surrounds a fan rotor 32, which carries a plurality of circumferentially spaced radially extending fan blades 34. The fan casing 26 also comprises a fan blade containment assembly 38, which is arranged substantially in the plane of the fan blades 34.
<Desc/Clms Page number 5>
The fan casing 26 and fan blade containment assembly 38 is shown more clearly in figures 2 to 7. The f an blade containment assembly 38 comprises an upstream flange 40 by which the fan blade containment assembly 38 is connected to a flange 46 on an intake assembly 44 of the fan casing 26. The fan blade containment assembly 38 has a downstream flange 42 by which the fan blade containment assembly 38 is connected to a flange 50 on a rear portion 48 of the fan casing 26.
The fan blade containment assembly 38, as shown more clearly in figures 3 and 4, comprises a relatively thin metallic cylindrical, or frustoconicai, casing 52 and a plurality of, two in this example, relatively thin corrugated metallic sheets 54 and 56. The thin corrugated metallic sheets 54 and 56 are wound together into spirals around the casing 52 and the circumferential ends of the thin corrugated metallic sheet 54 and 56 are joined to the casing 52 by for example welding, brazing, nuts and bolts or other mechanical fasteners. The thin corrugated metallic sheets 54 and 56 are wound around the thin metallic casing 52 more than once, preferably a number of times. The axial ends of the thin corrugated metallic sheets 54 and 56 are joined to each other and the casing 52 by welding or other suitable means or retained by band clamps. The axial ends of the casing 52 are provided with the flanges 40 and 42. The thin corrugated metallic sheets 54 and 56 are arranged to abut each other at circumferentially and axially spaced locations where the corrugations 62 contact. The thin corrugated metallic sheets 54 and 56 are spot welded, or seam welded, together at the spaced locations where the corrugations 62 contact to improve the rigidity or integrity of the fan blade containment assembly 38. The corrugations 62 of the thin corrugated metallic sheets 54 and 56 are shown more clearly in figures 4, 5, 6 and 7.
In some circumstances the welds between the corrugations 62 of the thin corrugated metallic sheets may not be required.
<Desc/Clms Page number 6>
The corrugations 62 of the thin corrugated metallic sheets 54 and 56 are arranged to extend with both axial and circumferential components. Additionally the corrugations 62 on the adjacent thin corrugated metallic sheets 54 and 56 are arranged at different angles. For example the corrugations 62 on metallic sheet 54 are arranged at an angle cc to the axis X of the gas turbine engine. The corrugations 62 on metallic sheet 56 are arranged at an angle cc2 to the axis X of the gas turbine engine. The angles oc and oc2 are the same, 450 in this example, but angles oc and oc2 are in the opposite directions. It would of course be possible to use any suitable combinations of angles x and oc2, but at least one of the thin metallic sheets 54 and 56 must be arranged to have a component in the circumferential direction.
The thin casing 52 and the thin corrugated metallic sheets 54 and 56 are provided with apertures 63 to provide acoustic attenuation of sounds generated in the gas turbine engine 10. The corrugations 62 of the thin corrugated metallic sheets 54 and 56 defines spaces 61 therebetween and the spaces 61 may be filled with an energy absorbing material, for example foam, to further increase the energy absorbing capability of the fan blade containment assembly 38. It may be desirable in some circumstances to provide a number of continuous layers of a strong fibrous material 64 wound around the thin corrugated metallic sheets 54 and 56 to further increase the energy absorbing capability of the fan blade containment assembly 38. The strong fibrous material may for example be woven aromatic polyamide fibres known as KEVLAR (KEVLAR is a registered trademark of Dupont Ltd) . There may also be a number of layers of discrete pieces of f lexible material woven from KEVLAR between the thin corrugated metallic sheets 54 and 56 and the continuous layers of fibrous material 64.
Figure 8 shows an attachment of the fan blade containment assembly 38 to the flanges 40 and 42. The edges
<Desc/Clms Page number 7>
of the thin metallic sheets 54 and 56 are mechanically fastened by nuts and bolts 47 to the flanges 40 and 42. However, welding, brazing or other suitable fastening may be used. It is to be noted that a fence, or hook, 45 is provided on the flange 40 to prevent forward movement of the tip of the fan blades 34 in the event of a fan blade off situation.
Figure 9 shows a corrugation 58 extending with a purely circumferential component at the upstream end of the fan blade containment assembly 38 to attach the fan blade containment assembly 38 to the flange 40.
Figure 10 shows a single thin corrugated metallic sheet wound into a spiral to form the fan blade containment assembly 38. The corrugations 62 extend with at least a circumferential component.
Figure 11 shows an alternative view of the fan blade containment assembly 38 in which the thin corrugated metallic sheet 56 has the corrugation 62 extending with a pure axial component, but the thin corrugated metallic sheet 52 has corrugations 62 extending with both circumferential and axial components.
Figures 13 and 14 show an alternative fan blade containment assembly 38B which comprises a relatively thin metallic cylindrical, or frustoconical, casing 52 and a plurality of, two, relatively thin corrugated metallic sheets 54 and 56. The thin corrugated metallic sheets 54 and 56 are wound around the thin metallic casing 52 at least once, preferably a number of times. The thin corrugated metallic sheets 54 and 56 are wound together into spirals around the casing 52 and the ends of the thin corrugated metallic sheet 54 and 56 are joined to the casing 52 by for example welding, brazing, nuts and bolts or other mechanical fasteners. The axial ends of the thin corrugated metallic sheets 54 and 56 are joined to each other and the casing 52 by welding or other suitable means or retained by band clamps. The axial ends of the casing 52 are provided with the flanges 40 and
<Desc/Clms Page number 8>
42. The thin corrugated metallic sheets 54 and 56 are arranged to abut each other at circumferentially and axially spaced locations where the corrugations 62 contact. The thin corrugated metallic sheets 54 and 56 are spot welded, or seam welded, together at the spaced locations where the corrugations 62 contact. This improves the rigidity or integrity of the fan blade containment assembly 38. The corrugations 62 of the thin corrugated metallic sheets 54 and 56 are shown more clearly in figures 14.
The corrugations 62 of the thin corrugated metallic sheets 54 and 56 are arranged to extend with pure circumferential components.
The thin casing 52 and the thin corrugated metallic sheets 54 and 56 are provided with apertures 61 to provide acoustic attenuation of sounds generated in the gas turbine engine 10. The corrugations 62 of the thin corrugated metallic sheets 54 and 56 defines spaces 61 therebetween and the spaces 61 may be filled with an energy absorbing material, for example foam, to further increase the energy absorbing capability of the fan blade containment assembly 38 . It may be desirable in some circumstances to provide a number of continuous layers of a strong fibrous material 64 wound around the thin corrugated metallic sheets 54 and 56 to further increase the energy absorbing capability of the fan blade containment assembly 38. The strong fibrous material may for example be woven aromatic poiyamide fibres known as KEVLAR (KEVLAR is a registered trademark of Dupont Ltd). There may also be a number of layers of discrete pieces of flexible material woven from KEVLAR between the thin corrugated metallic sheets 54 and 56 and the continuous layers of fibrous material 64.
The thin casing 52 and the thin corrugated metallic sheets 54 and 56 have a thickness of about 1-3mm, preferably 2mm, compared to the normal thickness of 12mm for a fan blade containment casing. This enables the weight of the fan blade
<Desc/Clms Page number 9>
containment assembly to be reduced. Additionally it may allow the use of the fibrous material containment to be dispensed with.
In operation of the gas turbine engine 10, in the event that a fan blade 34, or a portion of a fan blade 34, becomes detached it pierces the thin metallic casing 52, before encountering the thin corrugated metallic sheets 54 and 56. The thin corrugated metallic sheets 54 and 56 are impacted by the fan blade 34, or portion of the fan blade 34, and effectively remove energy from the fan blade 34, or portion of the fan blade 34.
Each of the turns of the thin corrugated metallic sheets 54 and 56 has relatively low mass and hence relatively low inertia. This allows the thin corrugated metallic sheets 54 and 56 to move with the detached fan blade 34, or portion of the fan blade 34. This movement spreads the impact energy over a larger area of the fan blade containment assembly 38 enabling the use of lower mass of material to contain the detached fan blade 34, or fan blade portion 34.
The detached fan blade 34, or portion of the fan blade 34, causes the corrugations 62 in the thin corrugated metallic sheets 54 and 56 to be straightened out and this process absorbs energy from the detached fan blade 34 or portion of a fan blade 34. As the corrugations 62 are straightened out in each turn of the thin corrugated metallic sheets 54 and 56, the adjacent thin corrugated metallic sheets 54 and 56 slide over each other and absorb more energy from the fan blade 34 by friction between the adjacent thin corrugated metallic sheets 54 and 56. As the corrugations 63 are straightened the welds between corrugations 62 on adjacent thin corrugated metallic sheets 54 and 56 are broken also absorbing energy. As each turn of the thin corrugated metallic sheets 54 and 56 straightens over the impact region it stiffens locally and transfers load to material further from the impact region, this increases the proportion of the fan blade assembly 36 contributing to energy absorption.
<Desc/Clms Page number 10>
The corrugations lead to a low-density structure with a greater stiffness to weight ratio than a solid casing of the same material.
The orientation of the corrugations relative to the axis of the gas turbine engine allows the elongation axially and circumferentially to be adjusted to an optimum for fan blade containment.
The use of a plurality of thin corrugated metallic sheets with the corrugations arranged at different angles to the axis of the gas turbine engine to increase the torsional rigidity of the fan blade containment assembly and/or to ensure consistent spacing between the thin corrugated metallic sheets. The use of a plurality of thin corrugated metallic sheets provides high integrity through the alternative load paths and hence damage tolerance.
The thin corrugated metallic sheets are easy to produce by passing thin metallic sheets through shaped rollers to form the corrugations.
The thin metallic sheet may be lower cost material because defects are easier to detect in thin metallic sheets and/or the defects have less significance due to the multiple turns of the thin corrugated metallic sheet(s).
The thin corrugated metallic sheets may be manufactured from titanium, titanium alloy, aluminium, aluminium alloy, nickei, nickel alloy, titanium aiuminide, nickel aluminide and steel.
The spacing between the corrugations and the radial height of the corrugations in the thin corrugated metallic sheets is selected to provide optimum energy absorption.
The invention has been described with reference to a fan blade containment assembly, however it is equally applicable to a compressor blade containment assembly and a turbine blade containment assembly.
The description has referred to the use of a thin metallic casing around which the thin corrugated metallic
<Desc/Clms Page number 11>
sheets are wound, the thin metallic casing may not be required in some circumstances.
<Desc/Clms Page number 12>
Claims (1)
- Claims: - 1. A gas turbine engine blade containment assembly comprising a generally cylindrical, or frustoconical, casing, the casing being arranged in operation to surround a rotor carrying a plurality of radially extending rotor blades, at least one corrugated metal sheet wound into a spiral surrounding the casing, wherein the corrugations of the at least one corrugated metal sheet wound into a spiral extend with axial and/or circumferential components. 2. A gas turbine engine blade containment assembly as claimed in claim 1 wherein the casing is a fan casing and the rotor blades are fan blades. 3. A gas turbine engine blade containment assembly as claimed in claim 1 wherein the casing is a turbine casing and the rotor blades are turbine blades. 4. A gas turbine engine blade containment assembly as claimed in claim 1, claim 2 or claim 3 wherein the corrugations are equi-spaced. S. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 4 wherein the corrugations in the 9t least one corrugated metal sheet wound into a spiral extend with purely circumferential components. 6. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 4 wherein the corrugations in the at least- one corrugated metal sheet wound into a spiral extend with both axial and circumferential components. 7. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 6 wherein the casing comprises a single metal sheet wound into a spiral. 8. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 6 wherein the casing comprises a plurality of metal sheets, each of which is wound into a spiral and each one of which is corrugated. 9. A gas turbine engine blade containment assembly as claimed in claim 8 wherein the corrugations in different metal sheets are arranged to extend at different angles.<Desc/Clms Page number 13>10. A gas turbine engine blade containment assembly as claimed in claim 9 wherein the corrugations in a first metal sheet are arranged to extend with purely circumferential components and the corrugations in a second metal sheet are arranged to extend with purely axial components. 11. A gas turbine engine blade containment assembly as claimed in claim 9 wherein the corrugations in a first metal sheet are arranged to extend with purely circumferential components and the corrugations in a second metal sheet are arranged to extend with both axial and circumferential components. 12. A gas turbine engine blade containment assembly as claimed in claim 9 wherein the corrugations in a first metal sheet are arranged to extend with both axial and circumferential components and the corrugations in a second metal sheet are arranged to extend with both axial and circumferential components. 13. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 12 wherein the at least one metal sheet wound into a spiral is provided with apertures therethrough to attenuate noise. 14. A gas turbine engine blade containment assembly as claimed in any of claims 9 to 12 wherein the plurality of metal sheets wound into spirals define spaces therebetween, the spaces are filled with a energy absorbing material to increase the blade containment capability of the casing. 15. A gas turbine engine blade containment assembly as claimed in claim 7 wherein the at least one metal sheet wound into a spiral defines spaces therebetween, the spaces are filled with a energy absorbing material to increase the blade containment capability of the casing. 16. A gas turbine engine blade containment assembly as claimed in any of claims 1 to 15 wherein the at least one metal sheet is formed from titanium, an alloy of tiltanium, aluminium or steel.<Desc/Clms Page number 14>17. A gas turbine engine blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 1 to 7 of the accompanying drawings. 18. A gas turbine engine blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 12 to 14 of the accompanying drawings. 19. A gas turbine engine comprising a blade containment assembly as claimed in any of claims 1 to 18.
Priority Applications (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0019803A GB2365926B (en) | 2000-08-12 | 2000-08-12 | A gas turbine engine blade containment assembly |
| GBGB0116988.7A GB0116988D0 (en) | 2000-08-11 | 2001-07-12 | A gas turbine engine blade containment assembly |
| US09/924,104 US6575694B1 (en) | 2000-08-11 | 2001-08-08 | Gas turbine engine blade containment assembly |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0019803A GB2365926B (en) | 2000-08-12 | 2000-08-12 | A gas turbine engine blade containment assembly |
Publications (3)
| Publication Number | Publication Date |
|---|---|
| GB0019803D0 GB0019803D0 (en) | 2000-09-27 |
| GB2365926A true GB2365926A (en) | 2002-02-27 |
| GB2365926B GB2365926B (en) | 2004-12-08 |
Family
ID=9897450
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB0019803A Expired - Fee Related GB2365926B (en) | 2000-08-11 | 2000-08-12 | A gas turbine engine blade containment assembly |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2365926B (en) |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1431522A2 (en) | 2002-12-20 | 2004-06-23 | Rolls-Royce Deutschland Ltd & Co KG | Containment ring for the fan casing of a gas turbine engine |
| US8186934B2 (en) | 2007-03-14 | 2012-05-29 | Rolls-Royce Plc | Casing assembly |
| GB2501918A (en) * | 2012-05-11 | 2013-11-13 | Rolls Royce Plc | Containment case for a gas turbine engine |
| US8591172B2 (en) | 2009-09-25 | 2013-11-26 | Rolls-Royce Plc | Containment casing for an aero engine |
| EP3139007A1 (en) * | 2015-09-07 | 2017-03-08 | MTU Aero Engines GmbH | Device for limiting a flow channel of a turbomachine |
| GB2513545B (en) * | 2010-08-12 | 2017-10-11 | Gen Electric | Fragment containment assembly and method for adding a fragment containment assembly to a turbine |
| CN110566293A (en) * | 2018-06-06 | 2019-12-13 | 曼恩能源方案有限公司 | Fracture protection device for turbomachinery |
| CN113446121A (en) * | 2020-03-26 | 2021-09-28 | 和谐工业有限责任公司 | Air turbine starter containment system |
| CN113446122A (en) * | 2020-03-26 | 2021-09-28 | 和谐工业有限责任公司 | Air turbine starter containment system |
| EP4265889A1 (en) * | 2022-04-21 | 2023-10-25 | Pratt & Whitney Canada Corp. | Multi-layered containment structure for a bladed rotor of a gas turbine engine |
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| GB1533017A (en) * | 1975-11-10 | 1978-11-22 | Caterpillar Tractor Co | Modular gas turbine engine assembly |
| US4762462A (en) * | 1986-11-26 | 1988-08-09 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) | Housing for an axial compressor |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4618152A (en) * | 1983-01-13 | 1986-10-21 | Thomas P. Mahoney | Honeycomb seal structure |
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| GB1500135A (en) * | 1973-02-23 | 1978-02-08 | Int Harvester Co | Seals |
| US3974313A (en) * | 1974-08-22 | 1976-08-10 | The Boeing Company | Projectile energy absorbing protective barrier |
| GB1533017A (en) * | 1975-11-10 | 1978-11-22 | Caterpillar Tractor Co | Modular gas turbine engine assembly |
| US4762462A (en) * | 1986-11-26 | 1988-08-09 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) | Housing for an axial compressor |
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1431522A2 (en) | 2002-12-20 | 2004-06-23 | Rolls-Royce Deutschland Ltd & Co KG | Containment ring for the fan casing of a gas turbine engine |
| EP1431522A3 (en) * | 2002-12-20 | 2006-07-19 | Rolls-Royce Deutschland Ltd & Co KG | Containment ring for the fan casing of a gas turbine engine |
| US8186934B2 (en) | 2007-03-14 | 2012-05-29 | Rolls-Royce Plc | Casing assembly |
| US8591172B2 (en) | 2009-09-25 | 2013-11-26 | Rolls-Royce Plc | Containment casing for an aero engine |
| GB2513545B (en) * | 2010-08-12 | 2017-10-11 | Gen Electric | Fragment containment assembly and method for adding a fragment containment assembly to a turbine |
| US9429039B2 (en) | 2012-05-11 | 2016-08-30 | Rolls-Royce Plc | Casing |
| GB2501918B (en) * | 2012-05-11 | 2014-06-18 | Rolls Royce Plc | Casing |
| GB2501918A (en) * | 2012-05-11 | 2013-11-13 | Rolls Royce Plc | Containment case for a gas turbine engine |
| EP3139007A1 (en) * | 2015-09-07 | 2017-03-08 | MTU Aero Engines GmbH | Device for limiting a flow channel of a turbomachine |
| CN110566293A (en) * | 2018-06-06 | 2019-12-13 | 曼恩能源方案有限公司 | Fracture protection device for turbomachinery |
| CN113446121A (en) * | 2020-03-26 | 2021-09-28 | 和谐工业有限责任公司 | Air turbine starter containment system |
| CN113446122A (en) * | 2020-03-26 | 2021-09-28 | 和谐工业有限责任公司 | Air turbine starter containment system |
| EP4265889A1 (en) * | 2022-04-21 | 2023-10-25 | Pratt & Whitney Canada Corp. | Multi-layered containment structure for a bladed rotor of a gas turbine engine |
Also Published As
| Publication number | Publication date |
|---|---|
| GB2365926B (en) | 2004-12-08 |
| GB0019803D0 (en) | 2000-09-27 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PCNP | Patent ceased through non-payment of renewal fee |
Effective date: 20180812 |