GB2361032A - A gas turbine engine blade containment assembly - Google Patents
A gas turbine engine blade containment assembly Download PDFInfo
- Publication number
- GB2361032A GB2361032A GB0008189A GB0008189A GB2361032A GB 2361032 A GB2361032 A GB 2361032A GB 0008189 A GB0008189 A GB 0008189A GB 0008189 A GB0008189 A GB 0008189A GB 2361032 A GB2361032 A GB 2361032A
- Authority
- GB
- United Kingdom
- Prior art keywords
- gas turbine
- turbine engine
- metal casing
- containment assembly
- rotor blade
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Withdrawn
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- 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
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
- B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
- B23K26/00—Working by laser beam, e.g. welding, cutting or boring
- B23K26/352—Working by laser beam, e.g. welding, cutting or boring for surface treatment
- B23K26/356—Working by laser beam, e.g. welding, cutting or boring for surface treatment by shock processing
-
- 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
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
- F04D27/02—Surge control
- F04D27/0292—Stop safety or alarm devices, e.g. stop-and-go control; Disposition of check-valves
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Optics & Photonics (AREA)
- Plasma & Fusion (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
A gas turbine engine fan blade containment assembly 38 comprises a cylindrical, or frustoconical metal casing 40 having a compressive residual stress in its outer 41 and/or inner 43 surfaces. The deep compressive residual stress provides a compressive pre-stress which has to be overcome by tensile force produced by impact of a fan blade 34, removing more energy from a blade than an un-prestressed casing of equivalent thickness. Provides improved blade containment or equivalent containment with reduced casing thickness and weight. The residual stress may be introduced by laser shock peening. The whole of either surface may have a compressive stress or the surfaces may have a stress pattern e.g. an isogrid. Containment enhancing fibrous material may be wound onto the casing which may have ribs, flanges or an acoustic lining.
Description
2361032 1 A GAS TURBINE ENGINE BLADE CONTAINMENT ASSEMBLY AND A METMOD OF
MANUFACTURING A GAS TURBINE ENGINE BLADE CONTAINMENT ASSEMBLY The present invention relates to gas turbine engine casings, particularly gas turbine engine fan cas"ings, 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, which is enclosed by a generally cylindrical, or frustoconical, 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 generally cylindrical, or frustoconical, turbine casing.
There is a remote possibility that 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 of, 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 also known to provide generally cylindrical, or frustoconical, locally thickened, isogrid, metallic containment rings. Furthermore it is known to provide strong fibrous material wound around relatively thin metallic casings or around the previously 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 2 manufacture. The relatively thick casings with fibrous material are both heavier and more expensive to manufacture.
Accordingly the present invention seeks to provide a novel gas turbine engine casing which overcomes the above 5 mentioned problems.
Accordingly the present invention provides a gas turbine engine rotor blade containment assembly comprising a generally cylindrical, or frustoconical, metal casing, the metal casing having an inner surface and an outer surface, and at least a portion of at least one of the inner surface, or the outer surface, having a compressive residual stress.
The whole of the outer surface of the metal casing may have a compressive residual stress. The whole of the inner surface of the metal casing may have a compressive residual stress.
The at least one portion of the inner surface, or the outer surface, may have a predetermined pattern of compressive residual stress.
compressive residual stress intersecting lines.
The predetermined pattern of may comprise a grid of The grid of intersecting lines may be an isogrid. The outer surface may have the predetermined pattern of compressive residual stress.
The metal casing may have ribs and/or flanges.
one or more continuous layers of a strong fibrous material may be wound around the metal casing.
The metal casing may comprise any metal or metal alloy.
Preferably the metal casing comprises a steel alloy, aluminium, an aluminium alloy, magnesium, a magnesium alloy, titanium, a titanium alloy, nickel or a nickel alloy.
An acoustic lining may be provided within the metal casing.
The containment casing may be a fan containment casing, a compressor containment casing or a turbine containment casing.
The compressive residual stress may have been applied by laser shock peening.
3 The at least one portion may be at least a part circumferential portion of the metal casing or may be at least a part axial portion of the metal casing. Preferably the at least one portion extends completely circumferentially around the metal casing. Preferably the at least one portion extends axially the full length of the metal casing.
The present invention also provides a method of manufacturing a gas turbine engine rotor blade containment assembly comprising forming a generally cylindrical, or frustoconical, metal casing, the metal casing having an inner surface and an outer surface, and introducing a compressive residual stress into at least a portion of at least one of the inner surface, or the outer surface, of the metal casing.
A compressive residual stress may be introduced into the whole of the outer surface of the metal casing. A compressive residual stress may be introduced into the whole of the inner surface of the metal casing.
A compressive residual stress may be introduced in a predetermined pattern into the at least one portion of the inner surface, or the outer surface.
The predetermined pattern of compressive residual stress may comprise a grid of intersecting lines. The grid of intersecting lines may be an isogrid.
The predetermined pattern of compressive residual stress may be in the outer surface.
Ribs and/or flanges may be formed on the metal casing.
One or more continuous layers of a strong fibrous material may be wound around the metal casing.
The metal casing may comprise any metal or metal alloy.
Preferably the metal casing comprises a steel alloy, aluminium, an aluminium alloy, magnesium, a magnesium alloy, titanium, a titanium alloy, nickel or a nickel alloy.
An acoustic lining may be provided within the metal casing.
The compressive residual stress may be introduced by laser shock peening.
4 The at least one portion is at least a part circumferential portion of the metal casing or at least a part axial portion of the metal casing.
Preferably the at least one portion extends completely circumferentially around the metal casing. Preferably the at least one portion extends axially the full length of the metal casing.
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 view of one embodiment of the fan blade containment assembly shown in figure 2.
Figure 4 is an enlarged sectional view of a part of a further embodiment of the fan blade assembly.
Figure 5 is an enlarged sectional view of a part of another embodiment of the fan blade assembly.
Figure 6 is an enlarged sectional view of an additional embodiment of the fan blade assembly.
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 (not shown). The turbine section 20 also comprises a turbine to drive the fan section 14 via a shaft (not shown).
The fan section 14 comprises a fan duct 24 defined partially by a fan casing 26. The fan 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 rotor 32 and fan blades 34 rotate about the axis X of the gas turbine engine 10, substantially in a plane Y perpendicular to the axis X. The fan casing 26 also comprises a fan blade containment assembly 38, which is arranged substantially in the plane of the fan blades 34.
The fan casing 26 and fan blade containment assembly 38 is shown more clearly in figure 2. The fan blade containment assembly 38 comprises a metal cylindrical, or frustoconical, casing 40. The metal casing 40 comprises an upstream flange 42 by which the fan blade containment assembly 38 is connected to a flange 48 on an intake assembly 46 of the fan casing 26. The metal casing 40 also comprises a downstream flange 44 by which the fan blade containment assembly 38 is connected to a flange 52 on a rear portion 50 of the fan casing 26.
The fan blade containment region A is substantially in the plane Y containing the fan blades 34. The f an blade containment region B is upstream of the plane Y of the fan blades 34 and downstream of the flange 42, where debris protection is required for the fan blade containment assembly 38. The fan blade containment region C is downstream of the plane Y of the fan blades 34, where a root of a fan blade 34 impacts the fan blade containment assembly 38.
The metal casing 40 provides the basic fan blade containment and provides a connection between the intake casing 46 and the rear casing 50.
The metal casing 40 is provided with a deep compressive residual stress in at least a portion of the inner surface 43 or the outer surface 41. The deep compressive residual stress provides a pre- stress into the metal casing 40.
The metal casing 40, as shown in figure 3, comprises a predetermined pattern 54 of deep compressive residual stress in the outer surface 41 of the metal casing 40. The predetermined pattern 54 of deep compressive residual stress 6 in this example is an isogrid of lines 55 and 57 arranged at 600 to each other. For example a plurality of lines 55 extend circumferentially, or axially, and a plurality of lines 57 extend at 600 to the circumferential or axial direction. Thus the predetermined pattern 54 extends fully circumferentially around the metal casing 40.
The metal casing 40, as shown in figure 4, comprises a deep compressive residual stress 56 in the outer surface 41 of the metal casing 40 and a deep compressive residual stress io 58 in the inner surface 43 of the metal casing 40.
The provision of a deep compressive stress 56 in outer surface 41 of the metal casing 40 reduces possibility of tensile failure of the metal casing 40.
the the The deep compressive residual stress 56 may be provided on the inner surface 43 of any one, any two or all three of the containment regions A, B and C. The deep compressive residual stress 56 in the outer surface 41 of the metal casing 40 extends fully circumferentially around the metal casing 40.
The provision of a deep compressive residual stress 58 in the inner surface 43 of the metal casing 40 provides improved impact resistance where the fan blade 34, or portion of a fan blade 34, strikes the metal casing 40. The deep compressive residual stress 58 may be provided on the inner surface 43 of any one, any two, or all three of t containment regions A, B and C. The deep compressive residual stress 58 in the inner surface 43 of the metal casing 40 extends fully circumferentially around the metal casing 40.
The metal casing 40, as shown in figure 5, comprises a deep compressive residual stress 60 in the outer surface 41 only of the metal casing 40. The provision of a deep compressive stress 60 in the outer surface 41 of the metal casing 40 reduces the possibility of further of tensile failure of the metal casing 40. The deep compressive residual stress 60 may be provided on the inner surface 43 of he 7 any one, any two, or all three containment regions A, B and C. The deep compressive residual stress 60 in the outer surface 41 of the metal casing 40 extends fully circumferentially around the metal casing 40.
The metal casing 40, as shown in figure 6, comprises a deep compressive residual stress 62 in the inner surface 43 only of the metal casing 40. The provision of deep compressive residual stress 62 in the inner surface 43 of the metal casing 40 provides improved impact resistance where the io fan blade 34 or portion of a fan blade 34, strikes the metal easing 40. The deep compressive residual stress 62 may be provided on the inner surface 43 of any one, any two, or all three containment regions A, B and C. The deep compressive residual stress 62 in the inner surface 43 of the metal casing 40 extends fully circumferentially around the metal casing 40. The metal casing 40 comprises one, or more, integral T-section ribs 63, which extend radially outwardly from the metal casing 40 and circumferentially around the metal casing 40. The ribs 63 stiffen the metal casing 40 to improve the ability of the metal casing 40 to contain a detached fan blade 34.
It may be desirable in some circumstances to provide a number of continuous layers of a strong fibrous material wound around the metal casing 40 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 3"5 registered trademark of Dupont Ltd). There may also be a number of layers of discrete pieces of flexible material woven from KWLAR between the metal casing 40 and the continuous layers of fibrous material.
An acoustic lining 64 may be provided on the inner surface of the first metal casing 40. The acoustic lining 64 comprises a honeycomb 66 and a perforate sheet 68. The honeycomb 66 and perforate sheet 68 are quite conventional.
8 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 encounters the metal casing 40. The metal casing 40 is impacted by the fan blade 34, or portion of the fan blade 34, and the metal casing 40 effectively removes energy from the fan blade 34, or portion of the fan blade 34. The impact of the fan blade 34, or portion of the fan blade 34, introduces a tensile force into the metal casing 40 which may lead to a tensile failure of the metal casing 34. The deep 10 compressive residual stress in the surface, or surfaces, of the metal casing 40 provides a compressively pre-stressed protection layer. The compressively pre- stressed protection layer requires that additional work must be done by the tensile force to overcome the compressively prestressed layer. Thus in the event that a fan blade 34, or portion of a fan blade 34, becomes detached more energy is removed by the compressively prestressed metal casing 40 from the fan blade 34, or portion of a fan blade 34, compared to a metal casing of equivalent thickness without the compressively prestressed layer.
The fan blade containment invention has several advantages.
assembly of the present The compressive residual stress in the metal casing provides improved dynamic crack propagation resistance during a fan blade off event. The compressive residual stress in the metal casing removes more energy than a conventional metal casing of equivalent thickness. This enables the thickness of the metal casing to be reduced and hence the metal casing is significantly lighter than conventional fan blade containment assemblies, especially on relatively large diameter turbofan gas turbine engines. Alternatively the thickness of the metal casing may be maintained to provide improved fan blade containment properties. The compressive residual stress in the metal casing reduces the possibility of puncturing and tearing of the metal casing. The first metal casing may be manufactured from any suitable metal or metal alloy.
9 Preferably the metal casing is manufactured from a steel alloy, aluminium, an aluminium alloy, magnesium, a magnesium alloy, titanium, a titanium alloy, nickel or a nickel alloy.
The deep compressive residual stress is introduced in the metal casing by laser shock peening or other suitable processes. The laser shock peening process produces a plurality of volumetrically spaced laser shock peened protrusions extending into the metal casing which have deep compressive residual stresses.
The laser shock peening process comprises painting the surface of the metal casing with a black paint, flowing water over the painted surface of the metal casing and repetitively firing a focused high energy laser beam, through the water, on the surface of the metal casing. The paint creates a peak power input into the metal casing.
The deep compressive residual stress extends from the surface of the metal casing to a depth of about 20 to 50 mils (or 0. 5 mm to 1. 3mm).
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.
Although the invention has been described with reference to the use of deep compressive stress in the surface of the metal casing throughout the whole circumferential extent of the metal casing it may be possible in some circumstances to use deep compressive stress in the surface of the metal casing over only a circumferential portion of the metal casing or only over an axial portion of the metal casing.
Claims (40)
1. A gas turbine engine rotor blade containment assembly comprising a generally cylindrical, or frustoconical, metal casing, the metal casing having an inner surface and an outer surface, and at least a portion of at least one of the inner surface, or the outer surface, having a compressive residual stress.
2. A gas turbine engine rotor blade containment assembly as claimed in claim 1 wherein the whole of the outer surface of the metal casing has a compressive residual stress.
3. A gas turbine engine rotor blade containment assembly as claimed in claim 1 or claim 2 wherein the whole of the inner surface of the metal casing has a compressive residual stress.
4. A gas turbine engine rotor blade containment assembly as claimed in claim 1 wherein the at least one portion of the inner surface, or the outer surface, comprises a predetermined pattern of compressive residual stress.
5. A gas turbine engine rotor blade containment assembly as 20 claimed in claim 4 wherein the predetermined pattern of compressive residual stress comprises a grid of intersecting lines.
6. A gas turbine engine rotor blade containment assembly as claimed in claim 4 or claim 5 wherein the grid of intersecting lines is an isogrid.
7. A gas turbine engine rotor blade containment assembly as claimed in claim 4, claim 5 or claim 6 wherein the outer surface has the predetermined pattern of compressive residual stress.
8. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 7 wherein metal casing has ribs and/or flanges.
9. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 8 wherein one or more continuous layers of a strong fibrous material are wound around the metal casing.
11
10. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 9 wherein the metal casing comprises a steel alloy, aluminium, an aluminium alloy, magnesium, a magnesium alloy, titanium, a titanium alloy, nickel or a nickel alloy.
11. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 10 wherein an acoustic lining is provided within the metal casing.
12. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 11 wherein the containment casing is a fan containment casing. a compressor containment casing or a turbine containment casing.
13. A gas turbine engine rotor blade containment assembly as claimed in any of claims 1 to 12 wherein the compressive residual stress has been applied by laser shock peening.
14. A gas turbine engine as claimed in any of claims 1 to 13 wherein the at least one portion is at least a part circumferential portion or at least a part axial portion.
15. A gas turbine engine as claimed in any of claims 1 to 14 20 wherein the at least one portion extends completely circumferentially around the metal casing.
16. A gas turbine engine as claimed in any of claims 1 to 15 wherein the at least one portion extends axially the full length of the metal casing.
17. A gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 2 and 3 of the accompanying drawings.
18. A gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 2 and 4 of the accompanying drawings.
19. A gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 2 and 5 of the accompanying drawings.
20. A gas turbine engine rotor blade containment assembly 35 substantially as hereinbefore described with reference to and as shown in figures 2 and 5 of the accompanying drawings.
12
21. A gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to and as shown in figures 2 and 6 of the accompanying drawings.
22. A method of manufacturing a gas turbine engine rotor 5 blade containment assembly comprising forming a generally cylindrical, or frustoconical, metal easing, the metal casing having an inner surface and an outer surface, and introducing a compressive residual stress into at least a portion of at least one of the inner surface, or the outer surface, of the 10 metal casing.
23. A method as claimed in claim 22 comprising introducing a compressive residual stress into the whole of the outer surface of the metal. casing.
24. A method as claimed in claim 22 or claim 23 comprising introducing a compressive residual stress into the whole of the inner surface of the metal casing.
25. A method as claimed in claim 24 comprising introducing compressive residual stress in a predetermined pattern into the at least one portion of the inner surface, or the outer 20 surface.
26. A method as claimed in claim predetermined pattern of compressive comprises a grid of intersecting lines.
27. A method as claimed in claim 25 or claim 26 wherein the grid of intersecting lines is an isogrid.
28. A method as claimed in claim 25, claim 26 or claim 27 comprising introducing the compressive residual stress in the predetermined pattern in the outer surface.
wherein the residual stress
29. A method as claimed in any of claims 22 to 28 comprising 30 forming ribs and/or flanges on the metal casing.
30. A method as claimed in any of claims 22 to 29 comprising winding one or more continuous layers of a strong fibrous material around the metal casing.
31. A method as claimed in any of claims 22 to 30 wherein 35 the metal casing comprises a steel alloy, aluminium, an aluminium alloy, magnesium, a magnesium alloy, titanium, a titanium alloy, nickel or a nickel alloy.
32. A method as claimed in any of claims 22 to 31 comprising providing an acoustic lining within the metal casing.
33. A method as claimed in any of claims 22 to 32 comprising introducing the compressive residual stress by laser shock peening.
34. A method as claimed in any the at least one portion is at portion or at least a part axial
35. A method as claimed in any the at least one portion extends of claims 22 to 33 wherein least a part circumferential portion. of claims 22 to 34 wherein completely circumferentially around the metal casing.
36. A method as claimed in any of claims 22 to 35 wherein the at least one portion extends axially the full length of the metal casing.
37. A method of manufacturing a gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to figures 2 and 3 of the accompanying drawings.
38. A method of manufacturing a gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to figures 2 and 4 of the accompanying drawings.
39. A method of manufacturing a gas turbine engine rotor blade containment assembly substantially as hereinbefore described with reference to figures 2 and 5 of the accompanying drawings.
40. A method of manufacturing a gas turbine engine rotor 30 blade containment assembly substantially as hereinbefore described with reference to figures 2 and 6 of the accompanying drawings.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0008189A GB2361032A (en) | 2000-04-05 | 2000-04-05 | A gas turbine engine blade containment assembly |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB0008189A GB2361032A (en) | 2000-04-05 | 2000-04-05 | A gas turbine engine blade containment assembly |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB0008189D0 GB0008189D0 (en) | 2000-05-24 |
| GB2361032A true GB2361032A (en) | 2001-10-10 |
Family
ID=9889114
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB0008189A Withdrawn GB2361032A (en) | 2000-04-05 | 2000-04-05 | A gas turbine engine blade containment assembly |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2361032A (en) |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8297912B2 (en) | 2008-07-29 | 2012-10-30 | Rolls-Royce Plc | Fan casing for a gas turbine engine |
| EP2096269A3 (en) * | 2008-02-27 | 2013-03-20 | Rolls-Royce plc | Fan track liner assembly for a gas turbine engine |
| EP2586999A1 (en) | 2011-10-25 | 2013-05-01 | MTU Aero Engines GmbH | Housing section and method for producing a housing section for a fluid flow engine |
| CN103133413A (en) * | 2011-11-25 | 2013-06-05 | 中国航空工业集团公司沈阳发动机设计研究所 | Engine blower multilayer receiver structure |
| US9714583B2 (en) | 2014-08-21 | 2017-07-25 | Honeywell International Inc. | Fan containment cases for fan casings in gas turbine engines, fan blade containment systems, and methods for producing the same |
Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2114233A (en) * | 1982-02-01 | 1983-08-17 | United Technologies Corp | Containment shell for a fan section of a gas turbine engine |
| US5403148A (en) * | 1993-09-07 | 1995-04-04 | General Electric Company | Ballistic barrier for turbomachinery blade containment |
| US5569018A (en) * | 1995-03-06 | 1996-10-29 | General Electric Company | Technique to prevent or divert cracks |
| US5742028A (en) * | 1996-07-24 | 1998-04-21 | General Electric Company | Preloaded laser shock peening |
-
2000
- 2000-04-05 GB GB0008189A patent/GB2361032A/en not_active Withdrawn
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2114233A (en) * | 1982-02-01 | 1983-08-17 | United Technologies Corp | Containment shell for a fan section of a gas turbine engine |
| US5403148A (en) * | 1993-09-07 | 1995-04-04 | General Electric Company | Ballistic barrier for turbomachinery blade containment |
| US5569018A (en) * | 1995-03-06 | 1996-10-29 | General Electric Company | Technique to prevent or divert cracks |
| US5742028A (en) * | 1996-07-24 | 1998-04-21 | General Electric Company | Preloaded laser shock peening |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP2096269A3 (en) * | 2008-02-27 | 2013-03-20 | Rolls-Royce plc | Fan track liner assembly for a gas turbine engine |
| US8297912B2 (en) | 2008-07-29 | 2012-10-30 | Rolls-Royce Plc | Fan casing for a gas turbine engine |
| EP2586999A1 (en) | 2011-10-25 | 2013-05-01 | MTU Aero Engines GmbH | Housing section and method for producing a housing section for a fluid flow engine |
| CN103133413A (en) * | 2011-11-25 | 2013-06-05 | 中国航空工业集团公司沈阳发动机设计研究所 | Engine blower multilayer receiver structure |
| US9714583B2 (en) | 2014-08-21 | 2017-07-25 | Honeywell International Inc. | Fan containment cases for fan casings in gas turbine engines, fan blade containment systems, and methods for producing the same |
Also Published As
| Publication number | Publication date |
|---|---|
| GB0008189D0 (en) | 2000-05-24 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| WAP | Application withdrawn, taken to be withdrawn or refused ** after publication under section 16(1) |