GB2442112A

GB2442112A - Gas turbine engine containment casing

Info

Publication number: GB2442112A
Application number: GB0718002A
Authority: GB
Inventors: David Lawrence Bedel; David William Crall; Stephen Craig Mitchell; Donald George Lachapelle; Ming Xie; Frank Worthoff; Leslie Louis Langenbrunner
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2006-09-25
Filing date: 2007-09-14
Publication date: 2008-03-26
Also published as: CA2602319A1; JP2008082332A; DE102007045138A1; GB0718002D0

Abstract

A gas turbine engine (10, figure 1) has a bladed rotor 44 surrounded by a casing (40, figure 1) having a blade containment system 62 which includes at least one layer of fibre metal laminate 78. The blade containment system 62 may also include a honeycomb layer 64 positioned radially exterior of an inner portion 66 of the annular casing (40), wherein the layer of fibre metal laminate 78 is positioned radially exterior or interior of the honeycomb layer 64. The blade containment system 62 may also include a radially exterior layer of Kevlar wrap (82, figure 4). The fibre metal laminate 78 may comprise aluminium and glass, aluminium and aramid, or titanium and graphite. In order to allow the fibre metal laminate 78 to be formed into a circumferential arrangement, a plurality of gore cuts may be provided therein. The fibre metal laminate 78 may be spliced together at its ends to form an annular shape.

Description

BLADE CONTAINMENT SYSTEM FOR A GAS TURBINE ENGINE

BACKGROUND OF THE INVENTION

The present invention relates to a system for containing any engine fragments in the event of a blade release in a gas turbine engine and, more particularly, to a blade containment system which utilizes a fiber metal laminate layer therein.

It is understood that foreign objects (e.g., birds, hailstones, sand, ice, etc.) may unavoidably be ingested during the operation of a gas turbine engine. As it is forced through the engine, the foreign object may impact a blade of the fan and cause a portion to be torn loose from the rotor. Without a system for containing this event, catastrophic damage can be caused to the engine casing. Accordingly, various types of blade containment systems have been employed to minimize foreign object damage.

A first type of system for containing fan fragments is known as a hardwall system, where the fan case is designed with enough structural integrity to provide complete containment. This type of case is usually made from monolithic steel or aluminum.

A second blade containment system is a softwall system, where engine fragments penetrate the structural case but are captured by a belt of woven or braided Kevlar.

While these systems are useful for their intended purpose, the search continues fbr a blade containment design which maximizes impact and crack propagation resistance and minimizes the weight added to the engine.

Fiber metal laminates are a class of material which have been developed recently for use in load bearing aircraft structures. Fiber metal laminates typically consist of relatively thin metallic sheets interleaved between glass or Kevlar tape in an epoxy matrix that is also used as an adhesive. One example of a glass-metal laminate is disclosed in U.S. Patent 5,039,571 to Vogelsang et al. and is sold under the trade name of GLARE . While fiber metal laminates are currently in use on load bearing aircraft structures, such as control surface flaps, cargo bay floors, and fuselage top skins, it has typically consisted of substantially planar sheets. ) I

In order to take advantage of the exceptional impact resistance properties of fiber metal laminates in a blade containment system, however, it needs to be formed into a cylindrical or conical structure. This requires appropriate splicing techniques to be developed and employed so that the essentially flat material maintains a desired strength and structural integrity.

Thus, there is a need to provide a blade containment system for a gas turbine engine which provides greater resistance to impact and crack propagation. In addition, it is desired that the weight of such blade containment system be minimized to decrease fuel consumption.

BRIEF SUMMARY OF THE [NVENTION

In a first exemplary embodiment of the invention, a gas turbine engine is disclosed as having a plurality of radially extending blades mounted on an annular disk, with the blades and disk being rotatable about a longitudinal axis of the engine. A blade containment system is further disclosed as including an annular casing positioned radially outward of the blades and in surrounding relationship therewith, the annular casing including at least one layer of fiber metal laminate utilized therewith. The blade containment system may also include a honeycomb layer, wherein the layer of fiber metal laminate is positioned either radially exterior of the honeycomb layer or radially interior thereof. For both cases, the blade containment system could include a radially exterior layer of Keviar wrap.

In a second exemplary embodiment of the invention, a gas turbine engine is disclosed as having a plurality of radially extending blades mounted on an annular disk, with the blades and disk being rotatable about a longitudinal axis of the engine. A blade containment system is further disclosed as including an annular casing made of fiber metal laminate positioned radially outward of the blades and in surrounding relationship therewith. The blade containment system may further include a honeycomb layer positioned radially exterior of the annular casing. In such case, the blade containment system could include a layer of Keviar wrap positioned radially exterior of the honeycomb layer. p

In a third exemplary embodiment of the invention, an annular member for use in a gas turbine engine is disclosed, where the annular member includes a first edge and a second edge. The annular member is made of a fiber metal laminate and includes a splice area connecting the first and second edges. Preferably, the annular member is shaped substantially cylindrically or substantially conically. Alternatively, the annular member may include a plurality of annular segmcnts. Each annular segment is made of a fiber metal laminate and includes a splice area connecting each edge to an adjacent annular segment.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings, in which: Fig. 1 is a diagrammatic view of a high bypass turbofan gas turbine engine; Fig. 2 is an enlarged, cross-sectional view of a blade containment system having a

prior art design;

Fig. 3 is an enlarged, cross-sectional view of a blade containment system in accordance with a first embodiment of the invention; Fig. 4 is an enlarged, cross-sectional view of a blade containment system in accordance with a second embodiment of the invention; Fig. 5 is an enlarged, cross-sectional view of a blade containment system in accordance with a third embodiment of the invention; Fig. 6 is a partial diagrammatic view of a fiber metal laminate ply; Fig. 7 is a front view of the annular member depicted in Figs. 3-5; Fig. 8 is a partial diagrammatic view of a splice area for the annular member depicted in Fig. 7; Fig. 9 is a partial diagrammatic view of an alternative splice area for the annular member depicted in Fig. 7; Fig. 10 is a partial diagrammatic view of an area for the annular member depicted in Fig. 7, where such area has been reinforced by internally added plies; Fig. 11 is a partial diagrammatic view of an area for the annular member depicted in Fig. 7, where such area has been reinforced by externally added plies; Fig. 12 is a partial, longitudinal, cross-sectional view of a fan casing for the gas turbine engine depicted in Fig. I having a blade containment zone therein; Fig. 13 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange has been formed at an end thereof; Fig. 14 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where the end thereof terminates in an all-metallic edge; Fig. 15 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange has been formed at an end thereof in an alternative manner; Fig. 16 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange is shown as being attached to an end thereof Fig. 17 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange is shown as being attached to an end thereof in a first alternative manner; Fig. 18 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange is shown as being attached to an end thereof in a second alternative manner; Fig. 19 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where a flange is shown as being formed at an end thereof in a third alternative manner; Fig. 20 is a partial diagrammatic view of the fan casing depicted in Fig. 12, where stiffening rings of various configurations are attached thereto; Fig. 21 is a partial diagrammatic view of a stiffening ring having a first configuration made of fiber metal laminate being attached to a fan casing; Fig. 22 is a partial diagrammatic view of a stiffening ring having a second configuration made of fiber metal laminate being attached to a fan casing; and, Fig. 23 is a partial diagrammatic view of a stiffening ring having a third configuration made of fiber metal laminate being attached to a fan casing.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures, Fig. I depicts in diagrammatic form an exemplary gas turbine engine 10 (high bypass type) utilized with aircraft having a longitudinal or axial centerline axis 12 therethrough for reference purposes. Engine 10 preferably includes a core gas turbine engine generally identified by numeral 14 and a fan section 16 positioned upstream thereof Core engine 14 typically includes a generally tubular outer casing 18 that defines an annular inlet 20. Outer casing 18 further encloses and supports a booster compressor 22 for raising the pressure of the air that enters core engine 14 to a first pressure level. A high pressure, multi-stage, axial-flow compressor 24 receives pressurized air from booster 22 and further increases the pressure of the air. The pressurized air flows to a combustor 26, where fuel is injected into the pressurized air stream to raise the temperature and energy level of the pressurized air. The high energy combustion products flow from combustor 26 to a first (high pressure) turbine 28 for driving high pressure compressor 24 through a first (high pressure) drive shaft 30, and then to a second (low pressure) turbine 32 fbr driving booster compressor 22 and fan section 16 through a second (low pressure) drive shaft 34 that is coaxial with first drive shaft 30. After driving each of turbines 28 and 32. the combustion products leave core engine 14 through an exhaust noz?'Ie 36 to provide propulsive jet thrust.

Fan section 16 includes a rotatable, axial-flow fan rotor 38 that is surrounded by an annular fan casing 40. It will be appreciated that fan casing 40 is supported from core engine 14 by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes 42. In this way, fan casing 40 encloses fan rotor 38 and fan rotor blades 44. Downstream section 46 of fan casing 40 extends over an outer portion of core engine 14 to define a secondary, or bypass, airflow conduit 48 that provides additional propulsive jet thrust.

From a flow standpoint, it will be appreciated that an initial air flow, represented by arrow 50, enters gas turbine engine 10 through an inlet 52 to fan casing 40. Air flow passes through fan blades 44 and splits into a first compressed air flow (represented by arrow 54) that moves through conduit 48 and a second compressed air flow (represented by arrow 56) which enters booster compressor 22. The pressure of second compressed air flow 56 is increased and enters high pressure compressor 24, as represented by arrow 58. After mixing with fuel and being combusted in combustor 26, combustion products 60 exit combustor 26 and flow through first turbine 28. Combustion products 60 then flow through second turbine 32 and exit exhaust nozzle 36 to provide thrust for gas turbine engine 10.

As seen in the prior art of Fig. 2, fan casing 40 includes an annular blade containment system 62 to prevent pieces of fan blades 44 from escaping through fan casing 40.

More specifically, it will be seen that blade containment system 62 includes a honeycomb section 64 located along a radially exterior surface 67 of an inner portion 66 of fan casing 40. In addition, blade containment system 62 includes a layer 68 of abradable material (e.g., Kevlar wrap) having a thickness 70 located adjacent a radially exterior surface 72 of honeycomb section 64. Layer 68 is then attached at an upstream end 74 and a downstream end 76 to fan casing inner portion 66.

In one embodiment of the present invention, blade containment system 62 replaces layer 68 of abradable material with an annular member 78 made of fiber metal laminate having a thickness 80 (see Fig. 3). It will be appreciated that annular member 78 is advantageous in terms of having less thickness and less weight than abradable material layer 68, while providing at least as much resistance to impact and crack propagation. As such, annular member 78 is positioned adjacent radially exterior surface 72 of honeycomb section 64. In an alternative configuration depicted in Fig. 4, blade containment system 62 may further include a layer 82 of abradable material having thickness 84 positioned radially exterior of annular member 78 to provide additional protection. Even so, it will be noted that inclusion of annular member 78 enables thickness 84 of abradable material layer 82 to be less than thickness 70 of abradable material layer 68 discussed above. As shown in Fig. 5, at least a designated part of fan casing inner portion 66 is made of fiber metal laminate.

The fiber metal laminate of annular member 78 may have various configurations and compositions, such as an aluminumlglass configuration sold under the product name GLARE , an aluminum/aramid configuration sold under the product name ARALL , and a titaniurnlgraphite configuration sold under the product name TIGRE.

As the term is utilized herein, however, it will be understood that fiber metal laminate generally is defined as one or more plies 89 including at least one metallic layer 90 and a synthetic layer 94 bonded together (see Fig. 6). Of course, a fiber metal laminate may include multiple plies of having alternating layers of metal and synthetic material. The metals utilized in metallic layer 90 preferably include aluminum or titanium, but any lightweight metal or metal alloy having high strength characteristics may be utilized. While it will be understood that metallic layer 90 is typically a solid sheet, it may be constructed from a mesh, such as a ribbon mesh or a wire mesh. The materials used for synthetic layer 94 preferably include glass reinlbrced epoxy resin, carbon reinforced epoxy resin, or aramid reinforced epoxy resin, but any other suitable fiber reinforced composite material may be utilized.

In order to form annular member 78 from fiber metal laminate, which is initially constructed of substantially planar sheets, a splice area 96 is required for joining first and second ends 98 and 100 thereof (see Fig. 7). This may be accomplished in any number of ways, but preferably the ends of individual plies in the fiber metal laminate are staggered so as to provide separation between weak areas. For example, splice area 96 is depicted in Fig. 8 as a staggered butt joint, where a break 1 02 between opposite ends 104 and 106 of a first metallic layer 108 is located a distance 110 from an adjacent break 112 between opposite ends 114 and 116 of a second metallic layer 118. Although not shown, it will be appreciated that similar breaks between opposing ends for metallic layers 120, 122, 124, 126 and 128 are likewise spaced apart a circumferential distance. Clearly, synthetic layers 130 and 132 assist in filling in break 102, as do synthetic layers 134 and 136 for break 112.

An alternative manner of forming annular member 78 is shown in Fig. 9, where splice area 96 is a staggered overlap joint. As seen therein, the fiber metal laminate includes at least a first ply 138 and a second ply 140 which extend adjacent to and substantially parallel one another except in splice area 96. Within splice area 96, it will be noted that first ply 138 has a first section 142 which angles in a first direction, a second section 144 which extends substantially parallel to its orientation outside splice area 96, and a third section 146 which angles in a second direction supplementary to the first direction. In this way, first ply 138 remains integral throughout splice area 96.

Second ply 140 similarly includes a first section 148 and a second section 150 in splice area 96 which extend in the same respective directions as first and second sections 142 and 144 of first ply 138. Instead of remaining integral throughout splice area 96, however, second ply 140 splits at approximately the point where second section 150 begins so that a first end 152 thereof overlaps a second end 154 and extends substantially parallel thereto through splice area 96.

It will also be appreciated that annular member 78 may include a plurality of annular segments (as identified by reference numerals 79 and 81, for example, shown in phantom in Fig. 7). Each such annular segment will be joined to an adjacent annular segment at each end thereof by means of the methods described herein. In this way, a plurality of splice areas are formed within annular member 78.

It will be appreciated that additional plies of fiber metal laminate for annular member 78 may be provided in specified areas for extra reinforcement. As seen in Figs. 10 and 11, respectively, such additional plies may be built up internally or externally.

Looking at Fig. 10, additional plies 156, 158, and 160 are added to the radial interior of main ply 162, where main ply 162 has a first section 164 which angles in a first direction, a second section 166 which extends substantially parallel to its original orientation apart from additional plies 156, 158 and 160, and a third section 168 which angles in a second direction supplementary to the first direction. In this way,

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space is created for additional plies 156, 158 and 160, which are oriented substantially parallel to second section 166 of main ply 162 and of a length accommodated by the distance between first and third sections of main ply 162 depending upon their radial position. Conversely, additional plies 157 and 159 are added to the exterior of main ply 162 in Fig 11.

Rather than substituting an annular member made of fiber metal laminate for one or more components of an existing blade containment system, it is also contemplated that the fan casing itself be made of such material. As seen in Fig. 12, an annular faii casing 172 having a blade containment zone 1 74 is made of fiber metal laminate. It will be noted that acoustic tiles 176 and 1 78 are located upstream and downstream of blade containment zone 174, respectively. A plurality of stiffening rings 180 are preferably positioned in axially spaced relation along an outer radial surface 182 of fan casing 172 to provide additional support thereto. An annular honeycomb layer 184 may optionally be positioned adjacent an inner radial surface 1 86 of fan casing 172, which is held in place with a face sheet 188. Although not shown, an abradable layer may also be employed within blade containment zone 174. As further seen therein with respect to acoustic tile 178, acoustic tiles 176 and 178 may also employ a layer 177 of fiber metal laminate as the perforated face sheet therewith.

Fan casing 172 may be required to incorporate a flange 190 at one or both of its upstream and downstream ends (see Fig. 13), especially if it is to be installed as a replacement on existing engine lines. Such a flange 190 may be formed by rolling an edge 192 of fan casing 1 72 after the curing cycle of the substantially cylindrical fan casing 172 takes place. Alternatively, such flange 190 may be formed integrally with fan casing 172, where the metallic layers in flange 190 are formed to the desired shape prior to curing the laminate. Once assembled, fan casing 1 72 is cured to form a seamless interface with flange 190.

Another approach is to provide an all-metallic edge 194 to fan casing 172 as depicted in Fig. 14, As seen therein, composite layers 196 and 198 are terminated short of edge 194 and replaced by metal layers 200 and 202 of the same thickness 204. An optional fastener 206 may be utilized to assist in consolidating metallic layers 200 and 202 with metallic layers 201, 203 and 205. Hardware can either be fastened to all-metallic edge 194 after fan casing 172 is cured or edge 194 can be drilled and attached to the hardware with fasteners and then cured into a single bonded and fastened structure.

A metal flange 208 may include a portion 210 which is matable with an end 212 of fan casing 172 having a reduced thickness. A metal ring 214 may be provided adjacent an outer surface 216 of fan casing 172, with a plurality of fasteners 218 securing flange 208 to fan casing 172 (see Fig. 15). Figs. 16-18 depict other options for attaching hardware 220, 222, and 224, respectively, directly to the fiber metal laminate of fan casing 1 72.

In order to permit forniing of fiber metal laminate sheets into fan casing 1 72, it is also preferred that a plurality of gore cuts 170 be provided therein (see Fig. 19). It will be appreciated that aluminum strips 171 and 173 may be positioned on either side of the fiber metal laminate to assist in forming flange 190.

As stated hereinabove, stiffening rings 1 80 of various configurations may be attached to outer radial surface 182 of fan casing 172 (see Fig. 20). Such stiffening rings 180 are generally constructed of a metal (e.g., aluminum, steel, or titanium). It will be appreciated that stiffening rings, as identified by numeral 183 in Figs. 21-23, may also be constructed of a fiber metal laminate or other composite laminate or sandwich structure, which are bonded to fan casing 172 either after fan casing 172 has been cured or during the same curing cycle as the remainder of fan casing 1 72.

Although particular embodiments of the present invention have been illustrated and described, it will be apparent to those skilled in the art that various changes and modifications can be made without departing from the spirit of the present invention.

* Accordingly, it is intended to encompass within the appended claims all such changes and modification that fall within the scope of the present invention.

Claims

CLAIMS: 1. In a gas turbine engine having a plurality of radially

extending blades mounted on an annular disk, the blades and disk being rotatable about a longitudinal axis of the engine, a blade containment system comprising an annular casing positioned radially outward of the blades and in surrounding relationship therewith, said annular casing including at least one layer of fiber metal laminate utilized therewith.
2. The blade containment system of claim 1, further comprising a honeycomb layer positioned radially exterior of an inner portion of said annular casing.
3. The blade containment system of claim 2, wherein said layer of fiber metal laminate is positioned radially exterior of said honeycomb layer.
4. The blade containment system of any one of the preceding claims, further comprising a layer of Kevlar wrap positioned radially exterior of said layer of fiber metal laminate.
5. The blade containment system of claim 2, wherein said layer of fiber metal laminate is positioned radially interior of said honeycomb layer.
6. The blade containment system of claim 5, further comprising a layer of Keviar wrap positioned radially exterior of said honeycomb layer.
7. The blade containment system of any one of the preceding claims, said layer of fiber metal laminate including at least one ply, said ply further comprising: (a) at least one metallic layer; and, (b) a synthetic layer bonded to said metallic layer.
8. The blade containment system of any one of the preceding claims, said fiber metal laminate layer including a plurality of gore cuts therein to enable expansion in a circumferential manner.

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9. The blade containment system of claim 7, said layer of fiber metal laminate including additional plies in a specified area.
10. The blade containment system of any one of the preceding claims, said layer of fiber metal laminate being spliced together at its ends to form an annular shape.
11. A blade containment system substantially as hereinbefore described with reference to the accompanying drawings.