DE102007045138A1 - Blade restraint system for a gas turbine engine - Google Patents

Abstract

In a gas turbine engine (10) having a plurality of radially extending blades (44) mounted on an annular disc (38), the blades (44) and disc (38) extending about a longitudinal axis (12) of the engine (10). a bucket restraint system (62) having an annular housing (40) positioned radially outwardly of and in surrounding relation to the rotor blades (44), the annular housing (40) comprising at least one layer of fibrous metal laminate (US Pat. 78). The blade retention system (62) may also include a honeycomb layer (64) disposed radially outward of an interior portion (66) of the annular housing (40), the layer of fiber metal laminate (78) extending radially outward from the honeycomb layer (64). or disposed radially inside thereof. For both cases, the blade retention system (62) could include a radially outer layer of a Kevlar wrap (82).

Description

BACKGROUND OF THE INVENTION

The The present invention relates to a retention system any engine parts in the event of a blade separation in a gas turbine engine, and in particular a blade restraint system, which uses a fibrous metal laminate therein.

It might understand that foreign objects (such as birds, hailstones, sand, Ice cream etc.) inevitably during the operation of a gas turbine engine can be sucked. As soon as It is sucked through the engine, the foreign object can on one Blowing blade of the blower and cause a part of the rotor to be torn off. Without a system for the containment This event can cause catastrophic damage to the engine casing. As a result, various types of blade restraint systems have already been used used to damage by minimizing foreign objects.

One first system type for retention of fan fragments is known as a hardwoven system in which the fan case with Sufficient structural strength is designed to provide complete retention sure. This housing type usually exists made of monolithic steel or aluminum. A second blade restraint system is a soft wall system in which engine parts penetrate the structural housing, but caught by a band of woven or braided kevlar become. Although these systems for are useful for their intended purpose, the search for a blade retention structure Continuing the impact and crack propagation resistance maximized and the extra Weight for the engine minimized.

Fiber metal laminates are a class of materials that has recently been developed for use in structural aircraft structures. Fiber metal laminates typically consist of relatively thin metallic plates sandwiched between glass or Kevlar tape in an epoxy matrix which is also used as an adhesive. An example of a glass / metal laminate is in the U.S. Patent 5,039,571 for Vogelsang et al. disclosed and sold under the trade name GLARE ^®. Although fiber metal laminates are currently in use on load-bearing aircraft structures such as control surface flaps, cargo compartment doors and tank skins, these typically consist essentially of flat panels.

Around but the extraordinary Impact resistance properties of fiber metal laminates in blade containment systems to use, these must however, are formed into a cylindrical or conical structure. This requires the development and use of suitable splicing techniques, such that the substantially planar material has a desired strength and structural integrity maintains.

Consequently there is a need to provide a blade retarding system for a gas turbine engine, which gives greater strength against blow and crack propagation offers. In addition, it is desirable that the weight of such a blade restraint system to reduce the Fuel consumption is minimized.

BRIEF SUMMARY OF THE INVENTION

In a first exemplary embodiment The invention is a gas turbine engine having a plurality of radially extending blades disclosed on an annular disc are mounted, wherein the blades and the disc about a longitudinal axis of the engine are rotatable. A blade restraint system is also known as a ring-shaped casing containing radially outside the blades and is positioned in a surrounding relationship thereto, wherein the annular housing at least contains a layer of a fiber metal laminate used therewith. The blade restraint system may also include a honeycomb layer, wherein the layer of the fiber metal laminate either radially outside the honeycomb layer or radially disposed therebetween. For both Cases could do that A blade retention system a radially outer layer a Kevlar wrap included.

In a second exemplary embodiment of the invention a gas turbine engine having a plurality of radially extending ones Revolving blades mounted on an annular disc, the blades and the disc being about a longitudinal axis of the engine are rotatable. A blade restraint system is further described as a annular casing comprising fiber metal laminate which is radial outside of the blades and in surrounding relation thereto is arranged. The blade restraint system may further include a honeycomb layer which is radially outward arranged by the ringför shaped housing is. In such a case could the blade restraint system include a layer of kevlar wrap that is radially outward the honeycomb layer is arranged.

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

BRIEF DESCRIPTION OF THE DRAWINGS

1 FIG. 12 is a schematic view of a high bypass ratio turbofan gas turbine engine; FIG.

2 Fig. 10 is an enlarged cross-sectional view of a blade restraint system of a conventional construction;

3 FIG. 10 is an enlarged cross-sectional view of a blade restraint system according to a first embodiment of the invention; FIG.

4 Fig. 10 is an enlarged cross-sectional view of a blade restraint system according to a second embodiment of the invention;

5 Fig. 10 is an enlarged cross-sectional view of a blade restraint system according to a third embodiment of the invention;

6 is a schematic partial view of a Fasermetalllaminatlage;

7 is a front view of the in the 3 to 5 illustrated annular element;

8th is a schematic partial view of a splice region for the in 7 illustrated annular element;

9 is a schematic partial view of an alternative splice region for the in 7 illustrated annular element;

10 is a schematic partial view of a splice region for the in 7 illustrated annular member in which such a region has been reinforced by internally added layers;

11 10 is a schematic partial view of a splice region for the in 7 illustrated annular member in which such a region has been reinforced by externally added layers;

12 is a longitudinal cross-sectional partial view of a fan case for in 1 illustrated gas turbine engine with a blade retention zone therein;

13 is a schematic partial view of the in 12 a fan case shown in which a flange is formed at one end thereof;

14 is a schematic partial view of the in 12 shown fan case in which its end terminates in a full metallic edge;

15 is a schematic partial view of the in 12 shown fan case, in which a flange is formed at the end in an alternative manner;

16 is a schematic partial view of the in 12 shown fan case, in which a flange is shown attached to one end thereof;

17 is a schematic partial view of the in 12 illustrated fan case, in which a flange is shown attached at one end thereof in a first alternative manner thereto;

18 is a schematic partial view of the in 12 illustrated fan case, in which a flange is shown attached at one end thereof in a second alternative manner thereto;

19 is a schematic partial view of the in 12 illustrated fan case, in which a flange is shown attached at one end thereof in a third alternative manner thereto;

20 is a schematic partial view of the in 12 shown fan case, in which stiffening rings with different configurations are attached thereto;

21 is a schematic partial view of a stiffening ring with a first made of a fiber metal laminate configuration, which is attached to a fan casing;

22 Figure 3 is a schematic partial view of a stiffening ring with a second configuration of a fiber metal laminate attached to a fan casing; and

23 is a schematic partial view of a stiffening ring with a third of a fiber metal laminate configuration, which is attached to a fan casing.

DETAILED DESCRIPTION THE INVENTION

Referring in detail to the drawings, in which identical reference numerals designate the same elements throughout the figures 1 in schematic form an exemplary gas turbine engine 10 (High bypass type) used in an aircraft and a longitudinal or axial centerline 12 as having by reference. The engine 10 preferably includes a core gas turbine engine, the whole with the reference numeral 14 and a fan section positioned upstream thereof 16 , The core engine 14 typically includes a substantially tubular outer housing 18 that has an annular inlet 20 Are defined. The outer case 18 encloses and also stores a booster compressor 22 to increase the temperature of the air entering the core engine 14 occurs at a first pressure value. A multi-stage high pressure axial flow compressor 24 receives compressed air from the compressor 22 and further increases the pressure of the air. The pressurized air flows to a combustion chamber 26 where fuel is injected into the pressurized air stream to increase the temperature and energy level of the pressurized air. The high energy combustion products flow out of the combustion chamber 26 to a first (high pressure) turbine 28 to the high pressure compressor 24 via a first (high pressure) drive shaft 30 and then to a second (low pressure) turbine 32 to the booster compressor 22 and the fan section 16 via a second (low pressure) drive shaft 34 To drive, which is coaxial with the first drive shaft 30 is arranged. After the drive of each turbine 28 and 32 The products of combustion leave the core engine 14 via an outlet nozzle 36 to generate a drive jet feed.

The fan section 16 includes a rotatable axial flow fan rotor 38 from an annular fan case 40 is surrounded. It can be seen that the fan case 40 from the core engine 14 over a plurality of substantially radially extending, in the circumferential direction in Ab arranged outlet guide vanes 42 is stored. In this way, the fan case includes 40 a fan rotor 38 and fan rotor blades 44 , A downstream section 46 the fan case 40 extends over an outer portion of the core engine 14 to a secondary or bypass airflow channel 48 to define which generates an additional drive jet feed.

From a flow point of view it can be seen that an initial air flow, represented by the arrow 50 into the gas turbine engine 10 over an inlet 52 in the fan case 14 entry. The airflow 50 happens the fan blades 40 and splits into a first compressed air stream (represented by the arrow 54 ), passing through the channel 48 moved, and a second compressed air flow (shown by the arrow 56 ), which in the booster compressor 22 enters, on. The pressure of the second compressed airflow 56 is increased and enters the high pressure compressor 24 as shown by the arrow 58 one.

After mixing with fuel and combustion in the combustion chamber 26 leave the combustion products 60 the combustion chamber 26 and flow through the first turbine 28 , The combustion products 60 then flow through the second turbine 32 and leave the outlet nozzle 36 to provide a boost to the gas turbine engine 10 to create.

As in the prior art of 2 to see, contains the fan case 40 an annular blade restraint system 62 To prevent parts of fan blades 44 through the fan case 40 pass through. In particular, you will see that the blade restraint system 62 a honeycomb section 64 which is along a radially outer surface 67 an inner section 66 the fan case 40 is arranged. In addition contains a blade restraint system 62 a layer 68 an abradable material (for example, a Kevlar wrap) having a thickness 70 which is adjacent to a radial outer surface 72 of the honeycomb section 64 is arranged. The layer 68 is then at an upstream end 74 and a downstream end 76 at the inner section 66 attached to the fan case.

In one embodiment of the present invention, the blade restraint system replaces 62 the layer 68 of the abradable material through an annular element 78 made of a fiber-metal laminate with a thickness 80 (please refer 3 ) consists. One will realize that the annular element 78 in terms of smaller thickness and lighter weight than the ablatable material layer 68 while providing at least as much resistance to impact and crack propagation. Thus, the annular element 78 radially adjacent to the outer surface 72 of the honeycomb section 64 positioned. In an in 4 illustrated alternative configuration, the blade restraint system 62 also a layer 82 made of abradable material with a thickness 84 contained, the radially outside of an annular element 78 is included to provide additional protection. Even so, one will realize that the inclusion of the annular element 78 , a smaller thickness 84 of the abradable material 82 when the fat 70 of the abradable material discussed above 68 allows. As shown in 5 is at least one indicated part of the inner section 66 the fan case made of a fiber metal laminate.

The fiber metal laminate of the annular element 78 can have various configurations and compositions, such as an aluminum / glass configuration, which is marketed under the product GLARE ^®, an aluminum / marketed under the product ARALL ^® aramid configuration, and a titanium / graphite configuration, sold under the product name TIGRE. However, as the term is used herein, it should be understood that the fibrous metal laminate is essentially one or more layers 89 is defined, which at least one metal layer 90 and a synthetic layer 94 contained, which are interconnected (see 6 ). Of course, a fiber metal laminate may include multiple layers of alternating layers of metal and synthetic material. The in a metal layer 90 metals used preferably include aluminum or titanium, but any lightweight metal or metal alloy having high strength properties may be used. Although it is understood that the metal layer 90 Typically a solid plate, it may be constructed of a grid such as a band-grid or a wire grid. The for the synthetic layer 94 used materials preferably include glass fiber reinforced epoxy resin, carbon reinforced epoxy resin or aramid reinforced epoxy resin, however, any other suitable fiber reinforced composite material may be used.

To ring-shaped material 78 fiber metal laminate, which is originally constructed from substantially flat plates, is a splice region 96 for connecting first and second ends 98 and 100 of it (see 5 ) required. This can be achieved by any number of means, but preferably the ends of the individual layers in the fibrous metal laminate are staggered to create a separation between weak areas. For example, a splice area 96 in 8th shown in a staggered butt joint, with an interruption 102 between opposite ends 104 and 106 a first metal layer 108 at a distance 110 from an adjacent interruption 112 between opposite 114 and 116 a second metal layer 118 are arranged. Although not shown, it will be appreciated that similar breaks occur between opposite ends for metallic layers 120 . 122 . 124 . 126 and 128 are arranged in the same way at a circumferential distance. Of course, synthetic layers help 130 and 132 in replenishing the support 102 such as the synthetic layers 134 and 136 for the interruption 112 ,

An alternative way of producing an annular element 78 is in 9 represented in which the splice region 96 is a staggered overlap connection. As seen therein, the fibrous metal laminate contains at least a first layer 138 and a second location 140 which are adjacent to and substantially parallel to each other except in the splice region 96 extend. Within the splice area 96 you will realize that the first location 138 a first section 142 has, which forms an angle in a first direction, a second section from 144 which is substantially parallel to its orientation outside the splice region 96 extends, and a third section 146 which extends in a second direction complementary to the first direction. In this way, the first situation remains 138 over the entire splice area 96 one piece. The second location 41 contains a first section 148 and a second section 150 in the splice area 96 which are in the same corresponding directions as the first and second sections 142 and 144 the first location 138 extend. But instead of the entire splice area 96 to stay in one piece, shares the second situation 140 approximated at the point where the second section 150 starts, so on, that his first end 152 a second end 154 overlaps and extends substantially parallel across the splice area 96 extends.

One will also recognize that the annular element 98 a plurality of annular segments (denoted by the dashed lines in FIG 7 represented reference numerals 79 and 81 ). Each such annular segment is joined to an adjacent segment at each of its ends by the methods described above. In this way, multiple splice areas within the annular element 78 generated.

It will be appreciated that additional layers of fibrous metal laminate are used for the annular member 78 may be provided in special areas for additional reinforcement. As in the 10 respectively 11 To see such additional layers can be applied inside or outside. In 10 are additional layers 156 . 158 and 160 to the radially inner or main position 162 added, the first location 162 a first section 164 which forms an angle in a first direction has a second portion 166 which is substantially parallel to its original orientation away from the additional layers 156 . 158 and 160 extends, and a third section 168 which extends in a second direction complementary to the first direction. In this way, there is room for additional layers 156 . 158 and 160 created, which is substantially parallel to the second section 166 the main situation 162 are oriented and of a length which is defined by the distance between the first and second sections of the main layer 162 is recorded depending on their radial position. Conversely, additional layers 157 and 159 the outside of the main situation 162 in 11 added.

Instead of replacing one or more components of an existing blade restraint system with an annular member made of fiber metal laminate, it is also contemplated that the blower housing itself may be made of such a material. As in 12 to see is an annular fan case 172 with a blade retention zone 174 made of a fiber metal laminate. It should be noted that acoustic tiles 176 and 178 upstream or downstream of the blade retention zone. Several stiffening rings 180 are preferably in axial spacing relationship along a radial outer surface 182 the fan case 172 positioned to provide additional support. An annular honeycomb layer 184 can optionally be adjacent to an inner radial surface 186 the fan case 172 be positioned with a surface plate 188 be held in their position.

Although not shown, there may also be an abradable layer within the blade retention zone 174 be used. As it further with respect to the acoustic tile 178 So see, they can use acoustic tiles 176 and 178 a layer 177 of a fibrous metal laminate as the perforated surface plate therewith.

It may be necessary for the fan case 172 a flange 190 at one or both of its upstream and downstream ends (see 13 ) especially if it is to be installed as a replacement on existing engine series. Such a flange 190 can be generated by an edge 192 the fan case 172 is rolled after the curing cycle of the substantially cylindrical fan casing 172 takes place. Alternatively, such a flange 190 in one piece with the fan case 172 be prepared, wherein the metallic layers in the flange 190 be formed into the desired shape prior to curing of the laminate. Once it is assembled, the blower housing becomes 172 Hardened to a seamless interface with a flange 190 train.

Another approach is to create a full metal edge 194 on the fan case 172 as shown in 14 , As you can see, composite layers end up 196 and 198 just before the edge 194 and are through metal layers 200 and 202 the same thickness 204 replaced. An optional fastener 206 Can be used to connect the metallic layers 200 and 202 with the metallic layers 201 . 203 and 205 to support. Components can either be on the completely metallic edge 194 be attached after the fan housing 172 cured, or the edge 194 can be drilled and fastened to the components with fasteners and then hardened into a single bonded and secured structure.

A metal flange 208 can a section 210 included, which with one end 212 the fan case 172 is connectable with a reduced thickness. A metal ring 214 may be adjacent to an outer surface 216 a fan case 172 be provided, with several fasteners 218 the flange 208 on the fan case 172 (please refer 15 ). 16 to 18 provide more options for mounting components 220 . 222 respectively 224 directly on the fiber metal laminate of the fan case 172 represents.

To form a fiber metal laminate panels in a fan case 172 To facilitate, are also preferred several miter cuts 170 provided therein (see 19 ). You will realize that aluminum strip 171 and 173 may be positioned on either side of the fibrous metal laminate in order to produce a flange 190 to support.

As stated hereinabove, stiffening rings 180 with different configurations on the outer radial surface 182 the fan case 172 (please refer 20 ) be attached. Such stiffening rings 180 are essentially made of a metal (eg aluminum, steel or titanium). It will be appreciated that stiffening rings, as denoted by the reference numeral 183 in the 21 to 23 may be constructed of a fiber-metal laminate or other composite laminate or egg ner sandwich structure, which with the fan case 172 either after the hardening of the fan case 172 or during the cure cycle of the rest of the fan case 172 get connected.

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 without departing from the spirit of the present invention can be led. Accordingly, it is intended in the appended claims to encompass all such changes and modifications that fall within the scope of the present invention.

In a gas turbine engine 10 with a plurality of radially extending blades 44 on an annular disc 38 are attached, with the blades 44 and the disc 38 around a longitudinal axis 12 of the engine 10 rotatable, a blade restraint system 62 with an annular housing 40 that is radially outside the blades 44 and positioned in surrounding relation thereto, the annular housing 40 at least one layer of a fiber metal laminate used therewith 78 contains. The blade restraint system 62 can also be a honeycomb layer 64 included radially outward of an inner section 66 of the annular housing 40 is arranged, wherein the layer of the fiber metal laminate 78 radially outside of the honeycomb layer 64 or disposed radially inside thereof. For both cases, the blade restraint system could 62 a radially outer layer of a Kevlar wrap 82 contain.

10

Aircraft turbofan engine (generally)

12

longitudinal axis

14

Core gas turbine engine (generally)

16

fan section

18

outer casing for core engine

20

annular core engine intake

22

booster compressor

24

High-pressure compressors

26

combustion chamber

28

first (High pressure) turbine

30

first (High pressure) drive shaft

32

second (Low pressure) turbine

34

second (Low pressure) drive shaft

36

outlet nozzle

38

fan rotor

40

annular fan case

42

Auslassführungs vanes

44

Fan rotor blades

46

downstream Section of the fan case

48

In addition to air flow channel

50

the Initial air flow arrow

52

inlet to the fan case

54

the first (outer) compressed Airflow depicting arrow

56

the second (inner) compressed air stream representing arrow

58

the compressed air flow to the high pressure compressor performing arrow

60

combustion products

62

A blade retention system (generally)

64

honeycomb section

66

internal Section of the annular fan case

67

radial inner surface of the honeycomb section

68

layer made of removable material

70

thickness of the abradable material

138

first location

140

second location

142

first Section of the first layer in the splice region

144

second Section of the first layer in the splice region

146

third Section of the first layer in the splice region

148

first Section of the second layer in the splice region

150

second Section of the second layer in the splice region

152

first End of the second situation

154

second End of the second situation

156

additional location

158

additional location

160

additional location

162

main location

164

first Section of the main situation

166

second Section of the main situation

168

third Section of the main situation

170

miter cuts in the fiber metal laminate

171

outer aluminum strips

172

Winding case (fiber metal laminate)

173

inner aluminum strips

174

Blade retention zone

176

upstream acoustic tile

177

Fiber metal laminate layer for acoustic tile

178

downstream acoustic tile

180

stiffening rings (Metallic)

182

outer radial surface the fan case

183

stiffening rings (Fiber metal laminate)

184

layer made of removable material

186

inner radial surface the fan case

188

outer plate

190

flange

192

edge the fan case

194

Total metal rim the fan case

Claims (10)

In a gas turbine engine ( 10 ) with a plurality of radially extending blades ( 44 ) mounted on an annular disc ( 38 ), the blades ( 44 ) and the disc ( 38 ) about a longitudinal axis ( 12 ) of the engine ( 10 ), a blade restraint system ( 62 ) with an annular housing ( 40 ) located radially outside the blades ( 44 ) and in surrounding relation thereto, the annular housing ( 40 ) at least one layer of a there with fibrous metal laminate used ( 78 ) contains. Blade restraint system ( 62 ) according to claim 1, further comprising a honeycomb layer ( 64 ), which radially outwardly of an inner portion ( 66 ) of the annular housing ( 40 ) is arranged. Blade restraint system ( 62 ) according to claim 2, wherein the layer of fibrous metal laminate ( 78 ) radially outside the honeycomb layer ( 64 ) is positioned. Blade restraint system ( 62 ) according to claim 3, further comprising a layer ( 82 ) has a Kevlar wrap which extends radially outward of the layer of the fiber metal laminate ( 78 ) is positioned. Blade restraint system ( 62 ) according to claim 2, wherein the layer of fibrous metal laminate ( 78 ) radially inside the honeycomb layer ( 64 ) is positioned. Blade restraint system ( 62 ) according to claim 5, which further comprises a layer ( 82 ) has a Kevlar wrap which extends radially outward of the layer of the fiber metal laminate ( 78 ) is positioned. Blade restraint system ( 62 ) according to claim 1, wherein the layer of fibrous metal laminate ( 78 ) at least one layer ( 89 ), the position ( 89 ) further comprises: (a) at least one metal layer ( 90 ); and (b) a synthetic layer ( 94 ) with the metal layer ( 90 ) connected is. Blade restraint system ( 62 ) according to claim 1, wherein the layer of fibrous metal laminate ( 78 ) several miter cuts ( 170 ) therein to allow expansion in a circumferential manner. Blade restraint system ( 62 ) according to claim 7, wherein the layer of fibrous metal laminate ( 78 ) additional layers ( 156 . 158 . 160 ) in a specified range. Blade restraint system ( 62 ) according to claim 1, wherein the layer of fibrous metal laminate ( 78 ) at their ends ( 98 . 100 ) is spliced together to form a ring shape.