Abstract:
A containment case for a turbine engine having a plurality of blades rotating therein has a first plurality of metallic layers each of the first plurality of layers being resistant to penetration of fragments therethrough and a second plurality of metallic layers, each of the second plurality of layers absorbing kinetic energy of fragments striking said second plurality of layers.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    This invention relates to turbine engines having rotatable blade arrays and particularly to a containment case for confining blade, disk and impeller fragments which may fail during engine operation. 
         [0002]    Gas turbine engines, such as those which power commercial aircraft, typically include multiple arrays of fans, compressors, turbine disks, and turbine blades. Each blade array comprises a multitude of blades that are attached to and extend radially outwardly from a hub. During engine operation each hub and associated blade array rotate about a longitudinally extending central axis. A non-rotating case, which is typically cylindrical or frustoconical in shape, circumscribes the tips of the blades and is radially spaced therefrom by a small amount. The case has a leading edge and a trailing edge, at least one of which is connected to an adjacent engine case. The case defines the outer boundary of a gas flow path that extends longitudinally through the engine. 
         [0003]    During engine operation, it is possible for a fragment of a fan, compressor, turbine disk, and turbine blade to crack and become separated. Separation of a fragment is rare and is usually attributable to failure of a component. Because the kinetic energy of a blade fragment is considerable (particularly if the fragment comprises substantially the part of the disk/compressor) the fragment is capable of damaging engine and aircraft components which lie along the fragment&#39;s trajectory. To prevent such damage, the case which circumscribes a blade array is designed to confine or contain a fragment and is commonly referred to as a containment case. 
         [0004]    One type of containment case is known as a softwall case. A softwall case comprises multiple layers of a light weight penetration resistant fabric wrapped around a rigid but penetrable support ring. A separated blade fragment will penetrate the support ring but be contained by the fabric. Softwall construction is expensive, but is also light weight, a distinct advantage in an aircraft application. A second type of case, known as a hardwall case, comprises a ring having sufficient radial thickness to resist penetration of a blade fragment. The choice of hard wall or soft wall construction depends largely on the case diameter and the temperature. For a large diameter case, hard wall construction is prohibitively heavy, and therefore soft wall construction, despite being expensive, is preferred. For a small diameter case, the radial thickness required for penetration resistance imposes only a modest weight penalty and so the less expensive hard wall construction is usually favored. 
         [0005]    Although hard wall construction is almost universally preferred for small diameter cases, it is not without several disadvantages. First, the thickness and rigidity of a hard wall case prevent it from deflecting readily when struck by a blade fragment. Consequently, the full force of the impact is concentrated over a very short time interval and therefore is quite damaging. The abruptness and resultant severity of the impact contribute to the required thickness of the case and therefore to its weight. In addition, the severe impact energy is transmitted to auxiliary components, such as engine control units and pneumatic lines which may be attached to the exterior of the engine (and especially to the exterior of the containment case), thereby exposing those components to potentially damaging forces. A second disadvantage of a conventional hard wall containment case is that it is typically machined from forgings, which adds cost to the containment ring. 
       SUMMARY OF THE INVENTION 
       [0006]    According to an embodiment disclosed herein, a containment case for a turbine engine having a plurality of blades rotating therein has a first plurality of metallic layers each of the first plurality of layers being resistant to penetration of blades and disk fragments therethrough and a second plurality of metallic layers, each of the second plurality of layers absorbing kinetic energy of blades and disk fragments striking said second plurality of layers. 
         [0007]    According to a further embodiment shown herein, a method for mounting layers that resist penetration of a blade or part thereof therethrough and absorb the kinetic energy thereof, includes providing a first plurality of metallic layers each of the first plurality of layers being resistant to penetration of blades therethrough, and providing a second plurality of metallic layers, each of the second plurality of layers absorbing kinetic energy of blades striking the any one of second plurality of layers. 
         [0008]    According to a still further embodiment disclosed herein, a method of mounting a containment ring upon a portion of an engine within which a rotating component that may break or fragment is disposed includes providing a first plurality of metallic layers each of the first plurality of layers being resistant to penetration of fragments therethrough, providing a second plurality of metallic layers, each of the second plurality of layers absorbing kinetic energy of fragments striking one of second plurality of layers, and, disposing the first plurality of metallic layers and the second plurality of metallic layers about the portion. 
         [0009]    These advantages and the features and operation of the invention will become more apparent in light of the following description of the best mode for carrying out the invention and the accompanying drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]      FIG. 1  is a schematic, cross sectional side view of an aircraft gas turbine engine. 
           [0011]      FIG. 2  is a cross sectional side view of an embodiment of a containment device used in the aircraft gas turbine engine of  FIG. 1 . 
           [0012]      FIG. 3  is a cross sectional side view of the containment device of  FIG. 2 . 
           [0013]      FIG. 4  is a perspective view of the containment device of  FIG. 2 . 
           [0014]      FIG. 5  shows an embodiment method of constructing the containment device of  FIG. 4   
           [0015]      FIG. 6  shows a further embodiment method of constructing the containment device of  FIG. 4   
           [0016]      FIG. 7  shows an embodiment method of constructing the engine of  FIG. 1 . 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       [0017]    Referring to  FIG. 1 , an aircraft gas turbine engine  10  includes a fan  12 , low pressure and high pressure compressors  14 ,  16 , a combustion chamber  18 , and low pressure and high pressure turbines  20 ,  22 . The gas turbine engine  10  may be a main engine or incorporated in an auxiliary power unit of an aircraft. A high pressure rotor comprises a high pressure compressor hub  24  and a high pressure turbine hub  26  connected together by a shaft  28 , and arrays of blades, such as representative compressor and turbine blades  30 ,  34 . The blades extend radially outwardly from their respective hubs, across a primary flow path  36 , and into close proximity with a primary or core case assembly  38 . Similarly a low pressure rotor comprises fan, low pressure compressor and low pressure turbine hubs  40 ,  42 ,  44  connected together by shaft  46  and arrays of blades such as representative fan, low pressure compressor and low pressure turbine blades  48 ,  50 ,  52 . Blades  48 ,  50 ,  52 , extend radially outwardly from their respective hubs, across the primary flow path and, in the case of the fan blades  48 , across a secondary flow path  49  as well, and into close proximity with the core case assembly  38 , or a fan case assembly  58 . The case assemblies define the outer flow path boundaries for the primary and secondary flow paths. 
         [0018]    During engine operation, the turbines rotatably drive the fan and compressors about a longitudinally extending central axis  31 . Since a fragment (not shown) may become separated from the rotor during engine operation, a containment case  60  has a particular structure, as will be discussed infra, to contain such a fragment. If a fragment, e.g., such as blade  52  or a fragment (not shown) thereof, breaks loose, that blade  52  or fragment has ballistic properties like a bullet and the containment case  60  deals with, as will be discussed infra, the penetration aspects associated with those ballistic properties. The case assembly  60  then absorbs the kinetic energy of the blade  52  to minimize damage to the engine  10 . 
         [0019]    Referring to  FIGS. 1-4 , containment case  60  circumscribes an array of fan blades  52  as shown in  FIG. 1 . The containment case has an impact zone  62  that is a region where a separated blade fragment may strike the containment case. The containment case  60  includes: a spool  64  that has a base  66 , shown as cylindrical in  FIGS. 2-6 , and a pair of rings  68  attached, by brazing or the like, to and extending from each end  70  of the cylindrical base  66  to define the spool  64 ; and, a plurality of layers  72  disposed around the spool  64  between the rings  68 . Each ring  68  has a flange  71  or the like through which a bolt  73  attaches the containment case  60  to the core case assembly  38  (see  FIG. 1 ). The base  66  also be contoured to adapt to the surface of a core case assembly  38  or the like. 
         [0020]    To attach the containment case  60  to the engine  10  (see  FIG. 7 ), the containment case  60  is slid over the core case assembly  38  during construction of the engine  10  and bolted via bolts  73  through flange  71  to the core case assembly  38 . Other attachment methods may be used such as welding or brazing or the like. 
         [0021]    The spool  64  including cylindrical base  66 , rings  68  and flange  71  are constructed of stainless steel or the like. The layers  72  are made up of a relatively high strength metallic layers  74  such as Inconel® 718 steel or others that address the ballistic penetrative properties of a separated blade, such as blade  52 , or a fragment thereof, and relatively ductile metallic layers  76 , such as Inconel® 625 steel to absorb the kinetic energy of a separated blade, such as blade  52 , or a fragment thereof. Each layer  74  is between about 0.1 inches (or 0.25 cm) and 0.01 (or 0.03 cm) thick and each layer  76  is similarly between about 0.1 inches (or 0.25 cm) and 0.01 (or 0.03 cm) thick though different thicknesses may be used for different applications and other materials. 
         [0022]    The layers  74 ,  76  may be interleaved or may be grouped depending on the required application (see  FIG. 2 ). Each layer  74 ,  76  may be attached by butt welding (see line  78  in  FIGS. 4-6 ) or the like to itself to form a strip upon the spool  64  or attached to itself and slid on the cylindrical base  66  and any already applied layer  74 ,  76 , before a ring  68  is attached to the cylindrical base  66 . Each subsequent layer  74 ,  76  gets larger in diameter. Additionally a layer  74  or  76  may be attached to an adjacent layer  74  or  76 , welded to the cylindrical base  66  and rolled or coiled up upon the spool  64 . The last layer  74 ,  76  is then attached to the rest of the spooled layers  72 ,  74 . 
         [0023]    In the event that a separated blade fragment strikes the containment case during engine operation, the containment case  60 , owing to its layers  74  that primarily resist the penetration of the separated blade  52  or fragment thereof, and the layers  76  that primarily absorb the energy of the separated blade  52  or fragment thereof. 
         [0024]    The invention has been described as a containment case for an array of compressor blades  52  in a turbine engine  10 . However the invention is equally applicable to the fan  48  and turbine blade  52  arrays of a turbine engine and to any other type of machinery where it is desirable to confine separated component fragments. These and other changes and modifications to the invention can be made without departing from the spirit and scope of the appended claims. 
         [0025]    Although the invention has been shown and described with respect to a best mode embodiment exemplary thereof, it should be understood by those skilled in the art that various modifications, changes, omissions and additions in the form and detail thereof may be made without departing from the spirit and scope of the invention. 
         [0026]    Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments. 
         [0027]    The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.