LINER FOR GAS TURBINE ENGINE

Abstract

An embodiment of a gas turbine engine having a liner and casing is disclosed. The liner is disposed between the casing and a rotatable turbomachinery component. In one form, the rotatable turbomachinery component is a turbofan engine. The liner can be thin relative to a distance between the liner and the casing. Protrusions can be used between the liner and the casing. The protrusions can be a rib and can extend from the casing or the liner. An insert, such as an abradable surface, can be used with the liner. A filler can be used in the space between the liner and the casing.

Description

TECHNICAL FIELD

The present disclosure generally relates to liner members for gas turbine engines. More particularly, but not exclusively, the present disclosure relates to configurations and orientations of liner members relative to casings of gas turbine engines.

BACKGROUND

Providing mechanisms to contend with blade out events, such as fan blade out events (FBO), remains an area of interest. Some liner systems employ honeycomb liners which can be used in the event of a blade rub or a blade out condition and, in these embodiments, a low density honeycomb can be used on the backside of an abradable lining that includes an epoxy filled honeycomb. Gas turbine engines can use these liners directly bonded to the inside of the fan case or in the form of a set of cassettes that are bolted into place. Fillers and/or sealants can be used between liner segments and at liner to casing interfaces. Some existing systems have various shortcomings relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.

SUMMARY

One embodiment of the present disclosure is a unique liner for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for liner systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

BRIEF DESCRIPTION OF THE FIGURES

FIG. 1 discloses one embodiment of a gas turbine engine;

FIG. 2 discloses an embodiment of a liner and casing;

FIG. 3 discloses an embodiment of a liner and casing;

FIG. 4 discloses an embodiment of a liner and casing;

FIG. 5 discloses an embodiment of a liner and casing; and

FIG. 6 discloses an embodiment of a liner and casing.

DETAILED DESCRIPTION OF THE ILLUSTRATIVE EMBODIMENTS

For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to the embodiments illustrated in the drawings and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the disclosure is thereby intended. Any alterations and further modifications in the described embodiments, and any further applications of the principles of the disclosure as described herein are contemplated as would normally occur to one skilled in the art to which the disclosure relates.

With reference to FIG. 1, a gas turbine engine 50 is illustrated having a compressor 52, combustor 54, and turbine 56 which together can be used to produce a useful power. Generally speaking, a working fluid such as air enters the gas turbine engine 50 whereupon it is compressed through action of the compressor before being mixed with a fuel and combusted in the combustor 54. The turbine 56 is arranged to receive a flow from the combustor 54 and extract useful work from the flow. The gas turbine engine can be used to provide power to, for example, an aircraft. As used herein, the term “aircraft” includes, but is not limited to, helicopters, airplanes, unmanned space vehicles, fixed wing vehicles, variable wing vehicles, rotary wing vehicles, unmanned combat aerial vehicles, tailless aircraft, hover crafts, and other airborne and/or extraterrestrial (spacecraft) vehicles. Further, the present disclosures are contemplated for utilization in other applications that may not be coupled with an aircraft such as, for example, industrial applications, power generation, pumping sets, naval propulsion, weapon systems, security systems, perimeter defense/security systems, and the like known to one of ordinary skill in the art.

Turning now to FIG. 2, an embodiment of the gas turbine engine 50 is shown wherein a turbomachinery component in the form of a fan section 58 of a turbofan engine is illustrated. A pressure of a flow stream 60 is changed via operation of a fan 62 before the flow stream is split between a core flow 64 and bypass flow 66. A casing 68 and a member 70 forming a liner surface (hereafter liner 70 whether or not the member is a single construction, unitary component, layered assembly of parts, etc. as is described further herein) are used near the fan 62 and are shaped to be useful during a Fan Blade Out (FBO) condition in which a portion of the fan 62 may penetrate the liner 70 and be contained by the casing 68. The casing 68 includes a portion 72 directed away from a flow path through the fan section 58 and the liner 70 is situated between the casing 68 and the flow path. The liner 70 forms a flow path surface over which the flow stream is conveyed by operation of the gas turbine engine 50. The portion 72 directed away from the flow path can take a variety of forms such as the c-shape disclosed in the illustrated embodiment, but other configurations of the recess that is formed are also contemplated herein.

Turning now to FIG. 3, a volume 74 is formed between the liner 70 and the casing that in some forms can be substantially annular in shape. The volume 74 can have substantially the same orientation/size/etc. circumferentially around the gas turbine engine 50, but some variations are contemplated herein. For example, some structure(s) can be located within the volume 74 for any variety of purposes. The volume 74 generally has a depth d that in some embodiments, such as the illustrated embodiment, can vary from a forward end 76 to an aft end 78 of the volume 74. The depth d is generally as large as or larger than a thickness t of the liner 70, but smaller ranges are also contemplated in any of the various embodiments herein. In some forms the depth d is several times the size of the liner thickness t. In those cases where the depth d varies widely between the forward end 76 and aft end 78, and/or in those cases where the thickness t varies widely between the forward end 76 and aft end 78, the depth d can be as large or larger than the size of the liner thickness t over a relatively large range of the axial length of the volume 74, and/or than the size of the thickness t in the area around the fan 62. Such a characteristic can sometimes be referred to as a base thickness. To set forth just a few non-limiting examples, the depth d can be twice as large as or larger than the thickness t, where such larger variation can be any multiple, or intermediate multiple.

Lugs 80 can be provided to mount the liner 70 to the casing 68. The lugs 80 can be formed in the casing 68 and in some forms are intermediate structure coupled to the casing 68. The lugs 80 can take any suitable form, including flanges, etc., that are used to provide a supporting surface to which the liner 70 is affixed. The lugs 80 can be circumferentially formed surfaces, and, in some embodiments, the lugs 80 can be localized features at select circumferential locations. The forward and aft lugs 80 can take a similar form (integral with casing, coupled with casing, flanges, etc.), but in some embodiments the forward and aft lugs 80 can be different.

The liner 70 can be affixed to the casing 68 and/or the lugs 80 using any variety of techniques. For example, the liner 70 can be affixed relative to the casing 68 using mechanical fasteners such as bolts, screws, rivets, etc., and in some forms can be bonded to the casing 68 using chemical and/or metallurgical bonding techniques, such as adhesion bonding or welding, to set forth just a few non-limiting examples of affixing the liner 70.

The manner in which the liner 70 is affixed, and alternatively and/or additionally the construction of the liner 70, can determine the manner in which the liner 70 reacts when contacted by the fan 62. It is generally contemplated that the liner 70 reacts by moving and/or removing at least a portion such that a space is created as a result of a trajectory of the fan 62. Such a space can be permanently formed such as, for example, when the liner 70 yields to a reaction with the fan 62. For example, the liner 70 can react to contact from the fan 62 by being frangible. In this example such a reaction can be determined by the construction of the liner 70, such as, for example, whether it is constructed of a single material having ductile properties. In some embodiments, the liner 70 can react by rupturing at a contact area with the fan 62, and in some forms can additionally and/or alternatively separate at one or more points where the liner 70 is affixed. In still further forms, the liner 70 can react by separating from the casing 68 and/or lugs 80 in lieu of permanently yielding. In still further forms, the liner 70 can be separated from the casing 68 through a combination of yielding and separating. Not all portions of the liner 70 need react when the fan 62 makes contact. For example, some portion of the liner 70 can remain behind after a contact. In short, any variety of dynamic impact reactions is contemplated.

The liner 70 of the illustrated embodiment generally extends around the circumferential annular flow space of the turbofan engine. The liner 70 can be segmented such that a series of liner constructions are distributed around the fan 62. In those embodiments in which segments are used, not all segments need react to a blade contact. For example, when a Fan Blade Out occurrence causes a rotor imbalance in which the fan 62 orbits about an axis, certain circumferential locations of the liner 70 can react with the fan while other circumferential locations do not react with the fan. The segments can be similarly shaped, but in some embodiments not all segments need be similar to each other. A seal, such as a filler or sealant, can be used between circumferential segments. Additionally and/or alternatively, a seal can be used in the forward and aft portions of the liner 70.

In one non-limiting embodiment, the volume 74 between the liner and the casing can be substantially empty. For example, the space can provide an annular volume of air. In some embodiments the volume 74 can be segregated in some fashion. In those embodiments, the segregated compartments can be substantially empty of any materials with the exception of air or other working fluid. In still further non-limiting embodiments, the volume 74 can include a low-density filler that could be used in some applications to increase stiffness and dampening. Such a filler could be located at one or more circumferential locations in the volume 74, or alternatively be located throughout the volume 74 distributed around the fan 62.

The liner 70 can be constructed in a variety of manners. For example, the liner 70 can be constructed as a single article or as an article that has portions fastened/bonded/etc. to one another. Such an article having portions connected to one another can take the form of a layered composition. The liner 70 can be cast, stamped, molded, or made in a composite construction. The liner 70, furthermore, can be made of one or more materials such as metallic, plastic, composite, etc. In short, the liner 70 can take on any variety of constructions.

Turning now to FIG. 4, one non-limiting embodiment is shown of a liner 70 having an insert 82 which can be used in the event of a rubbing event with the fan 62. Such an insert 82 can be an abradable material applied to the liner 70 using a variety of techniques whether mechanically fastened, cast, chemically or metallurgically bonded, spray coated, etc. The insert 82 can be any variety of depths and can extend any distance between the forward end 76 and aft end 78. As shown in the illustrated embodiment, the insert 82 extends only partially between the forward end 76 and aft end 78. In some forms the insert 82 is located in the area of the fan 62.

One or more protrusion(s) 84 can extend between the casing 68 and the liner 70, two non-limiting examples of which are shown in FIGS. 5 and 6. The protrusions 84 can take a variety of shapes, sizes, orientations, etc. and, in the illustrated embodiments, are shown generally as ribs that extend generally along a line. The protrusions 84 can be integrally formed with the structure from which it extends, but in some forms can be fastened using any variety of techniques. Furthermore, any number of protrusions 84 can be used in any given liner 70. FIG. 5 illustrates a number or protrusions 84 extending from the liner 70 toward the casing 68 and in which a gap is formed between the end of one or more protrusions 84 and the casing 68. The gaps can be, but need not be, the same for each of the protrusions 84. In some forms the gaps may not be present. A frictional interface can be used in some forms. FIG. 6 illustrates protrusions 84 extending from the casing 68 and in which no gap is present with respect to the liner 70. One or more gaps between the liner 70 and protrusions 84 could be formed in some embodiments.

In those embodiments that contain one or more protrusions from the backside of the liner 70, the thickness t discussed above in regard to the depth d of the volume 74 can be considered the thickness t on either side of the protrusion(s). In this way the liner 70 can be said to have a base from which the protrusions extend, where the base includes the thickness t. In those embodiments where the thickness t varies over the length of the liner 70, the protrusions can extend from locations other than the base. For example, the thickness t between the upstream end and downstream end of the liner 70 can vary, but generally, with the exception of intermediate structures such as the protrusions 84, is substantially less than the depth d of the volume 74 created between the liner 70 and the casing 68 as discussed herein.

The volume 74 formed between the liner 70 and the casing 68 can be considered the volume between the casing and the liner 70, whether or not the volume 74 includes a low density filler, etc. It will be understood that in those embodiments including protrusions 84, the depth d of the volume 74 is generally the depth associated between the casing 68 and the liner 70, and not the minimal distance, such as the gap shown in FIG. 5, between the protrusions 84 and the casing 68.

Any of the various embodiments disclosed herein can be combined with other embodiments. For example, the insert 82 can be used with any of the various embodiments. For that matter, the protrusions 84 can be used in any of the various embodiments. Other combinations are also contemplated herein.

One aspect of the present application provides an apparatus comprising a gas turbine engine having a casing and a flow path located radially inward of the casing within which is disposed a rotatable turbomachinery component. The flow path is bounded by a construction that provides a liner surface located between the rotatable turbomachinery component and the casing. The construction has a base thickness between a flow path side and a non-flow path side which is smaller than an offset between the construction and the casing. The offset is free of the construction. The construction that provides the liner surface forms a flow path surface during a nominal mode of operation of the rotatable turbomachinery component. The construction that provides the liner surface is constructed to be sacrificial during an off-nominal mode of operation of the rotatable turbomachinery component such that a portion of the turbomachinery component can penetrate into an area free from the construction.

A feature of the present application provides wherein the rotatable turbomachinery component is a fan, and wherein an annular liner is formed that is constructed from a plurality of construction segments. Another feature of the present application provides wherein the construction that provides the liner surface is a single article that forms the flow path surface and is substantially free of internal voids.

Yet another feature of the present application provides wherein the construction includes a base portion and an abradable material adjacent the base portion. Still another feature of the present application provides wherein the off-nominal mode of operation includes an orbiting motion of the turbomachinery component, and wherein the gas turbine engine also including protrusions disposed in the area free from the construction.

Yet still another feature of the present application provides wherein the construction that provides the liner surface includes the protrusions. Still yet another feature of the present application provides wherein the area free from the construction includes a low density filler material. A further feature of the present application provides wherein the construction includes a plurality of materials.

Another aspect of the present application provides an apparatus comprising a gas turbine engine including a fan section having a rotatable fan blade portion structured to change a pressure of a working fluid flowing through a flow path of the fan section, a casing radially outward of the rotatable fan blade portion and having a shape the diverges from the flow path of the fan section, and a component having a liner surface radially inward from the casing to separate the casing from the rotatable fan blade portion. The component is offset from the casing to create a space free of a honeycomb structure intermediate the liner and the casing. The space is larger than a thickness of the component axially coincident with the rotatable fan blade portion.

A feature of the present application provides wherein the component is fixed relative to the casing through one of mechanical fastening or being bonded in place. Another feature of the present application provides wherein a coupling interface between the component and the casing is structured to fail and release the component when the fan blade portion contacts the component.

Yet another feature of the present application provides wherein the component is stamped sheet metal. Still another feature of the present application provides wherein the component is molded, and wherein the component is one of plastic and composite.

Yet still another feature of the present application provides wherein the component includes a solidity that substantially lacks internal voids. A further feature of the present application provides wherein liner includes a base plate and which further includes elongate portions that extend from between the back of the component and the casing. A still further feature of the present application provides wherein the space free of a honeycomb structure is an empty space.

Yet another aspect of the present application provides an apparatus comprising a gas turbine engine having a compressor rotatingly coupled with a turbine and a casing configured to enclose a bladed component of the gas turbine engine, and reaction means for ingressing a portion of the bladed component into a space formed between the casing and the reaction means. The reaction means is disposed between the casing and the bladed component to create a non-flow path volume. The reaction means having a predominate thickness extending between an upstream end and a downstream end. The non-flow path volume having a depth measured between the casing to the predominate thickness. The thickness of the reaction means is substantially smaller than the depth measured between the casing and the predominate thickness.

Still yet another aspect of the present application provides a method comprising a number of operations. The operations including rotating a bladed member of a gas turbine engine having a casing and a liner situated between the casing and the bladed member to create an area therebetween, the bladed member travelling along a nominal arc of rotation, impacting at least a portion of the bladed member with a liner located between the bladed member and a casing of the gas turbine engine, and reacting the liner with the portion of the bladed member to expose a volume formed by a placement of the liner relative to the casing, the liner providing an offset flow path surface for a working fluid from the casing.

A feature of the present application further includes penetrating the liner with the portion of the bladed member. Another feature of the present application provides wherein the reacting includes breaking the liner.

Yet another feature of the present application provides wherein the breaking includes removing at least a portion of the liner. Still another feature of the present application provides wherein the removing includes breaching a manner in which the liner is affixed relative to the casing. Yet still another feature of the present application provides wherein the breaching includes destroying a bond between the liner and the casing.

While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiments have been shown and described and that all changes and modifications that come within the spirit of the disclosures are desired to be protected. It should be understood that while the use of words such as preferable, preferably, preferred or more preferred utilized in the description above indicate that the feature so described may be more desirable, it nonetheless may not be necessary and embodiments lacking the same may be contemplated as within the scope of the disclosure, the scope being defined by the claims that follow. In reading the claims, it is intended that when words such as “a,” “an,” “at least one,” or “at least one portion” are used there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. When the language “at least a portion” and/or “a portion” is used the item can include a portion and/or the entire item unless specifically stated to the contrary.

Claims

1. An apparatus comprising: a gas turbine engine having a casing and a flow path located radially inward of the casing within which is disposed a rotatable turbomachinery component, the flow path bounded by a construction that provides a liner surface located between the rotatable turbomachinery component and the casing, the construction having a base thickness between a flow path side and a non-flow path side which is smaller than an offset between the construction and the casing, the offset being free of the construction;wherein the construction that provides the liner surface forms a flow path surface during a nominal mode of operation of the rotatable turbomachinery component, and wherein the construction that provides the liner surface is constructed to be sacrificial during an off-nominal mode of operation of the rotatable turbomachinery component such that a portion of the turbomachinery component can penetrate into an area free from the construction.

2. The apparatus of claim 1, wherein the rotatable turbomachinery component is a fan, and wherein an annular liner is formed that is constructed from a plurality of construction segments, and wherein the construction that provides the liner surface is a single article that forms the flow path surface and is substantially free of internal voids.

3. The apparatus of claim 1, wherein the construction includes a base portion and an abradable material adjacent the base portion.

4. The apparatus of claim 1, wherein the off-nominal mode of operation includes an orbiting motion of the turbomachinery component, and wherein the gas turbine engine also includes protrusions disposed in the area free from the construction.

5. The apparatus of claim 4, wherein the construction that provides the liner surface includes the protrusions.

6. The apparatus of claim 1, wherein the area free from the construction includes a low density filler material.

7. The apparatus of claim 1, wherein the construction includes a plurality of materials.

8. An apparatus comprising: a gas turbine engine including a fan section having a rotatable fan blade portion structured to change a pressure of a working fluid flowing through a flow path of the fan section;a casing radially outward of the rotatable fan blade portion and having a shape that diverges from the flow path of the fan section; anda component having a liner surface radially inward from the casing to separate the casing from the rotatable fan blade portion, the component offset from the casing to create a space free of a honeycomb structure intermediate the liner and the casing, the space being larger than a thickness of the component axially coincident with the rotatable fan blade portion.

9. The apparatus of claim 8, wherein the component is fixed relative to the casing through one of mechanical fastening or being bonded in place.

10. The apparatus of claim 8, wherein a coupling interface between the component and the casing is structured to fail and release the component when the fan blade portion contacts the component.

11. The apparatus of claim 8, wherein the component is stamped sheet metal.

12. The apparatus of claim 8, wherein the component is molded, and wherein the component is one of plastic and composite.

13. The apparatus of claim 8, wherein the component includes a solidity that substantially lacks internal voids.

14. The apparatus of claim 8, wherein the liner includes a base plate and which further includes elongate portions that extend from between a back of the component and the casing.

15. The apparatus of claim 8, wherein the space free of a honeycomb structure is an empty space.

16. The apparatus of claim 8 further comprising a compressor rotatingly coupled with a turbine and wherein the casing is configured to enclose the fan blade; and reaction means for ingressing a portion of the fan blade into a space formed between the casing and the reaction means, the reaction means disposed between the casing and the fan blade to create a non-flow path volume, the reaction means having a predominate thickness extending between an upstream end and a downstream end, the non-flow path volume having a depth measured between the casing to the predominate thickness, and wherein the thickness of the reaction means is substantially smaller than the depth measured between the casing and the predominate thickness.

17. A method comprising: rotating a bladed member of a gas turbine engine having a casing and a liner situated between the casing and the bladed member to create an area therebetween, the bladed member travelling along a nominal arc of rotation;impacting at least a portion of the bladed member with a liner located between the bladed member and a casing of the gas turbine engine; andreacting the liner with the portion of the bladed member to expose a volume formed by a placement of the liner relative to the casing, the liner providing an offset flow path surface for a working fluid from the casing.

18. The method of claim 17, which further includes penetrating the liner with the portion of the bladed member.

19. The method of claim 17, wherein the reacting includes breaking the liner, and wherein the breaking includes removing at least a portion of the liner.

20. The method of claim 19 wherein the removing includes breaching a manner in which the liner is affixed relative to the casing, and wherein the breaching includes destroying a bond between the liner and the casing.