Patent Application
20190085856
The disclosure relates generally to fan blade containment cases in gas turbine engines.
Gas turbine engine fan blade containment cases may be designed to contain a liberated fan blade and associated high energy particles tending to result from separation of fan blade.
In various embodiments the present disclosure provides a hybrid composite-metallic containment system, comprising a cylindrical structure disposed radially outward of a fan blade and having a first composite layer extending about a circumference; a first metallic layer extending about the circumference; a second composite layer extending about the circumference, wherein the second composite layer is disposed radially inward of the first composite layer and the first metallic layer; and a binding component.
In various embodiments, the first metallic layer is disposed radially inward of the first composite layer. In various embodiments, a hybrid composite-metallic containment system further comprises a second metallic layer extending about the circumference radially inward of the first metallic layer. In various embodiments, the first metallic layer comprises a perforation. In various embodiments, the first composite layer comprises at least one of carbon, carbon fibers, glass fibers, aramid, para-aramid, polyethylene, or ultra-high-molecular-weight polyethylene. In various embodiments, the first metallic layer comprises at least one of steel, stainless steel, titanium, titanium alloy, aluminum, aluminum alloy, nickel, or nickel alloy. In various embodiments, the first metallic layer is at least one of sheet metal, spun metal, braided metal, metal mesh, metal fiber, woven metal, or metal fabric. In various embodiments, the binding component comprises at least one of a resin, an epoxy, a thermosetting polymer, a scrim supported adhesive, or pre-impregnated material. In various embodiments, the first composite layer and the second composite layer comprise an overwrapping composite layer about the first metallic layer. In various embodiments, the first composite layer is disposed radially inward of the first metallic layer.
In various embodiments, the present disclosure provides a gas turbine engine comprising a compressor section configured to compress a gas, a combustor section aft of the compressor section and configured to combust the gas, and a fan shroud wherein the fan shroud comprises an outer case having an outer diameter and an inner diameter, a flange, an aerodynamic surface proximate the inner diameter of the outer case, a fan blade proximate the aerodynamic surface, and a hybrid composite-metallic containment system coupled to the inner diameter of the outer case, wherein the hybrid composite-metallic containment system comprises a cylindrical structure disposed radially outward of the fan blade and having a first composite layer extending about a circumference; a first metallic layer extending about the circumference; a second composite layer extending about the circumference, wherein the second composite layer is disposed radially inward of the first composite layer and the first metallic layer; and a binding component.
In various embodiments, the first metallic layer is disposed radially inward of the first composite layer. In various embodiments, the gas turbine engine further comprises a second metallic layer extending about the circumference radially inward of the first metallic layer. In various embodiments, the first metallic layer comprises a perforation. In various embodiments, the first composite layer comprises at least one of carbon, carbon fibers, glass fibers, aramid, para-aramid, polyethylene, or ultra-high-molecular-weight polyethylene. In various embodiments, the first metallic layer comprises at least one of steel, stainless steel, titanium, titanium alloy, aluminum, aluminum alloy, nickel, or nickel alloy. In various embodiments, the first metallic layer is at least one of sheet metal, spun metal, braided metal, metal mesh, metal fiber, woven metal, or metal fabric. In various embodiments, the binding component comprises at least one of a resin, an epoxy, a thermosetting polymer, a scrim supported adhesive, or pre-impregnated material. In various embodiments, the fan blade comprises a metal and at the first metallic layer comprises the same metal.
In various embodiments, the present disclosure provides a method of manufacturing a hybrid composite-metallic containment system, the method comprising forming a first composite layer, a first metallic layer, a second composite layer into a cylindrical structure having the second composite layer radially inward of the first composite layer and the first metallic layer; infusing the cylindrical structure with a binding component; and curing the cylindrical structure to join the first composite layer, the first metallic layer, and the second composite layer.
The forgoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated herein otherwise. These features and elements as well as the operation of the disclosed embodiments will become more apparent in light of the following description and accompanying drawings.
The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosures, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements.
The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration and their best mode. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosures, it should be understood that other embodiments may be realized and that logical, chemical, and mechanical changes may be made without departing from the spirit and scope of the disclosures. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact.
In various embodiments and with reference to
Gas turbine engine 20 may generally comprise a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A-A′ relative to an engine static structure 36 via one or more bearing systems 38 (shown as bearing system 38-1 and bearing system 38-2 in
Low speed spool 30 may generally comprise an inner shaft 40 that interconnects a fan 42, a low pressure (or first) compressor section 44 (also referred to a low pressure compressor) and a low pressure (or first) turbine section 46. Inner shaft 40 may be connected to fan 42 through a geared architecture 48 that can drive fan 42 at a lower speed than low speed spool 30. Geared architecture 48 may comprise a gear assembly 60 enclosed within a gear housing 62. Gear assembly 60 couples inner shaft 40 to a rotating fan structure. High speed spool 32 may comprise an outer shaft 50 that interconnects a high pressure compressor (“HPC”) 52 (e.g., a second compressor section) and high pressure (or second) turbine section 54. A combustor 56 may be located between HPC 52 and high pressure turbine 54. A mid-turbine frame 57 of engine static structure 36 may be located generally between high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 57 may support one or more bearing systems 38 in turbine section 28. Inner shaft 40 and outer shaft 50 may be concentric and rotate via bearing systems 38 about the engine central longitudinal axis A-A′, which is collinear with their longitudinal axes. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The core airflow C may be compressed by low pressure compressor 44 then HPC 52, mixed and burned with fuel in combustor 56, then expanded over high pressure turbine 54 and low pressure turbine 46. Mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. Low pressure turbine 46, and high pressure turbine 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
Gas turbine engine 20 may be, for example, a high-bypass geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than about six (6). In various embodiments, the bypass ratio of gas turbine engine 20 may be greater than ten (10). In various embodiments, geared architecture 48 may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture 48 may have a gear reduction ratio of greater than about 2.3 and low pressure turbine 46 may have a pressure ratio that is greater than about 5. In various embodiments, the bypass ratio of gas turbine engine 20 is greater than about ten (10:1). In various embodiments, the diameter of fan 42 may be significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 may have a pressure ratio that is greater than about (5:1). Low pressure turbine 46 pressure ratio may be measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of low pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans.
In various embodiments, the next generation of turbofan engines may be designed for higher efficiency which is associated with higher pressure ratios and higher temperatures in the HPC 52. These higher operating temperatures and pressure ratios may create operating environments that may cause thermal loads that are higher than the thermal loads encountered in conventional turbofan engines, which may shorten the operational life of current components.
In various embodiments, HPC 52 may comprise alternating rows of rotating rotors and stationary stators. Stators may have a cantilevered configuration or a shrouded configuration. More specifically, a stator may comprise a stator vane, a casing support and a hub support. In this regard, a stator vane may be supported along an outer diameter by a casing support and along an inner diameter by a hub support. In contrast, a cantilevered stator may comprise a stator vane that is only retained and/or supported at the casing (e.g., along an outer diameter).
In various embodiments, rotors may be configured to compress and spin a fluid flow. Stators may be configured to receive and straighten the fluid flow. In operation, the fluid flow discharged from the trailing edge of stators may be straightened (e.g., the flow may be directed in a substantially parallel path to the centerline of the engine and/or HPC) to increase and/or improve the efficiency of the engine and, more specifically, to achieve maximum and/or near maximum compression and efficiency when the straightened air is compressed and spun by rotor 64.
According to various embodiments and with reference to
With reference now to
In various embodiments, the thickness of containment system 322 may be between one quarter inch (0.25″) and six inches (6″), or between one inch (1″) and five inches (5″), or between two inches (2″) and four inches (4″). In various embodiments, the length of containment system 322 may be defined by the chord line of a fan blade, such as blade 206, or may be between four inches (4″) and thirty-six inches (36″), or between eight inches (8″) and twenty-four inches (24″), or between ten inches (10″) and twenty inches (20″). In various embodiments an outer case may be made of carbon, carbon fibers, titanium, titanium alloy, aluminum, aluminum alloy, or sandwich-structured composite.
In various embodiments and with reference now to
In various embodiments and with reference now to
In various embodiments and with reference now to
In various embodiments a metallic layer, such as first metallic layer 406, may have a thickness between six thousandths of an inch (0.006″) and one inch (1″), or between one hundredth of an inch (0.01″) and one half inch (0.5″), or between five hundredths of an inch (0.05″) and one quarter inch (0.25″). In various embodiments a metallic layer, such as first metallic layer 406, may be made of a metal and/or an alloy. Suitable materials for first metallic layer 406 include steel, stainless steel, titanium, titanium alloy, aluminum, aluminum alloy, nickel, and/or nickel alloy. In various embodiments the metal may be in the form of sheet metal, spun metal, braided metal, metal mesh, metal fiber, woven metal, and/or metal fabric. In various embodiments a fan blade, such as blade 206 of
In various embodiments composite layers may be laid up with metal layers in molding or curing fixture. In various embodiments an outer case, such as outer case 402, a first composite layer, a first metallic layer, a second composite layer, a second metallic layer, and a third composite layer may be co-molded and/or simultaneously infused prior to curing. In various embodiments, any number of metallic layers, composite layers, or group of metallic and composite layers comprising features similar to, for example, hybrid composite-metallic containment system 400 as shown in
In various embodiments and with reference now to
Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosures.
The scope of the disclosures is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
Systems, methods and apparatus are provided herein. In the detailed description herein, references to “one embodiment”, “an embodiment”, “an example embodiment”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiment
Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element is intended to invoke 35 U.S.C. 112(f) unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
