Patent Application
20180202299
Exemplary embodiments of the present disclosure are directed to a fan blade for a gas turbine engine and methods for mitigating galvanic corrosion of the fan blade.
A gas turbine fan blade may be made out of aluminum, and to protect the leading edge from erosion, a titanium sheath is attached. Titanium and aluminum are galvanically incompatible materials, so they are isolated from each other as best possible, using non-conductive materials. However and in the event the isolation between them is defeated, galvanic corrosion could occur to the blade. In particular and in an aluminum/titanium coupling, with aluminum being the less noble element, the blade would become the anode in the galvanic couple and accordingly, corrosion may occur on the aluminum blade.
Accordingly, it is desirable to provide a fan blade with a sacrificial anode as a method of mitigating galvanic corrosion of the fan blade.
In one embodiment, a blade for a gas turbine engine is provided. The blade having: an airfoil formed from a first material; a protective sheath disposed on a leading edge of the airfoil, the protective sheath being formed from a second material, the first material being galvanically incompatible with the second material and the first material being less noble than the second material; a non-conductive material disposed between the protective sheath and the airfoil so that they are electrically isolated from each other; a sacrificial anode in contact with the blade, wherein the sacrificial anode is formed from a third material that is less noble than the first material such that it will corrode before the first material if the non-conductive material disposed between the protective sheath and the airfoil is compromised and the first material and the second material are no longer electrically isolated from each other.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the second material may be titanium.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the third material may be zinc or magnesium.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be zinc.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be magnesium and wherein the non-conductive material may be an epoxy adhesive.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to a root of the blade.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be zinc.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be magnesium.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to an end portion of the root of the blade.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to a side portion of the root of the blade.
In yet another embodiment, a gas turbine engine is provided, the gas turbine engine having: a disk; a plurality of blades secured to the disk, each of the blades having: a root, and an airfoil formed from a first material; a protective sheath disposed on a leading edge of the airfoil, the protective sheath being formed from a second material, the first material being galvanically incompatible with the second material and the first material being less noble than the second material; a non-conductive material disposed between the protective sheath and the airfoil so that they are electrically isolated from each other; and a sacrificial anode in contact with the blade, wherein the sacrificial anode is formed from a third material that is less noble than the first material such that it will corrode before the first material if the non-conductive material disposed between the protective sheath and the airfoil is compromised and the first material and the second material are no longer electrically isolated from each other.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be zinc and wherein the non-conductive material may be an epoxy adhesive.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be magnesium.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to a root of the blade.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be zinc.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the first material may be aluminum, the second material may be titanium and the third material may be magnesium.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to an end portion of the root of the blade.
In addition to one or more of the features described above, or as an alternative to any of the foregoing embodiments, the sacrificial anode may be secured to a side portion of the root of the blade.
In yet another embodiment, a method of protecting a fan blade of a gas turbine engine from corrosion is provided. The method including the steps of: forming an airfoil formed from a first material; locating a protective sheath disposed on a leading edge of the airfoil, the protective sheath being formed from a second material, the first material being galvanically incompatible with the second material and the first material being less noble than the second material; electrically isolating the protective sheath from the airfoil with a non-conductive material disposed between the protective sheath and the airfoil; and placing a sacrificial anode in contact with the blade, wherein the sacrificial anode is formed from a third material that is less noble than the first material such that it will corrode before the first material if the non-conductive material disposed between the protective sheath and the airfoil is compromised and the first material and the second material are no longer electrically isolated from each other.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by 1 bf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec),In a turbofan engine, lighter components generally lead to more efficient performance. If less energy is expended moving internal engine parts, more energy is available for useful work. At the same time, the components themselves must be strong enough to withstand forces typical for the operating environment and performance envelope.
In order to reduce weight, the fan blades in some gas turbine engines may be may be made out of aluminum, and to protect the leading edge from erosion, a titanium sheath is attached. As discussed above it is desirable to maintain isolation between galvanically incompatible materials of the fan blade and more particularly, it is desirable to prevent the blade from becoming the anode in the galvanic couple and thus, prevent galvanic corrosion of the aluminum blade and inhibit other forms of corrosion. Although aluminum and titanium are disclosed other equivalent materials are completed to be within the scope of the present disclosure.
Referring now to
The fan blade 70 may be solid or hollow. In the event the fan blade is hollow it will have at least one internal cavity (not shown) that is enclosed by a cover or shroud.
In one embodiment, a protective sheath 80 is disposed on a leading edge 82 of the fan blade 70. In one embodiment, the airfoil 72 may be made from an aluminum alloy material and the protective sheath 80 is formed from a titanium alloy. As mentioned above and since aluminum and titanium are galvanically incompatible a non-conductive material or insulator 84 is applied between the surface of the airfoil 72 and the protective sheath 80 to electrically isolate the two materials. In other words, the non-conductive material 84 electrically isolates the protective sheath 80 from the airfoil 72. There are many materials capable of electrically isolating the sheath 80 and the airfoil 72 some non-limiting examples include: adhesives, an epoxy adhesive, urethane; and equivalents thereof each of which are contemplated to be within the scope of the various embodiments of the present disclosure.
In accordance with one embodiment and referring now to
Thus, if the sheath-to-blade electrical isolation is defeated (e.g., non-conductive material 84), any corrosion would occur first on the least noble element in the system, which would be the non-structural zinc or magnesium piece or sacrificial anode 86. If these sacrificial anode(s) 86 are corroded away, the sacrificial pieces would also be easily identified and easily replaceable during engine overhaul. For example and as will be discussed herein the sacrificial anode(s) 86 may be located in areas that are easily viewable during service of the fan 42 of the gas turbine engine 20.
In one embodiment and referring now to at least
It is understood that while only a single blade 70 is illustrated in
The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” can include a range of ±8% or 5%, or 2% of a given value.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
